• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.5
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• pp.13-20
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• 2020

The effect of the relative distance between two cavity flame holders on the performance of a supersonic combustor was experimentally investigated. A rectangular cross-sectional combustor model with one cavity flame holder on each of two facing walls was used, with two difference distances between cavities of 135 mm and 220 mm. The fuel equivalence ratio was varied as 0.16 and 0.38. A direct-connected type test facility was used to provide Mach 2 flow condition. The test results revealed that the combustion pressure was higher for the shorter cavity distance case. But fuel equivalence ratio did not have large effect on the combustion pressure. It was concluded that, to get higher combustor pressure, there needs to be further combustor configuration change such as smaller cavity distance or tandem cavity installation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.108-111
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• 2017

High performance upper stage engine is necessary for space launch vehicles of geostationary orbit, and staged combustion cycle engine is suitable due to high specific impulse. Technology demonstration model for 9 tonf class staged combustion cycle engine, which is consisted of turbopump, preburner, combustion chamber and supply system, was assembled, and hot-firing test was conducted for three seconds in Upper-stage Engine Test Facility of Naro Space Center. Ignition, combustion and shut down of engine system was performed normally, and its performance parameters were evaluated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2002.04a
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• pp.4-5
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• 2002

로켓 엔진의 개발에 있어 성능과 수명(life time)에 가장 문제가 되는 것은 연소 안정성에 있다. 일반적으로 연소 불안정을 야기시키는 것은 3가지로서 축방향(longitudinal), 반경방향(radial)과 접선방향(longitudinal) 모드(mode)가 있다. 이와 같은 모드를 제어하는 방법에는 수동적 제어방법으로 음향공(acoustic cavity)과 배플(baffle)이 있으며, 음향공은 모드에 관계없이 특정한 주파수에 맞추어 체적을 조절하여 음향파(acoustic wave)를 감쇄시키는 것이고 배플은 주파수에 관계없이 접선방향 모드를 제어하는 것이 기본 목적이나 허브(hub)를 설치하여 반경방향 모드까지 제어할 수 있다. 음향공은 엔진의 성능 또는 연소장에는 영향을 주지 않고 작동하는 반면, 배플은 초기 엔진설계를 할 때 고려하지 않으면 후에 배플을 장착하였을 때는 초기 설계의 제한 때문에 장착의 어려움과 성능 및 연소장에 영향을 미쳐 원하지 않는 엔진의 시험 결과를 야기할 수 있다. 본 연구에서는 KSR-III와 동일한 조건의 연소기에서 다양한 배플을 장착하였을 경우에 대하여 성능과 연소장에 대하여 예측하였다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.4
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• pp.395-402
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• 2010

Baffle injectors are protruded into the combustion chamber and form an anti-pulsating baffle to prevent high-frequency combustion instabilities in transverse modes. Being exposed to a high heat-flux environment, the baffle injector has self-cooling passages through which kerosene is convected and heated. The baffle injector with 20 spiral cooling channels has been developed and successfully applied to 30 $ton_f$-class combustors without any performance loss due to an additional cooling. In this work, numerical analysis of conjugate heat transfer in baffle injectors with various cooling channel designs has been performed in order to reduce the fabrication cost which would be considerably increased for the 75 $ton_f$-class combustor. Prior to the application to a full-scale combustor, the thermal durability of the modified design has been verified through the subscale hot-firing tests.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.636-640
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• 2011

Film cooling technique has been applied to effectively reduce thermal load on liquid rocket combustion chambers by direct injection of a portion of propellant, which flows through the regeneratively cooling channels, into the chamber wall. This study developed a comprehensive model to quantitatively predict the effects of kerosene film cooling on propulsive performance and wall cooling at supercritical pressure conditions, and assessed the predictive capability against hot-firing tests of an actual combustor. The present model is expected to be utilized as a design and analysis tool to meet the conflicting requirements in terms of performance, cooling, pressure loss and weight.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.15-18
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• 2009

Technology demonstration for the development of a 75-tonf liquid rocket engine(LRE) thrust chamber for a space launch vehicle has been started on the basis of the previously acquired 30-tonf LRE technologies. For this purpose, a technology demonstration plan was established upon considering the currently available firing test facility in Korea and performance evaluation firing tests were performed on technology demonstration model thrust chambers under a restricted test condition. This paper describes the plan and current status of technology demonstration for a 75-tonf LRE thrust chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.107-110
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• 2008

The test range for turbopump and gas generator coupled test was determined considering the engine system test area which cover the qualification and development. Based on the test range, we determined the required loss coefficient for the throttle valves and lines.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.189-192
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• 2003

This paper summarizes a satellite program-specific performance requirements and test results for the verification of standard mono-propellant hydrazine thruster (MRE-1) producing 0.95 lbf (4.2 Newtons) nominal steady-state thrust at an inlet pressure of 350 psia (2.41 Mpa). Performance characteristics are shown in terms of thrust behavior at steady state and pulse mode firing. Hot firing test philosophy is briefly introduced, too.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.2
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• pp.144-149
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• 2014

A ground firing test for hot-fire performance evaluation according to the characteristic length($L^*$) variation of thrust chamber was carried out for the hydrazine thruster which may be employed in space launch vehicles. A scrutiny into the performance characteristics of each thruster is made in terms of thrust, specific impulse, response characteristics, and characteristic velocity at steady-state firing mode with propellant inlet pressure of 2.41 MPa (350 psia). Through the test results, it has been verified that performance of characteristic velocity and specific impulse degrades as the characteristic length deviates from that of the standard model. Thus, it is confirmed that the thrust chamber configuration of standard model was suitably designed for the requirement specified.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.2
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• pp.10-18
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• 2010

Scramjet engine is one of the core parts of hypersonic vehicle of next generation and being investigated by many countries. Korea Aerospace Research Institute performed a ground test of the model scramjet engine S1 in 2007. And, S2 model which is improved from S1 model in engine startability and thrust was tested with HIEST (High Enthalpy Shock Tunnel) at Kakuda Space Center of JAXA. Design condition of S2 model was Mach 6.7, however, it was tested at Mach 7.7 as an off-design condition test. As a test result, flow separation was found at the inside of the intake, but the engine showed stable combustion pressure distribution. Furthermore, compared to other test models, S2 model showed a good performance value in thrust and specific impulse.