• 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 for Aeronautical & Space Sciences
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• v.42 no.5
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• pp.437-443
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• 2014

Engine is one of the most important parts in a rocket for completing its mission successfully. In this paper, we provide a methodology for estimating reliability and development cost of a liquid rocket engine newly developed. To estimate reliability, a baseline engine is selected considering factors whose effects on reliability are unquantifiable. Then reliability of a baseline engine is adjusted to reflect the effect of factors that can be modeled quantitatively. Using the previous Transcost engine cost expressed in terms of mass and the number of hot firing tests, the engine development cost is reexpressed in reliability and thrust requirements. Finally, a numerical example is given to illustrate the application of the methodology to a turbopump rocket engine using staged combustion cycle with LOX/LH2 propellant.

• Journal of the Korean Society of Propulsion Engineers
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• v.5 no.2
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• pp.1-7
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• 2001

To protect combustion chamber from high temperature combustion gas, regenerative cooling is used for most liquid rocket engine. Although regenerative cooling is the most effective way to protect the chamber from high heat flux, realization of this system requires detail analysis, manufacturing technique and high cost. To demonstrate the possibility of applying regenerative cooling to a real rocket engine, the hot fire test has been carried out for the sub-scale liquid rocket with the water cooling system. The main purpose of the test is to identify the problem area of design, safety and cost effective manufacturing technique. The coolant passage was 3 mm in width and wall thickness was 1 mm with stainless steel. Maximum combustion time and pressure were 60 seconds and 400 psi, respectively. The flow rate of coolant was reduced gradually from 2 kg/s to 0.12 kg/s throughout firing test, combustion chamber was visually examined and no dwfect was observed.

• Journal of Mechanical Science and Technology
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• v.18 no.7
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• pp.1246-1257
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• 2004

The objective of this paper is to characterize dynamic pressure traces measured at self-excited combustion instabilities occurring in two combustion systems of different hardware. One system is a model lean premixed gas turbine combustor and the other a fullscale bipropellant liquid rocket thrust chamber. It is commonly observed in both systems that low frequency waves at around 300㎐ are first excited at the onset of combustion instabilities and after a short duration, the instability mode becomes coupled to the resonant acoustic modes of the combustion chamber, the first longitudinal mode for the lean premixed combustor and the first tangential mode for the rocket thrust chamber. Low frequency waves seem to get excited at first since flame shows the higher heat release response on the lower frequency perturbations with the smaller phase differences between heat release and pressure fluctuations. Nonlinear time series analysis of pressure traces reveals that even stable combustion might have chaotic behavior with the positive maximum Lyapunov exponent. Also, pressure fluctuations under combustion instabilities reach a limit cycle or quasi-periodic oscillations at the very similar run conditions, which manifest that a self-excited high frequency instability has strong nonlinear characteristics.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.2
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• pp.68-77
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• 2019

C/SiC composites were developed by a liquid silicon infiltration(LSI) method for use as heat-resistant parts of solid and liquid rocket propulsion engines. The heat resistance characteristics according to the composition ratio (carbon / silicon / silicon carbide) were evaluated by specimen test through arc plasma, supersonic torch test. An ablation equation for oxidation reactions was presented. Through the combustion test it was verified that various parts such as nozzle insert, exit cone and combustion chamber heat resistant parts for rocket propulsion can be manufactured and proved high ablation performance and thermal structure performance.

• Journal of the Korea Institute of Military Science and Technology
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• v.12 no.3
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• pp.396-405
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• 2009

The experimental research was conducted to setup a performance prediction logic for the regenerative cooling system on a small scale liquid rocket engine using kerosene and LOX. Total heat flux of the combustion gas side was determined for the flow rate of the coolant, combustion pressure using the calorimeter thrust chamber. Based on the experimental investigation, a performance prediction scheme for the regenerative cooling system is setup in our own way. A performance prediction logic for the regenerative cooling system has been developed by the correction scheme of the combustion gas side. The key parameters determining the temperature limitation of the coolant are the mass flow rate of the coolant and the length of the combustion chamber and the nozzle. And the parameters to control the limitation of the usable wall temperature are the number of channels and wall thickness.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.3
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• pp.19-27
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• 2002

The sequence of the propellant supply for ignition of a liquid rocket engine combustor is very important in the reliable and safe operation of the engine. The ignition sequence of KSR-III main engine was briefly described and the measuring parameters and their reliability determining ignition sequence were examined in this paper. The filling time of the engine propellant manifolds and the valve open/close time were reviewed to obtain the exact and reliable time of the propellant supply to the combustor. The combustion characteristics of the engine at starting were discussed at different supply lead of propellant. Finally, the hot firing test results with cyclogram determined by measuring parameters were presented.

• Aerospace Engineering and Technology
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• v.10 no.1
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• pp.63-69
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• 2011

In this study, we developed a dispersion analysis program of the gas-generator cycle liquid propellant rocket engine by expanding the mode analysis software(GEMAT). The performance dispersions of an engine that are arisen from the internal dispersion factors of engine's sub-components were formulated and solved to find the effects of each dispersion factor. We were also able to present the calculation method to find the required pressure margin for the compensation of those dispersion to satisfy the required performances of engine. Using this method, we could propose a novel procedure of compensating during the ground firing test which would induce the performance improvement by lessening the pumps discharge pressures or augmenting the combustion chamber pressure.

• 한국연소학회:학술대회논문집
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• 1999.10a
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• pp.209-217
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• 1999

A model for depicting the rocket engine combustion process is presented and basic experiments near a design point are provided with a FOOF type of unlike impinging injector for RP-I fuel and liquid-oxygen. The model is based on the assumption that the vaporization is the rate-controlling combustion process. The effects of initial drop size and initial drop velocity are systematically shown and discussed. It is seen that in the midst of considered parameters the change of initial drop size is more sensitive to the performance. The proposed model describes qualitative trends of combustion process well despite of its simplicity.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.4
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• pp.76-81
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• 2003

To investigate the convective heat transfer phenomena inside the Lox/Kerosene liquid rocket combustion chamber, hot fire tests were performed by using a water-cooled calorimetric chamber. The calorimetric chamber consists of one cylindrical section and nozzle section with independent cooling passage. To measure the heat flux, thermocouples were installed inlet and outlet of cooling passage of each section. The investigated range of combustion chamber pressure is from 100 psi to 300psi at fixed O/F ratio of 2.0 and radiation heat transfer from the hot gas to the surface is not considered. The measured heat flux was almost linearly depended on the chamber pressure.