• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.38 no.4
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• pp.395-402
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• 2010

Baffle injectors are protruded into the combustion chamber and form an anti-pulsating baffle to prevent high-frequency combustion instabilities in transverse modes. Being exposed to a high heat-flux environment, the baffle injector has self-cooling passages through which kerosene is convected and heated. The baffle injector with 20 spiral cooling channels has been developed and successfully applied to 30 $ton_f$-class combustors without any performance loss due to an additional cooling. In this work, numerical analysis of conjugate heat transfer in baffle injectors with various cooling channel designs has been performed in order to reduce the fabrication cost which would be considerably increased for the 75 $ton_f$-class combustor. Prior to the application to a full-scale combustor, the thermal durability of the modified design has been verified through the subscale hot-firing tests.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.5
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• pp.45-50
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• 2010

Thermal analyses have been performed to study the effect of location of fuel ring and thermal barrier coatings in regenerative cooling channels in a full-scale combustor. For the effective cooling, the fuel ring has better be installed near axial location of the low expansion ratio and low heat flux, and branching of cooling channels is preferable. Also, the radiative cooled nozzle extension is thought to be reasonable for the cooling of combustor walls. Among the possible coatings, $Y_2O_3$ stabilized $ZrO_2$ coating and Ni/Cr coating have been adopted. Compared with Ni/Cr coating which has high oxidation resistance, $Y_2O_3$ stabilized $ZrO_2$ coating, one of ceramic coatings is found to be much effective to sustain the thermal survivability of combustion walls.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.622-627
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• 2017

In this study, the numerical analysis of propellant supply manifold of a sub-scale model combustor with the same combustion and acoustic resonance conditions as the real combustor was carried out. The analysis of the results is based on the shape of the manifold and the number and type of inflow channels. The manifold form of the main propellant was rounded so that the recirculation area was small and easy to manufacture. The liquid oxygen mainfold included a distributor to uniformly supply the flow, and since the kerosene manifold was directly in contact with the flame side, the inflow channel was formed at a radial $360^{\circ}$ so as to minimize the recirculation region. The liquid nitrogen manifold was configured as a radial inflow channel to reduce the velocity difference near the injector.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.1
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• pp.88-97
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• 2022

In this study, a vitiated air heater was designed and manufactured to supply high-temperature and high-pressure air to the ground test apparatus of a direct-connected supersonic combustor, and an experiment was performed to verify the target design point. By installing wedges at the upper boundary, lower boundary and center of the nozzle exit of the vitiated air heater, it was confirmed that the Mach number satisfies the 2.0 level, and the pressure of the combustion chamber was also satisfactory compared to the design point. In the case of temperature, the measured temperature deviation was large due to the degree of exposure of the thermocouple and the slow response characteristics. After that, the isolator was connected to the rear of the vitiated air heater, and the Mach number was measured in the same method, and the Mach number at the center of the isolator eixt was slightly reduced to 1.8~1.9.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.8
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• pp.720-725
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• 2007

Cold acoustic tests have been performed to elucidate the effect of baffle gaps on the optimal damping characteristics in a liquid rocket combustor where coaxial injectors are installed. For several axial baffle lengths, an optimal acoustic damping capacitance has been achieved in a certain gap range. Cold acoustic tests for simulating fluid viscosity by changing the pressure in a model chamber have been done to study the main mechanism of optimal damping. Experimental data have shown that the optimal gap for high damping capacity exists mainly due to the viscosity near the gap of baffles. Therefore, axial baffle length can be reduced by using the optimal baffle gap, providing a possible solution of thermal cooling problems. Also, these optimum characteristics can be some guidelines for manufacturing and assembling injectors in full-scaled rocket combustors.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.50-54
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• 2011

This paper presents a two-phase model for hybrid rocket internal ballistics design using $N_2O$ as oxidizer The two-phase model results are compared with data obtained from static firing test. Two-phase model is suitable for blow-down type with saturated compressible fluid as $N_2O$, presented the result by Part 1. HDPE as Fuel, and $N_2O$ as oxidizer were used during the static firing test. The combustor were designed for an average thrust of 30 kgf where oxidizer tank pressure in set to 50 bar. The numerical results of internal ballistic showed good agreements with static firing test results where thrust, oxidizer tank pressure and chamber pressure are compared.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.45-49
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• 2011

The blow-down oxidizer feed system with self-pressurizing $N_2O$ has more advantages than the regulated system. However, it is difficult to predict the exhaust flow rate because there exist two phases in the $N_2O$ tank - liquid phase and gas phase, and the properties of $N_2O$ in storage tank are varied continuously during blow-down. In this paper, a method that can analyse simply the blow-down oxidizer feed system is studied. The properties of saturated $N_2O$ are found from the NIST data base, and mass flow through the orifice is modeled as NHNE. Cold flow test with hybrid rocket combustor is performed for the comparison where the results should found from the good agreement.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.114-118
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• 2007

Design and fabrication of a 30-tonf-class full-scale regenerative cooling combustion chamber of a liquid rocket engine for a ground hot firing test are described. It has chamber pressure of 60 bar and nozzle expansion ration of 12 and manufactured to have a single welded structure of· the mixing head and the chamber. The material of the mixing head is STS316L which has excellent mechanical property in cryogenic condition. The chamber comprise of the cylinder, nozzle throat, and 1st/2nd nozzle parts. The material of the inner jacket is copper alloy/STS329J1/STS316L and that of the outer jacket is STS329J1. The components of· the combustor were manufactured by mechanical processing including lathing, milling, MCT, rolling and pressing. The machined components were integrated to a single body by means of general welding, electron beam welding(EBW), and brazing.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.9
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• pp.82-88
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• 2006

In order to effectively reduce thermal loads on regenerative cooled walls, fuel cooling injectors and film cooling devices have often been employed. The present study has established a numerical methodology for prediction of performance and near-wall temperature distribution taking into account the nonuniform mixing due to these additional cooling devices. A correction procedure for main propulsive parameters has also been proposed based on comparison between prediction and experimental data. Under the computational framework of this study, the predicted results were in good agreement with hot-firing test data for a 30 tonf-class full-scale combustor at the design and off-design conditions. As a consequence, the present numerical method is expected to be useful for design and evaluation of regenerative cooled liquid rocket thrust chambers.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.5
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• pp.37-50
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• 2007

This study investigated the characteristics of spray generated by a liquid coaxial swirl injector used in a combustor of the liquid rocket engine. The linear stability analysis considered long and short wave was introduced in liquid sheet breakup. Through the hydrodynamic analysis the initial liquid sheet thickness spray angle and injection velocity were predicted. To evaluate the effect of turbulence model standard $k-{\varepsilon}$ and RNC $k-{\varepsilon}$ model were applied to numerical calculation and it was known that RNC $k-{\varepsilon}$ model was more applicable to predict spray characteristics. On the basis of this evaluation validation of the developed model was performed with swirl injector installed in LPRE and the predicted results of breakup length, spray angle, and SMD agreed well with experiments qualitatively and quantitatively.