Go to the main menu
Skip to content
Go to bottom
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 15, Issue 6 - Dec 2011
Volume 15, Issue 5 - Oct 2011
Volume 15, Issue 4 - Aug 2011
Volume 15, Issue 3 - Jun 2011
Volume 15, Issue 2 - Apr 2011
Volume 15, Issue 1 - Feb 2011
Selecting the target year
Effect of Scale and Geometry on the Performance of Heat-Recirculating Swiss roll combustors for Micro Power Generation Applications
Oh, Hwa-Young ; Huh, Hwan-Il ; Ronney, Paul D. ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 1~10
Combustion and extinction limits in heat-recirculating excess enthalpy reactors employing both gas-phase and catalytic reaction have been examined with an emphasis Reynolds number (Re) effects and possible application to microscale combustion devices. In this paper, geometrically similar reactors of different physical sizes and different numbers of turns were tested with the aim of estimating for combustor characteristics. Combustion efficiency is estimated by measuring exhausted gases through the gas chromatograph. From these results the effect of scale and number of turns are demonstrated and optimal operating conditions for Swiss roll combustors are identified.
Effect of Scale and Fuel Type on Heat-recirculating Swiss-roll Combustor Performance for Fuel Cell Reformer Applications
Kim, Youn-Ho ; Huh, Hwan-Il ; Ronney, Paul D. ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 11~18
The geometrically similar swiss roll reactors of different physical sizes were tested with the aim of independently determining the effects of Re and Da. It is found that the difference between catalytic and non-catalytic combustions extinction limits are narrowed as scale decreases. In addition to assess the importance of fuel chemistry, different families of fuels including alkanes and ethers were tested. From these results the effect of scale and fuel type on microscale reactor performance and implications for practical micro combustion devices are discussed.
Performance Characteristics of Velocity Compound Supersonic Impulse Turbine with the Rotor Overlaps
Cho, Jong-Jae ; Kim, Kui-Soon ; Jeong, Eun-Hwan ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 19~28
As a preview study, present research analysed the performance characteristics of a velocity compound supersonic impulse turbine with the rotor overlaps before adapting the overlap has the best turbine performance. This research was conducted for the turbine with square cross-section nozzles instead of axisymmetric nozzles and wrap around nozzles. Through 3-dimensional flow analysis for the turbine by a commercial flow analysis package, tip overlap case was more effective to improve the turbine performance than case hub overlap, and overlap case applied the hub and tip of the rotor had the largest improvement for the turbine performance in the cases. In case of overlap for the 2nd stage rotor, improvement of the turbine performance was not visibly large. Because, generated power in the 2nd stage is 22~23% of whole generated turbine power.
Impact Sensitivity of HTPE & HTPB Propellants using Friability Test
Kim, Chang-Kee ; Yoo, Ji-Chang ; Min, Byoung-Sun ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 29~34
Hydroxyl terminated polyether(HTPE) propellants have been developed recently as possible replacements for HTPB/AP propellants currently used in a number of tactical rocker motor. As analyzing friability of HTPE and HTPB propellants in this study, the following results could be derived. The friability of the tested propellants depended on its binder contents, mechanical property, and burning rate. It was decreased as burning rate was lowered and toughness was increased.
Modeling of Liquid Rocket Engine Components Dynamics at Transient Operation
Kim, Hyung-Min ; Lee, Kuk-Jin ; Yoon, Woong-Sup ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 35~44
Mathematical modelling for liquid rocket engine(LRE) main components were conducted to predict the dynamic characteristics when the LRE operates at the transient condition, which include engine start up, shut down, or thrust control. Propellant feeding system is composed of fuel and oxidizer feeding components except for regenerative cooling channel for the fuel circuit. Components modeling of pump, pipe, orifice, control valve, regenerative cooling channel and injector was serially made. Hydraulic tests of scale down component were made in order to validate modelling components. The mathematical models of engine components were integrated into LRE transient simulation program in concomitant with experimental validation.
Modeling of the Liquid Rocket Engine Transients
Ko, Tae-Ho ; Jeong, Yu-Shin ; Yoon, Woong-Sup ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 45~54
A program aiming at predicting dynamic characteristics of a Liquid Rocket Engine(LRE) was developed and examined to trace entire LRE operation. In the startup period, transient characteristics of the propellant flows were predicted and validated with hydraulic tests data. An arrangement of each component for the pipelines was based on an operating circuit of open cycle LRE. The flow rate ratio for the gas generator and the main chamber was determined to mimic that of real open cycle LRE. Individual component modeling at its transient was completed and was integrated into the system prediction program. Essential parameters of the component dynamic characteristics were examined in an integrated fashion.
