This work presents a novel computer graphics-based methodology for the efficient generation of realistic threedimensional micromodels of unidirectional carbon fiber-reinforced polymers (CFRPs) with stochastic fiber misalignment. Implemented in the open-source software Blender, the approach enables rapid creation of densely packed fiber architectures with controlled global and local misalignment while strictly preventing fiber interpenetration at high fiber volume fractions. The framework includes an automated workflow for transferring the generated fiber geometries into three-dimensional (3D) fiber-segmented finite element (FE) representative volume element (RVE) models for analysis in Abaqus, with explicit representation of the fiber, matrix, and fiber-matrix interface. To demonstrate the capabilities of the proposed method, a numerical study is conducted to investigate the influence of the local fiber misalignment phase distribution on fiber-direction compressive strength. Threedimensional fiber-segmented RVEs with identical misalignment amplitudes but different phase-angle distributions are analyzed under compressive loading. The results show that models with random misalignment phase angles exhibit delayed matrix shear localization and kink-band formation compared to models with uniform misalignment phase. On average, the predicted fiber-direction compressive strength for RVEs with random phase distributions was found 59 % higher than that obtained using uniform misalignment assumptions. The framework provides a scalable foundation for synthetic generation of realistic fiber architectures, supporting integration into Integrated Computational Materials Engineering (ICME) pipelines and machine-learning-based composite design workflows.

Introduction: Fatigue damage remains a major challenge in the design of composite structures, as cracks may initiate under low stress levels and evolve through multiple interacting failure mechanisms. To reduce reliance on extensive physical testing, finite element methods based on progressive damage analysis are increasingly used to predict fatigue-driven crack growth within structural substantiation workflows. However, most existing approaches require prior knowledge of the crack path, limiting their ability to represent complex, solution-dependent fracture trajectories. The eXtended Finite Element Method (XFEM) offers a promising alternative, as cracks can be modeled independently of the mesh, yet commercial implementations remain limited for fatigue loading.Materials and methods: This work introduces a fatigue-capable XFEM framework that combines the cohesive segments approach with a stress–life-based degradation law using only Abaqus built-in user subroutines and native elements. Crack initiation is governed by a fatigue endurance limit criterion, and damage evolution is driven by a user-defined internal variable that controls cohesive stiffness degradation along the crack interface. A dedicated bookkeeping strategy enables access and transfer of interfacial crack-opening quantities, ensuring consistent evaluation of the fatigue damage variable across the enriched interface.Results: The framework is verified against a classical Cohesive Zone Model (CZM) and applied to a two-dimensional single element model as well as Double Cantilever Beam (DCB) specimen under Mode I static and fatigue using the IM7/8552 carbon/epoxy material system. The results demonstrate excellent convergence between the proposed method and CZM approach for the single element model. Furthermore, the DCB results demonstrate accurate prediction of delamination growth and highlight the proposed formulation as a practical and more flexible alternative to the existing Virtual Crack Closure Technique (VCCT)-based XFEM fatigue capability in Abaqus, without requiring a pre-existing crack or inheriting the restrictions of linear elastic fracture mechanics.Conclusions: The proposed methodology effectively bridges cohesive zone modeling and XFEM, providing a tool for simulating progressive damage in composite structures without the limitations imposed by linear fracture mechanics. Future developments will extend the framework to mixed-mode loading and three-dimensional configurations, broadening its applicability to aerospace-grade laminates and structural components.

This work presents the development and application of a methodology for predicting fatigue life, implemented within the modern progressive failure analysis software tool CDMat, developed at the Advanced Materials and Structures Laboratory of the University of Texas at Arlington. CDMat is designed as an extension to the general-purpose finite element analysis program ABAQUS/Explicit. The set of user-defined subroutines for describing material behavior can be expanded by adding new subroutines. A recent development in CDMat is a computational model capable of predicting delamination crack growth under quasi-static and fatigue loading, based on a fracture mechanics approach using the J-integral. The J-integral is calculated by integrating stresses and displacements along a line defined by the negative gradient of displacements in the cohesive interface. Due to the large integration path, the J-integral allows for a highly accurate estimation of the energy release rate, which makes it possible to reasonably estimate the crack growth rate using Paris's law. A comprehensive verification of the J-integral-based fatigue prediction methodology was performed using a tapered structural element with internal ply drops. Experimental determination of the static and fatigue properties of the materials, static and fatigue tests of a tapered structural element, and numerical simulation of fatigue crack growth were performed. The predicted fatigue crack growth showed good agreement with experimental results.

