• International Journal of Aeronautical and Space Sciences
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• v.18 no.4
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• pp.788-796
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• 2017

This paper proposes a new measurement method to improve the shortcomings of an existing integral method for measuring heat flux in plug-type heat flux gauges in the high-temperature and high-pressure environments of liquid-rocket combustors. Using the existing integral measurement method, the calculation of the surface area for the heat flux in the gauge exhibits error in relation to the actual surface area. To solve this problem, transient profiles obtained from ANSYS Fluent were used to calculate unsteady heat flux as it adjusted to the measured temperature. First, a heat flux gauge was designed and manufactured specifically for use in the high-temperature and high-pressure conditions that are similar to those of liquid rocket combustors. A calibration test was performed to prove the reliability of the manufactured gauge. Then, a combustion experiment was conducted, in which the gauge was used to measure unsteady heat flux in a liquid rocket combustor that used kerosene and liquid oxygen as propellants. Reasonable heat flux values were obtained using the gauge. Therefore, the proposed measurement method is considered to offer significant improvement over the existing integral method.

• Journal of Mechanical Science and Technology
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• v.20 no.6
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• pp.896-904
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• 2006

In a liquid rocket engine, the role of gas-liquid scheme injector as an acoustic resonator or absorber is studied experimentally for combustion stability by adopting linear acoustic test. The acoustic-pressure signals or responses from the chamber are monitored by acoustic amplitude. Acoustic behavior in a rocket combustor with a single injector is investigated and the acoustic-damping effect of the injector is evaluated for cold condition by the quantitative parameter of damping factor as a function of injector length. From the experimental data, it is found that the injector can play a significant role in acoustic damping when it is tuned finely. The optimum tuning-length of the injector to maximize the damping capacity is near half of a full wavelength of the first longitudinal overtone mode traveling in the injector with the acoustic frequency intended for damping in the chamber. When the injector has large diameter, the phenomenon of the mode split is observed near the optimum injector length and thereby, the acoustic-damping effect of the tuned injectors can be degraded.

• Aerospace Engineering and Technology
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• v.7 no.1
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• pp.129-135
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• 2008

The 30 tonf-class liquid rocket combustor currently being developed is designed to connect the fuel feeding line at a middle position of the supersonic nozzle in order to reduce both pressure loss in the regenerative cooling passage and overall envelope of the thrust chamber in spite of increase in design complexity. To verify the design of cooling passages including fuel ring, connecting holes, two-directional cooling channels and collectors, numerical analysis has been performed.

• International Journal of Aerospace System Engineering
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• v.6 no.2
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• pp.11-16
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• 2019

A system analysis program has been developed for a gas generator cycle liquid rocket engine of 30 ton class. Numerical models have been proposed for a combustor, a turbopump, a gas generator and pressure drop through a regenerative cooling system. Numerical algorithm has been validated by comparing with the published data of MC-1. The major source of error is not the numerical algorithm but the imperfect performance models of subsystems. So the precision of the program can be improved by revising the performance models using experimental data. The sea level specific impulse and vacuum specific impulse have been demonstrated for a 30 ton class gas generator engine. The optimal condition of combustor pressure and mixture ratio for specific impulse which is a typical characteristic of a gas generator cycle engine has been illustrated.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.220-223
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• 2005

System design parameters of gas generator cycle liquid rocket engines were investigated and compared in the present study. Characteristic velocity of combustor, pressure drop of combustor injector, exit pressure of pump, pump efficiency and specific power of turbine were considered as a system design parameter. The result shows the characteristic velocity is in the range of 1700-1770 m/s, pressure drop of combustor injector, 4-10 bar, pump exit pressure ratio to combustion pressure, 120-230%, pump efficiency, 60-80%, specific power of turbine, $0.28-0.58MW{\cdot}s/kg$.

• Aerospace Engineering and Technology
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• v.8 no.1
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• pp.138-143
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• 2009

Thermal barrier efficiency and endurance of coatings in liquid rocket engine combustor were evaluated for air plasma spray coating and electro/electroless plating. The result of firing tests has revealed occasional occurrence of local delamination of $ZrO_2$, NiCrAlY coating obtained by the method of air plasma spray in the region of supersonic flow and it necessitated a new coating method as a substitution. It was found that Ni-Cr coating by means of electro/electroless plating can substitute $ZrO_2$, NiCrAlY coatings of air plasma spray in terms of thermal barrier efficiency and endurance.

• Journal of the Korean Society of Propulsion Engineers
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• v.11 no.3
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• pp.43-49
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• 2007

A ground firing test facility for Liquid Rocket Engine(LRE) combustor was built in Korea Aerospace Research Institute(KARI) in 2001 to support the development of the first Korean LRE for the KSR-III. About 170 tests were conducted up to date since its establishment and in the meantime a considerable improvements were made in the facility capability. This paper describes the outline, conducted tests and operation techniques which have been acquired through the operation of the test facility.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2007.04a
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• pp.128-134
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• 2007

Combustion performance and characteristics of high-pressure subscale liquid rocket combustors were studied experimentally. Four different models of combustor were considered in this paper. The high-pressure subscale combustor is composed of the mixing head, the water cooling cylinder and the nozzle. One model of the combustors employed regenerative cooling combustor in that the kerosene used for the chamber cooling is burned. This combustor was damaged due to a high frequency combustion instability occurred during a firing test. The results of the firing tests, comparison of performance, and characteristics of static and dynamic pressures of the combustors are described.

• 한국연소학회:학술대회논문집
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• 2000.12a
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• pp.92-100
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• 2000

One method to analyse acoustic modes is proposed to predict the characteristics of acoustic instability in liquid rocket engine. It is based on the similarity between transverse acoustic modes and adopts two-dimensional axisymmetric geometry. Using this method, the first tangential mode in the prototype combustor can be analysed through the analysis of the first radial mode in the model combustor with doubled chamber diameter. Sample numerical calculation is demonstrated applying this method to sample rocket engine and thereby acoustic instabilities of the engine are investigated. The present results show a good agreement with the previous findings. The numerical analysis based on the proposed method is cost-effective and serves as the first approximation to the true solution.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.163-167
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• 2008

This paper presents pressure fluctuation characteristics of a 30 ton-f class liquid rocket engine combustor. Combustion stability of the combustor was evaluated using the results 46 firing tests performed with a varying O/F ratio and chamber pressure. The RMS value of pressure fluctuation during the steady state combustion was less than 2.6% of the static chamber pressure, demonstrating static stability of the combustion phenomenon. The decay time of pressure fluctuation caused by forced disturbance of a pulse gun was found to be less than 3.5 msec verifying dynamic stability of the combustor.