• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.258-261
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• 2003

The optimal design and combustion analysis of the gas generator for Liquid Rocket Engine (LRE) were performed. A fuel-rich gas generator in open cycle turbopump system was designed for 101on1 in thrust with RP-1/LOx combination. The optimal design was done for maximizing specific impulse of main combustion chamber with constraints of combustion temperature and power matching in turbopump system. Results of optimal design show the dimension of length, diameter, and contraction ratio of gas generator. The configuration of the gas generator and the condition for performance which can maximize the objective function were determined and found to meet the design constraints. Also, the combustion analysis was conducted to evaluate the performance of designed chamber and injector of gas generator. And the effect of the turbulence ring was investigated on the mixing enhancement in the chamber.

• Clean Technology
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• v.10 no.3
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• pp.159-168
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• 2004

Benzene adsorption experiment was carried out for activated carbon and zeolite 13X adsorbents. Single column and dual column packed with two adsorbents were used to investigate the dynamic adsorptive characteristics. Effect of feed flow rate on the breakthrough curve was not significant. Specific adsorption amount of benzene for activated carbon was larger than that for zeolite 13X. On the contrary, adsorption amount per column volume was larger for zeolite 13X column because the density of zeolite 13X was larger. In the dynamic experiment using dual adsorbents column, length of mass transfer zone was changed by the feed direction. Breakthrough time was longer and breakthrough curve was sharper when activated carbon was packed in feed inlet and zeolite 13X was packed in column outlet. Also breakthrough time and breakthrough curve slope were affected by the packing ratio of the two adsorbents.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.38 no.3
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• pp.255-262
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• 2014

To investigate the effect of oil on evaporation heat transfer of plate heat exchanger, evaporation heat transfer experiment was carried out using experimental apparatus for micro gear pumped R134a-oil circulation. By varying oil circulation rate of POE oil from 0 to 5%, evaporation heat transfer performance of plate heat exchanger was investigated. As OCR(Oil Circulation Ratio) increases, the evaporation heat transfer coefficient of R134a decreases and pressure drop increases. When the evaporating temperature is $30^{\circ}C$ and the refrigerant mass flow rate is 80 g/s, evaporation heat transfer rate decreases by 10 % and pressure drop increases by 10% at 2% of OCR condition.

• Aerospace Engineering and Technology
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• v.1 no.1
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• pp.163-172
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• 2002

The turbine system composed of a nozzle and a rotor is used to drive turbopumps while gas passes through the nozzle, potential energy is converted to kinematic energy, which forces the rotor blades to spin. In this study, an aerodynamic design of a turbine system is investigated using compressible fluid dynamic theories with some pre-determined design requirements (i.e.,pressure ratio, rotational speed, required power etc.) obtained from a liquid rocket engine (L.R.E.) system design. For simplicity of a turbine system, impulse-type rotor blades for open type L.R.E. have been chosen. Usually, the open-type turbine system requires low mass flow rate compared to the close-type system. In this study, a partial admission nozzle is adopted to maximize the efficiency of the close-type turbine system. A design methodology of the a turbine system has been introduced. Especially, a partial admission nozzle has been designed by means of simple empirical correlations between efficiency and configuration of the nozzle. Finally, a turbine system design for a 10 ton thrust level of L.R.E is presented.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.1
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• pp.23-32
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• 2016

Various studies have been conducted for drag reduction of a high-speed vehicle by injecting counterflow jet from its nose cone. In this study, in order to obtain baseline data and key parameters for drag reduction method, the counterflow jet study of the USA is reviewed and summarized. The nose cone shapes of each study are hemisphere cylinder, truncated cone, and reentry capsule, and their test conditions are summarized accordingly. Key parameters for drag reduction are jet mach number, mass flow rate, and pressure ratio. Even though drag reduction effects show various results according to given test conditions, it is found that the drag reduction effect reaches up to 40~50%.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.457-460
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• 2012

