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REFERENCE LINKING PLATFORM OF KOREA S&T JOURNALS
> Journal Vol & Issue
Journal of the Korean Society of Propulsion Engineers
Journal Basic Information
Journal DOI :
The Korean Society of Propulsion Engineers
Editor in Chief :
Volume & Issues
Volume 18, Issue 6 - Dec 2014
Volume 18, Issue 5 - Oct 2014
Volume 18, Issue 4 - Aug 2014
Volume 18, Issue 3 - Jun 2014
Volume 18, Issue 2 - Apr 2014
Volume 18, Issue 1 - Feb 2014
Selecting the target year
Effect of Particle Size and Velocity Ratio on the Flow Mixing Characteristics in the Secondary Combustor
Park, Jung Shin ; Park, Soon Sang ; Han, Doo-Hee ; Shin, Jun-Su ; Sung, Hong-Gye ; Kwak, Jae Su ; Choi, Ho-Jin ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 1~7
DOI : 10.6108/KSPE.2014.18.1.001
In this study, the effect of velocity ratio and particle size on the flow mixing characteristics in the secondary combustor was investigated. Both PIV(Particle Image Velocimetry) technique and LES(Large Eddy Simulation) were applied. Two sizes of Polystyrene PIV seeding particle of 5 and
, and three velocity ratios of 5, 3, and 1.5 were considered. Results showed that the mixing of two air streams created reattachment and recirculation regions. The size of the recirculation region was decreased as the velocity ratio increased. For the larger particle cases, due to the increased momentum by the larger particles, the size of the recirculating regions were larger than that of the smaller particle cases and the effect of the velocity ratio was not as significant as in the smaller particle case.
The Effect of Pressure and Oxidation Mole Fraction on Ablation Rate of Graphite for Nozzle Throat Insert
Hahm, Heecheol ; Kang, Yoongoo ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 8~15
DOI : 10.6108/KSPE.2014.18.1.008
The ablation characteristics of graphite nozzle throat insert is analyzed for the use in solid rocket propulsion system. The propulsion system is composed of three types of conventional nozzles, such as De-Laval type, blast tube type, and submerged type. Various kinds of propellants are used in the thirteen kinds of propulsion system that has different shapes of each other. Total thirty seven tests are performed. From the results of the analysis, it is found that the ablation rate is higher for the higher average chamber pressure and the higher concentration of oxidizing species in combustion gas.
Development of System Analysis Program of Liquid Rocket Engine II
Lee, Sangbok ; Son, Min ; Seo, Jongcheol ; Lim, Taekyu ; Roh, Tae-Seong ; Koo, Jaye ; Kim, Kuisoon ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 16~25
DOI : 10.6108/KSPE.2014.18.1.016
The system analysis and design program of the liquid rocket engine has been developed for preliminary conceptual design process. The program consists of modular programs analyzing the main thruster, the gas-generator, turbo-pumps, the turbine, pipes, valves and so on. Each module has been developed in order to estimate performance, weight, and shape parameters of the components. The results of them have been verified with experimental data or other programs.
Temperature Field and Emission Spectrum Measurement of High Energy Density Steam Plasma Jet for Aluminum Powder Ignition
Lee, Sanghyup ; Lim, Jihwan ; Lee, Dohyung ; Yoon, Woongsup ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 26~32
DOI : 10.6108/KSPE.2014.18.1.026
In this study, DC (Direct current) type steam plasma igniter is developed for effective ignition of high-energy density metal aluminum and gas temperature is measured by emission spectrum of OH radical. Because of the ultra-high gas temperature, the DC plasma jet is measured by Boltzmann plot method which is the non-contact optical technique and spectrum comparison-analysis. And both methods were applied to experiment after accurate verification. As a result, we could identify that plasma jet temperature is 2900 K ~ 5800 K in the 30 mm range from the nozzle tip.
Optimal Design for the Rotor Overlap of a Supersonic Impulse Turbine to Improve the Performance
Cho, Jongjae ; Shin, Bong Gun ; Kim, Kuisoon ; Jeong, Eunhwan ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 33~41
DOI : 10.6108/KSPE.2014.18.1.033
In a supersonic turbine, A rotor overlap technique reduced the chance of chocking in the rotor passage, and made the design pressure ratio satisfied. However, the technique also made additional losses, like a pumping loss, expansion loss, etc. Therefore, an approximate optimization technique was appled to find the optimal shape of overlap which maximizes the improvement of the turbine performance. The design variables were shape factors of a rotor overlap. An optimal design for rotor overlap reduces leakage mass flow rate at tip clearance by about 50% and increases about 4% of total-static efficiency compared with the base model. It was found that the most effective design variable is the tip overlap and that the hub overlap size is the lowest.
Test & Evaluation for the Configuration Optimization of Thrust Chamber in 70 N-class N
Thruster (Part I: Pulse-mode Performance According to the Chamber Diameter Variation)
Kim, Jong Hyun ; Jung, Hun ; Kim, Jeong Soo ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 42~49
DOI : 10.6108/KSPE.2014.18.1.042
Performance evaluation was carried out for the 70 N-class hydrazine thruster whose design performance had been already verified. The pulse-mode firing test was conducted for the development model thrusters with various thrust chamber diameters. Evaluation was made by the performance parameters such as specific impulse, impulse bit, and characteristic velocity, etc: specific impulse and characteristic velocity were deteriorated as the thrust chamber diameter deviates from a standard model. Consequently, it is revealed that the performance characteristics of standard model is most superior among the test models.
