• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2009.11a
• /
• pp.125-129
• /
• 2009

The basic design of liquid rocket engine combustion chamber for a large space launch vehicle was described. It has vacuum thrust of 74.8 ton, vacuum specific impulse of 306.9 sec, chamber pressure of 60 bar, mass flow rate of 243.6 kg/s and combustion characteristic velocity of 1730 m/sec. The details of combustion performance and geometrical parameter were also given. The 75 ton combustion chamber consists of the combustor head with injector and the chamber/nozzle with regenerative cooling channels.

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

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2006.11a
• /
• pp.399-402
• /
• 2006

The purpose of this study is to propose a method for protecting injector face for prolonged combustion time and heat flux measurement technique at the injector face. To obtain basic design data and verify the performance of the proposed method, a regenerative cooling injector face was designed and manufactured for the hot firing test. Due to the safety reason, hot fire test were performed 3, 10, 30, 60 and 120 seconds time step. The discrepancy between analytical results adapting to combustion and nozzle and experimental results is believed due to the over estimation of the convection heat transfer calculation. for the injector face, flow velocity is almost negligible, therefore radiation is more important than convection. Consecutive hot firing test during 10, 30, 60 and 120 seconds combustion time shows good repeatability.

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

• Journal of the Korean Society of Propulsion Engineers
• /
• v.15 no.2
• /
• pp.44-49
• /
• 2011

A study of the fatigue life of copper alloy which was used in inner jacket of regenerative cooling chamber of liquid rocket engine has been performed. Generally used life prediction methods -original universal slopes method, modified universal slopes method, Mitchell's method, B$\"{a}$umel and Seeger's method, and Ong's method- have been used for predicting the fatigue data. It was found that the novel life prediction method which was modified from Ong's method was suggested since almost all data have not been predicted well with the widely used methods. The suggested modified Ong's method predicted well within 3X scatterbands.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2010.11a
• /
• pp.89-92
• /
• 2010

A study about the fatigue life of copper alloy which is used in inner jacket of regenerative cooling chamber of liquid rocket engine has been performed. Generally used life prediction methods-original universal slopes method, modified universal slopes method, Mitchell's method, Baumel and Seeger's method, and Ong's method-have been used for predicting the fatigue data. It was found that the novel life prediction method which is modified from Ong's method was suggested since almost all data have not been predicted well with the widely used methods. The suggested modified Ong's method predicted well within 3X scatterbands.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• v.y2005m4
• /
• pp.299-304
• /
• 2005

The design procedures of full-scale combustion chamber with chamber pressure of 53bara, mass flow rate of 90kg/s, combustion efficiency of $94\%$ and specific impulse at ground of 253sec were described. The details of combustion performance and geometrical parameters were also given. Full-scale combustion chamber consists of the combustor head with injector/baffle and the chamber/nozzle with regenerative cooling channels. The design results of combustion chamber with ablative materials, detachable injector head with SUS baffle or baffle injector and chamber body for ground hot firing tests were given in this paper.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.19 no.6
• /
• pp.64-71
• /
• 2015

Spinning process is generally used for manufacturing axisymmetrical, thin-walled thickness and hollow circular cross-section parts. Traditional spinning technology is classified to conventional spinning and power spinning(shear spinning and flow forming). Literature surveys of spinning application for regenerative cooling chamber and divergent nozzle of liquid propellent rocket thrust chamber have been conducted. Most spinning technology has been used mandel for manufacturing chamber and nozzle. Recently, hot spinning has been used much compared to traditional cold spinning.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2002.04a
• /
• pp.2-3
• /
• 2002

일반적으로 연소과정에서 발생한 고온고압의 연소가스로 인하여 액체추진기관의 연소실 및 노즐 벽면 그리고 추진기관 후방부위에 대류열전달(Convective heat transfer)과 복사열전달(Thermal radiative heat transfer)이 발생하는 것으로 알려져 있으며, 액체추진기관에서 발생하는 복사열전달 현상은 재생냉각장치의 열입력량 예측 및 발사체의 추진기관 후방부위에 탑재되는 전자장이 및 구조물의 열적환경(Thermal environmental phenomena)을 분석하는데 매우 중요하다. 이에 본 연구에서는 노즐 후방부위에서 발생하는 복사열전달량(Radiative heat transfer rate)을 측정하고 연소압(Chamber pressure)과 혼합비(Mixture ratio)에 따른 영향을 파악하였다.

• Journal of the Korean Society of Propulsion Engineers
• /
• v.25 no.4
• /
• pp.19-27
• /
• 2021

In this study, a high-pressure subsacle thrust chamber was developed to verify the core technology for KSLV-II performance enhancement. The core technologies are the design of an injector for high-pressure combustion, development of a combustion stabilization device using the additive manufacturing technique, and the design and fabrication of mixing head and regeneratively cooled combustion chamber. The core technologies, which have been verified through the development of high-pressure subscale thrust chamber, will be used to develop large engine liquid rocket engine thrust chamber in the future.