• 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.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.

• 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.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.892-897
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• 2011

The propulsion system of space launch vehicle generates thrust by supplying oxidizer and fuel to combustion chamber. KSLV-II 2nd stage engine, currently under development by KARI, is to use liquid oxygen as a oxidizer and JET-A1 as a fuel. The 2nd stage pump-fed engine is mainly composed of combustion chamber, turbo-pump and engine supply system. To develop liquid propulsion engine, the development of combustion chamber must be preceded. For performance validation of the combustion chamber, the designed and manufactured combustion chamber should be tested in combustion chamber test facility(CCTF). The detailed design for the planned CCTF in Naro Space Center was conducted. The fuel supply system modeling using AMESim was performed based on the results of the detailed design, and the fuel supply characteristics was analyzed in this paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.3
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• pp.9-17
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• 2002

This Study is to investigate 8 composite propellants including Butacene and ${Bi_2}{O_3}$ by Shotgun/RQ Bomb test. Burning rate and mechanical properity are known to be major factors in determining the deflagrability of propellant. Propellant including over 5.5% Butacene(Ferrocene grafted HTPB) burned out over 135 m/s of impact velocity during Shotgun/RQ Bomb test. It was blown that Butacene was very sensitive material under high velocity impact. In the test results, propellants under 25mm/s in burning rate at 1500 psia could meet the requirements for IM of UN Test Series 7c(ii). Propellant deflagrabillity depends on burning rate at performance in the results of the present.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.3
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• pp.109-117
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• 2006

Hot firing performance test of hydrazine decomposition catalyst used for monopropellant thruster of satellite and launch vehicle was performed on the ground. A test equipment for hot firing performance measurement of catalyst test was developed in collaboration with Hanwha Corp., and the catalyst firing performance were tested with the equipment. After a reaction delay time, a catalyst activity and a granule stability were measured for 2 times, satisfactory results were obtained such as 25msec, 2%, $704^{\circ}C$ for each test items on the average. In addition, the current development status of domestic prototype catalyst and its decomposition performance test results are presented.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.245-250
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• 2004

The objective of the present study is to conduct model combustion tests for various injectors to identify their combustion stability characteristics. Three different double swirl coaxial injectors with variation of a recess number have been tested for the comparative study of stability characteristic and flame structure. Gaseous oxygen and mixture of gaseous methane and propane have been employed for simulating actual propellants used for a fullscale thrust chamber. Upon test results, the direct comparison between various types of injectors can be realized for the selection of the best design among prospective injectors.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.133-136
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• 2008

The present study presents experimental results of combustion tests of a fuel-rich gas generator associated with a turbopump. Five combustion tests had been successfully executed. Static pressures of the gas generator promptly reacted to propellant supply variations from the turbopump. A closed-loop test for driving the turbopump revealed no flaw. Exit gas temperature results are very similar to previous ones. An orifice was effective on the suppression of pressure fluctuations although tests conducted below 45 bar showed the same dynamic behaviour as that of component-only tests.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.5
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• pp.19-26
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• 2011

In order to investigate the effect of fuel injection hole configuration within the scramjet combustor, experiment and quasi-one-dimensional analysis was performed. And the results were compared with experiment and analysis result which were performed in 2008 with same facility and test condition. Fuel injection hole size was decreased and quantity was increased. However the depth of fuel penetration and flow quantity of fuel were maintained. As a test result, combustion performance was increased significantly with no-cavity injector and slightly with plain-cavity. However, combustion performance with zigzag-cavity was decreased.