• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.4
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• pp.81-88
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• 2013

It was designed and tested at the design point that an oxidizer rich preburner for a staged combustion liquid rocket engine propelled by kerosene and LOx. The oxidizer rich preburner was designed as some of LOx injected from the mixing head was burned with kerosene and the rest of LOx injected from injection holes in the regenerative cooling chamber was vaporized by combustion gas. The preburner is operated at OF ratio of 60 and combustion pressure of 20 MPa. The Preburner has a honey-comb type mixing head with simplex swirl injectors, a turbulence ring improving combustion stability and uniformity of product gas temperature distribution, and a nozzle simulating the duct. With the combustion test results at the design point, the oxidizer rich preburner showed high combustion stability and uniformity of product gas temperature distribution.

• 한국연소학회:학술대회논문집
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• 2014.11a
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• pp.311-313
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• 2014

High pressure combustion with multiphase--liquid, gas, and supercritical phase--mixtures are widely used technology in the high efficiency liquid propellent rocket engine. This is the typical characteristics differentiate from the combustor of conventional air-breathing engines. Therefore, successful research of high pressure combustion at supercritical condition is essential to develope a high efficiency liquid rocket engine. Numerical studies have been carried out to explore capabilities of numerical method for LOx-CH4 non-premixed flames at high pressure. In this paper, corresponding numerical results are presented and compared with experimental result of MASCOTTE facility.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.475-478
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• 2005

Ducted rocket produces thrust by 2 steps, primary incomplete combustion in the gas generator, and secondary complete combustion reaction in combustion chamber mixed by air taken through duct. the range of a rocket is determined by the weight of propellant, especially the weight of fuel. So ducted rocket has more efficiency and high terminal speed compared to traditional solid rocket motor. This propulsion system expected to be applied to various kinds of missile for anti-aircraft, anti-ship

• International Journal of Aerospace System Engineering
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• v.2 no.2
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• pp.1-5
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• 2015

The powder propellant rocket which uses micron-sized particles as fuel is storable and costly. Functions like thrust control and multiple-ignition can be realized by changing powder mass flow rate. In this paper, we discuss the theoretical performance of bi-propellant and mono-propellant powder rocket. When used as the fluidization gas, helium can improve specific impulse dramatically. The stability of the powder feeding device is preliminarily quantified through metal/N2O powder rocket hot fire tests.

• 한국연소학회:학술대회논문집
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• 2003.05a
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• pp.41-50
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• 2003

Numerical analysis was carried out to investigate performance and combustion characteristics of KSR-III liquid rocket engine with several types of baffle. To evaluate the change of performance and combustion characteristics with several types of baffle, the first numerical calculations were performed about baffle tab, radial blade baffle, and hub-and-spoke baffle. Then radial blade and hub-and-spoke baffle were determined to design two types of the KSR-III engine with baffles. Also to investigate the effect of injector arrangements and baffle positions, two types of radial blade baffle were calculated then numerical calculations were carried out with changing axial length of radial blade I, II and hub-and-spoke baffle. While axial length of baffle effected to performance very small, injector arrangement effected to performance largely through calculations of radial blade I, II. From the viewpoint of combustion instability, hub-and-spoke baffle controlled combustion instability effectively and there was the performance of hub-and-spoke baffle between radial blade I and II.

• Journal of Mechanical Science and Technology
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• v.18 no.7
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• pp.1246-1257
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• 2004

The objective of this paper is to characterize dynamic pressure traces measured at self-excited combustion instabilities occurring in two combustion systems of different hardware. One system is a model lean premixed gas turbine combustor and the other a fullscale bipropellant liquid rocket thrust chamber. It is commonly observed in both systems that low frequency waves at around 300㎐ are first excited at the onset of combustion instabilities and after a short duration, the instability mode becomes coupled to the resonant acoustic modes of the combustion chamber, the first longitudinal mode for the lean premixed combustor and the first tangential mode for the rocket thrust chamber. Low frequency waves seem to get excited at first since flame shows the higher heat release response on the lower frequency perturbations with the smaller phase differences between heat release and pressure fluctuations. Nonlinear time series analysis of pressure traces reveals that even stable combustion might have chaotic behavior with the positive maximum Lyapunov exponent. Also, pressure fluctuations under combustion instabilities reach a limit cycle or quasi-periodic oscillations at the very similar run conditions, which manifest that a self-excited high frequency instability has strong nonlinear characteristics.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.147-150
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• 2007

In this study, cooling characteristics of combustion gas were investigated by injecting liquid nitrogen into liquid rocket combustion chamber. A injection ring of liquid nitrogen was installed between a combustion chamber and a mixing chamber which was designed for mixing of combustion gas and nitrogen. At first, a ignition test of liquid rocket engine was conducted to verify a stable combustion process and 10 second combustion tests were successfully conducted. The results showed that combustion gas of LRE could be cooled by using liquid nitrogen.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.61-64
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• 2017

The gas-generator open cycle is adapted for liquid rocket engine of Korea Space Launch Vehicle-II. The combustion instability can interfere with combustion performance and cause a noise and vibration or carry the potential for serious damage. This study introduces the experience cases of combustion instability in development of the gas generator for liquid rocket engine.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.1-9
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• 2020

Using kerosene and liquid oxygen, 1.5-tonf class liquid-liquid pintle injector with rectangular two-row orifice was designed and manufactured. The combustion test of the pintle injector was carried out to verify the combustion performance and combustion stability under a supercritical condition which is the actual operation condition of the liquid rocket engine. The combustion test result showed that the pintle tip was damaged by the high temperature combustion gas in the high-mixed ratio recirculation zone of the combustion chamber. To solve this problem, the insert nozzle was installed in the pintle injector to increase cooling performance at the pintle tip. As a result of the hot firing test, installation of the insert nozzle, AR and BF had a great effect on pintle tip cooling performance.

• Journal of Mechanical Science and Technology
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• v.17 no.10
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• pp.1572-1582
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• 2003

Hybrid rocket combustion has a manifestation of stable response to the perturbations compared to solid propellant combustion. Recently, it has revealed that the low frequency combustion instability about 10 Hz was occurred mainly due to thermal inertia of solid fuel. In this paper, the combustion response function was theoretically derived by use of ZN (Zeldovich-Novozhilov) method. The result with HTPB/LOX combination showed a quite good agreement in response function with previous works and could predict the low frequency oscillations with a peak around 10 Hz which was observed experimentally. Also, it was found that the amplification region in the frequency domain is independent of the regression rate exponent n but showed the dependence of activation energy. Moreover, the response function has shown that the hybrid combustion system was stable due to negative heat release of solid fuel for vaporization, even though the addition of energetic ingredients such as AP and Al could lead to increase heat release at the fuel surface.