Material Characterization of Tape, and 2-D/3-D Braided/Woven Compositesf

An important part of certification planning is to conduct complete material characterization to assess composite laminate response considering “as-built” and/or “as-is” material states: manufacturing anomalies (i.e., void size/shape, fiber waviness, interphase coating), design (i.e., ply orientation, thickness, 2D/3D architecture). Composite are being used in aerospace, automotive, wind turbine, ship building, and infrastructure industries. It is important to identify root cause for composite damage and failure thru comprehensive material characterization that involve;

• Reverse engineering of constituent fiber and matrix properties;
• Determination of material design envelope;
• Determination of material response under in-service environments including temperature and moisture loading;
• Generation of carpet plot to provide alternate laminates for design consideration;
• Generation of A- and B-Basis strength allowables;
• Accelerate material selection for use in certification;

Objective

Provide material and structural engineers with verified capability for rapid and accurate assessment of laminate response under in-service loading.

Approach

Use GENOA’s MCQ module integrating micro and macro-mechanics with damage tracking by assessing physics based damage and failure criteria encountered in composites. Material response is interrogated to determine loads that produce matrix cracking under tension, compression, shear, and fiber failure under tension and compression (including crushing, fiber-matrix debonding, and fiber micro-buckling) are evaluated. Analyst has a choice of using fiber/matrix or ply properties as input to the analysis. MCQ does not use finite element theory, rather it uses nit cell approach depicting ASTM specimen with uniform state of stress.

Tools Used

GENOA’s MCQ module are used for:

• Fiber, Matrix, and Lamina Calibration – Reverse engineer effective linear fiber/matrix properties from lamina or laminate test data (strength and stiffness). The effective properties accounts for the thermal residual stresses and interface due to curing process.
• Non-Linear Material Characterization Optimization (MCO) – Reverse engineers effective fiber, matrix, ply non-linear properties (stress strain curves) from ply or from laminate test data.
• Ply Level Analysis – Predicts equivalent ply properties (mechanical/thermal/electrical) using fiber matrix properties as input. Example of mechanical properties (Figure 1) is ply strength in 11, 22, 33, 12, 23, and 13 directions. Example of electrical properties is ply and laminate conductivity.
• Laminate Analysis – Predict equivalent laminate properties using fiber/matrix or ply properties as input. The properties calculated include laminate strength and stiffness, and electrical and thermal properties as well.
• Design Failure Envelope – Predicts design failure envelope for chosen failure criteria for laminates. Strength, strain, and interactive based failure mechanisms are available (Figure 2). Fiber failure under tension/compression including micro-buckling, matrix cracking under tension and compression and delamination (in-plane and out-of-plane) are determined for the ply and the laminate. Several Failure Criteria can be compared for better understanding and comparison against test data.
• Ply Characterization – Graphically shows variation in strength as a function of ply orientation and fiber or void volume ratio.
• A- & B- Basis Allowables – Predicts equivalent ply properties (mechanical/thermal/electrical) using fiber matrix properties as input. Example of mechanical properties (Figure 1) is ply strength in 11, 22, 33, 12, 23, and 13 directions. Example of electrical properties is ply and laminate conductivity.
• Ply Level Analysis – Rapid and accurate prediction of A- and B-basis strength allowables for un-notched uniformly stressed coupons. This module provides the option of predicting allowables from a minimal number of test replicates. With a dedicated sensitivity analysis one can determine the influence of manufacturing parameters and material properties on the laminate strength. This helps reduce the scatter and improve the performance of the material (Figure 3 & 4).
• Parametric Carpet Plots – Carpet plots identify laminate percentages (regions) that are governed by fiber failure dominance, by matrix cracking (transverse and shear) dominance and by combined fiber/matrix failure. The applications and benefits include reducing the number of laminate coupon tests required to verify the entire range of 0°, +/-45° and 90° percentages.

