Progressive Damage Analysis of Composite Laminates Under Compression

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Introduction

When repairing with adhesive, the repair method with scarf and externally bonded plasters is usually preferred [19]. However, the behavior of single-sided repairs is quite complex, as an additional bending effect occurs due to the shift in the neutral axis of the repaired panel.

Figure 1.1: Breakdown of materials used in Boeing 787 Dreamliner [1]

Literature Reviews

44] had studied the performance of the notched and double-sided repaired composite panel using progressive failure analysis. 48] conducted experimental and numerical investigations on progressive damage analysis in externally bonded patch repaired for the single- and double-sided CFRP laminates, respectively.

Motivation, Scope and Objectives

Most of the reported work is on progressive damage modeling and stress analysis of bonded repaired panels of composite laminates under tensile loading. The accuracy of the developed model is assessed by comparing the numerical prediction with experimental results.

Thesis layout

An extension of the research on open excisions is the repair of composite panels by means of adhesively bonded external patch. As the demand for the composite application increases significantly space and the composite material tends to deteriorate with time and service.

Introduction

Earliest test methods for Compression test

To improve this limitation, the next development came in the form of the Celenese device, shown in fig 2.1(f). This fixation introduced axial forces on the specimen by loading the sides of the specimen near the ends with shear.

Figure 2.1: Various Compression Test Fixtures [11] (a) TEI fixture; (b) Narmco Test Method 303; (c) Sandwich Stanilized fixture; (d) ASTM D695-69 fixture; (e) Federal Test Method 406 fixture; (f) Celanese fixture [3]

Present Methods of Composite Compression Testing

The first Type II test method involves the use of the Southwest Research Institute (SWRI) [8] sample setup as shown in Fig. The portion of the force initially transferred into the support is transferred into the test specimen by shear.

Figure 2.3: Exploded view of the IITRI test fixture (ASTM D 3410) [4]

Comparison of SACMA and UCSB Test Method

The SACMA and UCSB facilities were used to test quasi-isotropic laminated CFRP specimens without a notch. Once the unnotched specimen is mounted, the window in the center of the fixture can be used to view the final failure of the specimen. It can be observed that the failure strength of the quasi-isotropic specimens using the SACMA and UCSB clamp test method is approximately the same, i.e.

Open hole compression (OHC) testing with SACMA and UCSB fixture The SACMA and UCSB fixture were used to test unnotched CFRP quasi-isotropic S-samples. When an OHC specimen is installed, the window in the center of the fixture can be used to place strain gauges to determine the strain at break, as shown in Figure 1.16.

Figure 2.9: SACMA Test Fixture (ASTM D 6484) [10]

Closure

The experimental stress-strain curve up to ultimate damage for the OHC specimen under compressive load was obtained with the help of strain gauges. During the experiment, in each specimen the data is tracked at 2 different locations using strain gauges as shown in Fig. Strain gauge 1 is attached to the location in the transverse direction at a distance of 10 mm from the hole while strain gauge 2 is attached in the longitudinal direction at a distance of 15 mm from the hole as shown in Fig.1.16.

Longitudinal stress-strain behavior at two specific points is plotted using the strain gauge indicator; cross and triangle markers represent for the point 1 and 2 respectively. At the same point, strain data is also obtained from FEA model to final damage and comparison between stress-strain behavior of experiment and finite element analysis as in Fig.

Figure 2.16: Comparison of experimental and FEA of Stress-Strain Variation for SACMA and UCSB OHC test specimens

Introduction

Problem description

Experimental Study

Carbon black GOLDEN acrylic paint (#8040-NA Series) is applied to the surface of the white-coated sample randomly using an airbrush to obtain a random pattern of spots. Based on the previous study [12], a pressure of 0.15 MPa is selected in which adequate size and density of black dots are obtained. The dimensions of the anti-compression clamp are such that it prevents flexural failure and provides only in-plane static compressive load on the specimen.

The specimen with anti-hardening device is placed between the compression plate and aligned as shown in Fig. Two white light emitting diode (LED) light sources (capacity 30 W) are provided on both sides of the camera to ensure adequate illumination of the sample surface.

Figure 3.3: Speckle pattern applied over CFRP panel with different hole configurations (a) 1H (b) 2HL (c) 2HT (d) 2HD

Progressive Damage Model

This section focuses on the development of three-dimensional finite element model of the CFRP panel with cutout(s) using ANSYS 13. The degree of freedom (dof) along y-direction is limited on the underside of the laminate. This is because as the hole spacing decreases, the ineffective region of the laminate that carries no load increases and the stress flux redistributes within this zone.

