• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.3
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• pp.25-37
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• 2005

For pump-fed rocket propulsion system, the temperature of LOX to be supplied to turbopump inlet should be satisfied with pump inlet temperature requirement during all operating stages, as excessive temperatures can result in cavitation due to reduction in NPSH, thus either damaging the pump or adversely affecting pump performance rise. So exact estimation of LOX temperature rise is absolutely needed for developing reliable propulsion system. This paper presents systematic analysis scheme for estimating inner process of cryogenic propellant tank which is needed for LOX temperature rise. And this paper presents LOX temperature rise and thermal stratification for all rocket operating stages including cooling, filling, waiting, pre-pressurization and firing, with the application of buoyancy driven boundary layer theory.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.194-197
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• 2005

Pulse-mode performance evaluation is made for a set of monopropellant hydrazine thrusters producing $0.95 lb_{f}$ of nominal steady-state thrust at an inlet pressure of 350 psia. With a brief description on the hot-firing test matrix, a typical data obtained from pulse-mode firing is given directly showing the variational behavior of propellant supply pressure, vacuum condition, and thrust, in addition to the thermal response of the thruster. The performance features are successfully compared to the reference criteria of 1-lbf standard monopropellant rocket engine.

• 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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• 2006.11a
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• pp.122-126
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• 2006

Two types of pressurization system and low weight feeding piping system are developed. With sub-system tests, ullage pressure control performance was verified for 1 step and 2 step pressurization system and the feeding performance of feeding piping system was also verified. The weight of the feeding piping system is low enough for the application of launch vehicle. In addition, LOX conditioning system is developed for avoiding geysering and LOX temperature rise. Integrated performance was verified through integrated on-board feeding system performance tests.

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.3
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• pp.16-23
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• 2008

Flow analyses have been performed to investigate the uniformity of propellant flow through the oxidizer manifold of a 30 tonf full-scaled combustor. Injectors were simulated as porous medium layers of equivalent pressure drops. The uniformity of oxidizer propellant has been analyzed for various diameters of holes in vertical/horizontal distributors and configurations of oxidizer inlet to propose an improved design solution. It has been proven that the mass flow uniformity were improved by adjusting the holes in vertical/horizontal distributors.

• Journal of the Korean Society of Propulsion Engineers
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• v.13 no.6
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• pp.8-16
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• 2009

The development of high-pressure sub-scale combustion chambers is described. A total of four high-pressure sub-scale combustion chambers having either a detachable structure of the mixing head and the chamber or a single welded regenerative cooling structure have been developed. The sub-scale combustion chambers have a chamber pressure of 70 bar and propellant mass flow rate of 5.1~9.1 kg/s. The propellant mass flow rate and the recess number of the injector were changed for the improvement of combustion performance and they were validated through hot firing tests. The design and manufacturing techniques of regenerative cooling channel and film cooling to be applied to the full-scale combustion chamber were adopted through the present development and verified.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.10
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• pp.56-60
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• 2006

It is essential to predict thermodynamic properties of combustion gas with respect to a propellant mixture ratio for the development of a gas generator for a liquid rocket engine. The present study shows the temperature measurement of exit combustion gas as a function of a mixture ratio through the series of combustion tests of a fuel-rich gas generator with liquid oxygen and Jet A-1. The measurements of dynamic and static pressures, and combustion gas temperatures allowed the estimation of thermodynamic properties like a specific heat ratio, a gas constant, and a constant pressure specific heat of the combustion gas. The comparison of the experimental results with predictions made by interpolation parameters obtained from the modification of the chemical equilibrium code indicates that the interpolation method calibrated using the temperature measurements can be utilized as an effective tool for the initial design of a fuel-rich gas generator.

• 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.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.363-367
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• 2009

For the liquid rocket propulsion system using liquid oxygen as oxidizer, helium for pressurizing LOX is usually stored in the LOX tank with cryogenic temperature. For that kind of pressurizing system, cryogenic helium is discharged from the immerged pressurant cylinder and passes through the heat exchanger downstream of gas generator. During the process, helium pressurant is heated from cryogenic temperature to high one and supplied to the ullage of propellant tank. To develop the pressurizing system, a cryogenic heating apparatus is needed to simulate the heat exchanger. In this paper, the cryogenic heating apparatus for development of the pressurization system is presented along with its heating test results with cryogenic helium.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.1
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• pp.63-68
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• 2011

For the liquid rocket propulsion system using liquid oxygen as oxidizer, helium for pressurizing LOX is usually stored in the LOX tank with cryogenic temperature. For that kind of pressurizing system, cryogenic helium is discharged from the immerged pressurant cylinder and passes through the heat exchanger downstream of gas generator. During the process, helium pressurant is heated from cryogenic temperature to high one and supplied to the ullage of propellant tank. To develop the pressurizing system, a cryogenic heating apparatus is needed to simulate the heat exchanger. In this paper, the cryogenic heating apparatus for development of the pressurization system is presented along with its heating test results with cryogenic helium.