• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.3
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• pp.41-51
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• 2001

Combustion characteristics of a KSR-III liquid rocket engine with split-triplet (F-O-O-F) type injector elements are investigated numerically from the viewpoints of engine performance and combustion flowfield. To evaluate numerical analysis of liquid rocket engine with radial type injector arrangement, 2-D axisymmetric and 3-D calculations are carried out and the prediction of engine performance for design and off-design conditions is in a good agreement with hot-firing tests. According to 2-D axisymmetric and 3-D calculations, the prediction error is 3∼5 % from the standpoint of performance. Numerical results of combustion characteristics calculated through 3-D analysis agree well with hot-firing tests qualitatively at injector plate. Decreasing impinging angle and changing radial type injector arrangement to H type injector arrangement reduce effectively local high-temperature region. Also, it is examined that those affect the performance seriously. In conclusion, it is revealed that both injector arrangement and impinging angle are critical parameters to affect the performance and combustion characteristics of the liquid rocket engine.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.10
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• pp.102-111
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• 2004

The procedure of design and manufacture of sub-scale combustor using bipropellant swirl injector with external mixing for a liquid rocket engine are described. The results of cold flow test, ignition test and combustion test of the sub-scale combustor are also given 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 injector and 18 main swirl injectors. The cold flow, ignition and combustion tests were successfully performed without damage of combustor. Results of hot firing tests show that combustion efficiency meets the target of design and operations of start and stop cyclogram are stable and high frequency combustion instability does not occur.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.299-304
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• 2005

The design procedures of full-scale combustion chamber with chamber pressure of 53bara, mass flow rate of 90kg/s, combustion efficiency of $94\%$ and specific impulse at ground of 253sec were described. The details of combustion performance and geometrical parameters were also given. Full-scale combustion chamber consists of the combustor head with injector/baffle and the chamber/nozzle with regenerative cooling channels. The design results of combustion chamber with ablative materials, detachable injector head with SUS baffle or baffle injector and chamber body for ground hot firing tests were given in this paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.5
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• pp.22-30
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• 2015

A gas generator which generates turbine driving gas by burning a part of propellants is used in an open cycle liquid rocket engine and as a main component of an open cycle liquid rocket engine autonomous hot firing tests are required to investigate the combustion performance and characteristics of the gas generator. However, since the combustion gas generated by a gas generator is choked at the turbine nozzle in the turbine manifold, it is necessary to consider the internal volume of turbine manifold as well as that of the gas generator for correct investigation of the combustion performance, characteristics, and acoustic characteristics of the gas generator. Therefore, in the paper hot firing test results of a gas generator with a turbine manifold simulator are described and characteristic prediction using the autonomous test of a gas generator is explained.

• 한국연소학회:학술대회논문집
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• 2003.12a
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• pp.125-130
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• 2003

본 논문에서는 액체로켓엔진에서 터보펌프의 160kW급 터빈을 구동하고, 액체산소와 케로신을 추진제로 사용하는 가스발생기의 설계점 연소성능시험 결과에 대해 논의하였다. 충돌형 F-O-F 인젝터, 물냉각 채널을 가진 연소실, torch ignitor, turbulence ring 그리고 측정 링을 갖는 가스발생기에 대해 기술하였고, 점화, 연소, 종료 등의 시험 cyclogram에 대해 언급하였다. 설계점에서의 연소시험 및 turbulence ring 장착여부, 연소실 길이 변화에 따른 연소시험의 결과들에 대해 기술하였다. 연소시험 결과 가스발생기는 설계점에서 안정된 작동성을 보여주었고, 연소압력 및 온도 등의 성능이 예측치에 근접하는 결과를 보여 주었다. Turbulence ring은 출구에서의 가스온도를 균일하게 분포시켜 효과적인 혼합 장치임을 보여 주었고, 4-6msec 정도에서의 잔류시간에서는 연소효율의 차이가 크지 않음을 알 수 있었다. 가스발생기 출구에서의 온도는 공급되는 추진제의 O/F ratio에 따라 매우 민감하게 반응함을 알 수 있었다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.7
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• pp.593-600
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• 2015

