Failure Mechanism Study on Omega Stringer Debonding

SUN Jingjing; ZHANG Xiaojing; GONG Zhanfeng; WANG Hai

doi:10.7527/S1000-6893.2013.0279

ACTA AERONAUTICAET ASTRONAUTICA SINICA >

2013 , Vol. 34 >Issue 7: 1616 - 1626

DOI: https://doi.org/10.7527/S1000-6893.2013.0279

Solid Mechanics and Vehicle Conceptual Design

Failure Mechanism Study on Omega Stringer Debonding

SUN Jingjing ,
ZHANG Xiaojing ,
GONG Zhanfeng ,
WANG Hai

Expand

1. School of Aeronautics and Astronautics, Shanghai Jiao Tong University, Shanghai 200240, China;
2. Shanghai Aircraft Design and Research Institute, Commercial Aircraft Corporation of China, Ltd., Shanghai 201210, China

Received date: 2012-09-25

Revised date: 2012-12-17

Online published: 2013-01-05

Fold

Abstract

In the post-buckling mode, fatal failure of a composite stiffened panel is easily induced by the separation of a skin stiffener interface,because the local buckling of the skin will cause the stringer to withstand an out-plane bending load. In this paper, four-point bending tests are performed to simulate the interface properties between the skin and the stringer after post buckling, and a progressive damage model is presented to study the interface failure mechanism and the failure process of the composite hat-stiffened panel structure which takes into consideration both the glued interface and laminate failure with a material stiffness degradation model. The numerical results are in good agreement with the experimental data, which shows that the loading span greatly affects the failure of the structure. In the 90 mm loading conditions, the adhesive layer initially failed near the inner chamfer of the web and skin, and it is mostly influenced by the Ⅱ shear mode. In the 150 mm loading conditions, the adhesive layer initially failed near the free edge of the flange, and it is mostly influenced by both the tensile mode and the Ⅰ and Ⅱ shear mode when the stringer is tensile, and influenced by the Ⅱ shear mode when the stringer is compressed. After the interface was debonded, delamination occurred on the tensile stringer chamfer between the web and the flange with increasing load, which is consistent with the ultrasonic B-scan results.

Key words： omega stringer; skin/stiffener debonding; progressive damage; visual crack closure technique; energy release rate

Cite this article

SUN Jingjing , ZHANG Xiaojing , GONG Zhanfeng , WANG Hai . Failure Mechanism Study on Omega Stringer Debonding[J]. ACTA AERONAUTICAET ASTRONAUTICA SINICA, 2013 , 34(7) : 1616 -1626 . DOI: 10.7527/S1000-6893.2013.0279

References

[1] Yap W H, Scott M L, Thomson R S, et al. The analysis of skin-to-stiffener debonding in composite aerospace structures. Composite Structures, 2002, 57(1): 425-435.
[2] Bertolini J, Castanie B, Barrau J J, et al. An experimental and numerical study on omega stringer debonding. Composite Structures, 2008, 86(1): 233-242.
[3] Krueger R, Cvitkovick M K, O’Brien T K, et al. Testing and analysis of composite skin/stringer debonding under multi-axial loading. Journal of Composite Materials, 2000, 34(15):1263-1300.
[4] Krueger R, Minguet P J. Analysis of composite skin-stiffener debond specimens using volume elements and a shell/3D modeling technique. NASA/CR-2002-211947, 2002.
[5] Krueger R, Minguet P J. Analysis of composite skin-stiffener debond specimens using a shell/3D modeling technique. Composite Structures, 2007, 81(1): 41-59.
[6] van Rijn J C F N. Failure criterion for the skin-stiffener interface in composite aircraft panels. NLR-TP-98264, 1998.
[7] Degenhardt R, Rolfes R, Zimmermann R. Improved material exploitation of composite airframe structures by accurate simulation of post-buckling and collapse-the-projects POSICOSS and COCOMAT. Composite Structures, 2006, 73(2): 175-178.
[8] Orifici A C, Thomson R S, Herszberg I, et al. An analysis methodology for failure in postbuckling skin-stiffener interfaces. Composite Structures, 2008, 86(1):186-193.
[9] Liu C Y, Xu X W, Chen K. Buckling, post-buckling and collapse analysis of stiffened composite panels with debonding damage. Acta Materiae Composite Sinica. 2010, 27(6):158-166.(in Chinese) 刘从玉, 许希武, 陈康. 考虑脱粘的复合材料加筋板屈曲后屈曲及承载能力数值分析. 复合材料学报, 2010,27(6):158-166.
[10] Shi Y P, Zhou Y R. Detailed examples of finite element analysis in ABAQUS. Beijing: China Machine Press, 2006: 61. (in Chinese) 石亦平,周玉蓉. ABAQUS有限元分析实例详解. 北京: 机械工业出版社,2006: 61.
[11] Hibbit D, Karlsson B, Sorenson P. ABAQUS analysis user’s manual. 2010.
[12] Hashin Z. Failure criteria for unidirectional fiber composite. Journal of Applied Mechanics, 1980, 47(2): 329-334.
[13] Yeh H Y, Kim C H. The Yeh-Stratton criterion for composite materials. Journal of Composite Masterials, 1994, 28(10): 926-939.
[14] Yue Z F, Wang F S, Wang P Y, et al. Analysis and optimal design of aircraft composite structures. Beijing: Science Press, 2011: 31. (in Chinese) 岳珠峰,王富生,王佩艳,等. 飞机复合材料结构分析与优化设计. 北京: 科学出版社,2011: 31.

Options

Outlines

模态框（Modal）标题

Abstract

Cite this article

References