• 한국전산유체공학회:학술대회논문집
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• 한국전산유체공학회 2008년 추계학술대회논문집
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• pp.245-250
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• 2008

For simulation of a wing unfolding motion for the various aerodynamic conditions, equation governing unfolding motion and moments applying to the unfolding wing were modelled. Aerodynamic roll moment consists of the static roll moment and the damping moment, which were obtained through wind tunnel tests and numerical analyses respectively. Panel method was used to compute the roll damping coefficient with twisted wing, whose deflection angle was equivalent to angle of attack due to the deployment motion. Roll damping coefficient is a function of angle of attack, sideslip angle, and deployment angle but not of angular velocity of deployment. Simulation with aerodynamic damping model gave more similar deployment time compared to wing deployment test results.

• 대한기계학회:학술대회논문집
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• 대한기계학회 2008년도 추계학술대회A
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• pp.1441-1445
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• 2008

Insect flight is adapted to cope with each circumstance by controlling a variety of the parameters of wing motion in nature. Many researchers have struggled to solve the fundamental concept of insect flight, but it has not been solved yet clearly. In this study, to find the most effective flapping wing kinematics, we conducted to analyze CFD data on fixing some of the optimal parameters of wing motion such as stoke amplitude, flip duration and wing rotation type and then controlled the deviation angle by fabricating wing tip motion. Although all patterns have the similar value of lift coefficient and drag coefficient, pattern A(pear-shape type) indicates the highest lift coefficient and pattern H(pear-shape type) has the lowest lift coefficient among four wing tip motions and three deviation angles. This result suggest that the lift and drag coefficient depends on the angle of attack and the deviation angle combined, and it could be explained by delayed stall effect.

• Journal of Advanced Marine Engineering and Technology
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• 제39권1호
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• pp.13-18
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• 2015

본 연구에서는 ANSIS Workbench CFX-Mesh와 같은 기존의 컴퓨터 연산제어 기법과 수학적 공식을 적용하여 풍력 레저선박에 대한 복원력 특성을 검토하였다. 본 연구의 목적은 선박이 풍력을 받아 운항하고 있을 때 복원력과 힐모멘트를 구하는 것이다. 받음각의 조건은 5도에서 90도 범위로 설정하였다. 일반적으로 횡력은 받음각이 클 경우 상대적으로 크다고 알려져 있으나, 선박이 풍력을 받아 안전 운항을 하기 위해서는 임계각의 경우, 60도 범위에서 제한하는 것이 유리하다고 사료된다. 비항해시 대비, 항해시에 보다 큰 힐모멘트가 얻어지며, 이와 같이 우세한 힐모멘트로부터 보다 큰 받음각이 유기된다는 자연스러운 시험 결과물이 도출되었다.

• 항공우주시스템공학회지
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• 제3권2호
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• pp.1-11
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• 2009

A structural modelling for study on dynamic characteristics of tapered composite aircraft wings in the form of thin-walled beam is presented. The proposed structural model includes effects of transverse shear flexibility exhibited by the advanced composite materials and warping restraint characterizing elastic anisotropy and induced structural couplings. The complex effects of these factors could have a role in more efficient analysis on those structural models.

• 한국군사과학기술학회지
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• 제2권2호
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• pp.40-51
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• 1999

This paper reviews the combat survivability and supermaneuverbility which are principal factors in current and future high performance combat aircraft design. First of all, the fighter agility evaluation factors were presented. And then, emphasis was put on technologies associated with supermaneuverbility, such as vortex lift, high angle of attack aerodynamics, thrust vectoring and control system technologies that integrate each technology. The advanced nations' supermaneuverbility R/D programs were introduced as well.

