• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.196-200
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• 2005

A theoretical model for the calculation of chemical equilibrium composition of propellant combustion product is briefly presented for the performance analysis of monopropellant hydrazine rocket engine. Analysis result is compared to that of test and evaluation of 1-lbf class thruster and is scrutinized primarily from the view point of ammonia dissociation fraction. Chemical equilibrium composition and average molecular weight is additionally depicted according to the variation of propellant inlet pressures and the varying nozzle area ratio. The theoretical analysis is tried as a way of derivation of design parameters for mid- and large-thrust class of monopropellant rocket engines.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.3
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• pp.1-7
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• 2014

A set of preliminary design parameters for the bipropellant rocket engine using liquid methane-fuel as green propellant were derived through a theoretical performance analysis. Chemical equilibrium analysis utilizing CEA was conducted for the prediction of combustion performance: combustion characteristics according to the O/F ratio and chamber pressure variation were investigated. For a determination of chamber-characteristic length, the vaporization time of fuel-droplet with various performance parameters was calculated by applying Spalding's 1-D droplet vaporization model. Finally, the preliminary design specification of methane-bipropellant rocket engine, which is to be performance-tested under the ground firing condition, was proposed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 1998.10a
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• pp.1-1
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• 1998

로켓엔진 추력발생용으로 광범위하게 사용되는 액체추진제는 고성능, 대용량의 액체추진제 로켓엔진에서는 필연적으로 고주파 연소불안정의 문제를 수반하며, 이 연소불안정의 정도는 연소성능과 더불어 엔진개발의 성패를 좌우하는 중요한 여건이 된다. 따라서 안정한 로켓의 비행을 보장하기 위해서는 연소불안정의 문제가 선결되어야 한다. 연소불안정의 기본 메커니즘은 연소기에서 발생하는 압력섭동에 반응하여 불안정한 음향에너지를 되먹임하는 연소과정으로 설명된다. 연소불안정 현상이 발견된 이후 실험 및 이론적 접근에 의해 이와 같은 연소불안정 메커니즘 및 예측에 대한 체계적인 연구가 이루어져 왔으며, 현재까지의 다양한 고주파 연소불안정 예측방법 중에서 음향 및 기화 응답함수를 이용하는 방법은 직관적 고찰에 의존하는 단순한 연소모델을 적용하며 주로 음향적 섭동에 의한 연소의 반응을 연소안정성 평가의 파라메터로 사용한다. 이와 같은 음향적인 예측방법은 연소불안정 현상을 이론적으로 전개하므로 경제적으로 각종 설계변수에 대한 연소불안정의 변화를 구분할 수 있는 장점이 있어 성능 및 호환설계와 병행하여 로켓엔진 연소실의 초기 안정성 설계방법으로 주로 사용된다.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.3
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• pp.47-53
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• 2015

Design parameters required for Methane/oxygen bipropellant small-rocket-engine were derived through a theoretical performance analysis. The theoretical performance of the rocket engine was analyzed by using CEA and optimal propellant mixture ratio, characteristic length, and optimal expansion ratio were calculated by assuming chemical equilibrium. A coaxial-type swirl injector was chosen because of its outstanding atomization performance and high combustion efficiency compared to other types of injector and also a bell nozzle with 80% of its full length was designed. The rocket engine configuration with 1.72 MPa of chamber pressure, 0.18 kg/s in total propellant mass flow, and O/F ratio of 2.7 was proposed as a ground-firing test model.

• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.6
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• pp.85-92
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• 2020

Ground hot-firing tests were conducted to analyze the combustion performance according to the characteristic lengths 1.37 m, 1.71 m, and 2.06 m of the combustion chamber in 200 N-class GCH4-LOx small rocket engine. Thrust, specific impulse, and characteristic velocity at the steady-state could be obtained as the key performance parameters of the rocket engine. The performance characteristics acquired through the test were compared and analyzed with the theoretical performance calculated from CEA analysis. Observation of the influence of characteristic length on the combustion performance indicates that an optimal characteristic length shall remain between 1.71 m and 2.06 m.

• Journal of the Korean Society of Propulsion Engineers
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• v.8 no.2
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• pp.73-84
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• 2004

In general, static firing test is conducted before flight in order to obtain data such as thrust. pressure, temperature and strain, which show the characteristics of rocket motors. But the measured thrust of the obtained data is especially distorted by the effects of dynamic characteristics of thrust stand so that it is difficult for us to determine the exact value of peak thrust and rising time etc., which represent the performance of rocket motor. This paper, therefore. verified the causes of distortion of measured thrust, and proposed the theoretical method to estimate the true thrust from the distorted thrust. And also the proposed method was applied to virtual thrust stand using computer simulation, and showed good result. As a result of that, the proposed method was proven to be valid and applicable to estimate distorted thrust.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.393-397
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• 2006

In this study, combustion characteristics of solid propellants using sorbitol and potassium nitrate were found out. Burning rate was calculated with several combustion experiments, also specific impulse and characteristic exhaust velocity were compared with theoretical value. Thrust measured with thrust measurement system using plate spring. Mixture ratio of propellants was varied in experiments, also combustion characteristics of solid propellants which consulted experimental results was used micro solid rocket design having 1mm nozzle throat.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.211-217
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• 2003

A liquid rocket engine(LRE) Using LO$_2$/LNG(Liquefied Natural Gas) propellants was experimentally evaluated. The purpose of this study was to investigate the performance of the LO$_2$/LNG rocket combustor that is composed of three sect ions(igniter spacer, cylinder and nozzle section), especially focused on the influence of regenerative cool ing effect in association with the phase of regenerative coolant Series of tests were conducted under the conditions of water cool ing and regenerative cool ing with LNG in the cylinder section and independent cool ing with water in the igniter spacer and nozzle sections. Parametric studies on the variation of a chamber pressure and mixture ratio were undertaken. In addition, effect of propellant(LNG) composition and its enthalpy on the performance is examined.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.2
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• pp.72-82
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• 2017

Nozzleless booster is one of the applicable components for integral rocket ramjet (IRR). In order to predict nozzleless solid booster performance, the simplified theoretical analysis was applied for L/D=5, 6, 7, 9, 11, and 13. Al-HTPB and Zr-HTPB propellant with a high metal content were used to increase the hardness because of the combustion gas flow effect. It was found that the trends between the simplified theoretical analysis and experiments were similar.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.9
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• pp.805-815
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• 2007

In order to build a conceptual design program for a liquid rocket engine system, performance based sub-programs for each core component of the engine system were made. Parts included were the combustion chamber, supersonic nozzle, centrifugal pump, and impulsive turbine. Simple mathematical models based on classical thermodynamic and inviscid theories were adopted with proper tuning by empirical data. In Part I, aiming to validate each sub-program, we examined the results of each program qualitatively, and parametrically investigated the sensitivity due to the change in design parameters.