The Response of Textile Composites Subjected to Elevated Loading Rates.

The United States military is involved in a number of peacekeeping and combat operations throughout the world. In an effort to increase the deployability and agility of the military, a number of novel technologies are being developed. Lightweight combat vehicles have relied on thick structural members, typically made of metallic materials, to defend against specific threats including Improvise Explosive Devices (IEDs) and live fire munitions. While these techniques have proven successful in the past, the weight considerations greatly affect the mobility of these vehicles. Modern armoring techniques have been developed that include the use of Fiber Reinforced Polymer Composites (FRPCs) that offer high structural rigidity while reducing the weight of the vehicle. Because traditional composite laminates suffer from delamination which does not allow the material to reach its full mechanical performance, a new class of 3D composite materials, referred to as 3D textile composites, have been developed. 3D Textile composites involve the interweaving of fiber yarns in a number of preforms to achieve a desired mechanical performance. In this thesis, the high strain rate tensile response of 3D textile composites is addressed. Traditional 1-dimensional high strain rate tensile testing of composite materials is used to understand the failure mechanics of a single fiber tow and to develop constitutive models that are subsequently used to explain the results of textile composites under single and multi-axial load states at quasi-static and high rates of loading. Plain woven textile composites are subjected to shock loading in a shock tube facility to produce biaxial tensile load states. The deformation response is measured using digital image correlation. Experimental results are used as motivation to construct finite element models, including different length scales to understand and explain the observed deformation response and failure mechanics of textile laminates.PhDMechanical EngineeringUniversity of Michigan, Horace H. Rackham School of Graduate Studieshttp://deepblue.lib.umich.edu/bitstream/2027.42/113554/1/bjustuss_1.pd

An Overview of the NASA Advanced Composites Consortium High Energy Dynamic Impact Phase II Technical Path

Advanced composite structures are increasingly becoming the norm for use in military and commercial aircraft. Many of these structures are in places that are prone to high energy dynamic impact (HEDI) such as a wing or fuselage structures subjected to bird strike or a fan blade out event. Certification testing is expensive and industry currently lacks to the tools to perform reliable certification by analysis or smarter testing. As such, the NASA Advanced Composites Consortium HEDI team was formed with representatives from aerospace original equipment manufacturers, government research laboratories, and academia to advance the state-of-the-art in emerging progressive damage and failure analysis (PDFA) methods in a two phase program. These PDFA approaches have the ability to predict ply-by-ply level damage in composite structures, but to date, have not been thoroughly vetted for HEDI events. In this paper, the technical path that is used in Phase II of the program is presented

High Strain Rate Response of Adhesively Bonded Fiber-Reinforced Composite Joints A Computational Study to Guide Experimental Design

Adhesively bonded carbon fiber-reinforced epoxy composite laminates are widely used in aerospace applications. During a high energy impact event, these laminates are often subjected to high strain rate loading. However, the influence of high strain rate loading on the response of these composite joints is not well understood. Computational finite element (FE) modeling and simulations are conducted to guide the design of high strain rate experiments. Two different experimental designs based on split Hopkinson bar were numerically modeled to simulate Mode I and Mode II types loading in the composite. In addition, the computational approach adopted in this study helps in understanding the high strain rate response of adhesively bonded composite joints subjected to nominally Mode I and Mode II loading. The modeling approach consists of a ply-level 3D FE model, a progressive damage constitutive model for the composite material behavior and a cohesive tie-break contact element for interlaminar delamination

Comparison of Test Methods to Determine Failure Parameters for MAT162 Calibration

MAT162, a laminated composite failure material model developed by The Material Sciences Corporation for the commercial finite element software LS-Dyna, is widely used within the aerospace industry to predict damage events under a range of dynamic conditions. The material model involves numerous inputs consisting of both physical material properties and numerical calibration parameters. Due to the large number material card inputs, often there is a lack of uniqueness to MAT162 material cards that limits the predictive capability to only the directly calibrated space. To expand this space, MAT162 requires a prudent and robust calibration process in which significant parameters are calibrated to high confidence damage events observed in experiment. Critical to this success is fully defining the material properties correctly, namely the fiber crush (SFC) and fiber shear (SFS) values, prior to calibrating the numerical parameters. In this paper, the effect of the determination of SFS and SFC on subsequent calibration steps is examined using two different experimental techniques