Advanced structural analysis methods, known as progressive damage and failure analysis tools, are being developed to predict initiation and propagation of damage under repeated loading based on capturing individual and interacting damage modes. This work develops structural fatigue life prediction capability in state-of-the-art emerging progressive damage failure analysis tool CDMat developed at the University of Texas Arlington Advanced Materials and Structures Lab. While JIntegral, implemented in CDMat, appears as the most objective and rigorous approach to predict delamination growth-based fatigue life of composite structures, the key material properties of the J-Integral fatigue model have not been measured with the adequate accuracy. This work addressees a fundamental challenge of eliminating the established and routine assumptions and developed a methodology to determine the key material properties meeting the material input data requirements for the JIntegral based structural fatigue life prediction models. This work generated input data for fatigue crack growth propagation, including optimized input data parameters for the in-situ cohesive law and modeling as-manufactured specimen conditions. In addition, a methodology to account for and characterize the effects of fiber bridging in static tests is presented. It uses a simple standard unidirectional panel for testing, before a component is manufactured, fabricated from the same batch of prepreg to qualify the effects of fiber bridging. Also, this work attempted to determine a minimum conservative initial crack size to streamline the fatigue crack propagation prediction. Fatigue predictions have been demonstrated on a representative composite skin–hat stiffener sub-component section and compared with tests.

A state-of-the-art emerging progressive damage failure analysis tool CDMat has been successfully applied to multiple material systems on open-hole tension and compression, and double shear bearing laminate coupons under static and fatigue loading including simulation to ultimate failure. CDMat also successfully demonstrated component-level strength/fatigue analysis under the Air Force Composite Airframe Life Extension (CALE) and the Fail-Safe Technologies for Bonded and Unitized Composite Structures (FASTBUCs) Programs. Building on the success of CDMat an integrated software solution for certification and sustainment of rotorcraft primary composite structures is being developed. A method and an algorithm for fatigue crack growth simulation in laminated structures are proposed to improve the accuracy of CDMat fatigue predictions. The method is based on using cohesive material model, tracking material points at the crack front, and calculating the pointwise energy release rate employing the J-integral. The algorithm was implemented as a set of user material subroutines developed within the framework of explicit finite element formulation for ABAQUS. The effectiveness of the method is demonstrated on several examples of Mode I and II fatigue crack growth.

Composite materials are increasingly used in modern aircraft to enhance efficiency by reducing weight. However, the inherent complexity of failure mechanisms and susceptibility to manufacturing irregularities pose challenges, causing delays in the deployment of composite structures. To accelerate qualification and reduce reliance on physical tests, validated analysis techniques that capture initiation and progression of structural damage are needed. Fatigue delamination in primary aircraft composite structures is of particular concern, as it can initiate at low cyclic load levels and delamination growth can potentially lead to catastrophic failure. The development of accurate and efficient numerical models for predicting fatigue damage growth remains an active research area. This paper presents a non-local approach based on cohesive elements and the computation of the J-integral, utilizing a Paris law input for predicting the progression of fatigue-driven delamination. A Double Cantilever Beam (DCB) model is considered as benchmark for fatigue progressive damage analysis under mode I loading. Fatigue crack growth results obtained using the non-local approach are compared with those derived from a legacy local fatigue damage model. The robustness of the non-local approach, notably not requiring additional calibration parameters in contrast to the local approach, is demonstrated.

Adhesively bonded Polymeric Matrix Composite (PMC) structures could offer significant cost savings and increase production rates in the fabrication and assembly of future composite airframes. In particular, bonded PMCs may eliminate the use of thousands of mechanical fasteners and allow incorporation of multiple aircraft composite primary structures into a unitized structural design. However, increased susceptibility to defects and manufacturing irregularities, including voids, disbonds, and geometry variations, as well as complex interacting failure mechanisms, hinder a widespread use of bonded PMCs. Inherent susceptibility to manufacturing defects contributes to higher variability of part quality than found in structures manufactured using fasteners. The presence of microstructural defects at critical locations may significantly reduce strength and useful life of adhesively bonded structures. At the same time, existing methods for simulation of the performance of bonded PMCs are based on homogenization and similitude of material response and an incomplete physical understanding of the failure processes involved. Such approaches have proven unable to capture the effects of defects on structural performance and explain large data scatter commonly observed. Hence, these methods do not provide the level of confidence needed for qualification and certification of bonded interfaces. Certification of adhesively bonded interfaces currently used in primary aircraft PMC structures requires installation of fasteners to produce a redundant load path. This redundancy greatly increases production time and expense, and requires additional inspection and sealing. The need to develop a high-fidelity computational materials-based capability to support next-generation qualification and certification of bonded PMCs and safely reduce redundant fasteners has been identified under NASA Transformational Tools and Technologies (TTT) program. As part of such effort, this work presents the current work-in-progress towards the development of an integrated methodology to enable improved understanding of the failure mechanisms that govern performance along a bonded PMC interface in presence of manufacturing defects. The approach is envisioned to integrate (1) accurate 3D measurement of bondline defects and irregularities, including X-ray Computed Tomography (CT) data; (2) allow for input of material properties from experimental evaluation meeting the material input data requirements; and (3) conduct a finite element-based progressive damage and failure analysis (PDFA) capturing the multiple damage modes and their interaction at the bonded interface, and predicting the structural behavior. Preliminary results are presented using a continuum description of the adhesive material in combination with the eXtended Finite Element Method (XFEM) for prediction of cohesive failure within the adhesive. The approach is implemented in commercial FE software Abaqus for simulation of bondline fracture in adhesively bonded mode I Double Cantilever Beam (DCB) PMCs specimens.

