Stephen P. Engelstad

Comparison of composite damage growth tools for static behavior of notched composite laminates

This paper provides overall comparisons of the static results of an Air Force Research Laboratory exploration into the state of the art of existing technology in composite progressive damage analysis. In this study, blind and re-calibration bench-marking exercises were performed using nine different composite progressive damage analysis codes on unnotched and notched (open-hole) composite coupons under both static and fatigue loading. This paper summarizes the results of the static portion of this program. Comparisons are made herein of specimen stiffness and strength predictions against each other and the test data. Overall percent error data is presented, as well as a list of observations and lessons learned during this year-long effort.

Failure simulations of open-hole IM7/977-3 coupons subjected to fatigue loading using Autodesk Helius PFA

Finite element simulations of three laminates in open-hole configuration subjected to constant amplitude tension–tension fatigue loading are investigated as part of the Damage Tolerant Design Principles program organized by the Air Force Research Laboratory. All coupons were made from unidirectional IM7/977-3 plies, which are composed of intermediate modulus carbon fibers and a toughened epoxy matrix. Government furnished experimental data from an assortment of fatigue loaded unnotched coupons were used to characterize the behavior of the composite material in the simulations. The commercial software Autodesk Helius PFA was used to model the non-linear response of the material. Blind simulations of coupon stiffness and damage at several cycle numbers and residual coupon tensile and compressive strengths are benchmarked against experimental measurements and X-rays. Upon review of the experimental results, a second round of simulations was performed where the modeling strategy was updated to improve correlation to experiment.

High Fidelity Composite Bonded Joint Analysis Validation Study Part II: Analysis Correlation

The analysis/prediction phase of a high fidelity analysis validation activity conducted by the Composites Affordability Initiative (CAI) Program on composite bonded structure was presented in a prior paper and the correlation to test results are presented here. The CAI program has developed two improved approaches for the analysis of bonded joints, (1) high fidelity stress analysis combined with the incorporation of the new composite strain invariant failure criteria for the prediction of failure initiation, and (2) an interface fracture finite element (IFE) for predicting the delamination propagation and residual structural strength. The test articles are that of an all composite skin and stiffener run-out geometry with both bonded and co-cured configurations. The focus of this paper is on analysis/test correlation results and discussion of the findings. Extraordinary means were employed to detect the early failure events to be consistent with the fidelity of the analysis methods. The interpretation of sensors readings was not straight forward and the led to a question of how one defines “first failure” in a bonded composite joint where the scale of failure occurs on a level below the size of features that are being controlled during the manufacturing process. The IFE analysis displayed good correlation with test results in regards to the extent of delamination growth and prediction of the major load drop associated unstable delamination growth. Analysis of z-pin reinforced stiffener terminations using the CAI developed Interface Traction Elements (ITEs), a type of cohesive element, required modification of the analysis to get reasonable agreement. Both types of analyses, strength based and fracture based approaches, provided exceptional insight into the performance of the structure.

Failure simulations of open-hole and unnotched IM7/977-3 coupons subjected to quasi-static loading using Autodesk Helius PFA

Finite element simulations of three laminates in open-hole and unnotched configurations subjected to tension and compression quasi-static loading are investigated as part of the Damage Tolerant Design Principles program organized by the Air Force Research Laboratory. The coupons are made from unidirectional IM7/977-3 plies, which are a composite material composed of intermediate modulus carbon fibers and a toughened epoxy matrix. Blind simulations of coupon stiffness, nominal coupon stress at failure and damage evolution are benchmarked against experimental measurements and X-rays. The blind simulations are followed by a second round of simulations where the modeling strategy is modified to improve agreement between the simulations and experiments. In the present article, the commercial software Autodesk Helius PFA is used to model the non-linear response of the composite material. Within Helius PFA, failure is evaluated at the constituent level by extracting the fiber and matrix volume average stress state from the homogenized composite stress state. The relationships between the composite and constituents are developed using multicontinuum theory and a high-fidelity micromechanics model.

Crack Arrestment of Pi-Joined Bonded Composites

This paper describes current results of an ongoing program to develop a test verified analysis method for crack arrestment of bonded structures. Bonded specimens have been tested with and without fasteners, which include a bonded wide pull-off pi-joint specimen (Mode I), and a bonded double shear I-beam pi-joint specimen (Mode II). Finite element models utilizing the Virtual Crack Closure Technique (VCCT) for crack progression have been constructed to simulate the crack arrest capability of fasteners in the bondline of these composite joints. Promising comparisons were achieved between the models and tests, showing the ability of the damage modeling technique to predict the residual strength. However, the model and test comparisons have also shown that although the interlaminar fracture failure modes are necessary for model prediction of crack arrest behavior, additional failure modes occurred in the tests that will need to be included in the simulations.