• 한국운동역학회지
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• 제23권2호
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• pp.99-108
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• 2013

The purpose of this study was to analyze the kinematic characteristics of javelin to compare the characteristics between world elite athletes and local male athletes. The subjects selected 9 athletes out of total 13 athletes recorded more than 65 m in the preliminary and main competition participated in the 93rd National Sports Festival held in 2012. Three-Dimensional motion analysis using a system of 4 video cameras(Sony HXR-MC2000) at a sampling frequency of 60 fields/s was performed for this study. The Kwon3D 3.1 was used to obtain the three dimensional coordinates about the top, grip, end of javelin. And the kinematic factors of javelin calculated using Matlab2009a program. The resultant speed of javelin that affects directly to the record of performance showed 26.08 m/s indicated lower speed of about 2-3 m/s than world elite athletes. The release point appeared to have been made at the high of $1.79{\pm}0.07$ m of 99.8% of the height of the athletes. In terms of release angle, it was indicated average $33.0{\pm}3.81^{\circ}$ lower release angle compared to the world elite athletes. The attitude angle(up & down tilt angle, X axis) related to javelin indicated average $38.5{\pm}4.96^{\circ}$, its related attack angle average $5.5{\pm}5.11^{\circ}$, and yaw angle(sideslip, side attack angle) average $15.7{\pm}8.48^{\circ}$.

• Wind and Structures
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• 제32권3호
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• pp.239-248
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• 2021

In this study, three airfoil families, NACA, FX and S, in each case three from each series with different shapes were investigated at different angles of attack using Computational Fluid Dynamics (CFD) method. To verify the CFD model, simulation results of the NACA 0012 airfoil was compared against the available experimental data and k-ω SST was used as the turbulence model. Lift coefficients, lift to drag ratios and pressure distributions around airfoils were obtained from the CFD simulations and compared each other. The simulations were performed at three Reynolds numbers, Re=2×105, 1×106and 2×106, and angle of attack was varied between -6 and 12 degrees. According to the results, similar lift coefficient values were obtained for symmetric airfoils reaching their maximum values at similar angles of attack. Maximum lift coefficients were obtained for FX 60-157 and S 4110 airfoils having lift coefficient values around 1.5 at Re=1×106 and 12 degrees of angle of attack. Flow separation occurred close to the leading edge of some airfoils at higher angles of attack, while some other airfoils were more successful in keeping the flow attached on the surface.

• 한국군사과학기술학회지
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• 제10권1호
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• pp.33-43
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• 2007

Using the more common conventional chordwise aerodynamic approach, flapping a flat plate wing with zero degree chordwise pitch angle of attack and no relative wind should not produce lift. However, in hover, with no forward relative velocity and zero degree chordwise pitch angle of attack, flapping flat plate wings does in fact produce lift. In the experiments peformed for this paper, the flapping motion is considered pure(downstroke and upstroke) with no flapping stroke plane inclination angle. No changes in chordwise pitch angle are made. The total force is measured using a force transducer and the net aerodynamic force is determined from this measured total force by subtracting the experimentally determined inertial contribution. These experiments were repeated at various flapping frequencies and for various wing planform sizes for flat plate wings. The trends in the aerodynamic lift variation found using a force transducer have nearly identical shape for various flapping frequencies and wing planform sizes.

• International Journal of Aeronautical and Space Sciences
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• 제4권2호
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• pp.26-36
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• 2003

The objective of this study is to enhance neural-network guidance to consider the impact condition. The optimal impact condition in this study is defined as an head-on attack. Missile impact-angle error, which is a measure of the degree to which the missile is not steering for a head-on attack, can also have an influence on the final miss distance. Therefore midcourse guidance is used to navigate the missile, reducing the deviation angle from head on, given some constraints on the missile g performance. A coordinate transformation is introduced to simplify the three-dimensional guidance law and, consequently, to reduce training data. Computer simulation results show that the neural-network guidance law with the coordinate transformation reduces impact-angle errors effectively.

• International Journal of Aeronautical and Space Sciences
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• 제11권2호
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• pp.131-135
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• 2010

We report on the design of a three-axis missile autopilot using multi-objective control synthesis via linear matrix inequality techniques. This autopilot design guarantees $H_2/H_{\infty}$ performance criteria for a set of finite linear models. These models are linearized at different aerodynamic roll angle conditions over the flight envelope to capture uncertainties that occur in the high-angle-of-attack regime. Simulation results are presented for different aerodynamic roll angle variations and show that the performance of the controller is very satisfactory.