• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.2
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• pp.165-170
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• 2007

During the conceptual design phase of a wing-body-canard type fighter class aircraft, as a method of maneuverability enhancement for an aircraft, effects of canard-leading edge flap scheduling have been studied. In this study, corrected supersonic panel method has been used to predict the drag polar characteristics due to canard-leading edge flap deflections in the high speed regime. Utilizing the predicted drag polar curves, the canard-leading edge flap scheduling laws have been established. These scheduling laws are the relation of canard-leading edge flap deflections and the flight conditions to maximize the lift-drag ratio. Based on the results obtained from the canard-leading edge flap scheduling, the present method has shown to be useful to enhance the maneuverability of wing-body-canard type fighter class aircraft.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.6
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• pp.10-17
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• 2006

NACA23012익형에 대하여 synthetic jet을 이용하여 박리 제어를 수행하였다. 공력특성의 향상을 위해 앞전 droop과 plain flap의 박리 부근에 synthetic jet을 위치시켰다. 고 받음각에서 앞전 박리의 발생으로 인한 실속을 앞전 droop의 작동과 이때 발생하는 앞전 박리를 synthetic jet으로 효과적으로 지연시킬 수 있고, 또한 실속 특성을 개선 할 수 있음을 확인하였다. 양력의 향상을 위하여 plain flap을 장착하였고, 이때 발생하는 박리를 synthetic jet으로 지연시켜 제어면의 작동 효율을 증가 시킬 수 있음을 확인하였다. 앞전 droop과 plain flap으로 구성된 간단한 고양력 장치에 발생하는 박리를 synthetic jet으로 제어함으로 실속을 지연시킴과 동시에 실속 특성을 향상시키고, 최대 양력의 증가로 fowler flap에 버금가는 공력특성을 확보할 수 있음을 확인하였다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.10
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• pp.899-904
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• 2007

During the conceptual design phase of fighter class aircraft, the high speed wind tunnel test with 1/20 scale wing-body-tail model has been conducted to investigate the effects of leading edge flap deflection on the supersonic cruise performance of the aircraft. To select the proper leading edge flap deflection for the wind tunnel test, the aerodynamic characteristics due to various leading edge flap deflections have been analyzed by using corrected supersonic panel method. Based on the results obtained from the experimental and numerical approaches, the effects of leading edge flap deflection have shown to be useful to enhance the supersonic cruise performance of fighter class aircraft.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.7
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• pp.655-660
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• 2007

A canard-leading edge flap deflection scheduling laws have been established to enhance the maneuverability of the canard type fighter class aircraft. These scheduling laws are the relation of canard-leading edge flap deflections and flight conditions to maximize the lift-drag ratio. For these purposes, the corrected supersonic panel method has been used to predict the lift-drag characteristics due to canard-leading edge flap deflections. In addition, the high speed wind tunnel test has been conducted with 1/20 scale model to validate the predicted scheduling laws. Good agreements have been obtained compared with the results of high speed wind tunnel test. Based on the results obtained from the experimental and numerical studies, the corrected supersonic panel method has shown to be useful to establish the canard-leading edge flap deflection scheduling law for the aerodynamic design of canard type fighter class aircraft.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.4
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• pp.41-47
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• 2011

Turbopump, which is a part of 75 ton open cycle liquid rocket engine has a super sonic impulse turbine. This paper investigated the leading edge thickness effect on the turbine performance experimently. Two rotors were tested with the different leading edge thickness. The ratios (rotor thickness to Pitch) are 1.9 and 1.4 times to 30 ton turbine rotor. As a result, a rotor with 1.4 times ratio had a 1.5% higher efficiency gain than a rotor with 1.9 times ratio. The pressure ratio with the maximum efficiency on the same rotational speed was increased to the full expansion ratio of nozzle.