Test & Evaluation for the Configuration Optimization of Thrust Chamber in 70 N-class N
Thruster (Part II: Pulse-mode Performance According to the Chamber Length Variation)
Jung, Hun ; Kim, Jong Hyun ; Kim, Jeong Soo ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 50~57
DOI : 10.6108/KSPE.2014.18.1.050
A ground hot-firing test (HFT) was conducted to take out the optimal design configurations for the thrust chamber of 70 N-class liquid rocket engine under development. Monopropellant grade (purity:
) hydrazine was adopted as a propellant for the HFT, and three kinds of thrust chambers having characteristic lengths (
) of 2.79, 2.95, and 3.13 m were selected for their performance evaluation. It is revealed through the test and evaluation that the increase of the
leads to a performance degradation in the test condition specified, and pulse response performance of the development model shows superior characteristics to commercialized hydrazine thrusters.
Analysis on the Internal Flow of the Hydraulic Dual Chambers Applying Various Orifice
Cho, Kihong ; Park, Jungho ; Kim, Euiyong ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 58~64
DOI : 10.6108/KSPE.2014.18.1.058
Hydraulic dual chamber, as the simulator for a dual pulse rocket motor, was tested by a high pressure device with various orifice-hole size being applied. Pressure difference occurs between 1st chamber and 2nd chamber depending on area ratio of the orifice to nozzle throat. Studying a design configuration of the orifice is essential to the motor development because pressure difference severely affects the rocket motor performance. It is noticed in this study that energy dissipation is caused by the vortex flow originating from the orifice as the 2nd chamber is operated. The flow field is simulated by a commercial computational fluid dynamics program, ANSYS FLUENT V14.5.
Turbine Rotor-Pyrostarter Coupled Test for the Verification of Thermo-Structural Suitability of a Turbopump Turbine
Jeong, Eunhwan ; Kang, Sang Hun ; Hong, Moongeun ; Lee, Hanggi ; Lee, Soo Yong ; Kim, Jinhan ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 65~72
DOI : 10.6108/KSPE.2014.18.1.065
Turbine rotor-pyrostarter coupled test was performed for the verification of thermo-structural suitability of a turbopump turbine. Newly developed solid propellant and design concept were used in pyrostarter development. In case of turbine rotor, rotor configuration modification and post EDM machining process are adopted in rotor manufacturing respectively for the thermal stress relief and the surface integrity improvement on the blade surfaces. In the test, combustion gas of pyrostarter was directly ejected from the nozzles and impinged on the stationary turbine rotor specimen through the identically shaped flow passage of turbopump. Three kind of thermal load - design to extreme condition - test were performed and no damages were found on the turbine rotor specimens.
Recent Research Works on Chemiluminescence as Measures of Combustion Characteristics
Seo, Seonghyeon ; Moon, Insang ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 73~84
DOI : 10.6108/KSPE.2014.18.1.073
The present paper includes recent research works on the estimation of physical properties like equivalence ratio and heat release rate of flame through chemiluminescence measurement. Modern combustion devices require a precise control to increase combustion stability as well as to suppress pollutant emissions. The determination of combustion characteristics from chemiluminescence provides practical advantages over other techniques. However, the technique is dependent on equivalence ratio, combustion pressure, inlet temperature, turbulent intensity and fuel type. The intensity ratio of
has a strong relation with an equivalence ratio for methane/air premixed flames. The global measurement of chemiluminescence is accepted as a good indicator for a global heat release rate.
Design of Compressed Gas Supply System for Combustion Chamber Test Facility
Chung, Yonggahp ; Cho, Namkyung ; Han, Yeoungmin ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 85~90
DOI : 10.6108/KSPE.2014.18.1.085
To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility (CCTF). The CCTF is the test facility to develop the combustor of rocket engine, which uses liquid oxygen as a oxidizer and kerosene as a fuel. Present paper introduces the detailed design results of compressed gas supply system of CCTF, which is planned to be installed at Naro Space Center.
Brief Summary of KSLV-I Upper Stage Kick Motor Development
Lee, Hanju ; Lee, Jung Ho ; Oh, Seung Hyub ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 91~96
DOI : 10.6108/KSPE.2014.18.1.091
KSLV-I (Korea Space Launch Vehicle-I) upper stage KM (Kick Motor) is a solid propulsion system which consists of igniter, SAD (Safety Arming Device), composite case, and submerged nozzle capable of TVC (Thrust Vector Control) actuation. Each subsystem of KM fulfilled development requirements for achieving a flight mission successfully. We confirmed the successful development of KM from the
flight test results of NARO on January 30, 2013. This article deals with the requirements of KM and the results on configuration management, mass variation, thrust axis alignment, and major test results and so on.
Preliminary Thermal Sizing of Fuel Supply and Cooling System for High-speed Vehicles
Choi, Seyoung ; Park, Sooyong ; Choi, Hyunkyung ; Kim, Joontae ; Jeong, Haeseung ; Park, Jeongbae ;
Journal of the Korean Society of Propulsion Engineers, volume 18, issue 1, 2014, Pages 97~104
DOI : 10.6108/KSPE.2014.18.1.097
In this study, preliminary thermal sizing was performed with the aim of developing a fuel supply and cooling system design to solve the heating problems in high-speed vehicles. First, an analysis model was used to satisfy an optional mission profile. The heat loads were computed under boundary conditions. The results were verified using the precedent design case. Then, fuel consumption rates were estimated for the analysis trajectory. Accordingly, the cooling capacity in the system was calculated using the heat sink capacity of the endothermic fuel. Lastly, the fulfillment of the design requirements was confirmed in comparison to the cooling needs.