• Journal of the Korean Society of Combustion
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• v.18 no.3
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• pp.61-67
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• 2013

This study introduces the design methods for air flow distribution at the level of preliminary design, and reviews the typical combustion process and main functions of sub-components of aircraft gas turbine combustors. There are lots of design approaches and empirical equations introduced for air flow distributions at the combustors. It is shown that a decision on which design approaches work for the combustor development is totally dependent upon the objective of engine design, target performance, and so on. The current results suggested for preliminary air flow distributions need to be validated by combustor geometry checkups and performance evaluations for future works.

• Journal of the Korean Society of Combustion
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• v.18 no.3
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• pp.54-60
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• 2013

This paper shows a general development process for aircraft gas turbine combustors. As a first step for developing the preliminary combustor design program, several combustor sizing methodologies using reference area concepts are reviewed. There are three ways to determine the reference area; 1) combustion efficiency approach, 2) pressure loss approach, 3) velocity assumption approach. The current study shows the comparisons of the calculated results of combustor reference values from the pressure loss and velocity assumption approaches. Further works are required to add iterative steps in the program using more reasonable values of pressure loss and velocities, and to evaluate the sizing results using data for actual combustor performance and sizes.

• International Journal of Aeronautical and Space Sciences
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• v.15 no.4
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• pp.434-443
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• 2014

A turbo fan engine performance analysis and the optimization using particle swarm optimization(PSO) algorithm have been conducted to investigate the effects of major performance design parameters of an aircraft gas turbine engine. The FJ44-2C turbofan engine, which is widely used in the small business jet, CJ2 has been selected as the basic model. The design parameters consists of the bypass ratio, burner exit temperature, HP compressor ratio, fan inlet mass flow, and nozzle cooling air ratio. The sensitivity analysis of the parameters has been evaluated and the optimization of the parameters has been performed to achieve high net thrust or low specific fuel consumption.

• Journal of the Korean Society of Propulsion Engineers
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• v.26 no.3
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• pp.1-9
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• 2022

The aircraft targeted in this study is a vertical take-off and landing aircraft with 4 to 5 passengers, and the propulsion system for the aircraft is a distributed hybrid propulsion system that uses a gas turbine engine and a battery pack as the main power source to supply the power required by multiple motors. In this study, a design/analysis platform for a hybrid propulsion system was developed using the MATLAB/Simulink program based on the preliminary design results. Through simulation analysis, the output characteristics and operating range of each power source according to the mission profile were confirmed, and through this, the feasibility of the preliminary design result was confirmed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.829-832
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• 2010

For gas turbine engine starting, external power should be supplied with engine to accelerate to suitable rotational speed for air and fuel ignition conditions. Electric starting system for small gas turbine engine has simple system and light weight, so it is generally used for small aircraft. For system analysis of gas turbine engine electric starting system, Characteristics of battery, start motor, engine drag torque should be analyzed and theirs temperature effects should be considered. In this paper, preliminary design procedure of small gas turbine engine electric starting system and major design parameters were described.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2003.05a
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• pp.514-519
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• 2003

The aircraft engine is usually supported by rolling element bearings and has a small damping rate, which is vol y sensitive to external force. The high-performance requirement of the rotors leads to complex assembly designs and are more flexible. Squeeze film dampers (SFDs) are introduced to provide damping while crossing the critical speeds and stability to the rotor s :stem. Hence, the focus of the present investigation is on the decision of an optimal size of the flexible rotor system supported by the squeeze film dampers to minimize the maximum transmitted load and unbalance response over a range operating speeds. The enhanced genetic algorithm (EGA), which was developed by authors, is used in the optimization process. This algorithm is based on the synthesis of a modified genetic algorithm and simplex method. The results show significant benefits in using EGA when compared with nonlinear programming (NLP).

• Transactions of the Korean Society of Mechanical Engineers A
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• v.20 no.8
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• pp.2524-2539
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• 1996

A design process and an axisymmetric boundary integral equation method for precise structural analysis of the aircraft gas turbine rotor disk were developed. This axisymmetric boundary integral equation method for stress and steady-state thermal analysis was improved in solution accuracy by appling an implicit technique for Cauchy principal value evaluation, a subelement technique for weak singular integral evaluation and a double exponentical integral technoque for internal point solution near boundary surfaces. Stresses, temperatures, low cycle fatigue lifes and critical speeds for the turbine rotor disk of the thrust 1421 N class turbojet engine were analysed in a pratical calculation model problem.

• The Journal of Aerospace Industry
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• s.25
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• pp.77-101
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• 1992

• Journal of the Korean Society of Propulsion Engineers
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• v.12 no.3
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• pp.60-67
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• 2008

In this study, the genetic algorithm has been used for the real-time defect diagnosis on the operation of the aircraft gas-turbine engine. The component elements of the gas-turbine engine for consideration of the performance deterioration consist of the compressor, the gas generation turbine and the power turbine. Compared to the on-design point, the teaming data has been increased 200 times in case off-design conditions for the altitude, the flight mach number and the fuel consumption. Therefore, enormous learning time has been required for the satisfied convergence. The optimal division has been proposed for learning time decrease as well as the high accuracy. As results, the RMS errors of the defect diagnosis using the genetic algorithm have been confirmed under 5 %.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.285-288
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• 2006

Genetic Algorithms(GA) which searches optimum solution using natural selection and the law of heredity has been applied to teaming algorithms in order to estimate performance deterioration of the aircraft gas turbine engine. The compressor, gas generation turbine and power turbine are considered for estimation for performance deterioration of a complex component at design point was conducted. As a result of that, complex defect diagnostics has been conducted. As a result, the accuracy of diagnostics were verified with its relative error with in 10% at each component.