• Proceedings of the Korean Society For Composite Materials Conference
• /
• 1999.11a
• /
• pp.246-250
• /
• 1999

항공기 날개 및 동체의 보강재로 사용되는 스트링거가 압축하중을 받게되면 플렌지와 웹에서의 부분좌굴이 발생하고 이는 좌굴이 발생하지 않은 부분에 과도한 하중이 걸리게 하여 스트링거의 전체적인 하중지지능력을 현저히 감소시킨다. 이러한 손상의 형태가 크리플링(Crippling)이다. (중략)

• Journal of Korean Society of Steel Construction
• /
• v.26 no.6
• /
• pp.511-522
• /
• 2014

In this study, an innovative steel stair system is presented which enables rapid erection and high quality control in both residential and office building construction. This system features two lightweight steel stringers of box shape, bolted connections easy to absorb construction tolerance, and stair steps movable transversely (or sliding steps) such that the work space needed for concrete stairway wall could be easily provided. In this type of stairway system, other than providing robust connecting details, ensuring vibration performance is especially important since this system may be vibration-sensitive due to lightweight nature and/or probable low damping. To tackle these issues, a series of full-scale mock-up tests were conducted by using box-shape stringer members with or without concrete-fill. The connection system was shown to be sufficiently stiff and strong, or it remained elastic even under the 160% of service load level. Among the seven stringer alternatives, five exhibited satisfactory vibration performance according to the related North American and European acceptance criteria.

• Composites Research
• /
• v.14 no.3
• /
• pp.32-41
• /
• 2001

Z-section composite stringers with various lengths and flange-widths are tested in axial compression for the validation of a finite element algorithm to calculate the buckling and crippling stresses of composite laminated stringers. The stacking sequence considered is $[{\pm}45/0/90]s$. Strain gages are attached to each specimen, and deflection and end-shortening are obtained by two LVDTs. The buckling load is determined from the load vs. strain response, load vs. end-shortening curves, and load vs. out-of-plane deflection curves. The ultimate stress after local buckling is used as the crippling stress. Comparison between finite element and experimental results shows good agreement in the local buckling and crippling stresses.

• Composites Research
• /
• v.12 no.6
• /
• pp.65-73
• /
• 1999

Crippling stress and failure behavior of Z-section graphite/epoxy composite laminated stringers are investigated by the nonlinear finite element method. Stringers are idealized using 9-node laminated shell element. The complete unloading model is introduced into the finite element method for the progressive failure analysis. A modified Riks method is used to trace the post-failure equilibrium path after local buckling. Finite element results are validated with previous experimental results. The results show that the most important parameter affecting the crippling stress of Z-section stringers is the flange width. In terms of stacking sequence. the highest cripping stress is found at the stringer with $[{\pm}45/0/90]s$ lamination.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.33 no.12
• /
• pp.92-100
• /
• 2005

An aircraft wing or fuselage panel of skin-stringer assembly can fail in a variety of instable modes under compression loads. Instability modes can be buckling of the panel, local buckling of the stringer, flexure, torsion, wrinkle and combined flexural/torsional buckling of the panel assembly. Although researches on these buckling behaviors have been carried out for a long time, there are some difficulties to apply to the practice because of complex theoretical and empirical equations. Accordingly, It is well known that leading aerospace companies are using their own in-house programs for the convenience of practical usage, but our domestic situation is that no such program has been ever developed. In this study a comprehensive program has been developed, which can identify the instability modes and the magnitude of reserve factor of the modes for the skin/stringer panel assembly under compression load. The developed program is based on the theory manual of the Airbus program APA114. For the verification, calculation of the instable reserve factors for the A320 wing panel and A380 low wing panel sections were carried out and compared with results by APA114.

• Journal of Aerospace System Engineering
• /
• v.16 no.6
• /
• pp.8-17
• /
• 2022

This paper presents a novel geometric modeling technique to predict the mechanical properties of an aircraft wing's skin-stringer integrated panel. Due to mechanical and adhesive fastening, this panel is vulnerable to stress concentration and debonding, so we designed it to integrate the skin and stringer using three-dimensional woven composites. Geometric modeling was conducted by measuring the geometric parameters of the specimen and defining the pattern of the yarns as functions. We used a weighted average model with iso-strain and iso-stress assumptions to predict the mechanical properties of the panel parts. We then compared the results of a finite element analysis with a compression test to verify the accuracy of our model. Our proposed technique proved to be more efficient than the traditional experimental method for predicting the mechanical properties of skin-stringer integrated panels.

• Journal of the Korean Society for Aeronautical & Space Sciences
• /
• v.30 no.1
• /
• pp.65-74
• /
• 2002

Composite materials are used for main structural components of aircraft fuselage such as skin, stringer and frame to reduce weight. Failure and buckling analysis of the composite fuselage components have been done for structural design. The loads of MD90-30 are applied to each component. Various shapes of section such as I, Z and T-type are chosen as candidate composite stringer and frame. The analysis results of composite fuselage components are compared according to ply-angle and ply-number, and the section type. The numerical results shows that ply-angle and ply-number have important effects on failure caused by axial load for the frame are important design parameters of composite fuselage components. This study suggests several design tips for composite fuselage components.

• Proceedings of the Korean Society for Technology of Plasticity Conference
• /
• 2007.10a
• /
• pp.201-206
• /
• 2007

In this paper, we apply a three-dimensional rigid-plastic finite element method to simulate an unsteady-state roll forming process. A typical roll forming process is investigated from the standpoint of computer simulation and its realistic analysis model is proposed. The material is considered as bulk material and discretized into hexahedral finite elements. The presented approach is applied to simulating the roll forming process of straight stringer used for aircraft structure.

• Composites Research
• /
• v.13 no.1
• /
• pp.25-32
• /
• 2000

In this paper, postbuckling behavior of laminated composite-stringer stiffened-curved panels loaded in local compression is analyzed using the finite element program developed. Postbuckling Analysis is performed in dividing the panel behavior into three basic parts. The eight node degenerated shell element is used in modelling both panel and stiffeners, and the updated Lagrangian description method based on the 2nd Piola-Kirchhoff stress tensor and the Green strain tensor is used for the nonlinear finite element formulation. The progressive failure analysis is adopted in order to grasp the failure characteristics. The postbuckling experiment of the laminated composite-stiffened-curved panel had been done to verify the finite element analysis. The buckling load and the postbuckling ultimate load are compared in parametric study.

• Journal of the Korean Society for Precision Engineering
• /
• v.23 no.6 s.183
• /
• pp.136-142
• /
• 2006

Optimization of the aircraft panel assembly constructed by skin and stringers is investigated. For the design of panel assembly of the aircraft structure, it is necessary to determine the best shape of the stringer which accomplishes lowest weight under the condition of no instability. A panel assembly can fail in a variety of instability modes under compression. Overall modes of flexure or torsion can occur and these can interact in a combined flexural/torsion mode. Flexure and torsion can occur symmetrically or anti-symmetrically. Local instabilities can also occur. The local instabilities considered in this paper are buckling of the free and attached flanges, the stiffener web and the inter-rivet buckling. A program is developed to find out critical load for each instability mode at the specific stringer shape. Based on the developed program, optimization is performed to find optimum stringer shape. The developed instability analysis program is not adequate for sensitivity analysis, therefore RSM (Response Surface Method) is utilized instead to model weight and instability constraints. Since the problem has many local minimum, Genetic algorithm is utilized to find global optimum.