• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.2
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• pp.52-59
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• 2014

The effect of ceramic thermal barrier coating thickness on the cooling performance of a liquid-propellant rocket engine combustor has been investigated through combustion/cooling performance analysis whose results verified against measured data from hot-firing tests. Also have been confirmed the effects of film cooling amount near the face plate on the coolant temperature and on the thermal barrier coating surface temperature. Some important points to be considered for designing cooling schemes for regeneratively cooled rocket engine combustor have been drawn and reviewed from present study and further verification of the analysis tool should be performed in the future.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.2
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• pp.75-84
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• 2002

A numerical analysis of unsteady motion in solid rocket motors with a nozzle has been conducted. The numerical formulation including modified $\kappa$-$\varepsilon$ turbulence model treats the complete conservation equation for the gas phase and the one-dimensional equations in the radial direction for the condensed phase. A fully coupled implicit scheme based on a dual time-stepping integration algorithm has been adopted to solve the governing equations. After obtaining a steady state solution, pulse and periodic oscillations of pressure are imposed at the head-end to simulate acoustic oscillations of a travelling-wave motion in the combustion chamber. Various steady and unsteady state features in the combustion chamber of a rocket motor has been analyzed as results of numerical calculations.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.574-577
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• 2012

Thrust measuring is one of the crucial factor to decide the performance of a liquid rocket engine when the engine development test, especially for the combustion chamber, is implemented. Calculating the thrust from a combustion pressure is used when direct measuring the thrust is impossible, but direct measuring the thrust is necessary and various methods for doing it more precisely should be considered. This paper introduces the preliminary design concept about the new thrust measurement system for the vertical firing test stand, which is introduced domestically for the first time, of a liquid rocket engine combustion chamber.

• Journal of the Korean Society of Safety
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• v.13 no.4
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• pp.137-141
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• 1998

The combustion instability of a rocket motor can be predicted by the linear stability analysis. The most important input data in this analysis is the combustion response function of the solid propellant. In many cases, it is very difficult to measure the function. But, in that case, the combustion response function can be theoretically evaluated by properties of the propellant. In this study, the theoretical values were compared with measured values by T-burner. Data are relatively so well agreed that theoretical values are enough to be used in linear stability analysis of the rocket motor using a newly developed propellant.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.2
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• pp.179-185
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• 2008

Simulating combustion tests have been performed to elucidate the effect of baffle gaps on the optimal damping characteristics in liquid rocket combustors where coaxial injectors are installed. Amplitude of pressure oscillation in model combustion chamber and the combustion stability margin are used to quantify the damping capacitance of baffles. Satisfactory agreement has been achieved with the results of cold acoustic tests. Present results have shown that the optimal gap for high acoustic damping capacity has also the large combustion stability margin in simulating combustion tests. Therefore, the present results can be utilized to determine the baffle length and optimal gap in full-scaled rocket combustors.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.353-358
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• 2008

Inha Rocket Research Institute has made the igniter that is combination of black powder and PVA polymer for ignition small rocket. But recent igniter is not satisfy because of the performance of igniter is not identified. So, we confirmed requirement of igniter by comparing of ratio of black powder and PVA through experimental method. Especially we studied with ignition temperature for propellant and stable combustion pressure that is requirements of propellant. We can know the tendency of combustion properties by ratio of oxidizer and combustion catalyst through changing of temperature and pressure of exhaust gas of igniter.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.672-674
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• 2017

The addition of swirl is a common technique used in premixed combustors in order to gain stability of the combustion with the improvements in mixing characteristics. recent experimental studies have observed that the addition of swirl oxidizer flow can effectively reduce the combustion instability in hybrid rocket. Investigation was continued to analyze the effect of the swirl on the internal flow of hybrid rocket engine main combustion chamber. The flow influenced by wall blowing as a representation of fuel evaporation interacts with swirling flow. Swirl angle increases, the amplitude of the combustion pressure decrease as the unstable combustion processes. These results suggest that the oxidizer swirling flow by the swirl angle causes the change of the turbulent flow characteristics inside the combustion chamber and suppresses the factors causing the combustion instability.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.7
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• pp.573-580
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• 2021

Paraffin wax is attracting many attentions for promising solid fuel of hybrid rocket because of its higher regression than other fuels. However, even with paraffin fuel combustion, unsteady low-frequency oscillation of combustion pressure is still observed. And, this is related to the formation of liquid layer and the entrainment of fuel droplets entering the axial combustion gas flow. This study investigates the effect of additional combustion of fuel droplets on the occurrence of low-frequency combustion instability. On the other hand, the formation of fuel droplets depends on Weber Number (the ratio of the inertial force to the surface tension of the liquid) and Reynolds Number of the oxidizer flow. Therefore, a laboratory-scale hybrid rocket was used to monitor the occurrence of combustion instability while changing We number. A series of combustion tests were conducted to control We number by changing the oxidizer flow rate or adding LDPE (low density polyethylene) to base fuel. In the results, it was confirmed that there is a critical We number above which the low-frequency combustion instability occurs.

• International Journal of Aeronautical and Space Sciences
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• v.13 no.4
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• pp.507-514
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• 2012

The present study is the analyses of what has been attempted and what was understood in terms of improving the regression rate and enlarging the basic understanding of internal flow dynamics. The first part is mainly intended to assess the role of helical grain configuration in the regression rate inside the hybrid rocket motor. To improve the regression rate, a combination of swirl (which is an active method) and helical grain (which is a passive method) was adopted. The second part is devoted to the internal flow dynamics of hybrid rocket combustion. A large eddy simulation was also performed with an objective of understanding the origin of isolated surface roughness patterns seen in several recent experiments. Several turbulent statistics and correlations indicate that the wall injection drastically changes the characteristics of the near-wall turbulence. Contours of instantaneous streamwise velocity in the plane close to the wall clearly show that the structural feature has been significantly altered by the application of wall injection, which is reminiscent of the isolated roughness patterns found in several experiments.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.393-397
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• 2006

In this study, combustion characteristics of solid propellants using sorbitol and potassium nitrate were found out. Burning rate was calculated with several combustion experiments, also specific impulse and characteristic exhaust velocity were compared with theoretical value. Thrust measured with thrust measurement system using plate spring. Mixture ratio of propellants was varied in experiments, also combustion characteristics of solid propellants which consulted experimental results was used micro solid rocket design having 1mm nozzle throat.