Study of Thrust Control Performance Improvement for Hybrid Rocket Applications
Choi, Jae-Sung ; Kang, Wan-Kyu ; Huh, Hwan-Il ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 55~62
In this study, we tried to improve the thrust control performance through the thrust control combustion experiment of the hybrid rocket. We constructed the system which controls the oxidizer flow by combining a needle valve with a stepping motor and controlling the stepping motor drive according to the thrust control command order. Gas oxygen was used as the oxidizer for two different propellants, PE(Polyethylene), PC(Polycarbonate), respectively. To improve the slow response time and the oscillation phenomenon in the beginning stage of the thrust control combustion experiment, we measured and analyzed the change of the flow speed of the propellant pipe. The revised thrust control combustion experiment showed that the thrust was stably controlled with the margin or error from the thrust command within
Heating Apparatus Development and Tests for Cryogenic Gaseous Helium
Chung, Yong-Gahp ; Cho, Nam-Kyung ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 63~68
For the liquid rocket propulsion system using liquid oxygen as oxidizer, helium for pressurizing LOX is usually stored in the LOX tank with cryogenic temperature. For that kind of pressurizing system, cryogenic helium is discharged from the immerged pressurant cylinder and passes through the heat exchanger downstream of gas generator. During the process, helium pressurant is heated from cryogenic temperature to high one and supplied to the ullage of propellant tank. To develop the pressurizing system, a cryogenic heating apparatus is needed to simulate the heat exchanger. In this paper, the cryogenic heating apparatus for development of the pressurization system is presented along with its heating test results with cryogenic helium.
Low Pressure Firing Tests of 75-tonf-Class Channel Cooling Thrust Chamber
Lim, Byoung-Jik ; Han, Yeoung-Min ; Kim, Jong-Gyu ; Choi, Hwan-Seok ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 69~75
Firing tests have been carried out for a technology demonstration model of 75-tonf-class combustor which is to be used on the liquid rocket engine of a Korean space launch vehicle. Firing tests were done at 50% of the nominal flow rate because of incapability of the test facility and limit of the test bed strength. Through the low pressure firing tests of 75-tonf-class channel cooling thrust chamber, operability and stability at the ignition and combustion phases were confirmed. Additionally it was foreseen that the 75-tonf-class thrust chamber would satisfy the performance requirements.
A Study on the Thrust Axis Alignment of Kick Motor for KSLV-I
Jung, Dong-Ho ; Lee, Han-Ju ; Oh, Seung-Hyub ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 76~82
The thrust axis alignment of the launch vehicle is very important because the misalignment causes the unstable attitude control and results in mission failure. Generally, optical methods such as digital theodolite and laser tracker and mechanical method such as turn table method are used to align the thrust axis. This article deals with the simple method using inclinometer based on the gravitational direction. The inclinometer indicates zero degree when that is located on the perpendicular plate to gravitational direction. This method needs two inclinometer, such as standard and alignment ones and uses the angle difference as the reference data to adjust the TVC actuator offset.
Preliminary Design Plan for Determining Combustor Configuration of Regenerative-cooled Liquid Rocket Engine
Son, Min ; Seo, Min-Kyo ; Koo, Ja-Ye ; Cho, Won-Kook ; Seol, Woo-Seok ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 83~89
A design plan was proposed for determining combustor configuration of regenerative- cooled liquid rocket engine in the process of preliminary design. Rocket performance and regenerative cooling results were calculated using the properties of combustion gas estimated in CEA. For required thrust, chamber pressure, atmosphere pressure and propellant mixture ratio the mass flow rate of propellants and combustor performance were predicted by one-dimensional and experimental correlations. Finally, determinable plan for the contour of combustor were presented through Rao nozzle design method.
Overview of Propulsion System Performance for Lunar Orbiter and Recent Development Status
Lee, Kyun-Ho ;
Journal of the Korean Society of Propulsion Engineers, volume 15, issue 1, 2011, Pages 90~101
From 1990s, the lunar exploration programs, suspended over 20 years after the project Apollo's first successful human landing on the Moon in 1969, have been restarted according to a revived interest in Moon. In recent, several nations progress their own lunar exploration program successfully. In this report, to investigate the technical trends of the onboard propulsion system for the lunar orbiter, technical features related to the performance of the propulsion system of the lunar orbiters developed since 1990 are surveyed. In the future, it is expected that this technical report can provide a fundamental guideline for selecting a proper type of the onboard propulsion system for the domestic lunar orbiter.