• Journal of the Korean Society of Propulsion Engineers
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• v.1 no.2
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• pp.58-66
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• 1997

To investigate the effect of the oxidizer mass flow rate on the fuel regression rate of the hybrid rocket, a laboratory size rocket was designed and ground fire test were carried out. Oxidizer was gaseous oxygen and HTPB was used as a fuel. Following correlation was obtained from the experiment. $\dot{r}$=$0.183G_o^{0.605}$

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.203-207
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• 2005

An experimental study was carried out to investigate the effect of film cooling in the thrust chamber of liquid rocket using LOx and Kerosene as propellant. The heat fluxes were obtained from the measured wall temperature to the axial direction of thrust chamber for different type of coolant, the various O/F ratio, mass flow rate and the location of the film cooling injector. A thin wall combustion chamber and nozzle were used to obtain the heat flux.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.420-421
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• 2011

In this study, Instead of using urethane curative systems, which have long been used as solid propellants, a triazole curative system has been introduced into a new binder recipe in which azide groups in the polymer react with triple bonds of a dipolarophile curative.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.207-210
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• 2007

This paper describes the development of the grain design automation program using commercial CAD(computer-aided-design) software. This program allows to readily obtain output of burning area, volume, moments of inertial, motion of the center of gravity, and other geometric data as a function of regressed distance. These are utilized in performance analysis.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1997.11a
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• pp.15-15
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• 1997

본 연구는 전방점화 방식(the head-end ignition)을 채택하고 있는 composite 고체 추진기관(구룡 1모타)을 이용하여 후방점화 방식(the after-end ignition)에 의한 점화 가능성을 검토하였으며, 점화 방식 차이에 따른 추진기관의 초기 연소거동의 차이점을 고찰하고자 한 실험 연구로서, 후방 점화장치를 설계·제작하여 지상연소시험을 수행하였다. 점화장치는 착화장치(initiation system)와 에너지 방출장치(energy release system), 구조물(Hardware)로 구성되는데, 착화장치는 기존의 K2 squib를 사용하였고, 에너지 방출장치는 FRP튜브에 MTV pellet 점화제를 사용하였으며, 점화기를 후방에 부착시키는 방법으로는 flexible finger 형태의 locking sleeve를 설계하여 노즐목에 고정하였다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.283-287
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• 2008

The heat released during the external frictional motion is a factor responsible for initiating energetic materials under all types of mechanical stimuli including impact, drop, or penetration. We model the friction-induced ignition of HMX, RDX and AP/HTPB propellant using the BAM friction apparatus and one-dimensional time-to-explosion apparatus whose results are used to validate the friction ignition mechanism and the deflagration kinetics of energetic materials, respectively. The ignition times for each energetic sample due to friction are presented.

• Proceedings of the KSME Conference
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• 2001.06e
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• pp.288-293
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• 2001

This paper presents the water-hammer effect due to the rapid opening and closing of isolation valve and thruster valve in the spacecraft propulsion system. The single propellant feed system was modeled to investigate the maximum peak pressure due to the water-hammer effect. The test parameters are tank supply pressure, shape and throat length of orifice and line length. Kerosene was used as the inert simulant propellant liquid instead of hydrazine. As downstream line length after isolation valve increased from 1.5 to 2.5m, the maximum line-filling water-hammer peak pressure decreased, but the average time interval between peak pressures increased. The maximum line-filling water-hammer peak pressure with orifice was lower than without orifice, and the maximum line-filling water-hammer peak pressure with orifice at the back of isolation valve was lower than with orifice in front of isolation valve. Without orifice, the maximum water-hammer peak pressure due to the rapid opening and closing of the thruster valve was about 126% of tank supply pressure. With orifice, it decreased. As orifice throat length increased, it decreased. The maximum water-hammer peak pressure due to the rapid closing of the thruster valve with converging-diverging orifice was lower than normal orifice. It was found that the orifice as a means of pressure drop was very effective to reduce the water hammer peak pressure at the thruster valve. The results of this study can be used for the design of spacecraft liquid propulsion feed system.

• Journal of the Korea Society for Simulation
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• v.28 no.2
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• pp.1-14
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• 2019

A full scale hydrodynamic simulation that requires an accurate reproduction of shock-induced detonation was conducted for design of an energetic component system. A detailed hydrodynamic analysis SW was developed to validate the reactive flow model for predicting the shock propagation in a train configuration and to quantify the shock sensitivity of the energetic materials. The pyrotechnic device is composed of four main components, namely a donor unit (HNS+HMX), a bulkhead (STS), an acceptor explosive (RDX), and a propellant (BPN) for gas generation. The pressurized gases generated from the burning propellant were purged into a 10 cc release chamber for study of the inherent oscillatory flow induced by the interferences between shock and rarefaction waves. The pressure fluctuations measured from experiment and calculation were investigated to further validate the peculiar peak at specific characteristic frequency (${\omega}_c=8.3kHz$). In this paper, a step-by-step numerical description of detonation of high explosive components, deflagration of propellant component, and deformation of metal component is given in order to facilitate the proper implementation of the outlined formulation into a shock physics code for a full scale hydrodynamic simulation of the energetic component system.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.6
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• pp.11-18
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• 2014

The combustion flowfield at the near-injector region of a 50 N scale $N_2O/C_2H_5OH$ thruster was visualized using shadowgraph technique. The explosive ignition was occurred at the design spray condition, and the expanding combustion gas quenched the flame immediately. Approximately after 83 ms from the initial ignition, the propellant spray was re-ignited, and the flame was stabilized after 23 ms elapsed. In the increased oxidizer flow rate condition, the transient pressure at the moment of ignition was smoother than explosive ignition, and the blow down phenomenon was not appeared in the same operating sequence. In addition, the flame was stabilized within 17 ms, and it is caused by improved propellants mixing before ignition.

• Journal of the Korean Applied Science and Technology
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• v.36 no.1
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• pp.165-173
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• 2019

Hydrazine[$N_2H_4$] is a typical propellant for a rocket fuel in the field of aerospace. Since it is very toxic and harmful to the environment, various environmentally-friendly propellants have been developed. In this study, relatively a safe propellant, hydroxylammonium nitrate[$NH_3OHNO_3$], was prepared via a neutralization reaction of hydroxylamine[$NH_2OH$] and nitric acid[$HNO_3$]. FT-IR was used to analyze the chemical composition, chemical structure and functional groups of HAN. Thermogravimetric analysis showed the decomposition temperature of HAN. Ion chromatography was also used to evaluate the content of nitrate ions. It was proved that the peaks of FT-IR at $3161cm^{-1}$ and $1324cm^{-1}$ indicates the functionalities of N-H and N-O present in HAN. The decomposition temperature of HAN synthesized at pH 5 to 7 was $120-140^{\circ}C$, and pH 8 resulted in higher decomposition temperature than $140^{\circ}C$. Meanwhile, the sample obtained from pH 6-7 showed the concentration of nitric acid ion with 70%.