• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.5
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• pp.91-97
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• 2015

After-burner test facility for gas generators of 75 tonf class liquid rocket engines was designed, which was verified by the facility certification test of the Combustion Chamber Test Facility(CCTF). The purpose of the certification test of the after-burner test facility is to verify the combustion stability of gas torches equipped in the gas generator and the after-burner test facility by using methane and oxygen gases. In the case of the autonomous test, the supply system provided steadily methane and oxygen gases to the after-burner system without pressure drop. The combustion pressure of the gas torch approached the design requirement. In the case of the coupled test, the gas generator ignition and the fuel-rich exhaust gas combustion were successfully carried out, leading to the verification of the test facility.

• Journal of the Korean Society of Propulsion Engineers
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• v.14 no.5
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• pp.92-100
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• 2010

Combustion stability rating tests of 75-tonf-class thrust chamber for technology demonstration were carried out at a low pressure. Two kinds of mixing heads were used in this study. One of them has injectors of 631 and the other has injectors of 721. Mixing head with injectors of 631 showed a self-oscillation instability at the chamber pressure of 30 bar. Mixing head with injectors of 721 showed that a high frequency combustion stability was maintained under the same pressure and the same mass flow rate. But mixing head with injectors of 721 generated a self-oscillation instability at the chamber pressure of 20 bar and it was found that stability boundary region was changed due to the configuration of a mixing head from these results.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.2
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• pp.55-64
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• 2015

In this study, a 3-D flame cooling analysis is conducted to examine thermal safety for the flame deflector of the 75 tonf class propulsion test facility, and the safe discharge of the exhaust gas is assessed by using numerical results. The Mixture multiphase model is adopted for the simulation of heat transfer and phase exchange process between flame and cooling water, and the computational study using the single species unreacted model for the exhaust plume is carried out for the flame cooling. Numerical analysis predicts maximum temperature on the flame deflector wall for different water flow rates, and evaluates the safe minimum flow rate of water corresponding to the fire-resistant temperature for concrete.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.479-483
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• 2012

This paper describes the development status of rocket engine test facility for the performance evaluation of liquid rocket engine of KSLV-II 1st stage. Design specification and composition of rocket engine test facility are suggested based on the design requirements. The results of the basic design of rocket engine test facility will be used as base data for the detail design and construction of rocket engine ground test facility of KSLV-II 75tonf liquid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.608-612
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• 2009

Technology demonstration models(TDM) of a 75-$ton_f$ liquid rocket engine(LRE) thrust chamber were manufactured on the basis of development technologies of 30-$ton_f$ LRE. It was confirmed that some machining and welding technologies which were aimed to be verified through the manufacturing of demonstration models could be applied to the thrust chamber 75-$ton_f$-class. New designed mixing head part was manufactured by means of new process. The manufacturing process and technologies established through TDM's will improve the reliability of manufacturing process of large LRE thrust chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.885-891
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• 2011

This paper describes the results of preliminary design of rocket engine test facility for the performance evaluation of liquid rocket engine. Design specification and composition of rocket engine test facility are suggested based on the design requirements. The results of the preliminary design of rocket engine test facility will be used as base data for the detail design and construction of rocket engine ground test facility of KSLV-II 75tonf liquid rocket engine.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.5
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• pp.67-74
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• 2019

In this paper, mathematical models of a simulation code are presented. The simulation code was developed for the startup analysis of an open cycle liquid rocket engine (LRE). Most of the components comprising an LRE, including the priming process in the propellant feeding line, were considered. A startup simulation of a 75-tonf LRE, which was used for the KSLV-II test launch vehicle (TLV), was performed. The simulation results showed good agreement with the engine acceptance test results, thus proving the validity of the startup simulation code.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.1
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• pp.88-95
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• 2015

This paper considers the effects of propellant mixture ratio and loading errors on the performance of a liquid rocket. Propellant residuals generated by error sources are analyzed for a launch vehicle model whose first stage consists of a cluster rocket of four 75-tonf class engines using a statistical Monte-Carlo approach and then the optimal fuel biases minimizing residuals are computed. The results are validated through comparison with analytic method using approximate formula, which have been applied for other space launch vehicles.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.1084-1087
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• 2017

When a liquid rocket engine is started the oxidizer and fuel must be flowed into combustion chamber and gas generator with time differences. The wrong time difference between propellants or malfunction of ignition device can occur the explosion of combustion chamber due to detonation by energized premixed-propellants. Therefore it is important to observe the transient characteristic of propellants or to measure the inflow time of propellants into combustion chamber and gas generator. The measurement of static pressure is not enough to observe the propellants inflow time into combustion chamber and gas generator. By measuring dynamic pressure of main flow passage of propellants the accurate propellants inflow time could be investigated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.4-10
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• 2012

Torque characteristics of a 75-tonf turbopump turbine was analyzed using the turbine performance test result. Specific torque of the subject turbine could be expressed as a linear function of corrected rotor speed at a fixed pressure ratio and it has been confirmed by the test result. It also found that corrected rotor speed-specific torque characteristics does not change anymore if the turbine pressure ratio is set bigger than a certain value. Using the turbine torque characteristics and pyro-starter performance test results, rotational speed development behavior of the turbopump was predicted. Prediction revealed that the lap time reaching 50% of turbopump design speed is less than 0.7 second. Effect of the thermal loss between pyro-starter and turbopump was negligible.