• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.6-10
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• 2003

Accurate description of starting transient history allows and justifies the use of small margin of safety for the engine parts, resulting in high motor mass ratio in addition to satisfying the control and guidance requirements of the vehicle. Studies have been carried out for the prediction and reduction of ignition peak and pressure-rise rate during the starting transient of solid rocket motors. Numerical studies have been carried out using a two dimensional Navier-Stokes solver. It has been inferred through the parametric studies that, in the case of solid rocket motors with uniform port, high ignition peak is observed at high spread rate and low pressure-rise rate. In the case of the port with sudden expansion configuration, high ignition peak is observed at relatively high average spread rate and high-pressure rise rate. These studies are expected to aid the designer in reducing the ignition peak by altering the propellant properties or igniter characteristics without sacrificing the motor performance.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.36 no.5
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• pp.563-569
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• 2012

An energy balance analysis is conducted for a 30 t thrust level liquid rocket engine. The relations between thrust and combustion pressure, between thrust and propellant flow rate, and between combustion pressure and fuel pump pressure rise are compared against those indicated by a published database of the existing rocket engines. A combustion pressure higher than the old design value is obtained, implying that the present design is high-performance oriented. The thrust to propellant flow rate ratio is the same as that of the existing engines, indicating that the specific impulse performance is at the usual level. The fuel pump pressure rise is found to be slightly high when the combustion pressure is considered, and it is attributed to the pressure budget of the present ground test engine not being optimized.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.4
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• pp.369-375
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• 2010

The study on the characteristics of nitrous oxide catalytic decomposition was carried out to utilize the nitrous oxide as a propellant. The Pt, Ir and Ru were synthesized to select a high performance catalyst for the nitrous oxide decomposition reaction. The respective catalyst precursors were loaded in the $Al_2O_3$ support using an wet impregnation method. The $N_2O$ conversion as a variation of space velocity and reaction temperature was measured using a tubular reactor. The catalyst loss was measured to evaluate the durability of catalysts after the reaction at $800^{\circ}C$ for 2 hours. The $N_2O$ conversion was increased at the decrease of space velocity and at the increase of temperature. The Ru/$Al_2O_3$ catalyst had the highest $N_2O$ conversion at low temperature and the best durability.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.4
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• pp.74-80
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• 2014

This paper describes the development of the dual-thrust flight motor for enhancing the hit probability of unguided rockets. We designed dual-thrust flight motor by shape modification of the double base propellant with high burning rate, and confirmed the dual-thrust performance by static firing tests. The test results showed the thrust ratio of about 1:7.6 between sustaining phase and boosting phase, and had a quietly normal dual-thrust characteristics. And the results showed that there was not the fire extinction phenomenon of propellant due to the pressure drop.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.74-77
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• 2011

A mixing characteristics on recess length change of Gas-centered swirl coaxial injector using high-performance staged combustion rocket engine carry out study through CFD(Computational fluid dynamics). propellant phase that combined gas-liquid simulate gas-gas. In order to measure spreading angle, velocity distribution to injector exit and spray structure of propellant analyzed. Axial velocity increase by increasing recess length, but tangential velocity decrease. The result confirmed qualitative characteristics that the spreading angle decreases.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.406-408
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• 2017

Ammonium dinitramide (ADN) Low toxicity monopropellant based 1N class thruster and test equipment were developed. Compared with the hydrazine which used in existing satellite thruster, ADN is easy to handle and has excellent physical characteristics such as density and specific impulse. Due to these characteristics, ADN is attracting attention as an eco-friendly propellant. In this paper, 1N class thruster and thrust measurement system was designed for performance testing of ADN monopropellant. The composition of the propellant for the design and experiment was set at 11.2: 25.4: 63.4 for each of Methanol: $H_2O$: ADN.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.866-870
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• 2008

In this study, we designed cold gas propulsion system with minimum 0.25 mm nozzle and micro-thrust measurement system to analyze flow characteristic of micro propulsion system in ambient and vacuum condition. Argon and Nitrogen are used for propellant and the result of experiments is compared with CFD analysis and theory. But there is a point where reduced scale versions of conventional propulsion systems will no longer be practical. Therefore, a fundamentally different approach to propulsion systems was taken. That is thermal transpiration based micro propulsion system. It has no moving parts such as lubricants, pressurizing system and can pump the gaseous propellant by temperature gradient only(cold to hot). We are advancing basic research of propulsion system based on thermal transpiration in vacuum conditions and had tried experiment process and theoretical access in advance. To characterize membrane of Knudsen pump, we select Polyimide material that has low thermal conductivity(0.29 W/mK) and can stand high temperature($300^{\circ}C$) for long time. And we fabricated hole diameter 1, 0.5, 0.2, 0.1 mm using precision manufacturing. Experimental results show that pressure gradient efficiency of Knudsen pump is increased to maximum 82% according to Knudsen number and thick membranes are more effective than thin membranes in transition flow regime.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.3
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• pp.112-118
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• 2019

The development of an underwater rocket propulsion system was described in this paper. Throttle able liquid propellant and hybrid rocket propulsion systems were selected for underwater propulsion. A subscale liquid rocket combustion chamber and it's portable stand were developed and their applicability was examined. 1.5-ton.f and 1.8-ton.f hybrid rockets were developed for underwater applications. The test results indicated that the 18-ton.f hybrid rocket fully complies to the performance and underwater cruise stability requirements; thus, its development was concluded to be successfully complete.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.03a
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• pp.519-524
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• 2008

A numerical study is conducted to investigate propulsion performance enhancement based on S225 experiment case of ISL(French-German Research Institute of Saint-Louis)'s superdetonative ram accelerator. For govern equation, multi-species Navier-Stokes equation coupled with Baldwin-Lomax turbulence modeling is used. Govern equation is discretized by Roe's FDS and integrated by LU-SGS time integration. Detailed chemical reaction about $H_2/O_2/CO_2$ for high pressure is considered. $2H_2+O_2+2.5CO_2$ mixture was used for propellant gas. For the same over-driven factor, the launching speed of computation was faster than one of S225. Another configuration and condition of S225 was applied. A flame structure is very different from S225. For strong mixture case, it shows ignition by viscous effect. Acceleration and speed increment is higher than S225 computation and experiment. By using more strong mixture, propulsion performance was enhanced.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.5
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• pp.31-36
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• 2007

Combustion performance characteristics of subscale high-pressure combustor were investigated at 70 bar combustion pressure. All tests were successfully performed without any damage on the combustor. The mixing characteristics and distribution pattern of the injectors were found to have considerable influence on the combustion performance. The characteristic velocity of the combustor was higher in the injector with internal mixing than that of external mixing and in the injector with smaller mass flowrate. The pressure fluctuations at the propellant manifolds and the combustion chamber were measured to be less than 3% of the mean combustion pressure to meet the combustion stability criterion and to prove stable combustion characteristics of the combustor.