• Journal of the Korean Society of Propulsion Engineers
• /
• v.19 no.5
• /
• pp.62-70
• /
• 2015

A liquid rocket engine system can cause rapid pressure and temperature variations during the startup period. Thus the startup analysis is required to reduce time and expense for successful development of liquid rocket engine through the startup prediction. In this study, a startup analysis simulator is developed for a staged combustion cycle engine powerpack. This simulator calculates propellant flow rates using pressure and flow rate balances. In addition, a rotational speed of turbopump is obtained as a function of time by mathematical modeling. A startup analysis result shows that the time to reach a steady-state and a rotational speed at the steady-state are 1.3 sec and 27,500 rpm, respectively. Moreover it can indicate proper startup sequences for stable operation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2017.05a
• /
• pp.1021-1025
• /
• 2017

Numerical heat/flow analysis was performed on a liquid rocket engine head with the cooling water manifold to ensure the durability of a ground test facility for heat exchanger. Through these studies, the shapes of the injector and the flow path were determined and applied to the head of the engine under development. Firing tests were conducted to verify the designed coolant manifold and no thermal damage was found on the engine-head-face. Comparing the combustion test results with the numerical analysis, the outlet temperature of coolant showed a difference of about $15^{\circ}C$. This trend is reasonable considering existence of LOX manifold, thermal barrier coating, and the actual location of flame.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2005.11a
• /
• pp.240-246
• /
• 2005

A series of combustion tests of a 30-tonf-class full scale liquid rocket thrust chamber under development has been carried out to verify its design. The test results revealed decent performance in the aspects of efficiency. The combustion stability is one of the most important parameters of liquid rocket engine in addition to the efficiency. Assessment tests of combustion stability must be accomplished to confirm the possibility of combustion instability due to spontaneous or external disturbances. The combustion stability rating tests of the full scale thrust chamber with temporary baffles made of stainless steel were carried out utilizing pulse guns to estimate combustion stability characteristics. The tests results show highly stable combustion stability characteristics. The outcome acquired from the present experimental study will be used to design an actively cooled baffle that can survive for the life time operation of the thrust chamber.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.31 no.10
• /
• pp.66-72
• /
• 2003

The cooling mechanism for a liquid rocket engine of 10tf-thrust using kerosene as a fuel was studied from the viewpoint of both the regenerative and curtain cooling. Based on the concept of a highly-stratified gas flow in the combustion chamber, the cross section of the combustion chamber was spilt into 2 independent parts, core and exterior part. Additional fuel is injected into the exterior section and gas temperature can be reduced in the exterior section. Consequently, the heat flux into the coolant and wall temperature are reduced and the thermal stability of a liquid rocket en g i.ne could be improved.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2005.11a
• /
• pp.341-345
• /
• 2005

This paper is for the combustion characteristics of gas generator which drive 1.5MW-class turbo pump and runs in fuel-rich combustion regime with LOx/kerosene as propellant. The outline of development procedure of real scale high pressure gas generator is introduced and the relation between O/F ratio and outlet temperature and the molecular weight and specific heat ratio of combustion gas are described. The relation between O/F ratio and temperature is newly obtained at higher pressure and the molecular weight and specific heat ratio is modified and their validity is confirmed by the mass relation equation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2008.05a
• /
• pp.163-167
• /
• 2008

This paper presents pressure fluctuation characteristics of a 30 ton-f class liquid rocket engine combustor. Combustion stability of the combustor was evaluated using the results 46 firing tests performed with a varying O/F ratio and chamber pressure. The RMS value of pressure fluctuation during the steady state combustion was less than 2.6% of the static chamber pressure, demonstrating static stability of the combustion phenomenon. The decay time of pressure fluctuation caused by forced disturbance of a pulse gun was found to be less than 3.5 msec verifying dynamic stability of the combustor.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.130-133
• /
• 2009

The sequence of the propellant supply is very important for the reliable and safe operation of a LRE combustion test. So combustion performance tests were performed to find an optimum test sequence by changing supply time of propellants and purge gas in the moment of ignition and extinction. The liquid rocket engine consisted of a catalytic ignitor and six swirl-coaxial injectors which used hydrogen peroxide and kerosene. Conclusively, an optimum sequence was found for stable combustion in the moment of ignition and extinction.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.18 no.6
• /
• pp.88-95
• /
• 2014

Spinning process to inner wall has been applied for reducing the weight of regenerative cooling chamber of liquid propellent rocket engine. The fractures of the blanks of cylinder part and nozzle throat part have been observed during spinning processes. In order to overcome the problem, the mandrel and the blank shape have been modified, and the inner wall was successfully manufactured through the modifications. The manufactured spinning prototype of nozzle throat part was successfully bulged without cracking and necking, and it was confirmed to secure sufficient formability necessary for fabricating thrust chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.10a
• /
• pp.259-264
• /
• 2004

The critical component of combustor having high combustion efficiency for high performance liquid rocket engine is injector. The results of design and hot firing tests of six sub-scale combustors which have respectively an impinging type injector(1ea.), an bi-propellant swirl closed injector(1ea.), and hi-propellant swirl mixed injector(4ea.) were described in this paper. The combustion test were successfully performed. The combustion efficiency have higher value than predicted value and high frequency combustion instability does not occur.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.636-641
• /
• 2010

A preburner for the staged combustion in the high performance liquid rocket engine is being developed and strength experiments and finite element analyses on specimens, which simulate the brazing joint of the preburner, were performed and the results were compared. One specimen simulate the joints near oxygen injectors of head by the funance brazing and two specimens the joints of the combustion chamber cooling channel by vacuum brazing. The experiments were burst ones with strain gauges.