• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.2
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• pp.39-45
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• 2006

In liquid rocket engine turbopump, it is difficult to evaluate turbine performance for high pressure, high temperature circumstance. Turbine test is often done by using air at similarity condition so that the turbine can be tested at lower risk. This paper describes an air similarity test program of liquid rocket engine turbopump turbine. A test facility has been built to evaluate aerodynamic performance of turbines. The test facility consists of high pressure air supply system, mass flow rate measuring nozzle, test section, hydraulic break, exit orifice for pressure control, instrumentation and control system. This paper also presents how to decide the similarity conditions of the turbine test and describes how to control test conditions. Relative standard deviation of measurement parameter was less than 1% and measured turbine efficiency corresponded with analysis result within 2%.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.43-46
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• 2006

Storage air heater(SAH) is a general purpose facility that is used to simulate the high altitude condition of supersonic ground test facility, thurst compensation test of rocket engine nozzle and gas turbine engine combustor test. SAH in KARI is built to simulate the total temperature of the supersonic ground test facility which has a wide flight envelope from altitude 0km, Mach 2 to altitude 25km, Mach 5 and operates up to 1300K, 3.5MPa. In this paper, we introduces the SAH in JAXA which is model of SAH in KARI and summarizes the design process and manufacture of ours.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.7-11
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• 2011

A velocity-compounded(VC) turbine for the 75ton turbopump was developed as an improved performance backup for the single-rotor baseline turbine. Curvic coupling was adopted for the power transmission between the rotors and shaft. High temperature torsion test and spin test was performed for the curvic coupling design validation. Aerodynamic performance test revealed that VC turbine can generate 20.5% higher specific power than the baseline turbine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.353-358
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• 2009

For the successful development of 75 tonf class liquid rocket engine, a plenty of tests on each engine component has to be performed and this is equally true for a combustor. However the test facility which is in operation at Korea Aerospace Research Institute lacks its capacity to perform fire tests of a 75 tonf class combustor at its nominal thrust. Since the test facility has to be ready prior to the start of development tests, it is very urgent to establish the test facility. The preliminary design of a test facility for a 75 tonf class combustor which was performed according to the urgent necessity is described in the paper.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.4
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• pp.1-11
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• 2017

TPTF (Turbopump Real Propellant Test Facility) at Naro Space Center has used alcohol burner system to simulate the gas flow of gas generator of liquid rocket engine. During the test at TPTF, the temperature and pressure at turbine inlet were smoothly increased while those of the gas generator of engine were constant. Present research developed a simulation code for the burner and the piping system and applied to the system. The calculation results were in good agreement with the test, and confirmed quantitatively that the non-steadiness is due to the heat transfer of the pipe. While the insulation of the pipe is ineffective, the length has a large impact on the turbine inlet condition. The present research clarified the empirically estimation of test condition, and can be applied to determination of the following test conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.228-231
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• 2011

Burning rate of a propellent is an essential factor when designing a propulsion system. In order to come up with burning rate, first we need to design and build propellent grain to get neutral pressure curve. Then check the pressure with ground test and calculate the burning rate using burning rate equation. This burning rate is then compared to the burning rate of a propellent which was resulted from making a standardized specimen and combusting it using a strand burner. An accurate burning rate is calculated after comparing those two burning rates. For this study, compact propulsion system was designed, produced, tested and analyzed in order to get burning rates, an essential factor in propulsion system design, in an effective way.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.4
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• pp.38-43
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• 1999

An experimental study was carried out, in order to set up the procedure for evaluation of hot fire test, to investigate the effect of mixture on combustion performance and combustion stability , and to determine the optimum design condition for designing the liquid rocket engine. $HNO_3$/Kerosene uni-element liquid rocket engine(thrust 24 $\iota{b}_f$, chamber pressure 200 psia) using impinging streams doublet injector was designed, and ground hot-fire test was carried out. To prevent or reduce the hard start during ignition period, two step ignition method was used. This was accomplished by maintaining about 25% of the designed operating pressure doting transient period, then chamber pressure was built up to the designed operating pressure. Maximum combustion efficiency was at O/F ratio 3.6, and combustion efficiency is decreased with increasing momentum ratio.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.4
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• pp.46-51
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• 2007

The irregular fuel surface was observed by the visualization of hybrid rocket combustion. Even though the test condition maintained oxidizer rich environment, the irregular dark fuel surface was formed as the result of incomplete combustion. In order to investigate the correlation of the characteristics of oxidizer flow and the irregular fuel surface, various flow conditions were imposed such as swirl flow, induced swirl flow by helical fuel configuration and straight flow. Test results revealed no correlation was found between oxidizer flow condition and irregular fuel surface. And this can be a commonly observed phenomena in the tests with different fuel/oxidizer combination. Thus, the irregular fuel surface can be a result of the interaction of blowing flow of vaporized fuel and the boundary layer of oxidizer flow. A further study will be required to confirm the assumption for the formation of irregular fuel surface.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.45-49
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• 2011

The blow-down oxidizer feed system with self-pressurizing $N_2O$ has more advantages than the regulated system. However, it is difficult to predict the exhaust flow rate because there exist two phases in the $N_2O$ tank - liquid phase and gas phase, and the properties of $N_2O$ in storage tank are varied continuously during blow-down. In this paper, a method that can analyse simply the blow-down oxidizer feed system is studied. The properties of saturated $N_2O$ are found from the NIST data base, and mass flow through the orifice is modeled as NHNE. Cold flow test with hybrid rocket combustor is performed for the comparison where the results should found from the good agreement.

• 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.