• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• v.y2005m4
• /
• pp.313-317
• /
• 2005

The facility improvement for hot firing test of combustion chamber having thrust of 30-tonf class and chamber pressure of 60bara were performed at ReTF in KARI. The KSR-III main engine having combustion pressure of 13bara and thrust of 12.5tonf had been successfully tested in this facility. To increase the capability of the facility, the feeding and the trust measurement system have been modified. The modification of the feeding system plays also a role of ensuring the stability of propellant supply and two step ignition sequence of combustion chamber. The one-axis thrust measurement system of up to 60tons has been newly manufactured and installed in test stand and the water/kerosene supply lines with high pressure vessel of $4m^3$ and gas nitrogen vessel of $10m^3$ have been designed for regenerative cooling system. The results of cold flow test show that this facility has been successfully improved to satisfy the requirement for hot firing test of high performance combustor.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2007.11a
• /
• pp.271-274
• /
• 2007

In the fuel of the solid fuel ramjet there are metal particles in order to improve the Isp like as solid rocket propellants. Because of the short combustion residence time these metallized fuels have low combustion efficiencies. Therefore it is necessary to increase the combustion efficiency and the inlet air temperature does an important role to this. The main factors to affect the inlet air temperature is the free stream temperature and the flight Mach number. Also the flow velocity in the combustor does an important role, therefore entire range of the air flow; from the stagnation to the sonic velocity in the ramjet combustor is considered.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.11 no.5
• /
• pp.78-84
• /
• 2007

In the regeneratively cooled combustion chambers of liquid rocket engine using hydrocarbon fuels, coking occurs as the wall temperature increases which results in compounds deposition on the wall of cooling channels. This phenomenon reduces cooling capability of the coolant which finally causes damage to the combustor by overheating of the chamber wall. In this paper, experiment results using an electrical heating equipment to simulate the regeneratively cooled channel are introduced and based on the results the compatibility of copper alloy with hydrocarbon fuel Jet A-1 is investigated.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.33 no.5
• /
• pp.79-86
• /
• 2005

The role of the injector as an acoustic resonator is studied for the high performance rocket engine adopting the gas-liquid scheme injector. Acoustic behavior in the combustor with single injector is investigated numerically adopting linear acoustic analysis for cold condition. Acoustic-damping effect of the injector is evaluated by damping factor as a function of the injector length. From the numerical results, it is found that the injector can play a significant role in acoustic damping and the optimum length of the injector corresponds to half of a full wavelength of the longitudinal mode with the acoustic frequency to be damped in the chamber. In baffled chamber, the optimum lengths of the injector are calculated as a function of baffle length for both cold and hot conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.05a
• /
• pp.103-106
• /
• 2008

In this work it was studied that the effect on sub-component of engine considering change of engine inlet pressure caused by variable acceleration during flight of launcher. Also the transitional performance was predicted according to variable acceleration. Engine inlet pressure was defined as summation of propellant head in tank, ullage pressure and pressure difference of line, etc. Therefore consumption of propellant and acceleration of launcher led to change of engine inlet pressure, which affected on discharge pressure of pumps. This effect changed mass flow rate of gas generator and main combustor hence it was observed that engine performance was changed definitely.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.17 no.3
• /
• pp.67-75
• /
• 2013

The numerical analysis for the verification of preburner's cooling characteristics applying to kerosene-LOx rocket engine has been fulfilled. The distribution of combustion gas properties in primary combustion zone was calculated by the mixture ratio based on head injector arrangement, the properties of oxygen flowing in wall channels as coolant were applied under real-gas conditions, and multi-phase mixing model was employed to calculate the mixing process of primary combustion zone with liquid oxygen which was used for wall cooling. The results of numerical analysis were compared with the experimental results, hence thermo-physical properties in cooling channels and a combustor could be quantitatively identified.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.05a
• /
• pp.66-70
• /
• 2010

The performance of 75 ton liquid rocket engine combustion chamber for a space launch vehicle was predicted through firing tests at low pressure. In low pressure tests of 75 ton LRE combustor chamber, the combustion characteristic velocity of 1750 m/sec and the specific impulse of 240 sec were obtained which are higher than the low pressure performance of 30ton combustion chamber. The combustion characteristic velocity of 1770 m/sec and the specific impulse of 278 sec at design point for 75 ton LRE combustion chamber were predicted by using the low/high pressure performance correlation of 30ton LRE combustion chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.05a
• /
• pp.61-65
• /
• 2010

To research and develop a liquid rocket engine injector, it needs empirical studies about the hydrodynamic and spray characteristics such as pressure drop, mixing and atomization. In this study, the design and implementation of lab-scale cold-flow/hot fire test stand which can supply cryogenic propellant and be controlled by time-critical LabVIEW cyclogram logic has been done. In order to visualize the spray of a liquid-centered swirl coaxial injector in cryogenic condition, LN2-GN2 cold-flow test has been done, and combustor assembly and thrust bed for LOX-$GCH_4$ hot-fire test have been fabricated.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.25 no.3
• /
• pp.1-13
• /
• 2021

Effects of the nozzle wall angles on the supersonic flow field in a thrust optimized nozzle were numerically investigated. The combustor and operating condition of 30-tonf rocket engine was selected to study the optimum nozzle shape. The nozzle flow of combustion products was realized by the shifting equilibrium calculation for the propellant of kerosene-LOx. The change of nozzle wall angles induced different developing patterns of the internal and secondary shock wave. The optimum nozzle was obtained when the internal shock was in a specific position at the nozzle outlet. The nozzle wall angles of the optimum nozzle were very similar to those of the optimum nozzle which does not consider the shock wave.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.35 no.12
• /
• pp.1108-1114
• /
• 2007

The effect of oxidizer injection angle on the combustion characteristics of end-burning hybrid combustor is numerically investigated. Besides the previously studied parameter(injector arrangement, port diameter and O/F ratio), three different injection angle are considered: parallel angle to fuel surface(Case 1), +30 degree inclined angle toward the fuel(Case 2) and 30 degree inclined angle toward the nozzle(Case 3). It is found that Case 2 has the best mixing pattern in the upstream area but has the worst combustion efficiency since non negligible amount of unburned fuel is expelled from the nozzle. In contrast, though Case 1 and Case 3 showed relatively low mixing effect than the Case 2, they had high combustion efficiency. The comparison of numerical results between Case 1 and Case 3 demonstrate that no major difference is encountered, however, Case 1 is expected to have the best combustion efficiency due to the low residence time of the Case 3 injector which heads toward the nozzle.