• 한국항공우주학회지
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• 제35권3호
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• pp.238-247
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• 2007

유전자 알고리즘은 자연선택과 유전법칙을 적용하여 최적해를 탐색하는 방법으로, 본 연구에서 항공기용 가스터빈 엔진의 결함 진단을 위한 학습 알고리즘으로 사용되었다. 성능 저하를 고려한 구성요소는 압축기, 가스발생기 터빈, 동력 터빈이며, 설계점에서 엔진의 단일 구성요소에 대하여 각각 성능 저하 예측을 수행한 후, 이를 바탕으로 결함 진단을 수행하였다. 학습데이터 수의 증가가 유전자 알고리즘을 이용한 성능 저하 예측 및 결함 진단에 미치는 영향을 분석하였으며, 결과적으로 결함치에 대한 RMS 오차율이 모두 3% 이내로 예측됨을 확인하였다.

• Advances in aircraft and spacecraft science
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• 제10권3호
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• pp.245-256
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• 2023

In order to improve aerodynamic performance of multi-stage axial flow turbines used in aircraft engines, a one-dimensional aerodynamic design and optimization framework is constructed. In the method, flow path is generated by solving mass continuation and energy conservation with loss computed by the Craig & Cox model; Also real gas properties has been taken into consideration. To obtain an optimal result, a multi-objective genetic algorithm is used to optimize the efficiencies and determine values of various design variables; Final design can be selected from obtained Pareto optimal solution sets. A three-stage axial turbine is used to verify the effectiveness of the developed optimization framework, and designs are checked by three-dimensional CFD simulation. Results show that the aerodynamic performance of the optimized turbine has been significantly improved at design point, with the total-to-total efficiency increased by 1.17% and the total-to-static efficiency increased by 1.48%. As for the off-design performance, the optimized one is improved at all working points except those at small mass flow.

• 한국추진공학회지
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• 제24권2호
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• pp.61-72
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• 2020

국외에서 개발이 진행되고 있거나 완료된 UAV에 적용된 추진기관의 성능 특성을 분석하였다. 본 연구에서 총 10종의 UAV의 임무 및 성능 특성을 검토하고 민간항공기 및 군용항공기와 성능 특성을 비교 검토하였다. 또한, UAV 추진기관의 성능 특성을 정리하고 엔진 설계 파라미터에 대한 분석을 수행하였다. 추진기관의 추력, SFC 및 설계변수인 압축비, 바이패스비에 대해 민간 및 군용항공기 엔진과 비교하여 검토하였으며, 본 연구를 통해 UAV에 따른 설계 파라미터를 보다 잘 이해할 수 있었다.

• 한국정밀공학회지
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• 제31권1호
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• pp.15-20
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• 2014

Thermal barrier coating (TBC) is used to protect the substrate and extend the operating life of the gas turbine for a power plant and an aircraft. The major cause of failure of such a coating is the spallation of coating, and it results from the thermal stress between top coating and bond coating. To improve the durability of TBC system, the dense vertical cracked (DVC) coating method to insert vertical cracks is applied to a gas turbine blade. In this study, a criterion for the design of vertical crack in the DVC coating was presented using the finite element analysis.

• 대한기계학회논문집A
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• 제26권2호
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• pp.365-373
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• 2002

For gas turbine engines, the safe life methodology has historically been used fur fatigue life management of failure critical engine components. The safe retirement limit is necessarily determined by a conservative life evaluation procedure, thereby many components which have a long residual life are discarded. The objective of this study is to introduce the damage tolerant design concept into the life management for aircraft engine component instead of conservative fatigue life methodology which has been used for both design and maintenance. Crack growth data were collected on a nickel base superalloy which have been subjected to combined static and cyclic loading at elevated temperatures. Stress analysis fur turbine disk was carried out. The program for computing creep-fatigue crack growth was developed. The residual lifes of turbine disk component under various temperatures and conditions using creep-fatigue crack growth data were estimated. As the result of analysis, it was confirmed that retirement fur cause concept was applicable to the evaluation of residual life of retired turbine disk which had been designed based on the conventional fatigue life methodology.