Crack Arrestment of Bonded Composite Joints

This paper describes current results of an ongoing program to develop a test verified analysis method for crack arrestment of bonded structures. Bonded specimens have been tested with and without fasteners, which to-date include a bonded double cantilever beam (DCB) (Mode I), double shear (Mode II), and a DCB-like pi-joint (Mode I) specimen. Finite element models utilizing the Virtual Crack Closure Technique (VCCT) for crack progression have been constructed and correlated to the tests, utilizing fracture toughness data measured using all zero and bonded layup fracture toughness tests. Utilizing the VCCT capability within ABAQUS, fasteners have been shown to arrest a crack growing in the DCB and double shear specimens, as well as cracks in the DCB pi-joint specimen. Tests of the DCB and Double Shear specimens verified these results, but further investigations are needed to resolve fastener pre-load force and crack face friction effects, especially for the Double Shear. Promising comparisons were achieved between the pi-joint DCB model and test data. This project is ongoing, and additional pi-joint Mode I and Mode II specimens are planned for future evaluations.

Benchmarking of composite progressive damage analysis methods: The background

This article introduces an Air Force Research Laboratory study, which performed static and fatigue benchmark exercises for nine composite progressive damage analysis methods. Air Force Research Laboratory is interested in exploring the feasibility of these progressive damage analysis methods to predict composite damage growth for the purposes of improved durability and damage tolerance analysis of composite aircraft structure. This article gives the background, goals, motivation, and guiding principles of the study and provides brief descriptions of the teams that participated and the tools that were utilized.

Integrated Computational Materials Engineering for Airframe Composite Structure Applications

Introduction Integrated Computational Materials Engineering (ICME) is a game-changing AFRL and industry vision to reduce the material and process development cycle time and cost, simultaneously bringing optimized material systems to the war fighter tailored to the needs of both the airframe and propulsion systems [Ref: National Materials Advisory Board]. The nearterm path to achieving these goals is through integration of material modeling capabilities. AFRL is currently working on two Foundational Engineering Problems (FEPs), one for metallic aircraft applications, and one for composites. GE Aviation and Lockheed Martin Aeronautics (LM Aero) have teamed to work the composites FEP, called “Integrated Computational Methods for Composite Materials (ICM2)” specifically targeting integration of composites processing, micromechanics, and damage progression modeling codes to address composite material development and application issues. GE is focused on engine applications, whereas LM Aero is focused on airframe applications. For the airframe specific ICM2 FEP, LM Aero is targeting the fundamental issues that drive the design of acreage composite materials on the next generation airframes. In order to meet composite airframe future needs, large scale airframe manufacturing will target larger, unitized composite assemblies with increased use of bonding and reduced part-count. Process automation will be utilized to reduce costs through reductions in touch labor. Improvements in composite design allowables are critical to optimal airframe weight (and hence performance) and must be obtained through use of higher performance resins and fibers along with reduced variability in key sizing properties. The ICM2 program intends to integrate composite process and design modeling codes to streamline the development cycle time and reduce the cost to implement such new high performance materials on next generation aircraft. For the ICM2 program’s demonstration purposes LM Aero is studying the IM7/M65 bismaleimide (BMI) system for application to large acreage wing skin and web applications. M65 is an established BMI system (MRL ≥5) well suited to manufacturing using high speed automated fiber placement (AFP). BMI systems have experienced increased usage on fighter aircraft due primarily to key structural design properties such as open hole compression (OHC) and compression-strengthafter-impact (CSAI), the values for which exceed epoxies at max service temperature and moisture conditions [Rousseau et al.] These key properties often “size” the acreage of the aircraft composite skins. Bolted joint strength and acreage repair criteria are most closely related D ow nl oa de d by U ni ve rs ity o f M ic hi ga n D ud er st ad t C en te r on D ec em be r 14 , 2 01 7 | h ttp :// ar c. ai aa .o rg | D O I: 1 0. 25 14 /6 .2 01 501 98 56th AIAA/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference 5-9 January 2015, Kissimmee, Florida 10.2514/6.2015-0198