American Institute of Aeronautics and Astronautics 1 Simulation of Discrete Damage Evolution in Laminated Composite Compact Tension Specimens David H. Mollenhauer 1 Air Force Research Laboratory, Wright Patterson AFB, OH, 45433-7750, USA Endel V. Iarve 2 , Sirina Putthanarat, Kevin Hoos University of Dayton Research Institute, Dayton, OH, 45469-0060, USA and Stephen R. Hallett 3 and Xiangqian Li University of Bristol, Bristol, BS8 1TR, United Kingdom Damage progression in laminated Overheight Compact Tension specimens was modeled using discrete representations of individual cracks and delaminations. Matrix cracking and delamination initiation, propagation, and interaction, without any prior knowledge and/or meshing of matrix cracking surfaces, is accomplished by combining stress and fracture mechanics-based constitutive modeling within a mesh independent crack-modeling framework. Simulation results including only matrix damage for specimens with [45 2 /90 2 /- 45 2 /0 2 ] s and [0 4 /90 4 ] 2s stacking sequences were compared with load-displacement curves and 3D X-ray micro computed tomography results from tested specimens. Excellent correlation was shown between the simulated and experimental load-displacement curves including statistical variations and proper representation of both the curve non-linearity and peak load. Similarly, remarkable correlation between simulated and experimental damage extent was shown. Additionally, a [45/90/-45/0] 2s specimen exhibiting significant fiber fracture was modeled and results compared with experiment. Fiber fracture was simulated using a continuum damage mechanics approach in addition to the discrete cracking and delamination damage representations of matrix damage. The simulated load displacement curve and damage extent compared favorably with experimental results. Nomenclature POD = pin opening displacement MIC = mesh independent cracking CT = computed tomography BSAM = B-Spline Analysis Method CDM = continuum damage mechanics I. Introduction ailure in laminated composite materials is generally dominated by extensive evolution of discrete damage in the form of matrix cracking between fibers and delamination between plies in various combinations depending on stacking sequence and ply thickness 1 . Often this matrix damage occurs well before ultimate failure and tends to cause redistribution of stress in the failing composite. In notched composites, it has been shown that splits within 0° plies, matrix cracking in other plies, and delamination between various plies can modify the notch stress field and improve the notched strength of the composite 2-5 . With most of these notched composite test specimens, when 0° fiber fracture begins to occur, final specimen failure is rapid and catastrophic, providing the researcher with no means of investigating the effects of sub-critical damage on the progression of fiber failure. To overcome this experimental dilemma, the Overheight Compact Tension (OCT) specimen was conceived at the University of British Columbia 6 . It was designed to represent laminate behavior of large, aircraft structures. It has the advantage, depending on stacking sequence, of allowing stable formation of matrix damage and 1 Senior Research Engineer, Composites & Hybrids Branch, 2941 Hobson Way/AFRL/RXBC, Non-Member. 2 Distinguished Research Engineer, 300 College Park Avenue, Non-Member. 3 Senior Lecturer, Aerospace Structures, University Walk, Non-Member. F 52nd AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics and Materials Conference<BR> 19th 4 - 7 April 2011, Denver, Colorado AIAA 2011-1794 This material is declared a work of the U.S. Government and is not subject to copyright protection in the United States.