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REFERENCE LINKING PLATFORM OF KOREA S&T JOURNALS
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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 17, Issue 6 - Dec 2013
Volume 17, Issue 5 - Oct 2013
Volume 17, Issue 4 - Aug 2013
Volume 17, Issue 3 - Jun 2013
Volume 17, Issue 2 - Apr 2013
Volume 17, Issue 1 - Feb 2013
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Requirement Analysis of Propulsion System for Active Anti-Ship Missile Decoy
Moon, Yongjun ; Kwon, Sejin ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 1~9
DOI : 10.6108/KSPE.2013.17.4.001
An active anti-ship missile decoy system was designed conceptually to analyze propulsion system requirements and feasibility to use a liquid bi-propellant rocket engine. Overall mass, size, and shape were assumed referring to specifications of Nulka which was developed by US and Australia in 1990s. The propulsion system was assumed to be a 1,000 N-class
/kerosene rocket engine with a pressurized feed system. A three-degree-of-freedom optimal trajectory was calculated based on the assumptions, and mass budget was designed from the calculation results. It was found that the requirements for the propulsion system is that it shall be operated more than 100 sec; it shall be re-ignitable; it shall have a throttle capability of a range from 35% to 100% when the maximum thrust at sea level is 1,000 N.
Optimization of Turbofan Engine Design Point by using Seven Level Orthogonal Array
Kim, Myungho ; Kim, Youil ; Lee, Kwangki ; Hwang, Kiyoung ; Min, Seongki ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 10~15
DOI : 10.6108/KSPE.2013.17.4.010
For design optimization, engineers should require the accurate information of design space and then explore the design space and carry out optimization. Recently, the total design framework, based on design of experiments and optimization, is widely used in industry areas to explore the design space above all. For optimizing turbofan engine design point, the response surface model is constructed by using the 7 level orthogonal array which satisfies the statistical uniformity and orthogonality and gets the dense design space information. The multi-objective genetic algorithm is used to find the optimal solution within the given constraints for finding global optimal one in response surface model. The optimal solution from response surface model is verified with GasTurb simulation result.
Analysis of Elements Influencing on Performance of Interior Ballistics
Sung, Hyung-Gun ; Yoo, Seung-Young ; Lee, Sang-Bok ; Choi, Dong-Whan ; Roh, Tae-Seong ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 16~24
DOI : 10.6108/KSPE.2013.17.4.016
The analysis of performance and internal flow according to various numerical models for interior ballistics has been conducted. The initial flow has been mainly affected by the drag model of propellants and their drag degradation reduces oscillations of differential pressure between the breech and the shot base. Models of Nusselt number haven`t influenced the major performance of interior ballistics. The negative differential pressure isn`t generated in the case without the heat transfer of propellants.
Effect of Shroud Split on the Performance of a Turbopump Turbine Rotor
Lee, Hanggi ; Jeong, Eunhwan ; Park, Pyungoo ; Yoon, Sukhwan ; Kim, Jinhan ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 25~31
DOI : 10.6108/KSPE.2013.17.4.025
A blisk with rotor shroud is usually adopted in LRE turbine to maximize its performance. However it experiences the severe thermal load and resulting damage during engine stating and stop. Shroud splitting is devised to relieve the thermal stress on the turbine rotor. Structural analysis confirmed the reduction of plastic strain at the blade hub and tip. However, split gap at the rotor shroud entails additional tip leakage and results performance degradation. In order to assess the effect of shroud split on the turbine performance, tests have been performed for various settings of shroud split. For the maximum number of shroud splitting, measured efficiency reduction ratio was 2.65% to the value of original shape rotor.
Study of Supersonic Flame Acceleration within AN-based High Explosive Containing Various Gap Materials
Lee, Jinwook ; Yoh, Jai-Ick ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 32~42
DOI : 10.6108/KSPE.2013.17.4.032
We study the gap effect on detonating high explosives using numerical simulation. The characteristic acoustic impedance theory is applied to understand the reflection and transmission phenomena associated with gap test of high explosives and solid propellants. A block of charge with embedded multiple gaps is detonated at one end to understand the ensuing detonation propagation through pores and non uniformity of the tested material. A high-order multimaterial simulation provides a meaningful insight into how material interface dynamics affect the ignition response of energetic materials under a shock loading.
