• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.35 no.5
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• pp.390-396
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• 2007

In the present study, we have developed a flapping wing using a smart material to mimic the nature's flyers, birds. The wing consists of composite frames, a flexible PVC film and a surface actuator, and the main wing motions are flapping, twisting and camber motions. To change the camber, a Macro-Fiber Composite(MFC) is used as the surface actuator, and it's structural response is analyzed by the use of piezoelectric-thermal analogy. To measure the lift and thrust simultaneously, a test stand consisting of two load cells is manufactured. Some aerodynamic tests are performed for the wing in a subsonic wind tunnel to evaluate the dynamic characteristics. Experimental results show that the main lift is mostly affected by the forward velocity and the pitch angle, but the thrust is mostly affected by the flapping frequency. The effect of the camber generated by the MFC actuator can produce the sufficient lift increment of up to 24.4% in static condition and 20.8% in dynamic condition.

• Journal of the Computational Structural Engineering Institute of Korea
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• v.30 no.5
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• pp.445-452
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• 2017

The wing leading edge skin in this research is an essential structural factor for improving wings' aeromechanical functions, protecting the interior elements of the wings from external damage including birds, and navigating planes safely. The study compared and reviewed models manufactured for optimal light-weight wings of composite UAVs. It compared and investigated displacement forms of torsion loads through finite element analysis using MSC. Patran/Nastran. By confirming the improvement of light-weighting performance according to lamination type, thickness change and shape through torsion strength tests of each model, the research suggested the optimal light-weight wing leading edge skin for small composite UAVs.

• International Journal of Aeronautical and Space Sciences
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• v.12 no.1
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• pp.84-91
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• 2011

A finite element analysis for the wing and landing gear of a composite target-drone air vehicle was performed. For the wing analysis, two load cases were considered: a 5g symmetric pull-up and a -1.5g symmetric push-over. For the landing gear analysis, a sinking velocity of 1.4 m/s at a 2g level landing condition was taken into account. MSC/NASTRAN and LS-DYNA were utilized for the static and dynamic analyses, respectively. Finite element results were verified by the static test of a prototype wing under a 6g symmetric pull-up condition. The test showed a 17% larger wing tip deflection than the finite element analysis. This difference is believed to come from the material and geometrical imperfections incurred during the manufacturing process.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.23 no.1
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• pp.96-102
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• 2015

Composite materials are widely used as structural materials for manufacturing an aircraft, due to their : low weight, low thermal expansion coefficient, production efficiency, anisotropy, corrosion resistance and long fatigue life. The range of using composite materials has been extended from the fuselage and the wings to the entire aircraft structure. In this paper, by analyzing the problems which were generated while designing and fabricating aircraft structures using composite materials, the differences between metallic structures and composite structures are described. In addition, the methodological improvement directions on design and fabricating are described.

• Proceedings of the Korean Society For Composite Materials Conference
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• 1999.04a
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• pp.270-273
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• 1999

• Proceedings of the Korean Society For Composite Materials Conference
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• 2000.11a
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• pp.187-190
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• 2000

Cyclocopter is air vehicle to vertically take-off and land like a helicopter. This is an efficient and quiet means of being able to direct thrust compared to a helicopter. The rotor consists of several blades rotating about a horizontal axis perpendicular to the direction of normal flight. The direction of blade span is parallel to rotating axis and both end roots are connected to the hub to resist centrifugal force and to transmit the power. The pitch of the individual blades to the tangent of the circle of the blade's path is varied cyclically to gain thrust. In the paper, the design and manufactures of cyclocopter rotor blades are presented. Stress at the roots of cyclocopter blades is great due to centrifugal and aerodynamic forces and aeroelastic instabilities appear. The blades consist of main spar, front spar, polyurethan foam, weight, and skin and spars and skin are made of glass/epoxy composite.

• Composites Research
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• v.19 no.5
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• pp.12-19
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• 2006

In this paper, conceptual structural design of the main wing for a small scale WIG(Wing in Ground Effect) among high speed ship projects, which will be a high speed maritime transportation system for the next generation in Rep. of Korea, was performed. The Carbon/Epoxy material was selected for the major structure, and the skin-spar with a foam sandwich structural type was adopted for improvement of lightness and structural stability. As a design procedure for the present study, firstly the design load was estimated through the critical flight load case study, and then flanges of the front and rear spars from major bending loads and the skin and the spar webs from shear loads were preliminarily sized using the netting rule and the rule of mixture. Stress analysis was performed by a commercial FEA code, NASTRAN. From the stress analysis results for the first designed wing structure, it was confirmed that the upper skin between the front spar and the rear spar was unstable fer the buckling. Therefore in order to solve this problem, a middle spar and the foam sandwich type structure at the skin and the web were added. After design modification, the structural safety and stability for the final design feature was confirmed. In addition to this, the insert bolt type structure with eight high strength bolts to fix the wing structure to the fuselage was adopted for easy assembly and removal as well as in consideration of more than 20 years fatigue life.

• Transactions of the Korean Society of Mechanical Engineers
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• v.19 no.3
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• pp.615-622
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• 1995

It may be inefficient to conduct the aeroelastic analysis by using full-scale conventional finite-element analyses or experiments, from the initial design phase, for an aircraft wing which can be considered as the discontinuum complex structure with composite laminated skins. In this paper, therefore more efficient aeroelastic analysis has been conducted for a box-beam typed aircraft wing by using the equivalent continuum beam-rod model which is derived from the concept of energy equivalence. Equivalent structural properties of the continuum beam-rod model are obtained from the direct comparison of the finite-element matrices of continuum beam-rod model with those of box-beam typed aircraft wing. Numerical results by the continuum beam-rod model approach are compared with those by the conventional finite-element analysis approach to show that the continuum beam-rod model proposed herein is quite satisfactory as a simplified model of aircraft wing structure for aeroelastic analyses.

• Proceedings of the Korean Society For Composite Materials Conference
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• 2005.04a
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• pp.253-256
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• 2005

Nonlinear aeroelastic analyses of composite wing with flap are performed considering free-play and dynamic stiffness of actuator. Doublet-Hybrid method is used for the calculation of subsonic unsteady aerodynamic forces. Free-play is modeled as an asymmetric bilinear spring and is linearized by using the describing function method. The linear and nonlinear flutter analyses show that the flutter characteristics are significantly dependent on the free-play and dynamic stiffness. From the nonlinear flutter analysis, various types of limit cycle oscillations are observed in a range of air speeds below or above the linear divergent flutter boundary.

• Proceedings of the Korean Society For Composite Materials Conference
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• 2005.11a
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• pp.173-178
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• 2005

In this paper, an efficient and robust analysis system for the flutter optimization of laminated composite wings has been developed using the coupled computational method based on the genetic algorithm. General three-dimensional doublet-lattice method is efficiently used to compute generalized aerodynamic forces of T-tail configuration in the frequency domain. Structural dynamic analyses of laminated composite T-tail models are conducted using finite clement method. The classical P-k flutter analysis technique is applied to effectively solve the aeroelastic governing equations in the frequency domain. Optimum design studies using genetic algorithm have been conducted in order to obtain maximum flutter stability of a composite T-tail configuration. The results show that flutter stability can be significantly increased using composite materials with proper optimum design concepts even for the same weight and shape condition. In the view point of engineering design, it is also importantly shown that the optimization of the vertical wing part is highly effective comparing to the optimization of horizontal wing part.