• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.3
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• pp.95-101
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• 2004

Stability rating of KSR-III rocket engine is conducted based on stability rating tests in the course of development of KSR-III rocket engine. Rocket engine is approved to have combustion stabilization ability when it can suppress the external perturbation or pressure oscillation with finite amplitude and recover the original stable combustion. Rocket engine in flight may be perturbed by unexpectedly large-amplitude pressure oscillation and thus a designer should not only assure combustion stabilization ability of the engine but also quantify the stabilization capacity. For this, principal quantitative parameters and their evaluation are introduced. To verify dynamic stability of KSR-III rocket engine, six stability rating tests have been conducted. Based on these test results, such parameters are quantified and thereby, the stabilization capacity of KSR-III rocket engine is evaluated.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.2
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• pp.10-17
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• 2004

The results of the combustion performance tests of gas generator which supplies hot gas into the turbine of turbo-pump for liquid rocket engine and uses LOx and kerosene as propellant are described. The gas generator consists of a injector head with F-O-F impinging injector, a water cooled combustion chamber, a gas torch igniter, a turbulence ring and an instrument ring. The effect of turbulence ring and combustion chamber length on performance of gas generator are investigated. The ignition and combustion at design point are stable and the pressure and gas temperature at gas generator exit meets the target. The turbulence ring installed at middle of chamber effectively mixes hot gas with cold gas and the effect of residence time of hot gas in gas generator on combustion efficiency is small. Test results show that the main parameter controlling the gas temperature at gas generator exit is overall O/F ratio.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.401-405
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• 2010

An efficient performance model for scramjet engines has been proposed for scramjet performance design. In supersonic air intake design, the compression angles of the wedge were determined to maximize the total pressure recovery of the intake based on Oswatisch criterion. Both combustion models of chemical equilibrium and finite-rate chemistry model are implemented, and compared each model with the results by Starkey for Waverider engine configuration. Finally, the performance model of concern has been confirmed by conducting performance analysis with hypothetical mission profile and design conditions.

• 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.38 no.6
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• pp.596-604
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• 2010

In order to investigate the operating conditions and major design parameters of a dual ramjet propulsion system, an theoretical analysis of ramjet and scramjet propulsion systems was performed. The performance characteristics of each engine are delivered by thermo-dynamical cycle analysis, considering loss effects in a real engine. The performance sensitivity analysis is conducted by investigating various performance parameters, such as an intake efficiency, combustor inlet Mach number, configuration of the combustor, fuel flow rate, and exhaust nozzle efficiency. Based on these analysis results, the processes of application to dual ramjet cycle engines are specified.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.3
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• pp.1-7
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• 2014

A set of preliminary design parameters for the bipropellant rocket engine using liquid methane-fuel as green propellant were derived through a theoretical performance analysis. Chemical equilibrium analysis utilizing CEA was conducted for the prediction of combustion performance: combustion characteristics according to the O/F ratio and chamber pressure variation were investigated. For a determination of chamber-characteristic length, the vaporization time of fuel-droplet with various performance parameters was calculated by applying Spalding's 1-D droplet vaporization model. Finally, the preliminary design specification of methane-bipropellant rocket engine, which is to be performance-tested under the ground firing condition, was proposed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.71-75
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• 2009

Test results of combustion chamber to verify the operation and the combustion performance at low pressure, design and off-design conditions for 30ton-class liquid rocket engine were described. The combustion chamber has nominal chamber pressure of 60 bar, propellant mass flow rate of 89 kg/s, and nozzle expansion of 12. Effects of chamber pressure on combustion characteristic velocity are largely affected by mixture ratio. The specific impulse of combustion chamber is proportional to the chamber pressure regardless of the mixture ratios. The present results can be used as the base to predict the combustion performance of large sized chamber at high pressure while demonstrating the possibility of low pressure firing test of large sized chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.04a
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• pp.26-26
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• 2000