• 한국유체기계학회 논문집
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• 제15권5호
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• pp.27-31
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• 2012

Turbine inlet temperature is steadily increasing to achieve high specific thrust and efficiency of gas turbine engines. Turbine cooling technology is essential to increase turbine inlet temperature. For this study, a small or medium sized aircraft engine of 10,000 lbf class with the turbine inlet temperature of $1,400^{\circ}C$, the engine overall pressure ratio of 32.2, and the bypass ratio of 5 was set as the baseline model and its performance analysis was performed at the design point. The engine has the performance of 10,013 lbf thrust and the specific fuel consumption of 0.362 lbm/hr/lbf. The thrust and the specific fuel consumption of the baseline model were compared with those of similar class engines. Based on these results, the turbine design requirements were assigned. In addition, the parametric analysis of the engine, related to aerodynamic and cooling design of the high pressure turbine, was performed. Based on the baseline model engine, the influence of turbine inlet temperature, cooling flow ratio, and high pressure turbine efficiency variations on the engine performance was analyzed.

• 한국추진공학회:학술대회논문집
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• 한국추진공학회 2007년도 제29회 추계학술대회논문집
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• pp.350-353
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• 2007

본 연구에서는 항공기용 가스터빈 엔진의 운용에 있어 실시간 결함 진단을 위해 유전 알고리즘을 사용하였다. 탈 설계 영역에서 성능 저하를 고려한 가스터빈 엔진의 구성요소는 압축기, 가스발생기 터빈, 동력 터빈이다. 지상정지 상태인 설계점에 비해 고도, 비행 마하수, 연료유량에 대한 탈 설계 진단의 경우 학습 데이터는 약 200배 이상으로 증가하였으며, 요구 수렴도를 만족시키기 위해 방대한 학습시간이 요구된다. 탈 설계 영역에서 단일결함에 관한 진단오차를 만족시키고 학습시간을 단축시키기 위해 최적분할을 사용하였고 그 결과, 오차범위 5% 이내로 진단됨을 확인하였다.

• 한국항공우주학회지
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• 제43권4호
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• pp.326-333
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• 2015

최적화 프로그램과 연동시키기 위한 터보팬 엔진 성능해석 프로그램을 개발하고, 최적화 기법인 Particle Swarm Optimization을 이용하여 전투기 엔진의 주요 설계변수인 바이패스비, 팬 압축비, 고압압축기 압축비 및 버너출구온도에 대한 성능 최적화를 수행하였다. 최적화 목표는 순추력과 비연료소모율을 다목표 함수로 설정하였으며, 두 개의 목표에 대해 가중치를 주어 각 가중치별 최적 설계점을 도출하였다. 기본 모델은 F-18 전투기와 T-50 고등훈련기에 쓰이고 있는 F404 터보팬 엔진을 선정하여 분석을 수행하였다. 본 연구 결과로 네 개의 변수에 대한 최적 조건을 도출하고, 다양한 설계조건에 대한 최적 설계점 추이를 분석하였다.

• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2003년도 추계 학술대회논문집
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• pp.200-205
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• 2003

The Canard rotor/wing (CRW) aircraft concepts offer great potential for application by allowing the use of a common propulsion system for high-speed cruise and low-speed powered lift. Using the rotor for lift in both flight modes increases its utility. In the hovering mode, the exhausted gas from an gas turbine engine is accelerated through the duct system and it provides the tipjet power for rotor system enough to lift the aircraft. In the cruise mode, the rotor is fixed and the exhausted gas is extracted through the main nozzle, such that the aircraft is able to flight with high speed. The duct system was designed using 1-D fanno line flow theory and empirical data. However, the empirical data of the pressure loss coefficient for various bending and dividing ducts were not enough to design our duct system adaptively. Therefore, using 3-D CFD analysis we obtained the pressure loss coefficient for our duct models and chose the appropriate bending or diving duct type. In this paper, we used the CFD-ACE+ software package for the CFD analysis and the modeling of duct system. Through the 3-D CFD analysis, we investigated also the pressure loss and the velocity distributions of the designed whole duct system as well as the blade duct. Comparing the 3-D CFD result with 1-D analysis result, we lessened the uncertainty of the designed duct system and speculated the problem that was not concerned in design state.

• International Journal of Aeronautical and Space Sciences
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• 제2권1호
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• pp.93-102
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• 2001

Flight control design for the autonomous waypoint navigation of aircraft is presented in this study. The waypoints are defined in terms of desired longitude and latitude. The control design is conducted in longitudinal and lateral directions, respectively. The lateral control is based upon coordinated turn strategy for which no sideslip is allowed under the turning maneuver. The longitudinal control is mainly focused on altitude hold during navigation. Neural network control approach is applied to the altitude-hold mode control. Simulation of the proposed control strategy has been performed under various conditions. A graphical simulation tool was developed to visually demonstrate the control technique developed in this study. A method to simulate the gas turbine transient behavior is developed. The basic principles of the method.