Microscale residual stress may develop during the manufacturing of Carbon Fiber-Reinforced Polymer (CFRP) composites and negatively affect apparent macroscale mechanical properties. Accordingly, accurately capturing residual stress may be essential in computational methods used for composite material design. This work presents a new data-driven methodology for the evaluation of microscale residual stress in CFRPs using fiber push-out experiments with in situ scanning electron microscopy (SEM) imaging. SEM images reveal significant through-thickness matrix sink-in deformation in resin-rich areas after nearby fibers are pushed out, which is attributed to the release of microscale process-induced residual stress. The sink-in deformation is measured experimentally, and a Finite Element Model Updating (FEMU) method is used to retrieve the associated residual stress. The finite element (FE) analysis includes simulation of the curing process, test sample machining, and fiber push-out experiment. Significant out-of-plane matrix deformation larger than 1% of the specimen thickness is reported and associated with a high level of residual stress in resin-rich areas. This work emphasizes the importance of in situ data-driven characterization for integrated computational materials engineering (ICME) and material design.

The rotorcraft industry has shown a strong interest in High-modulus (HM) carbon fiber-reinforced polymers (CFRPs) due to their potential to create lightweight airframes and rotor components, resulting in significant weight reduction. However, a major drawback of HM CFRPs has been their very low compressive strength in the fiber direction compared to the currently used intermediate-modulus (IM) CFRPs in primary structures. This weakness has been delaying the implementation of HM CFRPs in aircraft structures. Microstructural tailoring may provide an innovative means for breaking through the fiber-direction compressive strength barrier of the HM CFRPs. The primary failure mechanism in both HM and IM CFRPs under fiber-direction compression is shear microbucking, which is significantly influenced by fiber-matrix interface strength. In-depth analysis using in-situ scanning electron microscopy (SEM) experiments has revealed significant differences in the surface characteristics of the carbon fibers, leading to a much stronger interface for IM fibers compared to HM fibers. These findings have prompted a microstructural tailoring strategy involving the reinforcement of HM fibers with IM fibers, which enhances the overall stability of the microstructure governing the fiber-direction compressive strength performance of the material. A laboratory-scale production-quality manufacturing system has been delivered, and promising experimental results enabling HM CFRPs with adequate fiber-direction compressive strength have been achieved through hybridization of IM and HM fibers at the filament level in HM CFRP toughened with nano-silica. This new material solution not only approaches the compressive strength of IM CFRPs but also provides more than 30% higher axial modulus.

This work is a part of an on-going effort to develop effective methods and algorithms enabling automated conversion of X-ray Computed Tomography (CT) scan data into finite element (FE) models for as-built composite parts. In particular, this paper is focused on the automated methods of segmentation of composite plies in X-ray CT scans and finding resulting ply waviness as well as local fiber directions and ply orientations, including average fiber misalignment and fiber direction variability. Sample applications of the proposed algorithm are provided for as-manufactured tape composite laminates as well as additively manufactured composites. Effects of X-ray CT scan quality on the algorithm performance are demonstrated.

