• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.4
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• pp.59-66
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• 2007

In this paper, experimental investigation results of the effect of partial admission ratio on the performance of axial turbine was presented. A supersonic impulse turbine of gas generator cycle liquid rocket engine turbopump was used for the test. for experimental purpose, a nozzle block, in which total 14 number of axi-symmetric convergent-divergent nozzles are arranged circumferentially, was designed and manufactured. Partial admission ratio was controlled by changing the number of active nozzles. High pressure air was used as working medium for the test. The experimental result revealed that the performance of the supersonic impulse turbine does not much affected by the partial admission ratio for supersonic impulse turbine.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.4
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• pp.1-7
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• 2015

A cavitation venturi is a device that allows a liquid flow rate to be fixed or locked independent of a downstream pressure and has been successfully used in a liquid rocket engine system which requires a stable propellant flow rate. In the present research, four cavitation venturis which have same dimensions except for converging inlet angle and diverging outlet angle, were designed and manufactured. Flow rates through each venturi and upstream/downstream pressures were measured by changing the pressures. From the experimental data, the discharge coefficients and critical pressure ratios were calculated for each venturi. It was found that the inlet and outlet angles of the cavitation venturi affected the discharge coefficient, and the outlet angle influenced on the critical pressure ratio.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.641-644
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• 2011

The performance tests of secondary throat supersonic exhaust diffusers were carried out by using scaled down model and gas nitrogen. It was performed to find the performance characteristics according to diffuser inlet length(Ld), secondary throat length(Lst), divergence length(Ls). There was few change by diffuser inlet length(Ld), but starting pressures of the diffusers were effected by secondary throat length(Lst), divergence length(Ls). It was confirmed that starting pressure was not changed over 8 Lst/Dst.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.505-509
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• 2011

The combustion chamber and nozzle of a liquid rocket engine should be protected from the high temperature combustion gas generated by the chamber. An upper-stage nozzle extension has a large expansion ratio, therefore, The light-weight refractory materials have been used since the weight impact on the launcher performance is crucial. Gas film cooling and ablative cooling methods were used before, but were not applicable nowadays. Radiative cooling method with niobium alloy, Ni-based superalloy and ceramic based composite has been used to this day.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.74-77
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• 2011

A mixing characteristics on recess length change of Gas-centered swirl coaxial injector using high-performance staged combustion rocket engine carry out study through CFD(Computational fluid dynamics). propellant phase that combined gas-liquid simulate gas-gas. In order to measure spreading angle, velocity distribution to injector exit and spray structure of propellant analyzed. Axial velocity increase by increasing recess length, but tangential velocity decrease. The result confirmed qualitative characteristics that the spreading angle decreases.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.968-969
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• 2011

The characteristics of second throat exhaust diffuser were investigated by using CFD. Because the second throat exhaust diffuser(STED) is known as the effective device for simulating high-altitude circumstance more than a cylindrical supersonic diffuser STED was analyzed. The back pressure around nozzle was reduced by entrance size of STED and it was observed that the initial strong shock was the weak oblique shock along the diffuser. Therefore the static pressure at nozzle exit was recovered as the ambient pressure and the STED worked well.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.830-834
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• 2011

In this study, we performed three-dimensional CFD analysis to investigate the effect of the rotor blade sweep of a partial admission supersonic turbine on the stage performance and the flow field. The computations are conducted for three different sweep cases, No sweep(NSW), Backward sweep(BSW), and Forward sweep(FSW), using flow analysis program, $FLUENT^{TM}$ 6.3 Parallel. The results show that BSW model give the effect on the reducing of mass flow rates of tip leakage and the increasing of t-to-s efficiency.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.175-178
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• 2006

A numerical study for spray combustion of fluctuated fuel and oxidizer droplets injected into combustion chamber has been conducted for the analysis of spray combustion considering characteristics of injector. The 2 dimensional unsteady state flow fields have been calculated by using QUICK Scheme and SIMPLER Algorithm. As the spray model, DSF model and Euler-Lagrange Scheme have been used. The sine Auction has been used for droplet fluctuation model of fuel and oxidizer, while the coupling effects of the droplets between gas phase and evaporated vapor have been calculated by using PSIC model.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.216-220
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• 2006

In this study, it was investigated that the characteristics of startup and compatibility using several type hot and cold gases. The characteristics of starting LRE by pyro starter was compared with that by a He spinner. The compatibility of pyre gas, a gaseous He, H2+N2 mixture gas, and air was investigated by a simple 1D turbine analysis considered the properties of each gases and turbine efficiency. Most of them were compatible to start up the LRE but air was properly used only when the turbine was low power mode.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.121-124
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• 2008

This paper presents experimental results of dynamic characteristics of fuel-rich gas generators. Pressure fluctuation measurements in the chamber and manifolds have been analyzed. Gas-generator-alone tests revealed stable combustion regardless of a chamber pressure but low-frequency combustion instabilities occurred for cases of turbine-manifold tests at chamber pressure conditions below 50 bar. The instabilities are considered as an axial resonant mode and acoustic intensity increases along with a chamber pressure.