Verification and Validation Process for Progressive Damage and Failure Analysis Methods in the NASA Advanced Composites Consortium

Abstract

The Advanced Composites Consortium is a US Government/Industry partnership supporting technologies to enable timeline and cost reduction in the development of certified composite aerospace structures. A key component of the consortium's approach is the development and validation of improved progressive damage and failure analysis methods for composite structures. These methods will enable increased use of simulations in design trade studies and detailed design development, and thereby enable more targeted physical test programs to validate designs. To accomplish this goal with confidence, a rigorous verification and validation process was developed. The process was used to evaluate analysis methods and associated implementation requirements to ensure calculation accuracy and to gage predictability for composite failure modes of interest. This paper introduces the verification and validation process developed by the consortium during the Phase I effort of the Advanced Composites Project. Specific structural failure modes of interest are first identified, and a subset of standard composite test articles are proposed to interrogate a progressive damage analysis method's ability to predict each failure mode of interest. Test articles are designed to capture the underlying composite material constitutive response as well as the interaction of failure modes representing typical failure patterns observed in aerospace structures.

An experimental and numerical investigation into damage mechanisms in tapered laminates under tensile loading

“…4b) and the damage sensitive interface (i.e. the continuous plies interface of ply drop 1 in front of the pre-insert) is mode I loaded, thus its mesh convergence trend is consistent with the mode I numerical cohesive zone length as derived by a DCB model and the requirement of at least two to three elements existing in the cohesive zone [21,23,26,29,[32][33][34][35][36][37]. A DCB model with a very fine mesh (0.025 mm) was run following the procedures detailed in Refs.…”

“…[30,31]. There exists discrepancy in the literature regarding the cohesive strength pair [21,23,26,29,[32][33][34][35][36][37] and the TTC enhancement factor [28,38,39]. The reason why [60 MPa, 90 MPa] is used here as the baseline cohesive strength pair is simply because that it has been proved to give satisfactory predictions for many of the previous IM7/8552 laminate models [21,23,26,31,40,41].…”

A Methodology to Investigate Skin-Stringer Separation in Postbuckled Composite Stiffened Panels

“…Bertolini et al [7][8] showed that this can be used both for symmetrical and skewed buckling. Recently, NASA performed seven-point bending tests as part of the Advanced Composite Consortium (ACC) program [9][10][11][12].…”

Study of Skin-Stringer Separation in Postbuckled Composite Aeronautical Structures

“…The three-point bending test, in which out-of-plane deformation is applied to a specimen with a skin and a co-bonded or co-cured doubler, is representative of the opening mode of a stringer flange disbond [1][2] that results from a combination of the high deflection of the skin and the mismatch in flexural stiffness of the skin and stringer. A similar type of damage has also been investigated using a seven-point bending (7PB) test, which adds the complex interaction of the buckling deformation of the skin [3][4][5]. The 7PB test consists of a single-stringer specimen supported at five points and loaded at two points such that the out-of-plane deformation of the skin is similar to a half-wave section of a postbuckled multi-stringer panel.…”

“…In all of these tests, debonding of the skin-stringer is dominated by mode I loading, representing the location where the deflection due to the buckling wave is at its maximum. However, research on skin-stringer separation indicates that a possible combined mode II + III failure can also be a critical mode of failure due to the twisting of the skin at the inflection point, located in between two buckling halfwaves [5,9]. Inflection points can be critical locations, for example, in structures with a tapered stringer flange termination, where a relatively low bending stiffness mismatch between the skin and stringer flange reduces the likelihood of separation in mode I.…”