• International Journal of Aerospace System Engineering
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• v.6 no.2
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• pp.11-16
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• 2019

A system analysis program has been developed for a gas generator cycle liquid rocket engine of 30 ton class. Numerical models have been proposed for a combustor, a turbopump, a gas generator and pressure drop through a regenerative cooling system. Numerical algorithm has been validated by comparing with the published data of MC-1. The major source of error is not the numerical algorithm but the imperfect performance models of subsystems. So the precision of the program can be improved by revising the performance models using experimental data. The sea level specific impulse and vacuum specific impulse have been demonstrated for a 30 ton class gas generator engine. The optimal condition of combustor pressure and mixture ratio for specific impulse which is a typical characteristic of a gas generator cycle engine has been illustrated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.38-41
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• 2009

For the technology development of LOx/kerosene liquid rocket engine, turbopump-gas generator closed-loop coupled tests using 30tonf main engine components such as turbopump and gas generator except combustion chamber are performed. In the engine system operation environment, simulating combustion chamber by flow control units, the chill-down procedure, startup characteristics, nominal operability and smooth shutdown of turbopump+gas generator closed-loop coupled Test Plant are successfully confirmed. The serviceability of the turbopump and gas generator are evaluated. The feed-back control system for the turbopump rotational speed and gas generator mixture ratio are also verified. The results of closed-loop coupled test will be used as the technology development for the liquid rocket engine.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.2
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• pp.73-82
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• 2020

Additively manufactured, small rocket engines are perhaps the focal activities of space startups that are developing low-cost launch vehicles. Rocket engine companies such as SpaceX and Rocket Lab in the United States, Ariane Group in Europe, and IHI in Japan have already adopted the additive manufacturing process in building key components of their rocket engines. In this paper on technology trends, an existing valve housing of a rocket engine is chosen as a case study to examine the feasibility of using additively manufactured parts for rocket engines.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.22-29
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• 2011

Critical speed of high thrust liquid rocket engine turbopump is obtained through a rotordynamic analysis and a unloaded turbopump test is peformed for validation of the numerical model. The first critical speed predicted by the numerical analysis is correlated well with the test result for the bearing unloaded rotor condition only considering mass unbalance load. Using the previous rotordynamic model, critical speed variation is estimated as a function of varied bearing stiffness due to pump and turbine radial loads with relative angle difference. From the numerical analysis, it is found that the relative angle difference of pump and turbine radial loads greatly affects the critical speed. However, additional axial load reduces the effect derived from the relative angle difference of radial loads.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.4
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• pp.42-49
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• 2012

Critical speed of high thrust liquid rocket engine turbopump is obtained through a rotordynamic analysis and a unloaded turbopump test is peformed for validation of the numerical model. The first critical speed predicted by the numerical analysis is correlated well with the test result for the bearing unloaded rotor condition only considering mass unbalance load. Using the previous rotordynamic model, critical speed variation is estimated as a function of varied bearing stiffness due to pump and turbine radial loads with relative angle difference. From the numerical analysis, it is found that the relative angle difference of pump and turbine radial loads greatly affects the critical speed. However, additional axial load reduces the effect derived from the relative angle difference of radial loads.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.8
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• pp.805-811
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• 2009

An oxidizer pump of a turbopump for a 30-ton class gas generator cycle engine was tested in the medium of liquid oxygen. The turbine was driven by cold hydrogen gas in the test. The oxidizer pump was operated stably at both design and off-design conditions, satisfying the performance requirements. The pump head coefficient from the liquid oxygen test was 2~3% lower than that from the water test. The power required to run the oxidizer pump was well balanced with the power produced by the turbine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.409-412
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• 2009

A 75ton thrust turbopump system for liquid rocket engine was analyzed thermally and mechanically. A 2D axisymmetric model of the turbopump assembly was created. In the analysis operation cycle including chill-down, operation and post operation steps were considered. Appropriate heat transfer conditions for each step were modeled and applied. Transient temperature distribution was calculated, consequent mechanical analysis was conducted to predict stress and deformation. Effects of external heat insulators and heat dissipation at the bearings were considered in the heat transfer analysis.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.05a
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• pp.11-17
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• 2009

Hot-fire test of a turbopump for a gas generator cycle rocket engine of 30 ton class was carried out in real propellant environment. Liquid oxygen and kerosene were used for the oxidizer pump and the fuel pump, respectively, while hot gas produced by the gas generator was supplied to the turbine. A part of the propellant discharged from the pumps was provided to the gas generator. The turbopump was run stably at both on-design and off-design conditions, satisfying all the performance requirements. This paper describes one of the test cases, where the turbopump was run for 120 seconds at three different operating modes in one test. In terms of performance characteristics of pumps and turbine, the results from turbopump assembly test using real propellant showed a good agreement with those from the turbopump component tests using simulant working fluid.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.1
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• pp.64-70
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• 2007

A test facility was developed where an inducer for a liquid rocket engine turbopump can be tested using liquid nitrogen as a working fluid. At the facility, a hydrodynamic performance test and a cavitation performance test for an oxidizer turbopump were carried out. Head-flow relation at liquid nitrogen test was similar to the case at water test. However, cavitation performance at the liquid nitrogen was superior to the case at water test, which results from the thermodynamic effect of cavitation.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.93-99
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• 2006

A test facility was developed where an inducer for a liquid rocket engine turbopump can be tested using liquid nitrogen as a working fluid. At the facility, a hydrodynamic performance test and a cavitation performance test for an oxidizer turbopump were carried out. Head-flow relation at liquid nitrogen test was similar to the case at water test. However, cavitation performance at the liquid nitrogen was superior to the case at water test, which results from the thermodynamic effect of cavitation.