In this study, a combined hybrid rocket system is newly introduced which has characteristics of both gas generators and afterburner type hybrid rockets. In particular, a combined gas generator utilizing solid fuel and liquid/gas oxidizer was designed as a primary combustor of the system. Combustion tests were carried out with various equivalence ratio affected by parameters such as fuel length, oxidizer flow rate, fuel port diameter and fuel type. In general, fuel-rich gas generator produces low combustion gas temperature to meet the temperature requirement and the target temperature was transiently set less than 1600 K. Since it was found that controlling parameters showed limited effects on the change of equivalence ratio, mixture of $O_2$ and $N_2$ as an oxidizer was additionally introduced. As a result, a combined gas generator successfully produced combustion gas temperature of less than 1600 K Future studies will carry out more combustion tests to attain fuel-rich combustion gas temperature less than 1200 K, which was a temperature requirement of a gas generator system in the previous studies.

• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.6
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• pp.8-16
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• 2009

The development of high-pressure sub-scale combustion chambers is described. A total of four high-pressure sub-scale combustion chambers having either a detachable structure of the mixing head and the chamber or a single welded regenerative cooling structure have been developed. The sub-scale combustion chambers have a chamber pressure of 70 bar and propellant mass flow rate of 5.1~9.1 kg/s. The propellant mass flow rate and the recess number of the injector were changed for the improvement of combustion performance and they were validated through hot firing tests. The design and manufacturing techniques of regenerative cooling channel and film cooling to be applied to the full-scale combustion chamber were adopted through the present development and verified.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.235-239
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• 2005

A 30-tonf-class fullscale thrust chamber for the application to a LEO SLV has been combustion tested over the wide ranges of a mixture ratio and a chamber pressure. The thrust chamber designed for an open cycle engine with a turbopump was tested with a ablative combustion chamber instead of a regenerative chamber to first evaluate its performance and function. The test results revealed stable combustion characteristics. The hardware survived the harsh environment and showed very sound functional characteristics. The estimated combustion efficiency of the chamber turned out to be 95% and a specific impulse at sea level was estimated as 254sec, which are comparable to or above the predicted design values.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1999.04a
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• pp.2-2
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• 1999

액체추진제 로켓엔진에서 분사기의 미립화 및 혼합 특성과 그에 따른 연소 특성은 성능과 안정성을 결정하는 중요한 파라미터이며 분사기는 제한된 설계 조건하에서 최대의 열방출율을 발휘하도록 설계되어야 한다. 여기서 연소효율은 연료와 산화제의 혼합특성과 충돌 분무의 미립화의 정도에 의해 결정되므로 충돌 분무 유동성의 혼합, 미립화 특성과 이에 따른 인조성능 특성을 명확하게 밝힘으로써 최대 엔진성능을 위한 설계가 가능하게 된다. 분사기의 설계에는 분사요소형태, 분사공의 형상 및 유동시스템 등이 포함되며 특히 분사요소 형태의 선택에는 추진제, 연소실냉각방법, 연소실 형상, 자동조건 및 엔진의 수명 등이 중요한 제한조건으로 고려된다. 이런 형태의 분사 요소들 중, 충돌형 분사기는 저장성 추진제를 사용하는 중, 저추력의 액체추진제 로켓엔진에 주로 사용된다. 이 분사형태는 미립화 성능이 높지 않고, 분사공 직경 및 운동량비에 따른 혼합성능이 만감하며 blow apart 등에 의한 열부하 혹은 안정성에 대한 문제가 있으나 양호한 혼합효율, 신뢰성과 제작의 용이함으로 인하여 광범위하게 사용된다.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.407-410
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• 2005

Acoustic coupling between combustion chamber and gas-liquid scheme injectors are studied numerically in liquid rocket engine adopting linear acoustic analysis. The injectors can play a role as half-wave resonators. The combustion chamber with numerous injectors shows peculiar acoustic coupling with the injectors. As the injector length approaches a half wavelength or the original tuning length, new injector-coupled acoustic modes show up in the chamber and thereby, the acoustic-damping effect of the tuned injectors is appreciably degraded.