• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.87-91
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• 2004

It is definitely required to control dispersion of the rocket engine performance in order to accomplish the mission of launch vehicle successfully. We performed the dispersion analysis of gas generator cycle LRE (liquid rocket engine) accompanied with ANASYN. As a result, the vacuum thrust dispersion of the engine was $+5.34\%,\;-5.27\%$ and the mixture ratio deviated $+9.07\%,\;-9.82\%$ from the nominal value due to the errors of components and engine inlet condition of propellants. By applying the gas generator regulator only, the dispersion of the engine performance increases. Error in turbine efficiency is the most influential factor to the dispersion of engine performance.

• Journal of the KSME
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• v.56 no.9
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• pp.44-48
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• 2016

실제 크기의 로켓엔진은 수많은 예산이 소요되기 때문에 새롭게 설계된 엔진에 대하여 비행모델의 제작 및 실험연구는 예산과 위험성 때문에 매우 제한적으로 시행된다. 따라서 비행모델의 설계를 확정하기 위한 개발단계에서는 실제 크기의 로켓엔진의 특성을 나타낼 수 있는 소형 액체로켓을 제작하여 시험평가를 수행하게 된다. 이 글에서는 이러한 축소형 연소기 관련 연구동향을 알아보고자 한다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.10a
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• pp.1-1
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• 1998

로켓엔진 추력발생용으로 광범위하게 사용되는 액체추진제는 고성능, 대용량의 액체추진제 로켓엔진에서는 필연적으로 고주파 연소불안정의 문제를 수반하며, 이 연소불안정의 정도는 연소성능과 더불어 엔진개발의 성패를 좌우하는 중요한 여건이 된다. 따라서 안정한 로켓의 비행을 보장하기 위해서는 연소불안정의 문제가 선결되어야 한다. 연소불안정의 기본 메커니즘은 연소기에서 발생하는 압력섭동에 반응하여 불안정한 음향에너지를 되먹임하는 연소과정으로 설명된다. 연소불안정 현상이 발견된 이후 실험 및 이론적 접근에 의해 이와 같은 연소불안정 메커니즘 및 예측에 대한 체계적인 연구가 이루어져 왔으며, 현재까지의 다양한 고주파 연소불안정 예측방법 중에서 음향 및 기화 응답함수를 이용하는 방법은 직관적 고찰에 의존하는 단순한 연소모델을 적용하며 주로 음향적 섭동에 의한 연소의 반응을 연소안정성 평가의 파라메터로 사용한다. 이와 같은 음향적인 예측방법은 연소불안정 현상을 이론적으로 전개하므로 경제적으로 각종 설계변수에 대한 연소불안정의 변화를 구분할 수 있는 장점이 있어 성능 및 호환설계와 병행하여 로켓엔진 연소실의 초기 안정성 설계방법으로 주로 사용된다.

• Journal of Aerospace System Engineering
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• v.14 no.5
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• pp.91-99
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• 2020

A liquid rocket engine goes through many tests to prove its performance before liftoff. It means the tests for setting ignition and start-up conditions or a test on design condition, which verifies the design performance. However, the development process requires verification of performance under off-design conditions through tests involving different operating conditions, which affects the duration of engine development. The off-design performance test is performed by altering the conditions of the propellant supplied to the engine in conjunction with the engine performance test that varies the opening of the control valves in the engine. This paper is based on the results of the engine tests performed at the KSLV-II engine test facilities in the Naro Space Center and describes the operations of the test facility for off-design condition test that changes the inlet conditions of the turbo-pump due to changes in the pressure and temperature of the propellant supplied to the test engines.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.5
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• pp.37-50
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• 2007

This study investigated the characteristics of spray generated by a liquid coaxial swirl injector used in a combustor of the liquid rocket engine. The linear stability analysis considered long and short wave was introduced in liquid sheet breakup. Through the hydrodynamic analysis the initial liquid sheet thickness spray angle and injection velocity were predicted. To evaluate the effect of turbulence model standard $k-{\varepsilon}$ and RNC $k-{\varepsilon}$ model were applied to numerical calculation and it was known that RNC $k-{\varepsilon}$ model was more applicable to predict spray characteristics. On the basis of this evaluation validation of the developed model was performed with swirl injector installed in LPRE and the predicted results of breakup length, spray angle, and SMD agreed well with experiments qualitatively and quantitatively.

• Aerospace Engineering and Technology
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• v.4 no.2
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• pp.135-141
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• 2005

The procedure of conceptual and detailed design of sub-scale combustor using bipropellant swirl or impinging injector with external or internal mixing for a liquid rocket engine are described in this paper. The sub-scale combustor uses liquid oxygen(LOx) and kerosene as propellants and has a injector head, an ablative material combustor wall and a water cooled nozzle. The injector head has LOx manifold, fuel manifold, fire face plate, one center swirl or impinging injector and 18 main swirl or impinging injectors.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.139-143
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• 2005

Combustion Test Facility for Liquid Rocket Engine using kerosene and liquid oxygen has been developed for the purpose of cooling and performance study. Test engine of thrust 0.5 KN(design thrust) is tested to confirm the normal operation. Therefore, water-cooled firing tests using kerosene engine with injectors of fuel-centered coaxial type are conducted. With the viewpoint of characteristic velocity, and specific impulse at sea level, and chamber pressure on OF mixture ratio are analyzed.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.8
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• pp.69-76
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• 2003

For stable combustion and safety of a structure of the propulsion system, a cooling system to the liquid rocket engine should be incorporated. In this study, Eulerian-Lagrangian scheme for two phase combustion, nongray radiation and soot formation effect, and film-wall interaction have been introduced to study the effect of film cooling. After briefly introducing the governing equation, combustion characteristics with change of wall temperature has been investigated by varying such parameters as fuel mass fraction for film cooling, diameter of the fuel droplet, overall mixture fraction of oxygen to fuel. Also, radiative heat flux is compared with the conductive one at the combustor wall.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.341-345
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• 2005

This paper is for the combustion characteristics of gas generator which drive 1.5MW-class turbo pump and runs in fuel-rich combustion regime with LOx/kerosene as propellant. The outline of development procedure of real scale high pressure gas generator is introduced and the relation between O/F ratio and outlet temperature and the molecular weight and specific heat ratio of combustion gas are described. The relation between O/F ratio and temperature is newly obtained at higher pressure and the molecular weight and specific heat ratio is modified and their validity is confirmed by the mass relation equation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.213-219
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• 2005

Small liquid rocket engine (SLRE) with calorimeter were developed and tested to evaluate cooling characteristics in the liquid rocket engine. Therefore, cooling performance analysis was performed to predict the heat transfer coefficient on gas side wall in 10 calorimeter channel. A heat transfer empirical formula was determined by results of firing test and computational simulation.