• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.1-9
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• 2020

Using kerosene and liquid oxygen, 1.5-tonf class liquid-liquid pintle injector with rectangular two-row orifice was designed and manufactured. The combustion test of the pintle injector was carried out to verify the combustion performance and combustion stability under a supercritical condition which is the actual operation condition of the liquid rocket engine. The combustion test result showed that the pintle tip was damaged by the high temperature combustion gas in the high-mixed ratio recirculation zone of the combustion chamber. To solve this problem, the insert nozzle was installed in the pintle injector to increase cooling performance at the pintle tip. As a result of the hot firing test, installation of the insert nozzle, AR and BF had a great effect on pintle tip cooling performance.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.4
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• pp.81-88
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• 2013

It was designed and tested at the design point that an oxidizer rich preburner for a staged combustion liquid rocket engine propelled by kerosene and LOx. The oxidizer rich preburner was designed as some of LOx injected from the mixing head was burned with kerosene and the rest of LOx injected from injection holes in the regenerative cooling chamber was vaporized by combustion gas. The preburner is operated at OF ratio of 60 and combustion pressure of 20 MPa. The Preburner has a honey-comb type mixing head with simplex swirl injectors, a turbulence ring improving combustion stability and uniformity of product gas temperature distribution, and a nozzle simulating the duct. With the combustion test results at the design point, the oxidizer rich preburner showed high combustion stability and uniformity of product gas temperature distribution.

• Aerospace Engineering and Technology
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• v.3 no.2
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• pp.81-90
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• 2004

This paper describes the results of combustion performance test of fuel-rich gas generator(GG) using LOx and kerosene as propellant at off-design conditions. The chamber pressure is thought to be a function of O/F ratio and total propellant mass flow rate. The test shows that the spatial temperature deviation at the exit of gas generator remains within 7.5K and that the average gas temperature at the exit is a function of propellant O/F ratio. The results of firing test of gas generator at off-design conditions, especially the relation between gas temperature and O/F ratio, can provide useful data for the design of future gas generator and for the development of low-O/F ratio reaction analysis code.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2017.05a
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• pp.374-377
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• 2017

A cold-flow test equipment was designed to carry out the performance verification of TDACS. For that purpose, the pressure dynamics in the solid rocket motor combustor and the cold-flow test was modeled, and the response time showing the dynamic characteristics of each model was obtained. In this paper, the system response time of the cold-flow test was designed to be equal to that of the motor, making the dynamic response in cold-flow and hot gas condition to be similar.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.15 no.6
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• pp.3403-3409
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• 2014

Hot-firing tests were carried out using a mixing head with 19 swirl coaxial injectors and a combustion chamber with internal cooling channels. The propellants of liquid oxygen and kerosene(Jet A-1) were burned in a range of chamber pressures (59~82 bar) and mixture ratios (2.0~3.0). The temperature of water used as the cooling fluid was measured at the inlet and outlet of the cooling channels, and the heat flux was calculated. The aim of this study was to examine the effect of combustion instability on heat transfer in a subscale thrust chamber, and detect the temperature variation of cooling water. During several hot-firing tests, combustion instability was encountered which caused a 5~20% increase in heat flux. The peak heat flux took place in the initial stages of combustion instability.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.5
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• pp.72-77
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• 2007

The combustion test results of the sub-scale combustor having dual swirl injector with internal mixing for a liquid rocket engine are described. The sub-scale combustor uses liquid oxygen(LOx) and kerosene as propellants and has an injector head, an ablative material combustor wall and a water cooled nozzle. The injector head has LOx manifold, fuel manifold, fire face plate, one center swirl injector and 18 main swirl injectors of internal mixing. The combustion tests were successfully performed at design and off-design points without any damages on the injectors. Combustion characteristics velocity of 1756m/s was measured at design point. High frequency combustion instability was not observed but low frequency pulsations occurred at off-design conditions.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.193-198
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• 2010

In solid propellant rocket motors, $Al_2O_3$, one of combustion products, can be accumulated inside a combustion chamber. A special rocket motor was designed and tested to simulate thermal reaction of rubber insulator affected by the deposited slag. We successfully demonstrated through a dynamic radioscopy that the slag was deposited at the location as designed. In this paper we present a new test method which can simulate a high temperature and pressure environment in combustion chamber to evaluate material characteristics of rubber insulator and can provide design data to decide its thickness for a new solid rocket motor. The solid rocket motor, which has an average chamber pressure of 770 psia and a burning time of 50 seconds, was tested. The results show that erosion of EPDM insulator is more affected by a gas velocity rather than by the thermal reaction of slag with a high thermal capacity.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.169-174
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• 2006

This paper is related to a design of high pressure sub-scale combustor with regenerative reeling. As a previous step for the evaluation of thermal heat flux, a similar combustor with cooling water was manufactured. Design conditions with high combustion efficiency and cooling performance were verified through the hot firing tests of the water-cooled high pressure combustor. Finally the regeneratively cooled high pressure combustor has been designed based on these data. After manufacturing it, its practical utility will be tested and verified through hot firing tests.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.1
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• pp.9-16
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• 2020

A performance evaluation of the 200 N-class GCH₄/LOx small rocket engine was performed through ground hot-firing test. As a result, the combustion pressure and thrust raised with the increase of the oxidizer supply pressure, and thus the specific impulse, characteristic velocity, and their efficiency increased. The characteristic velocity was measured at about 90% performance efficiency. The change of chamber aspect ratio did not affect the performance of the rocket engine in the test condition specified. In addition, uncertainty evaluation was conducted to ensure the reliability of the test results.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.5
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• pp.77-83
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• 2020

The operating characteristics of a main oxidizer shut-off valve have been examined by life-cycle tests in which the valve operates as many times as is required at room and cryogenic temperatures. It is shown that there is no significant change of the operating characteristics in the course of the life-cycle tests. The critical pressures and principal forces, which represent the operating characteristics, differ at each operating time by less than about 3%. These qualified operating characteristics of the valve are expected to play an indispensible role in the stable startup and shutdown of a liquid rocket engine.