Benefits Provided

• Facilitates material selection and material qualification
• Provides alternate laminate designs
• Generate composite input for commercial FEA solvers (NASTRAN, ABAQUS, ANSYS and others)
• Determine strength allowables with reduced testing
• Assess effect of manufacturing defects and environment on material response
• Reduced testing
• Accelerate certification

Past Experience

Capability was used in qualifying several classes of carbon and glass composites for aerospace and automotive applications.

MCQ Generated Ply Properties

Design Failure Envelope Progression for a [0,90,45,-45] Symmetric Layup

Material Performance Envelope Generated by the Software

Reference
DOT/FAA/AR-03/19, Final Report, “Material Qualification and Equivalency for Polymer Matrix Composite Material System: Updated Procedure” Office of Aviation Research, Washington, D.C. 20591, U.S. Department of Transportation Federal Aviation Administration, September, 2003.

Carpet Plot of Laminate Strength as Function of 0, 45, and 90 degrees Ply Angles (Red = Fiber Failure, Green = Matrix Failure, Blue = Matrix Shear Failure)

Generation of Composite Strength Allowables

A- and B-basis allowable strength values are essential for reducing risk in aircraft structural components made from fiber reinforced polymer composite materials. Risk reduction is achieved by lowering the probability of failure of critical aircraft structures through the use of A- and B-basis design values. For A-basis, at least 99% of the population of material strength values is expected to equal or exceed this tolerance bound with 95% confidence. A-basis strength is applied to single members within an assembly whose failure would result in loss of structural integrity. For B-basis, 90% of the population of material strength values is expected to equal or exceed that strength value with 95% confidence. B-basis strength is applied to redundant structures where failure would result in safe load redistribution. Current practices for determining allowables follow procedures recommended by FAA and working draft of the composite materials handbook CMH-17 Rev G (formerly military handbook Mil-HDBK-17-1F). Determining allowables solely by means of testing [1] is costly and time consuming as large number of composite coupons must be tested under various environments: cold, ambient and elevated temperatures (with and without moisture). AlphaSTAR devised technical approach for calculating allowables with reduced testing. The methodology eliminates unnecessary conservatisms that can contribute to increased weight.

Reference
DOT/FAA/AR-03/19, Final Report, “Material Qualification and Equivalency for Polymer Matrix Composite Material System: Updated Procedure” Office of Aviation Research, Washington, D.C. 20591, U.S. Department of Transportation Federal Aviation Administration, September, 2003.

Objective

Provide material and structural engineers with verified capability for rapid and accurate assessment of laminate response under in-service loading.

Approach

Use lamina level testing per CMH-17 guidelines to reverse engineer sources of uncertainties that produce scatter in material response. The variability is generally caused by: (1) scatter in constituent mechanical properties; (2) variability in composite manufacturing parameters; and (3) manufacturing defects such as void, waviness, and gaps. Then use advanced multi-scale multi-physics progressive failure analysis (MS-PFA) coupled with probabilistic analysis and Bayesian statics update to produce scatter at the laminate level. Lamina level scatter can be used to generate laminate level scatter as indicated in [1]. Data from testing of fewer replicates of composite laminates can be used to guide the determination of allowables. AlpahSTAR developed MCQ (Material Characterization and Qualification) software to determine allowables with reduced testing.

Reference
J. Tomblin, and W. Seneviartne, “Laminate Statistical Allowable Generation for Fiber-Reinforced Composite Materials: Lamina Variability Method”. Report number DOT/FAA/AR-06/53, January 2009. US Department of Transportation, Federal Aviation Administration (Office of Aviation Research and Development), Washington, DC.

Tools Used

MCQ is used to determine allowables for laminate level un-notched coupons following ASTM standards. MCQ is a mesh-less unit cell concept that does not use FEA as long as the state of stress is uniform in the coupon. For all other ASTM coupons (open hole, filed hole, etc.) or general structural component design allowables are determined using MS-PFA invoking finite element solution with progressive failure analysis.

Benefits Provided

• Reduce laminate level testing for the first level of FAA’s building block where testing of a large number of coupons becomes inevitable; test reduction can be on the order of a minimum of 50% as compared to current standards.
• Determine A-basis from B-basis.
• Identify root cause for scatter in material strength.
• Provide guidelines for improving material behavior.
• Applicable to determination of B-basis at any level of FAA’s building block (coupon, element, subcomponent, etc.)

