• Journal of Aerospace System Engineering
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• v.14 no.5
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• pp.73-82
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• 2020

Recently, there has been an increase in additive manufacturing for the fabrication of liquid rocket engines. This technology can innovate conventional fabrication methods to reduce the lead time and manufacturing cost and can enhance the performances such as weight reduction. In this study, a literature survey is presented that includes types, advantages, disadvantages, and foreign government-based projects of the technology related to liquid rocket engine manufacturing. The present survey focuses on the technology that has been applied to various components such as turbopumps and valves while much larger efforts are made for combustion chambers with regenerative cooling channels and diverging nozzles, as the advantages of the technology are maximized for the applications.

• Proceedings of the KSME Conference
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• 2007.05a
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• pp.1143-1148
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• 2007

LRE(Liquid propellant Rocket Engine) is one of the important parts to control the motion of rocket. For operation of rocket in error boundary of the set-up trajectory, it is necessarily to control the thrust of LRE according to the required thrust profile and control the mixture ratio of propellants fed into combustor for the constant mixture ratio. It is not easy to control thrust and mixture ratio of propellants since there are co-interferences among the components of LRE. In this study, the dynamic model of LRE was constructed and the dynamic characteristics were analyzed with control system as PID control and PID+Q-ILC(Iterative Learning Control with Quadratic Criterion) control. From the analysis, it could be observed that PID+Q-ILC control logic is more useful than standard PID control system for control of LRE.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.3
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• pp.109-118
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• 2018

Korea Aerospace Research Institute has developed a staged combustion cycle rocket (SCCR) engine with high specific impulse to send a 3-ton class satellite into geostationary orbit while conducting a Korean Space Launch Vehicle (KSLV) II project. The SCCR engine is different from the KSLV-II engine, which is an open cycle engine using a gas-generator. The SCCR engine with a closed cycle engine is composed of a pre-burner, a turbo pump, and a main combustor. The technology demonstration model (TDM0) was assembled and tested in the 7ton-class engine combustion test facility of Naro Space Center, and the combustion test was successfully conducted.

• Aerospace Engineering and Technology
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• v.13 no.1
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• pp.70-75
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• 2014

Programs of energy balance, mode analysis and transient analysis for a liquid rocket engine have been introduced. The analysis methods have been verified through comparison between the present results, and the results of the other program and experimental data. An energy balance analysis is used for engine system design at the early development phase. A mode analysis is used for decision of engine operation conditions and test conditions, and studying deviation of an engine performance. A transient analysis can predict a propellant flow rate, thrust, impulse at transient phase. It is essential to establish a startup/shut down sequence. The analysis programs will be used to develop the engines of KSLV-II.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.3
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• pp.19-27
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• 2002

The sequence of the propellant supply for ignition of a liquid rocket engine combustor is very important in the reliable and safe operation of the engine. The ignition sequence of KSR-III main engine was briefly described and the measuring parameters and their reliability determining ignition sequence were examined in this paper. The filling time of the engine propellant manifolds and the valve open/close time were reviewed to obtain the exact and reliable time of the propellant supply to the combustor. The combustion characteristics of the engine at starting were discussed at different supply lead of propellant. Finally, the hot firing test results with cyclogram determined by measuring parameters were presented.

• Aerospace Engineering and Technology
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• v.10 no.1
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• pp.63-69
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• 2011

In this study, we developed a dispersion analysis program of the gas-generator cycle liquid propellant rocket engine by expanding the mode analysis software(GEMAT). The performance dispersions of an engine that are arisen from the internal dispersion factors of engine's sub-components were formulated and solved to find the effects of each dispersion factor. We were also able to present the calculation method to find the required pressure margin for the compensation of those dispersion to satisfy the required performances of engine. Using this method, we could propose a novel procedure of compensating during the ground firing test which would induce the performance improvement by lessening the pumps discharge pressures or augmenting the combustion chamber pressure.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.47 no.7
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• pp.517-524
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• 2019

A system design has been performed for a vacuum thrust 88 ton staged combustion cycle rocket engine. Previous research has been used to estimate the performance of the engine components. And the algorithm has been proposed to evaluate the converged engine system performance. The present methodolgy has been verified by comparing the published data for RD-180. The present work adopts the most of the previous KSLV-II engine heritage for both performance improvement and cost competitiveness. The combustion pressure has been decided as 12MPa considering manufacturing difficulty, cost and performance improvement, and as a result the vacuum specific impulse has increased by 23.4s.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.10
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• pp.93-101
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• 2004

A method to calculate stability limits is investigated to predict the characteristics of high-frequency combustion instability in liquid-propellant rocket engine. It is based on the theory of linear stability analysis proposed in previous works and useful to predict combustion stability at the beginning stage of engine development. The system of equations governing reactive flow in combustor has the simplified and linearized forms. The overall equation expressing stability limits is adopted. The procedures to evaluate quantitatively each term included in the equation are proposed. The thermo-chemical properties and flow variables required in the evaluation can be obtained from calculation of thermodynamic equilibrium, CFD results, and experimental test data. Based on the existent data, stability limits are calculated with actual rocket engine (KSR-III rocket engine). The present calculations show the reasonable stability limits in a quantitative manner and the stability characteristics of the engine are discussed. The prediction from linear stability analysis could be serve as the first approximation to the true prediction.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.7
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• pp.576-584
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• 2016

In this study, AP added propellant has been proposed as a method of enhancing the low specific impulse performance found for staged hybrid rocket engine. Experimental tests were carried out to analyze and evaluate the effect of AP added propellant on specific impulse performance as well as fuel-rich combustion characteristics in a staged hybrid rocket engine. Upper limit of AP content in propellant was set to be 15 wt% to maintain the hybrid rocket engine advantages. As a result, 15 wt% AP added propellant showed 3% higher specific impulse performance compared to 0 wt% AP added propellant. Moreover, AP addition proved to offer less injected oxidizer mass flow, less O/F variation, and less combustion pressure while producing fuel-rich gas of the same combustion temperature. Future studies will carry out more combustion tests with metal additives to further enhance specific impulse.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.6
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• pp.120-130
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• 2013

A study was conducted for the performance characteristics of methane taking recently the limelight in the world as a next-generation propellant, with the survey for state of the art in the development of methane/oxygen rocket engine being accompanied. Liquid methane as a rocket fuel has the favorable characteristics such as non-toxic, low cost, regenerative cooling capability, and potential for in-situ resource utilization (ISRU). The combination of liquid methane and liquid oxygen also provides the excellent performance including high specific impulse and low system mass. For these reasons, many researches have been actively carried out on the methane/oxygen engine, nevertheless, its technology readiness level is not that high enough just yet. Therefore, it is judged that it is the time to mitigate the technical gap with the space technology of advanced countries through a swift onset of the development of methane rocket engine.