• Journal of the Korean Society of Propulsion Engineers
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• v.20 no.6
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• pp.38-45
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• 2016

A study was conducted on the effect of leading edge location for the performance characteristics of a centrifugal compressor impeller. Five impellers with different leading edge location were selected for numerical analysis. The impeller with leading edge located 10% away from the inlet about meridional distance from entrance to exit showed the best total pressure ratio and efficiency. Also, this case showed relatively uniform flow distribution because of a weak intensity of the separation region at impeller exit. The impeller with leading edge located far from this location showed lowest total pressure ratio and efficiency. Performance of compressor also decreased due to non-uniform flow distribution at impeller exit.

• Journal of computational fluids engineering
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• v.6 no.2
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• pp.9-21
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• 2001

초음속 여객기의 천음속 순항 성능을 개선하기 위하여 날개의 플랩 꺽임각을 최적화하였다. 이를 위하여 3차원 Euler 코드와 adjoint 코드를 이용한 최적설계기법을 적용하였다. 설계변수로서, 앞전플랩 5개, 뒷전 플랩 5개 등 총 10개의 플랩의 꺽임각이 사용되었다. 설계과정중에 격자계 내부격자점의 수정을 위해 타원형방정식법을 이용하였다. 계산 시간의 단축을 위해 내부격자의 민감도는 무시하였다. 또한 본 설계문제에 근사구배기법의 적용가능성 여부를 조사하였다. 충격파가 없는 경우 앞전 플렙에 한하여 근사구배기법을 적용할 수 있음을 알았다. 최적설계기법으로 BFGS기법을 적용하여 항력을 최소화하였으며, 양력 및 날개 표면 마하수에 대한 제약조건을 적용하였다. 앞전 플랩의 최적화 및 앞전과 뒷전 플랩의 최적화 등 두 개의 설계 문제를 고려하였다. 성공적인 결과를 얻음으로써 본 설계방법의 타당성 및 효율성을 확인하였다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.1
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• pp.11-19
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• 2005

In order to improve dynamic stall characteristics of an oscillating airfoil, optimal design has been performed for fixed nose droop and Gurney flap. Fixed nose droop is known to be very effective to improve pitching moment characteristics but may cause degeneration of aerodynamic lift at the same time. On the other hand, Gurney flap has the opposite characteristics. For fixed nose droop, location and angle are chosen as design variables, while length is defined as design variable for Gurney flap. Higher order response surface methodology and sensitivity based optimal design method are employed to handle highly nonlinear problem such as dynamic stall. Optimal design has been performed so that lift and pitching moment are simultaneously improved. The design results show that aerodynamic characteristics can be remarkably improved through present design approach and the present passive control method is as good as active control method which combines variable nose droop and Gurney flap.

• Journal of Aerospace System Engineering
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• v.13 no.5
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• pp.32-38
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• 2019

The flying wing configuration with high sweep angles and rounded leading edge represent a complex flow of structures by the leading edge vortex. For control of the tailless flying wing configuration with unstable directional stability, flaperon is used. In this study, we conducted numerical simulations for a non-slender flying wing configuration with a rounded leading edge and analyzed the effect of the sideslip angle and flaperon. Through aerodynamic coefficient analysis, it was found that the effect of AoS on lift and drag coefficient was minimal and the side force and moment coefficient were markedly influenced by AoS. As the sideslip angle increased, the pitch break, which is related to the pitching moment coefficient, was delayed. Through stability analysis, the directional and lateral static stability of the flying wing configuration were increased by flaperon. Also, the structure and behavior of the leading edge vortex were analyzed by observing the contour of the pressure coefficient and the skin friction line.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.257-265
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• 2009

The present study deals with the leading edge shape on a wing-body junction to decrease a horseshoe vortex, one of the main factors to generate the secondary flow losses. The shape of leading-edge is optimized with design variables form the leading-edge shape. Approximate optimization design method is used for the optimization. The study is investigated using $FLUENT^{TM}$ and $iSIGHT^{TM}$. As the result, total pressure coefficient of the optimized design case was decreased about 9.79% compare to the baseline case.