KM(Kick Motor)는 항우연에서 주관하여 개발하고 있는 3단형 과학관측로켓(KSR-III)이다. 본 연구는 이 KM에 적용되는 추진제 개발로서 추진기관에서 요구하는 성능, 밀도, 연소특성, 기계적특성, 점화성, 추진제/라이너/EFDM 접착력을 달성하고, 장기저장시 추진제에 작용하는 온도, 중력등 하중에 대한 추진제의 내구성을 확인하는 수면예측시험을 통해 KM이 사용하는 기간중 요구성능을 발휘할 수 있는 추진제 개발을 목표로 하고 있다. 본 논문에서는 KM 추진제 개발중 1차적으로 추진기관 요구성능을 달성하기 위해서 추진제 성능분석을 통한 기본조성을 설정하고, 이 기간조성을 토대로 밀도, 연소특성, 기계적 특성, 추진제/라이너/EPDM 접착력 실험 결과등을 정리하였고, 추진제 성능을 확인하는데 일반적으로 널리 이용하는 있는 표준모타(ST-8)에 적용하여 이론적 성능분석 및 실제 연소시험을 실시하여 그 결과를 비교 분석하였다.

• Journal of the korean Society of Automotive Engineers
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• v.16 no.4
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• pp.51-61
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• 1994

기관의 출력성능은 기관으로 공급되는 연료공기의 혼합기량에 따라서 크게 달라진다. 이것은 기관의 출력성능은 기관으로 공급되는 흡기 용량에 따라서 변화하기 때문이다. 고출력을 얻기 위하여는 동일한 조건의 경우 흡기량을 증가시켜 기관 실린더 내에서 많은 연소 열에너지를 생성하는 것이 필요하다. 이러한 관점에서 기관의 체적 효율(volumetric efficiency)을 증가시킬 목적으로 여러가지 흡기 계통의 개서을 도모하고 있으나 흡기 용량을 증가시키는 방법의 하나는 과급기(supercharger)를 이용하는 과급 방식이다. 이와같은 과급방식은 기관의 출력성능의 향상을 가져오지만 기관 내부의 노크(knock), 연소 압력 및 열부하의 증가, 연비 문제등에 관한 여러가지 문제점이 제기되고 있다. 여기서는 과급에 적용되는 과급기의 종류와 과급 성능 특성 등에 대하여 살펴보고 과급기관의 성능에 대하여 다루기로 한다.

• Journal of the korean Society of Automotive Engineers
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• v.3 no.2
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• pp.43-49
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• 1981

연료 연소를 수반하는 장치는 공업계에서 널리 사용되고 있으나, 그 설계에 있어서 확립된 이 론은 거의 없고 대체로 과거의 설계에 따라 설계하고 있는 것이 현상이라 하겠다. 변조장치의 형 식이 정해져 있고, 성능도 거의 알려져 있는 경우에는 위와같은 방법도 좋겠지만 예를 들어 새로 운 형식 변조기나 연소실을 계획할 경우에는 실기와 동규모의 실험장치를 몇 종류 제작하여 그 성능을 실험적으로 비교 검토하여 좋은 것을 찾는 방밥을 취하게 된다. 이러한 방법은 많은 비용이 소요되며 기종이 바뀌면 새로위와 같은 일을 반복하여야 한다. 이와같은 불합리를 없애기 위하여 많은 연구자는 이론적으로 또 실험적으로 연구를 계속해왔다. 그리하여 현재 많은 사 실을 구명하였으나 아직 구체적 이론을 확립하기까지는 이르지못한 실정이다. 이것은 연소란 현상이 아주 복잡하고 또 관계된 인자가 많은 까닭이다. 본 강좌에서는 열체분우연소의 기본을 이루는 액적의 증발과 연소 문제에 대해서 현상적으로 해설하여 기초적 해석에 도움이 되기를 바라는 바이다.