Numerical Study on the Adverse Pressure Gradient in Supersonic Diffuser
Kim, Jong Rok ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 43~48
DOI : 10.6108/KSPE.2013.17.4.043
A study is analyzed on the adverse pressure gradient and the transient regime of supersonic diffuser with Computational Fluid Dynamic. The flow field of supersonic diffuser is calculated using Axisymmetric two-dimensional Navier-Stokes equation with
turbulence model. The transient simulation is compared in terms of mach number and static temperature of vacuum chamber according to pressure variation of rocket engine combustion chamber. Combustion gas flow into the vacuum chamber during operation of the supersonic diffuser. According to this phenomenon, the pressure and the temperature rise in the vacuum chamber were observed. Thus, the protection system will be necessary to prevent the pressure and temperature rise in the transition process during operation of the subsonic diffuser.
Flight Test of Hybrid Propulsion System for Electrically Powered UAV
Park, Poomin ; Kim, Keunbae ; Cha, Bongjun ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 49~55
DOI : 10.6108/KSPE.2013.17.4.049
This paper deals with the flight test of propulsion system of middle size electrically powered UAV (EAV2, Electric Aerial Vehicle 2) which is under development in KARI. EAV2 is low speed endurance type UAV whose wing span is 6.9 m, and weight is 18 kg. The UAV has flown for 22 hours in June of 2012. The flight test result showed that the propulsion system worked well suppling power for any circumstances during the test flight. Each power source worked according to the design purpose.
Development of System Analysis Program of Liquid Rocket Engine I
Lee, Sang-Bok ; Son, Min ; Seo, Jongcheol ; Lim, Taekyu ; Roh, Tae-Seong ; Koo, Jaye ; Kim, Kuisoon ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 56~62
DOI : 10.6108/KSPE.2013.17.4.056
The system analysis and design program of the liquid rocket engine has been developed for preliminary conceptual design process. The program analyzes the engine system and obtains optimal design variables by optimization methods such as genetic algorithm for the higher specific impulse and thrust to weight ratio using given input parameters and requirements. For the users` convenience, the GUI has been offered. The 3-dimensional model for the visualization of results has been interconnected with the CATIA program. The results are expected to be applied to the design process of the space launch vehicle for the analysis and selection of the propulsion system.
A Starting Characteristics Study of the Scramjet Engine Test Facility with a Mach 5.0 Nozzle
Lee, Yang-Ji ; Yang, In-Young ; Yang, Soo-Seok ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 63~72
DOI : 10.6108/KSPE.2013.17.4.063
A Mach 5 nozzle and a diffuser of the Scramjet Engine Test Facility (SETF) were made for a hydrocarbon-fueled scramjet engine. SETF, attached with a diffuser guide, started with a model of 60% blockage, though the model engine could not start by over expansion of the facility nozzle. The model was moved into the nozzle to escape the shock generated from the nozzle exit, both SETF and the engine could start. The pitot rake experiments (blockage of 2.3%) were done for measuring the core flow in the test section. From the pitot experiments, the core flow was expanded by an under expansion. It means that the core flow in the test section was related with a model blockage. SETF and the engine with a blockage of 33% work normally. From a series of experiments, SETF started with a normal shock efficiency of 58%, regardless of a blockage ratio.
Research Activities about Characteristics of Fuel Injection and Combustion Using Endothermic Fuel
Choi, Hojin ; Lee, Hyungju ; Hwang, Kiyoung ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 73~80
DOI : 10.6108/KSPE.2013.17.4.073
Endothermic fuel utilizing technology is considered as a unique practical method of hypersonic vehicle for long distance flight. Research activities about characteristics of fuel injection and combustion using cracked by endothermic reaction are reviewed. Studies on characterization of supercritical fuel injection and mixing within supersonic flow field are surveyed. Researches on combustion characteristics such as ignition delay time, laminar burning velocity and combustion efficiency at supersonic model combustor are reviewed. In addition, domestic research activities on endothermic fuel are surveyed.
Design Point Operating Characteristics of an Oxidizer Rich Preburner
Moon, Ilyoon ; Moon, Insang ; Kang, Sang Hun ; Ha, Seong-Up ; Lee, Soo Young ;
Journal of the Korean Society of Propulsion Engineers, volume 17, issue 4, 2013, Pages 81~88
DOI : 10.6108/KSPE.2013.17.4.081
It was designed and tested at the design point that an oxidizer rich preburner for a staged combustion liquid rocket engine propelled by kerosene and LOx. The oxidizer rich preburner was designed as some of LOx injected from the mixing head was burned with kerosene and the rest of LOx injected from injection holes in the regenerative cooling chamber was vaporized by combustion gas. The preburner is operated at OF ratio of 60 and combustion pressure of 20 MPa. The Preburner has a honey-comb type mixing head with simplex swirl injectors, a turbulence ring improving combustion stability and uniformity of product gas temperature distribution, and a nozzle simulating the duct. With the combustion test results at the design point, the oxidizer rich preburner showed high combustion stability and uniformity of product gas temperature distribution.