• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.237-240
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• 2006

Hybrid propulsion systems provide many advantages in terms of stable operation and safety. However, classical hybrid rocket motors have lower fuel regression rate and combustion efficiency compared to solid propellant rocket motor. Accordingly, the recent research efforts are focused on the improvement of engine efficiency and regressionrate in the hybrid rocket engine. The present study has numerically investigated the combustion processes and the flame structure in the hybrid rocket engine. The turbulent combustion is represented by the flamelet model and Low Reynolds number $k-{\varepsilon}$turbulent model is employed to reduce the uncertainties for convective heat transfer near solid fuel surface having strong blowing effect. Numerical results suggest that the present approach is capable of realistically simulating the combustion characteristics of the hybrid rocket engines.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2001.11a
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• pp.61-67
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• 2001

Hybrid propulsion systems provide many advantages in terms of stable operation and safety. However, classical hybrid rocket motors have lower fuel regression rate and combustion efficiency compared to solid propellant rocket motor. The recent research efforts are focused on the improvement of volume limitation and regression rate in the hybrid rocket engine. The present study has numerically investigated the combustion processes in the hybrid rocket engine. The turbulent combustion is represented by the eddy breakup model and Hiroyasu and Nagle and Strickland-Constable model are used for soot formation and soot oxidation. Radiative heat transfer is modeled by finite volume method. To reduce the uncertainties for convective heat transfer near solid fuel surface having strong blowing effect, the Low Reynolds number k-$\varepsilon$ turbulent model is employed. Based on numerical results, the detailed discussion has been made for the turbulent combustion processes in the vortex hybrid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.424-428
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• 2012

Air-breathing Propulsion System is one of the promising propulsion systems because of low cost, easy storage, compactness and simplicity. A study of gas generator propellant for air-breathing propulsion system was performed in this paper. Amorphous Boron Powder was applied in propellant with various kinds of additives to determine combustion characteristics. And boron beads were made to apply them to the propellant. Combustion characteristics of propellant using amorphous boron powder and boron beads was compared.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.293-298
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• 2011

The combustion and flow-rate control characteristics of fuel-rich gas generator which could be adopted to Ducted Rocket propulsion system are investigated. The gas generator is designed considering the design requirements of propulsion system and solid fuel for fuel-rich combustion is developed then adopted to ground test. The results of combustion test show the necessity of the special analysis tool for estimating the gas generator performance where multi-phase flow of fuel-rich gas exists. During the flow-rate control test, characteristics of gas generator pressure with the angle of valve are analyzed and, method to estimate the pressure of gas generator is suggested using the relation between the valve exit area and discharge coefficient.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.5
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• pp.53-59
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• 2019

Solid fuel reacts with an oxidizer during combustion of a propellant to increase performance. Representative solid fuels are aluminum, cyclotrimethylenetrinitramine (RDX) and octahydro-1, 3,5,7-tetra nitro-1,3,5,7-tetrazocin (HMX). During combustion, these compounds generate white smoke by reacting with moisture and produce materials that are harmful to the environment, such as carbon monoxide, carbon dioxide, and methane gas. This study prepared a high-nitrogen-containing energetic material, hydrazinium 5-aminotetrazolate (HAT), which could be applied as a solid fuel. The compound was characterized by nuclear magnetic resonance (NMR) spectroscopy, and a thermal analysis was measured by differential scanning calorimetry (DSC). Also, the specific impulses and volumes of detonation gases were calculated using the EXPLO5 program.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.886-898
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• 2008

A survey is offered of the present status of microaluminized propellants industrially used worldwide in most space applications, but new directions are also pointed out making profitable use of the nanoaluminized propellants currently tested in many laboratories. Different industrial- and research-type of solid rocket propellants, mainly but not only, of the well-known family oxidizer/Al/HTPB(oxidizer being AP, AN or a mixture of the two) were experimentally analyzed at the Space Propulsion Laboratory of Politecnico di Milano. In general, they feature the same nominal composition but implement different grain size distributions of the oxidizer or metal fuel. The basic properties of all formulations were compared to that of a standard propellant already certified for flight.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.331-334
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• 2009

An air-independent propulsion (AIP) system based on fuel cell technologies was developed for space and underwater applications in the present study. Hydrogen peroxide was selected as an oxidizer for space and underwater power applications where air independence is a must. Catalytic decomposition of hydrogen peroxide was used to generate oxygen and water. The pure oxygen was provided to a fuel cell and the water was stored separately. Sodium borohydride in the solid state was used as a hydrogen source in the present study. Pure hydrogen can be generated by a catalytic hydrolysis reaction. A fuel cell system was fabricated to validate the fuel cell based air-independent power system and was evaluated at the various conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.957-967
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• 2017

Firing test of the fuel grain for solid fuel ramjet with additives and ignition support material was conducted. Fuel grain consist of HTPB mixed with AP particle 15 wt.%, Boron particle 5 wt.%. To cause the short ignition delay, ignition support consist of $NC/BKNO_3$ and composite propellant was coated to the fuel grain. An oxidant gas having a controlled temperature, pressure and oxygen composition close to the air condition in the ramjet combustor was supplied using the Ethanol blended $H_2O_2$ gas generator. Gas was set to flow at a mass flow rate of 150 g/s and mass flux of $200kg/m^2s$ in the grain port. Through the test, ignition support operated well and ignition delay of 0.5. During the test, stable chamber pressure with 8 bar and high combustion efficiency of 0.86 was confirmed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.451-454
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• 2011

Hydro-Reactive metal Fuel (HRF) which has more fuel than general solid propellant reducing oxidizing agent is suitable for ultrahigh speed rocket motor in the water. However, burning rate of HRF has not been studied yet. Through the earlier studies, we established ultrasonics measurement system measuring burning rate of solid propellant as a function of pressure in a single test and verified its reliability. In this paper, we will measure burning rate of HRF using ultrasound with previous development measurement system.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.469-473
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• 2010

In this study, combustion tests using liquefying fuels with fast regression rate were performed. The chamber pressure oscillation was analyzed and hazards of combustion instabilities were examined. In case of Liquefying fuel with fast regression rate, the amplitude of chamber pressure oscillation was increased compared to the polymeric fuels. However, the critical combustion instability can hardly occur in liquefying fuel. This is because the rapid change of inner chamber diameter limits the amplification of chamber pressure oscillation. The chamber pressure oscillation due to the large increase of fuel production and the vortex shedding in pre-chamber violently occurs during combustion using single-port axial injector.