Implementation of a Matrix Crack Spacing Parameter in a Continuum Damage Mechanics Finite Element Model

Continuum Damage Mechanics (CDM) based progressive damage and failure analysis (PDFA) methods have demonstrated success in a variety of finite element analysis (FEA) implementations. However, the technical maturity of CDM codes has not yet been proven for the full design space of composite materials in aerospace applications. CDM-based approaches represent the presence of damage by changing the local material stiffness definitions and without updating the original mesh or element integration schemes. Without discretely representing cracks and their paths through the mesh, damage in models with CDM-based materials is often distributed in a region of partially damaged elements ahead of stress concentrations. Having a series of discrete matrix cracks represented by a softened region may affect predictions of damage propagation and, thus, structural failure. This issue can be mitigated by restricting matrix damage development to discrete, fiber-aligned rows of elements; hence CDM-based matrix cracks can be implemented to be more representative of discrete matrix cracks. This paper evaluates the effect of restricting CDM matrix crack development to discrete, fiber-aligned rows where the spacing of these rows is controlled by a user-defined crack spacing parameter. Initially, the effect of incrementally increasing matrix crack spacing in a unidirectional center notch coupon is evaluated. Then, the lessons learned from the center notch specimen are applied to open-hole compression finite element models. Results are compared to test data, and the limitations, successes, and potential of the matrix crack spacing approach are discussed

Progressive Damage and Failure Analysis of Bonded Composite Joints at High Energy Dynamic Impacts

Both wing and fuselage structures utilize bonded composite joints for structural efficiency in modern commercial and military aircraft. To ensure compliance with certification requirements mechanical fasteners are typically used as a failsafe mechanism for appropriate strength in the event of complete stiffener disbond. However, the use of fasteners decreases the structural efficiency of the structure by adding weight. This establishes the requirement to better exploit the efficiency of bonded structures and fully understand the failure behavior of adhesively bonded composite structures, particularly when subjected to elevated loading rates due to high energy dynamic impacts (HEDI). For this reason, the NASA Advanced Composite Consortium (ACC) HEDI team developed an experimentation and numerical modeling program for high rate loading of composite joints. In the present work, the response of adhesively bonded composite joints subjected to elevated loading rates is studied numerically and validated against experimental results. Due to dynamic considerations of experiments, the idea of wedge insert was extended to use with Split Hopkinson Pressure Bar (SHPB) testing techniques. Mode-I and Mode-II test configurations were simulated to evaluate the capability of two continuum damage material (CDM) models in LS-DYNA, namely MAT162 and MAT261. Three different levels of fidelity were considered to investigate the level of detail required to numerically predict the failure behavior and the results from high fidelity analysis are presented

Assessment of Intralaminar Progressive Damage and Failure Analysis Using an Efficient Evaluation Framework

Reducing the timeline for development and certification for composite structures has been a long standing objective of the aerospace industry. This timeline can be further exacerbated when attempting to integrate new fiber-reinforced composite materials due to the large number of testing required at every level of design. computational progressive damage and failure analysis (PDFA) attempts to mitigate this effect; however, new PDFA methods have been slow to be adopted in industry since material model evaluation techniques have not been fully defined. This study presents an efficient evaluation framework which uses a piecewise verification and validation (V&V) approach for PDFA methods. Specifically, the framework is applied to evaluate PDFA research codes within the context of intralaminar damage. Methods are incrementally taken through various V&V exercises specifically tailored to study PDFA intralaminar damage modeling capability. Finally, methods are evaluated against a defined set of success criteria to highlight successes and limitations

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

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 Overview of NASA ACC High Energy Dynamic Impact Methodology for Prediction of Ballistic Limit

This presentation gives an overview of the NASA Advanced Composites Project and a summary of the progress and plans of the High Energy Impact Dynamics Team