• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.135-138
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• 2006

We developed an engine system design program by balancing the pressure-mass-power relation which can be acquired from each component's specification. In gas generator type open-cycle rocket engine system it is possible to distinguish the variables into two categories, which are input variables and requirement variables. We define 11 design variables corresponding to the 11 balance equations as functions of pressure, mass and power of target engine system. We solved these equations by Newton method. As an example we designed gas generator cycle engine system and finally we could conclude that this developed program is well suited to the engine system design.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.5
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• pp.18-25
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• 2008

A performance analysis program has been developed for the gas generator cycle liquid rocket engine. This program predicts the system performance with the performances of subsystems which are evaluated by the models based on another analyses or experiments. The analysis method has been validated by comparing the engine performance against the published conceptual design. The performance models of the subsystems have been verified to give reasonable results by comparing with the MC-1 engine design and the system analysis of 10 ton thrust engine. The system performance of the 30 ton thrust rocket engine using LOx/Jet-A1 has been presented as an application example.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.366-370
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• 2008

Mode analysis is very important for the development of liquid rocket engine in various applications. We developed a mode analysis program for the gas-generator cycle liquid rocket engine by proposing 13 independent equations with 13 independent variables which can be solved by Newton method. As an example we calculated the change of engine operating mode according to the control valve's loss coefficient change located in the gas-generator oxidizer supply line. And we concluded that this program can give basic idea for the mode analysis of gas-generator cycle liquid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.61-64
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• 2017

The gas-generator open cycle is adapted for liquid rocket engine of Korea Space Launch Vehicle-II. The combustion instability can interfere with combustion performance and cause a noise and vibration or carry the potential for serious damage. This study introduces the experience cases of combustion instability in development of the gas generator for liquid rocket engine.

• Journal of the Korean Institute of Gas
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• v.3 no.1
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• pp.14-20
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• 1999

This study describes the results of Coefficient of Performance(COP) analysis by cycle simulation for two types of absorption-compression hybrid cycle using the Water/Lithium Bromide solution pair. These types are basic hybrid systems introducing a mechanical compression process into the refrigerant vapor phase of the single effect absorption cycle. In absorption-compression hybrid cycles, coefficient of performance is improved compared with absorption cycle. Hybrid cycle Type 2 is considered as a key technology to support energy utilization system, given its capability of utilizing waste heat to drive system with a high level of efficiency.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.34 no.4
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• pp.443-448
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• 2010

In this paper, the changes in performance parameters, e.g., the combustor pressure, turbine power, engine mixture ratio, temperature of gas generator, and product gas, of a liquid rocket engine employing gas generator cycle with the variations in propellant-supply pressure have been described. Engine performance is numerically calculated using the 13 major system-level variables of the rocket engine. The combustor pressure and turbine power increase with an increase in the oxidizer-supply pressure and decrease with an increase in fuel-supply pressure. The lower mixture ratio of gas generator for increased fuel mass flow rate decreases the gas generator gas temperature and deteriorates the gas material properties as the turbine working fluid. The turbine power decreases with an increase in fuel-supply pressure; this results in a decrease in the main-combustor pressure, which is directly proportional to engine thrust.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.191-195
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• 2006

It is definitely required to control dispersion of the rocket engine performance in order to accomplish the mission of a launch vehicle successfully. A performance dispersion analysis was conducted for a gas generator cycle liquid rocket engine and the required pressure drops were estimated for engine tunning. As a result, the vacuum thrust dispersion of the engine was from +9.1% to -8.7% and the mixture ratio deviated from +9.7% to -9.6% from the nominal value due to the errors of components and the engine inlet condition of propellants. The required pressure drop in the LOx line to the combustor is higher than in the fuel line for same mixture ratio change.

• Aerospace Engineering and Technology
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• v.6 no.1
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• pp.120-127
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• 2007

Performance dispersion in the engine should be considered to predict the flight accuracy of a launch vehicle. A dispersion estimation method was presented with a LOx/Kerosene gas generator cycle engine. The orifices in the propellant supply lines in the engine were assumed to be used for calibration of the performance and the required pressure drops were acquired. The dispersions after calibration were quantified also.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.6
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• pp.91-97
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• 2011

A test range for a 75tf class gas generator cycle liquid propellant rocket engine is defined. The engine system test range is defined by the performance variation during flight, the dispersion after engine calibration, and additional margin. The component development test range includes the operation range corresponding to the engine system test range and the component performance margin.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.51-56
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• 2011

A test range for a 75tf class gas generator cycle liquid propellant rocket engine is defined. The engine system test range is defined by the performance variation during flight, the dispersion after engine calibration, and additional margin. The component development test range includes the operation range corresponding to the engine system test range and the component performance margin.