Past Experience

Technology was applied successfully to carbon and glass composites (tape and fabric). The approach consistently provided accurate allowable values. Additionally, AlphaSTAR published extensively on the topic with Joint Publications with key Aerospace Industries and University Partners:

Reference
[2] G. Abumeri, M. Garg, F. Abdi, A. McCloskey and R. Bohner, “Validation of a Computational Approach for Composite Material Allowables Using Sealed Envelope Predictions for Reduced Testing,” SAMPE Journal, September/October 2009.

Objective

It is highly desirable to be able to simulate structural components made of these advanced materials subjected to static, fatigue or impact loading conditions. It is also desirable that the simulation be done without any extensive finite element modeling or several expensive coupons level test data. Recall that usual finite element approach requires extensive modeling efforts whereby the engineer models all the constituents without knowledge of complete set of mechanical material properties, or homogenized model ignoring any possible fiber matrix level root cause issues.

Approach

Since in advance complex material systems the root cause for damage initiation is the constituents, it is vital that constituent level damage evolution be accounted accurately during the load increment without extensive modeling efforts. We recommend the following approach for simulating such problems:

• Characterize the in-situ linear/nonlinear constituent material properties obtained using simple flat coupon test data
• Perform D&DT analysis using less extensive finite element models combined with progressive failure and advanced 3 dimensional micro-mechanics theories at each incremental loading to evaluate the damage progression at the constituent level.

Tools Used

Material characterization is done using the following tools:

• Conventional composites (PMC, CMC, MMC): Material Characterization & Qualification (MCQ)
• Nanocomposites: MCQ Nano
• Metals: Fracture Toughness Determination (FTD) and Fatigue Crack Growth Curve (FCG)

Benefits Provided

• Reduce constituent level testing using expensive scientific tests, for example, Atomic Force Microscope.
• Generate scatter at laminate level by virtually introducing the variation in material properties
• Reverse engineer the constituent properties.
• Provide guidelines for improving material behavior.
• Identify root cause problems for D&DT analysis without extensive FE models

Past Experience

Joint Publications with key Aerospace Industry and University Partners:

Honeycomb Structure

Ceramics

Challenge

Modeling manufacturing defects by means of finite element analysis (FEA) is impractical. The modeling of voids and fiber waviness is very time consuming. Additionally, use of FEA will introduce other challenges as some voids, when modeled, can cause stress concentration issues. Micro-mechanics based approach is the most efficient approach to model and analyze effects from composite manufacturing defects.

Implications

Not accounting for the deterministic and probabilistic nature of composite manufacturing defects can increase the risk associated with the use of composite sin critical structures such as those used in aerospace. For example, manufacturing defects can reduce the stability of composite structures and reduce their impact resistance. To account for such defects, designers tend to reduce the strength allowables used in design, resulting in increased weight and cost.

How we address it

AlphaSTAR devised a numerical approach combining multi-scale progressive failure analysis to predict the in-situ degraded ply properties containing voids of different sizes and shapes and gaps.

GENOA’s PFA module allows the insertion of manufacturing defects in composite laminates for use in progressive failure analysis.

Benefits Provided

• Simplified practical approach for modeling of manufacturing defects
• Determination of equivalent ply properties with defects for use in multi-scale progressive failure analysis
• Capture damage and fracture evolution in presence of manufacturing defects
• Determine loss in strength and stiffness due to manufacturing defects
• Reduce number of tests by validating coupons tests with defects then use approach to assess effects at component level

Testimonials

AlphaSTAR assessed risks associated with composite manufacturing defects for aerospace and automotive industries.

Reference
J. Sawicki, P. J. Minguet, “The Effect of Interply Overlaps and Gaps upon the Compression Strength of Composite Laminates” 39th AIAA/SDM Conference Paper, AIAA-98-1786, Long Beach, CA, 1998

Damage marked in red in open hole laminate under compression with gap defects obtained from multi-scale PFA