• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.39 no.4
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• pp.297-305
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• 2011

In this study, the aeroelastic analysis of rotorcraft in forward flight has been performed using dynamic inflow model to handle unsteady aerodynamics. The quasi-steady airload model based on the blade element method has been coupled with dynamic inflow model developed by Peters and He. The nonlinear steady response to periodic motion is obtained by integrating the full finite element equation in time through a coupled trim procedure with a vehicle trim for stability analysis. The aerodynamic and structural characteristics of dynamic inflow model are validated against other numerical analysis results by comparing induced inflow and blade tip deflections(flap, lag). In order to validate aeroelastic stability of dynamic inflow model, lag damping are also compared with those of linear inflow model.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.9
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• pp.88-93
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• 2003

This Paper described the simulation program development for helicopter. In the design of flight control system to accomplish some special missions like UAV, it is important to minimize the execution time obtaining a linear model from nonlinear model that is used for design of controller. The first step for this kind of purpose is to complete a nonlinear model that contains full dynamic characteristics. The second step is to get the trim values that are obtained from the nonlinear model by solving an algebraic equation. And then stability and control derivatives are derived through hovering to forward flight by numerical perturbation that will be used for linear model for a specified flight condition. The software program(HeliSim) is developed by using MATLAB GUI and will provide easy modeling procedure. The suggested method in this paper is much more simpler than any other method like a fully scale helicopter model. The advantage of our suggested method will reduce the computational time due to simple formula to extract a linear model from nonlinear model that will be beneficially used for flight control system of unmanned helicopter by some reduction of computational load.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.49 no.9
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• pp.729-738
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• 2021

Feedback flow control using an artificial neural network was numerically investigated for NACA0015 Airfoil to suppress flow separation on an airfoil. In order to achieve goal of flow control which is aimed to reduce the size of separation on the airfoil, Blowing&Suction actuator was implemented near the separation point. In the system modeling step, the proper orthogonal decomposition was applied to the pressure field. Then, some POD modes that are necessary for flow control are extracted to analyze the unsteady characteristics. NARX neural network based on decomposed modes are trained to represent the flow dynamics and finally operated in the feedback control loop. Predicted control signal was numerically applied on CFD simulation so that control effect was analyzed through comparing the characteristic of aerodynamic force and spatial modes depending on the presence of the control. The feedback control showed effectiveness in pressure drag reduction up to 29%. Numerical results confirm that the effect is due to dramatic pressure recovery around the trailing edge of the airfoil.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.42 no.10
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• pp.884-892
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• 2014

In this paper, a meta model using neural network substituting for the simulator aerodynamic database is proposed to improve simulation fidelity near the critical flight area and real-time performance. It is shown that the accuracy of the meta model is relatively higher than the existing table lookup methods for arbitrary nonlinear database and the calculation speed is also improved for a specific F-16 maneuver simulation. The increase in the number of hidden nodes in the meta model for better accuracy of database representations causes a delay in function generation due to increased time required for computing exponential functions. In order to make up this drawback, we additionally study the fast exponential function method.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.4
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• pp.25-32
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• 2006

An advanced method of Relaxed Static Stability (RSS) is utilized for improving the aerodynamic performance of modern version supersonic jet fighter aircraft. The flight control system utilize RSS criteria in both longitudinal and lateral-directional axes to achieve performance enhancements and improve stability. The T-50 advanced trainer employs the RSS concept in order to improve the aerodynamic performance and the flight control law in order to guarantee aircraft stability. The longitudinal two modes are the short period with high frequency and the phugoid mode with low frequency. The design goals of longitudinal control laws is concerned with the short period damping and frequency optimization using lower order equivalent system and utilizing the requirement of MIL-F-8785C. Analysis of short period mode has been and continues to be performed This paper addresses the analysis of aircraft phugoid node characteristics such as damping, natural frequency, and analysis of aircraft pitch motion that impacted by angle of attack limiter and auto pitch attitude control law.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.6
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• pp.83-91
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• 2006

An advanced method of Relaxed Static Stability (RSS) is utilized for improving the aerodynamic performance of modern version supersonic jet fighter aircraft. The flight control system utilizes RSS criteria in longitudinal axis to achieve performance enhancements and improve stability. The flight control law of T-50 advanced trainer employs RSS concept in order to improve the aerodynamic performance and guarantee aircraft stability. The longitudinal center of gravity(X-c.g) varies as a function of external stores, fuel state and gear position. Shifts in X-c.g relate directly to longitudinal static margin in aircraft stability. This paper deals the maximum aft X-c.g for critical aircraft loadings and checks static margin limits using sensitivity such as damping, natural frequency, gain and phase margin. And nonlinear analysis was conducted for such as short period input. And also, this paper shows the T-50 aircraft stability based on the result of high angle of attack flight such as upright and inverted departure.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.7
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• pp.575-580
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• 2007

During the conceptual design phase of a canard type high maneuverable aircraft, the low speed small scale wind tunnel test was conducted to investigate the high angle-of-attack static stability of the aircraft. In this study, 1/50th scale generic canard-body-wing model was used for the small scale wind tunnel test. For the analysis of static stability including high angle-of-attack nonlinear characteristics, the vertical tail effects were studied due to canard deflections. In addition, the nose chine effects were studied at high angle-of-attack. Based on the results obtained from the experimental study, the configuration change effects for canard type aircraft on high angle-of-attack static stability have been able to analyze.