High-modulus (HM) carbon fiber-reinforced polymers (CFRPs) have attracted strong demand by the rotorcraft industry as such materials can potentially enable lightweight airframes and rotor components with significant weight savings. However, low fiber-direction compressive strength, compared to intermediate-modulus (IM) CFRPs currently used in primary structures, has been a well-recognized weakness of HM CFRPs, prohibiting their implementation in rotorcraft platforms. Microstructural tailoring provides an innovative means for breaking through the fiber-direction compressive strength barrier of the HM CFRPs. Microbuckling, the fiber-direction compressive failure mechanism of the subject HM and IM CFRPs, is driven by fiber-matrix interface shear strength. Assessment of the interface properties using in-situ SEM-based experiments reveal substantial difference in surface topology between IM and HM fibers, which is related to higher interface strength in HM fibers. This instigates a microstructural tailoring approach of reinforcing material surrounding HM fibers with IM fibers to improve microstructural stability. A manufacturing system has been developed, and promising results enabling HM CFRPs with adequate fiber-direction compressive strength have been achieved through hybridization of IM and HM fibers at the filament level in HM CFRP toughened with nano-silica. A new material achieving compressive strength of IM CFRPs but with >30% higher modulus has been developed.

High-resolution X-Ray computed tomography have demonstrated a unique ability to accurately detect and quantify critical subsurface irregularities, such as voids, waviness, and service damage in composites. While significant progress has been made in the respective areas of characterization and failure predictions, integrating 3D material characterization and non-destructive inspection of composites at the coupon scale into a comprehensive methodology applicable to failure prognosis of composite structures has yet to be developed. Such integration is especially important for the additively manufactured continuous fiber-reinforced thermoplastics that are often built with specific microstructural irregularities, such as voids, ply and fiber waviness, geometric errors, and variability of individual layers. The objective of this work is to provide a comprehensive description of novel modeling techniques used for the automated generation of structured mesh FE models for composite laminates that include as-designed and as-manufactured conditions while maintaining high-fidelity definition of model properties needed for the accurate predictions of structural performance.

Adhesively bonded Polymeric Matrix Composite (PMC) structures could offer significant cost savings and increase production rates in the fabrication and assembly of future composite airframes. In particular, bonded PMCs may eliminate the use of thousands of mechanical fasteners and allow incorporation of multiple aircraft composite primary structures into a unitized structural design. However, increased susceptibility to defects and manufacturing irregularities, including voids, disbonds, and geometry variations, as well as complex interacting failure mechanisms, hinder a widespread use of bonded PMCs. High-fidelity computational materials-based analysis methods that can capture the effect of defects and improve the understanding of failure mechanisms that govern bondline failure are needed to support qualification and certification of next-generation bonded PMCs and safely reduce redundant fasteners. The objective of this work is to contribute to such development. A continuous-discontinuous framework is considered for capturing quasi-brittle failure within the epoxy adhesive and predicting growth of arbitrary solution-dependent crack paths in presence of bondline defects. In this approach, an elasto-plastic constitutive model with isotropic damage is enhanced with the introduction of extended finite element method (XFEM) enrichments that incorporate non-local features. The implementation of the approach in commercial FE software Abaqus using built-in XFEM capabilities in combination with a set of user-defined subroutines is presented and illustrated in two-dimensional FE models representative of adhesively-bonded double-cantilever beam (DCB) specimens.

Carbon-fiber-reinforced polymers (CFRPs) enable lightweight, strong, and durable structures for many engineering applications including aerospace, automotive, biomedical, and others. High-modulus (HM) CFRPs enable the most significant improvement in mechanical stiffness at a lower weight, allowing for extremely lightweight aircraft structures. However, low fiber-direction compressive strength has been a major weakness of HM CFRPs, prohibiting their implementation in the primary structures. Microstructural tailoring may provide an innovative means for breaking through the fiber-direction compressive strength barrier. This has been implemented by hybridizing intermediate-modulus (IM) and HM carbon fibers in HM CFRP toughened with nanosilica particles. The new material solution almost doubles the compressive strength of the HM CFRPs, achieving that of the advanced IM CFRPs currently used in airframes and rotor components, but with a much higher axial modulus. The major focus of this work has been understanding the fiber-matrix interface properties governing the fiber-direction compressive strength improvement of the hybrid HM CFRPs. In particular, differences in the surface topology may cause much higher interface friction for IM carbon fibers compared to the HM fibers, which is responsible for the interface strength improvement. In situ Scanning Electron Microscopy (SEM)-based experiments were developed to measure interface friction. Such experiments reveal an approximately 48% higher maximum shear traction due to interface friction for IM carbon fibers compared to the HM fibers.

