• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.11a
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• pp.165-169
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• 2007

Qualification test range for Lox/Kerosene gas generator cyle liquid rocket engine was determined by considering engine dispersion and flight inlet conditions. With various pump characteristics, the operation range of components and system was investigated through dispersion analysis. The variation of engine performance shows opposite trends in calibration and dispersion.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.34 no.4
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• pp.391-397
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• 2010

Performance tests of a liquid-oxygen pump were carried out using liquid nitrogen (LN2) as a working fluid in a cryogenic turbopump test facility in Korea Aerospace Research Institute (KARI). The tests were performed at 30-55% of the design rotational speed, and the results were compared with those from a water test. The experimental results confirmed the similarity of the hydraulic performance, which allows the prediction of the pump performance at a design rotational speed of 20,000 rpm. The overall cavitation performance of the pump in the cryogenic environment was better than that in the water environment for all ranges of flow rates and rotational speeds. Critical cavitation number at the design flow rate was determined as 0.012 from the cryogenic test, and as 0.024 from the water test. The improved cavitation performance is due to the thermodynamic effect in cryogenic fluids.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.61-68
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• 2020

7-tonf turbopump real-propellant tests in Naro Space Center were conducted and high-frequency signals from an accelerometer and pressure sensors installed on the casing and the inlet/outlet pipeline of LOX pump were analyzed to estimate the structural and hydrodynamic stabilities. Waterfalls, frequency spectrums and RMS(Root Mean Square) values of the measured signals were calculated and characteristic instability frequencies by the rotating cavitation and the rear floating ring seal(F.R.S) were investigated. Static pressures of the inlet/outlet pipeline and an acceleration of the pump casing are strongly affected on pressure fluctuation induced by the rear floating ring seal in the leakage path. Despite the acceleration RMS value seems totally small, the rotating-speed-related synchronous frequency affecting the shaft instability is distinctly observed in the frequency contour.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.10
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• pp.1038-1047
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• 2009

The development process of a gas generator for a 30-tonf pump-fed space liquid rocket engine is described. Starting from the development of an injector, followed by subscale and full-scale test specimens, the development of LOx/kerosene fuel-rich gas generator has been concluded successfully. Various analytical methods have been utilized in the course of design and the performance requirements have been verified experimentally through ignition tests, combustion performance and stability assessment tests and duration tests. The gas generator has proven its workability and stability within a defined operation window of varying chamber pressure and mixture ratio and demonstrated compliance to the performance and life time requirements.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.2
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• pp.86-93
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• 2016

Performance tests of a 75-tonf liquid rocket engine turbopump were conducted. The performance of sub-components - two pumps and a turbine - and their power matching were measured and examined firstly near the design speed under the LN2 and kerosene environment. In the real propellant - LOX and kerosene - environment tests, design and off-design performance of turbopump were fully verified in regime of the rocket engine operation. During the off-design performance tests, turbopump running time was set longer than the engine operating time to verify the pump operability and set the pump inlet pressure close to design NPSHr to investigate pump suction capability in parallel. It has been found that developed-turbopump satisfied all of the engine required performances.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.36 no.5
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• pp.563-569
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• 2012

An energy balance analysis is conducted for a 30 t thrust level liquid rocket engine. The relations between thrust and combustion pressure, between thrust and propellant flow rate, and between combustion pressure and fuel pump pressure rise are compared against those indicated by a published database of the existing rocket engines. A combustion pressure higher than the old design value is obtained, implying that the present design is high-performance oriented. The thrust to propellant flow rate ratio is the same as that of the existing engines, indicating that the specific impulse performance is at the usual level. The fuel pump pressure rise is found to be slightly high when the combustion pressure is considered, and it is attributed to the pressure budget of the present ground test engine not being optimized.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.341-345
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• 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
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• 2004.10a
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• pp.178-183
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• 2004

Thermal stratification phenomena in the liquid oxygen tank of launch vehicle is caused by heat influx from ambient and non-equilibrium heat and mass transfer in the cryogenic tank. The thermal stratification study is needed for designing vent system, tank insulation, pump inlet. In this paper by investigating buoyancy driven boundary layer flow by side wall heating, one dimensional analysis of thermal stratification is peformed. thermal gradient is described with time.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.447-450
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• 2012

An analysis of chill-down process was performed for 30 tonf Turbopump-Gas generator coupled tests. The chill-down process must be fulfilled before liquid rocket engine test using cryogenic propellant. Cavitation, damage and/or combustion instability due to bubble of propellant must be eliminated by chill-down process in a test specimen, especially cryogenic pump. The analysis of test data obtained by 30 tonf TP-GG coupled tests was performed in order to be based on the test process of KSLV-II liquid propellant rocket engine which will be developed. To macroscopically understand the process of chill-down from the viewpoint of test procedure the temperatures of important part and total time of chill-down process were analyzed.

• Journal of Aerospace System Engineering
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• v.14 no.5
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• pp.107-121
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• 2020

The performance analysis of the small satellites launch vehicle using the electric pump cycle as the upper stage engines was performed. The first stage is the launch vehicle that uses the test launch vehicle of the Korea Space Launch Vehicle II and the second stage employs elecpump cycle engine that uses liquid methane and kerosene (RP-1) as fuel. A model for the mass estimation was presented and the analysis was conducted for the range of thrust of 20 to 40 kN and combustion pressure of 3 to 6 MPa with a nozzle expansion ratio of 60 to 100. The mixture ratio with the maximum velocity increment was calculated and the performance of the LEO and SSO payload were calculated from the stage mass estimation. In both the cases, liquid methane, and RP-1 showed maximum payload for 20 kN thrust, 3 MPa combustion pressure, and the nozzle expansion ratio of 100, with a mixture ratio of 3.49 for liquid methane and 2.75 for RP-1. In addition, the ditching points of the first stage and the fairing in the LEO mission were analyzed using ASTOS.