Design and fabrication of a sub-scale thrust chamber for verification of 7 tonf-class thrust chamber injectors were described in this paper. The 7 tonf-class thrust chamber consists of mixing head with 90 coaxial swirl injectors and regeneratively combustion chamber cooled by kerosene. The coaxial swirl injectors with different pressure drop and recess number were designed for 7 tonf full-scale thrust chamber. By applying the designed injectors to the sub-scale thrust chamber before applying them to the full-scale thrust chamber, the injector performance and functioning were verified. The sub-scale thrust chamber consists of 19 injectors, has chamber pressure of 70 bar, total propellant mass flow rate of 4.3 kg/s, mixture ratio(O/F) of 2.45.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.2
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• pp.9-15
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• 2013

An effect of a spiral turbulent ring, so-called Shchelkin spiral, on a detonation performance was studied experimentally for acetylene and oxygen mixture. A couple of dynamic pressure transducers were used to calculate a detonation wave velocity by a time difference between two pressure peaks. In addition, impulse was measured by a load cell and the impulse was used to analyze the spiral effect on the detonation performance. A CFD analysis was adopted to calculate mass flow rates of the propellants and the minimum filling time. The maximum velocity and pressure were measured at the equivalence ratio of 2.4, and the measured values showed similar trend to C-J conditions calculated from CEA. For the shorter chamber with the short spiral, the maximum detonation velocity was appeared. In contrast, the longer chamber without the spiral showed the maximum thrust performance.

• Design & Manufacturing
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• v.13 no.2
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• pp.37-43
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• 2019

This study conducted a research as to condensation heat transfer friction loss headby using three types of flat micro multi-channel tubes with different processing of micro-fin and number of channels inside the pipes and different sizes of appearances. In addition, identical studies were conducted by using smoothing circular tubes with 5mm external diameter to study heat enhancement factor and pressure drop penalty factor. 1) The friction head loss showed an increase as the vapor quality and mass flux increased. In case of saturation temperature, it shows an increase as it gets lower. These factors are the reason occurring as the lower the saturation temperature is, the higher the density of refrigerant vapor gets. The influence of heat flux is similar as the dryness is low, but as it gets higher, it lowers in heat flux, and as the high temperature of high heat flux, it is a factor that occurs as the density gets lower. 2) RMS error of the in case of friction head loss, it showed to be predicted as 0.45~0.67 by Chisholm, Friedel, Lockhart and Martinelli. 3) As forfriction head loss penalty factor, the smaller the aspect ratio is, the larger the penalty factor gets, and as for the effect of micro-fin, the penalty factor increased because it decreases to the gas fluid the way groove for the refrigerant's flow.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.3
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• pp.58-66
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• 2019

In this study, the start-up characteristics of a staged combustion engine were analyzed numerically based on relational equation modeling of the entire engine components. The start-up characteristics were extensively analyzed considering the transient period of the total engine system from the start-up sequence till the steady-state of the engine. The performance characteristics of the engine components such as RPM of engine power-pack, chamber pressure and O/F ratio of pre-burner, and mass flow of propellants in the start-up period were investigated. Furthermore, the calculated engine data were compared satisfactorily with the experimental data. Through the comparison of data, successful validation of present engine start-up analysis has been obtained.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.5
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• pp.21-33
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• 2020

In this study, the program for system analysis of an expander cycle rocket engine using liquid hydrogen as a fuel was developed. The properties of hydrogen were considered by the ratio of isomers with temperature. The analysis procedure was established with the open and closed types of the expander cycle engine and the simulation methods were suggested for each component. To validation of the analysis program, we compared the performance of the engine operating point and the analysis results performed overseas for Vinci and SE-21D, which are expander cycle engines. As a result of the analysis, the main performance factors of the system, such as the mass flow of the propellant, specific thrust, and power, except for some of the inaccurate input information, showed high accuracy with an error of around 1-2%.