Aircraft composite parts are commonly manufactured from resin‐saturated thermoset pre‐impregnated plies laid‐up over a rigid tool and consolidated and cured in an autoclave. In many applications, autoclave consolidation alone is not sufficient to remove the air, or bulk, that has been entrapped within the laminate during layup. Entrapped bulk can lead to the formation of defects, including wrinkles and voids, which can significantly affect structural performance. Vacuum consolidation, or “debulking,” has been a standard practice extensively used by aircraft manufacturers to reduce the amount of bulk and mitigate defect formation. Yet, the underlying physical principles governing formation of defects during these early stages are not well understood. This work presents the development of a new finite element‐based method for simulation of debulking with the objective to contribute to a better understanding of the key physics involved. The model includes pore‐pressure cohesive elements inserted at ply interfaces for discrete representation of entrapped air pockets and modeling air flow during debulking; and cohesive contact between the plies for simulating the typical tacky behavior of uncured thermoset tape prepregs. The debulking of a two‐ply laminate with an initial seeded wrinkle is considered for illustration of the approach and comparison with experimental results.

Aircraft composite parts are commonly manufactured from resin-saturated thermoset pre-impregnated plies laid-up over a rigid tool and consolidated and cured in an autoclave. In many applications, autoclave consolidation alone is not sufficient to remove the air, or bulk, that has been entrapped within the laminate during layup and may lead to the formation of voids in the final part. Entrapped air at ply interface can also participate in the formation of ply wrinkles, especially in curved laminates. High bulk content in curved sections is associated with an excess of ply length material that builds up during the layup. During compaction, this excess ply length may be squeezed into a tighter radius, which might lead to ply buckling and formation of wrinkles. Wrinkles and voids may significantly affect structural integrity of composite structures and increase rejection rates in the production cycle. Debulking, or vacuum consolidation, has been a common practice extensively used by aircraft manufacturers to reduce the amount of air entrapped during the layup of composite parts prior to curing. Yet, the underlying physical principles governing formation of defects during debulking are not well understood. This work is part of the Office of Naval Research (ONR) project “Physics-Based Composite Process Simulation” that seeks to fill the gaps in understanding the physical principles governing the formation and evolution of manufacturing defects. Recently, the authors of this work have introduced an approach for modeling debulking in resinsaturated prepregs where cohesive elements enriched with pore-pressure degree of freedom are inserted at ply interfaces, using simulation concepts originally developed for the analysis of hydraulic fracturing in geomechanics problems. Further improvement of the methodology combined the pore-pressure cohesive zone model with cohesive contact to capture the tacky behavior of the interface. This work presents a verification of the ability of the method to capture the physics involved during debulking at the coupon scale using a two-ply carbon/epoxy IM7/8552 debulk test specimen with a seeded wrinkle. The methodology is then further extended and verified at the element scale for the first time for simulation of debulking in a rotor blade grip test element.

Accurate assessment of fiber-matrix interface shear strength may play key role in understanding structural strength behavior of polymer-matrix composites, and provide additional flexibility for enabling high-performing material designs. This work presents a method for assessment of fiber-matrix interface strength in carbon fiber reinforced polymers (CFRPs). The method uses in-situ scanning electron microscopy (SEM) imaging while pushing individual fibers out of 20–30 μm thick membranes. Femtosecond laser machining has been utilized to address standing challenges associated with manufacturing high-quality fiber push-out specimens. To ensure that the specimen preparation method used did not affect the properties of the composite, a thorough study has been conducted using benchmark specimens. The benchmark specimens are prepared using a conventional lapping/polishing process. Results for two CFRPs compare fiber-matrix interface shear strength generated using free-standing samples (conventional lapping/polishing preparation) and femtosecond laser-machined specimens. Both specimen types exhibit similar strength. However, load-displacement response of the free-standing samples is overly compliant, representing specimen response rather than fiber-matrix interface material behavior. The excessive compliance is caused by inherent imperfections in the contact of the sample and the test fixture. The femtosecond laser-machined samples are free of such imperfections. Finite element analysis has been accomplished to better understand the load-displacement response.

Advanced structural analysis methods, known as progressive damage and failure analysis tools, are being developed to predict initiation and propagation of damage under repeated loading based on capturing individual and interacting damage modes. This work shows the ability of the progressive damage and failure analysis method implemented in CDMat software developed at the University of Texas Arlington Advanced Materials and Structures Lab to predict strength and fatigue failure of an advanced mechanically fastened aerospace structural joint, the common feature test component (CFTC)— representative of flight-critical structural attributes and failure modes—without a priori knowledge of the test result. The CFTC structural features include a composite tape skin, a composite fabric stiffener, and an aluminum rib applying pullthrough load through multiple countersunk bolts combined with the axial compression of skin and stiffener. Failure and damage predictions under static and constant-amplitude cyclic loading are compared with tests. Developed by Boeing under the Air Force Research Laboratory Composite Airframe Life Extension Program, the CFTC has been the most complex progressive damage and failure analysis validation article to date.