• Composites Research
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• v.15 no.1
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• pp.1-8
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• 2002

In this study, the flutter analysis for a rectangular box wing and an actual fighter wing with composite shin, aluminum spar and aluminum rib has been conducted. A conservative 3D wing-box model of an actual wing is modeled by MSC/PATRAN and the corresponding free vibration analysis has been performed by MSC/NASTRAN. The finite elements of membrane, rod and shear panel are used. Using the practical ply angles, various composite laminates are composed and analysed. The DLM code which is linear aerodynamic theory in frequency domain is applied to calculate unsteady aerodynamic pressure in subsonic flow region and the V-g and p-k methods are applied to obtain the solution of aeroelastic governing equation in frequency domain.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.10
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• pp.984-991
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• 2009

In this study, transonic aeroelastic response analyses have been conducted for the DLR-F4(wing-body) aircraft configuration considering shockwave and flow separation effects. The developed fluid-structure coupled analysis system is applied for aeroelastic computations combining computational structural dynamics(CSD), finite element method(FEM) and computational fluid dynamics(CFD) in the time domain. It can give very accurate and useful engineering data on the structural dynamic design of advanced flight vehicles. For the nonlinear unsteady aerodynamics in high transonic flow region, Navier-Stokes equations using the structured grid system have been applied to wing-body configurations. In transonic flight region, the characteristics of static and dynamic aeroelastic responses have been investigated for a typical wing-body configuration model. Also, it is typically shown that the current computation approach can yield realistic and practical results for aircraft design and test engineers.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 1996.10a
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• pp.65-70
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• 1996

엔진소음을 소음특성에 따라 분류하면 공력소음(Aerodynamic Noise), 연소소음(Combustion Noise), 기계적인 소음(Mechanical Noise)으로 나눌 수 있으며 소음원의 종류에 따라 분류하면 배기계소음(Exhaust System Noise)으로 나눌 수 있으며 소음원의 종류에 따라 분류하면 배기계소음(Exhaust System Noise), 흡기계소음(Intake System Noise), 냉각계소음(Cooling System Noise), 엔진표면소음(Engine System Noise)등으로 분류할 수 있다. 이러한 여러소음중 엔진 내부의 유동에 의한 흡배기계통으로의 소음방출은 자동차 실 내외 소음의 중요한 문제로 대두되는데, 이를 줄이기 위해 그 동안 소음기 등의 서브시스템의 형태와 그 위치조정에 관한 연구가 수행되어 왔다. 그러나 이것이 비용 또는 성능에 영향을 미치므로 본질적인 소음원을 규명해 내는 것이 필요하게 되었다. 흡배기계의 소음은 엔진의 흡입, 배기행 정시 피스톤의 운동에 의해 팽창 및 압축파 형태의 압력파(pressure wave)로 발생하게 되고, 밸브근방에서는 유동의 박리(separation)에 의해 발생하게 된다. 소음기 등의 서브시스템에서도 유동의 박리에 의해 발생하게 되며 특히 배기행정시 발생하는 압력파는 비선형영역에 있게된다. 흡기소음은 배기에 비해 그 크기가 작아서 그동안 등한시 되어왔으나 이것이 소비자의 불평요인으로 작용하므로써 이에 대한 연구도 활발히 수행되어야 한다. Bender, Bramer[1]는 흡배기계 소음의 외부 방사에 관하여 전반적으로 기술하였고 Sierens등[2]은 흡기계에서 1차원 MOC(Method of Characteristics)방법으로 비정상 유동해석을 하고 실험결과와 비교하였다. J.S.Lamancusa 등[3]은 흡기 소음원을 실험을 통해 예측하였고, 흡기소음도 비선형 거동을 보인다고 밝혔다. Yositaka Nishio 등[4]은 새로운 흡기실험장치를 고안하여 공명기(resonator)의 위치 변화에 의한 저소음 흡기계를 설계 초기단계에서부터 적용하려 하였다. 일반적으로 흡배기계의 복잡한 형상 때문에 대부분 실험을 통해 문제를 해결하려 하였고, 수치해석은 피스톤의 운동을 배제한 단순화한 흡배기계의 정상상태 유동해석이 주를 이루어왔다. Taghaui and Dupont 등[5]은 KIVA코드를 사용하여 흡기포트와 연소실 그리고 밸브의 움직임을 동시에 고려한 수치해석을 도입하였다. 하지만 이들이 밸브의 운동을 고려하기 위해 사용한 이동격자는 격자점은 시간에 따라 변화하지만 그 격자의 수가 일정하게 유지되어 있어서 밸브의 완전개폐를 해석할 수가 없다. 강희정[6]은 단일 실린더와 단일 배기밸브를 갖는 문제로 단순화하여 피스톤과 밸브의 움직임을 고려하므로써 배기행정 후 소음이 어떻게 전파해 나가는가를 연구하였다. 본 연구에서도 최소밸브간격과 최대밸브간격 사이에서만 계산이 가능하나 흡기의 경우는 밸브가 닫힐 때 생기는 압력파가 중요하므로 실린더와 밸브사이에 벽면조건을 주어 밸브의 개폐를 모사하였다.