• Journal of the Korean Society of Propulsion Engineers
• /
• v.21 no.2
• /
• pp.26-31
• /
• 2017

Environmental problem of the solid propellants is an issue of growing importance in solid rocket. For examples, ammonium perchlorate (AP) as an solid propellants oxidizer could create a poisonous gas and atmospheric pollutions, such as HCl. Among the several oxidizers, N-guanylurea dinitramide (GuDN) is an effective candidate substance for eco-friendly oxidizer, which has high performance, pressure exponent, and eco-friendly smog during combustion for solid propellant of gas generator. In this paper, the theoretical analysis of characteristics as a gas generator propellant, propellant manufacturing processability, propellant hardness properties and combustion characteristics were studied.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.05a
• /
• pp.257-262
• /
• 2009

Elastic-Plastic structural analysis was performed to evaluate structural characteristic and integrity for rocket motor case of solid propulsion system. The structural analyses were compared and evaluated using the simplified 2-D axisymmetric model and 3-D full model for rocket motor case with torispherical dome type. And pre-tension load for bolt model was considered in structural analysis. The results of displacement and stress for the simplified 2-D axisymmetric model and 3-D full model were in an good agreement with each other. Therefore, the simplified 2-D axisymmetric model for rocket motor case was recommended to verify quickly the structural integrity and save the modeling and calculating time in initial design stage. Also, the structural characteristic and integrity for rocket motor case according to 5 dome types was evaluated to select the optimal dome shape.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.22 no.2
• /
• pp.96-107
• /
• 2018

Flow analysis using computational fluid dynamics was conducted to investigate the effect of the igniter flame caused by the gap between the igniter and the propellant grain in a solid rocket motor. Two propellant grain types were assumed; namely cylinder type (1 mm, 3 mm, and 5 mm gap) and the slot type. The slot type had two igniter hole locations. One was located at the small gap of the propellant grain, and the other one was located at the large gap. In the case of the cylinder type, the pressure in the igniter zone was higher with a thinner gap. Additionally, in the case of the cylinder type, the pressure difference between the igniter installed zone and the free volume was also higher as the gap became lower. The cylinder types were affected by the gap distance, but the slot types were not. Moreover, the results of the slot types were similar to the 5-mm gap case of the cylinder type.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.17 no.5
• /
• pp.132-139
• /
• 2013

Liquid-propellant rocket engines are widely used all over the world, thanks to their high performances thrust, in particular high thrust-to-weight ratio. The sucess rate of the launching of the liquid propulsion is similar to the solid one even though it has more complex mechanical system. In general, liquid propulsion is seemed as a mature technology, the requirements of a renewed interest for space exploration has led to the development of a family of new engines, with more design margins, simpler to use and to produce associated with a wide variety of thrust and life requirements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.05a
• /
• pp.246-250
• /
• 2006

In general, combustion characteristic of hybrid propulsion was shown with the regression rate depending on only massflow rate of oxidizer But this empirical relation was not represented well effect of the thermo-chemical properties of solid fuel. So, in this study, the combustion characteristics was studied with the mass transfer number(B number) of solid fuel instead of regression rate with various fuel. The PMMA, PP, and PE were used as fuel, and gas oxygen as oxidizer in this experiment. The mass flowrate of gas oxigen was controlled by the several chocked orifices that have different diameter, and the oxidizer supply range was $3.66\sim45.3g/sec$. As result, the empirical relation for mass flux of solid fuel was obtained with mass transfer number, and mass flux of oxidizer as follow; $\dot{m}^{'}_f\;=\;0.0175G^{0.55}B^{0.4}$.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.03a
• /
• pp.593-596
• /
• 2004

We are developing a micro-solid propellant rocket array thruster for simple attitude control of a 10 kg class micro-spacecraft. The prototype has ø 0.8 mm solid propellant micro-rockets arrayed at a pitch of 1.2 mm on a 22 x 22 mm substrate. In previous studies, an impulse thrust of 4.6 x 10$^{-4}$ Ns was obtained in vacuum, but we found the problems of unacceptably low ignition success rate and incomplete combustion. This paper describes experiments to improve the ignition rate. In order to achieve this goal, we tried to solidify paste-like ignition aid (RK) on the ignition heaters with strong adhesion. To make the paste-like RK, isoamyl acetate was added to RK powder. We tested 9 rockets, but only 2 rockets were ignited with huge ignition energy. This is because the heat con-duction between the ignition heater and the RK was too low to ignite the RK, since dried RK had a lot of pores. Also, a large cavity was sometimes found just above the ignition heater.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.5 no.2
• /
• pp.16-23
• /
• 2001

For liquid rocket engine test, one of most important design parameters to be measured is thrust. However, not like solid rocket motor, a liquid rocket engine is attached to the propellant feed system, control valve and many other safety systems. Without considering these effects, thrust data measured from firing test is not reliable and sometimes almost meaningless. In this research, new thrust measurement system, which includes all these side effects, was designed and fabricated.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.6 no.2
• /
• pp.75-84
• /
• 2002

A numerical analysis of unsteady motion in solid rocket motors with a nozzle has been conducted. The numerical formulation including modified $\kappa$-$\varepsilon$ turbulence model treats the complete conservation equation for the gas phase and the one-dimensional equations in the radial direction for the condensed phase. A fully coupled implicit scheme based on a dual time-stepping integration algorithm has been adopted to solve the governing equations. After obtaining a steady state solution, pulse and periodic oscillations of pressure are imposed at the head-end to simulate acoustic oscillations of a travelling-wave motion in the combustion chamber. Various steady and unsteady state features in the combustion chamber of a rocket motor has been analyzed as results of numerical calculations.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.05a
• /
• pp.129-132
• /
• 2010

The objective of IM rocket motor is to minimize the probability of inadvertent initiation and severity of subsequent collateral damage, hence it is important to define personnel and equipment survivability to a rocket motor accident. The violent response probability associated with shock, impact and thermal effects be minimized. And during production, transportation/storage and stack of rocket motor, sympathetic detonation, giving severe effects of the propagation of adverse reaction on its surroundings, be reduced. Hence Reaction type also based on reaction results of the overpressure, fragment throw and heat flux.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.10a
• /
• pp.251-254
• /
• 2004

Hybrid rockets have many advantages over solid and liquid rockets. Hybrid rockets put forth high $I_{sp}$ like liquid rockets in spite of simple structure and low cost. As oxidizer flow rate is increased, thrust of hybrid rocket is increased accordingly. In this study, lab-scale hybrid rocket is designed, fabricated and tested. This system consists of lab-scale hybrid rocket motor, ignition system, flow system and data aquisition system. In order to control oxidizer flow rate, we construct flow rate control system by using needle valve and stepping motor.