• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.2
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• pp.68-74
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• 2006

An experimental study was carried out to investigate the combustion characteristics of the slinger combustor. A combustion test rig was manufactured and installed in KARI combustor test facility. From the ignition test results, we found that there were two major factors influencing the ignition limits; by increasing the rotational speed and the air mass flow rate, a better ignition performance was attained. From the combustion test results, we obtained 99.6% combustion efficiency, 15% pattern factor, and 3% profile factor. The results in this work indicate that the ignition and combustion characteristics of a slinger combustor are markedly different from those of a conventional annular combustor.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.12
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• pp.35-40
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• 2006

The combustion instability in a model dump combustor with an exhaust nozzle and the possibility of combustion control using a loudspeaker to these instabilities were studied. By changing inlet velocity, combustor length and equivalence ratio, dynamic pressure signals and flame structures were simultaneously taken. Because inlet velocity and combustor length affect the life time of vortex in the dump combustor, the results showed that as the combustor length increased and the inlet velocity decreased, the instability frequency decreased and the maximum power spectral density of the dynamic pressure generally decreased. Also, instability frequency and maximum power spectral density of the dynamic pressure increased with the increment of equivalence ratio. From the data of close-loop control, the optimum time-delay control using a loudspeaker was confirmed to be able to reduce the vortex shedding induced from the mixed acoustic-convective mode combustion instability.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.6
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• pp.79-87
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• 2003

Application of combustion stabilization devices such as baffle and acoustic cavity to liquid propellant rocket engine is investigated to suppress high-frequency combustion instability, i.e., acoustic instability. First, these damping devices are designed based on linear damping theory. As a principal design parameter， damping factor is considered and calculated numerically in the chambers with/without these devices. Next, the unbaffled chambers with/without acoustic cavities are tested experimentally for several operating conditions. The unbaffled chamber shows the peculiar stability characteristics depending on the operating condition and it is found to have small dynamic stability margin. As a result, the acoustic cavity with the present design has little stabilization effect in this specific chamber. Finally， stability rating tests are conducted with the baffled chamber, where evident combustion stabilization is observed, which indicates sufficient damping effect.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.187-192
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• 2006

The combustion performance tests of a 30 tonf-class full-scale combustion chamber performed with water as a coolant were described. The combustion chamber has chamber pressure of 53bara and propellant flow mass rate of 90kg/s. Since it was first firing test for 30tonf-class combustion chamber using channel cooling, water coolant mass flow .ate of 35kg/s and 18kg/s were performed which correspond to 110% of kerosene design volume flow rate and equivalent cooling performance of kerosene. The test results are described and the results showed that the water cooling performance of this combustion chamber is sufficient and the firing test is feasible using the kerosene as a coolant.

• 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.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.40-46
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• 2006

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

• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.5
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• pp.15-20
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• 2009

Hot-firing tests were performed on two 30 tonf-class regenerative cooling combustion chambers, with different injector distribution and wall cooling method. In the paper, the combustion stability test results were analyzed and presented. The pressure fluctuation and stability rating test(SRT) results of the combustion chambers were examined to evaluate combustion stability. The combustion chambers exhibited satisfactory results on combustion stability. The RMS values of the chamber pressure fluctuation were less than 3% of the chamber pressure and the decay time of artificial pressure peaks was measured to be around 10% of the reference decay time. It is interesting that the RMS values of pressure fluctuation in the combustion chamber with film cooling are smaller than those in the chamber with cooling injectors at the periphery row.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.231-235
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• 2009

The mainly objectives of this study was a combustion dynamics and instability characteristics in a model dump type combustor which is scale down of GE 7FA+e DLN 2.6 gas turbine combustor with running at Seo-Inchon combined cycle power plant. Model gas turbine injector has 2-stage swirl vane and it's reduced 1/3 size of the original one. The shape of plenum and combustor were designed for similar acoustic characteristics. As the result, this research have been shows the peak frequency of model combustor was changed quarter-wave mode to Helmholtz resonator mode in plenum and longitudinal mode in dump combustor at unstable flame conditions caused by the different of combustor temperature and fuel-air mixture distributions.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.3
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• pp.52-60
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• 2010

Combustion stability tests of 30 $ton_f$-class LRE thrust chamber with double swirl coaxial injector were carried out in domestic ground combustion test facility by means of artificial perturbation method. In these tests, thrust chambers with varying design factors like recess number of injector, baffle length, types of film cooling and chamber diameter were used and test results showed that these design factors are closely related with high frequency combustion stability. By using the oscillation decrement instead of the decay time in the combustion stability analysis of artificially perturbed LRE thrust chamber, it was confirmed that increment of damping factor results in the improvement of high frequency combustion stability of LRE thrust chamber.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.10
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• pp.1027-1037
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• 2009

The development of a combustion chamber for a 30-$ton_f$ regeneratively-cooled space liquid rocket engine is described. Starting from the development of bi-propellant swirl coaxial injectors, essential technologies were verified through subscale combustion chambers and afterwards applied to the full-scale combustion chambers. A total of 5 full-scale combustion chambers have been utilized to verify ignition, combustion efficiency and stability, cooling, and duration requirements. A total of 46 combustion tests were performed among which 23 tests were parallely performed with stability rating tests using a pulse gun device. The test results have revealed that the 30-$ton_f$ regeneratively-cooled combustion chamber fully complies to the performance and combustion stability requirements and thus concluded that the development is successfully completed.