In the case of the 2HD configuration plate, as the hole spacing increases, the SCF continues to decrease due to less stress interaction between the two holes (see Figure 3.6). In this study, the optimal pinhole spacing of 2.5 D is chosen as the average of the two.

Table 3.1: Material properties of the carbon/epoxy laminate [11, 12]

Results and Discussions

The failure in these specimens occurs along web section AA' in a direction transverse to the loading axis, further most of the delamination and micro-bending of the fiber is concentrated near the hole and little damage is observed away from the hole as shown in Fig. of linear fit and regression coefficient for data set point 1 are 81.101 MPa and 0.9914 respectively as shown in Fig. The location of point 1 is far away from the hole, so we note that the slope of the stress-strain curve at location 1 is found to be in close agreement with the Youngs modulus value (yy) of the CFRP laminate as predicted as shown in Table 3.1.

In all these configurations, the location of point 1 is far away from the hole, and therefore we observe that the slope of the stress-strain curve at location 1 is in close agreement with the Young's modulus (yy) value of the CFRP laminate as expected as shown in Table 3.1. In Table 3.3, we present failure initiation load (obtained through FEA) where failure initiates (from any condition) in any element of the panel predicted by PDM.

Table 3.2: Compressive Strength of for different hole configuration Configuration Compressive Strength (kN) Failure Strain (%)

Closure

45◦ layers initiate failure, the failure modes are mainly fiber matrix shear and matrix failure. Overall, the final failure mechanism observed on the +45◦ ply (surface) from the experiment is in good agreement with the PDM prediction, as shown in the figure. The damage mechanism predicted by PDM is also in good agreement with the experimental observations there, confirming the accuracy of the developed PDM algorithm.

The model-predicted damage propagation results show that the majority of damage in open-hole composite laminates under compressive loading consists of fiber-matrix shear fracture, matrix fracture, and delamination. Eight observed failure modes are linked together, such that initiating one failure mode of damage triggers other failure modes of damage that then lead to ultimate failure.

Figure 3.19: Illustration of damage propagation predicted by the PDM with increasing load for [+45/0/-45/90] 2S laminate having 1H configuration.

Introduction

Problem Description

The 1H configuration contains an open cutout in the middle, the SSR contains a single-sided bonded repair, and the DSR contains a double-sided bonded repair. A rounded composite patch of parent CFRP with a thickness of tp = 3 mm was bonded using an adhesive material (Araldite 2011) as shown in Figure 3b.

Experimental Study

Typical geometry and dimensions of the open cut and repaired specimen are shown in Figs. A circular hole with a diameter of 5 mm is drilled in the center of the panel (see Fig. 4.1) to simulate the effect of damage removal. The dimensions of the anti-stiffening compression fastener are such that it prevents flexural failure and provides only in-plane static compressive load on the specimen.

The specimen with the anti-bend clamp is placed between the compression plate and properly aligned as shown in the figure. All specimens are loaded in compression and the test is performed in displacement control mode at a rate of 2 mm/min.

Finite Element Modeling

The zoomed view of the finite element model of an open cutout and repaired panel is shown in the figure. The degree of freedom (dof) along the x-direction is limited on the bottom side of the laminate. The degrees of freedom along the y direction of all the nodes in the top face of the specimen are coupled and u-displacement is applied at the master node, which is located in the center.

For better understanding, the schematic representation of the applied boundary condition of the FEA model is shown in the figure. Material properties of the carbon/epoxy laminate Longitudinal modulus, Exx (GPa) 84.16 Transverse modulus, Eyy =Ezz (GPa) 7.12 Shear modulus, Gxy =Gxz (GPa) 3.30.

Figure 4.2: Experimental Setup of Compression Anti-buckling fixture with Specimen loaded between Compression Platen

Progressive Damage Model

Once the failure is detected in any of the elements, the damage modeling is performed to simulate the loss in the bearing capacity of the defective element. In the third step, once the damage is detected by a failure theory, a damage modeling technique is incorporated to account for the effect of the damage on the load-bearing capacity of the laminate, and further post-damage analysis is performed. This is achieved by degrading the elastic property of the failed elements, and this method is termed Material Property Degradation Method (MPDM), which assumes that the damaged element can be replaced by an equivalent element with degraded material properties.

The proposed PDM is implemented using the ANSYS parametric macro routine as shown in the flow chart shown in the figure. Patch detachment is mainly influenced by the presence of high shear stress/deformation in the adhesive layer [19, 42].

Figure 4.3: Finite element model (a) open cutout panel, (b) repaired panel and (c) (e) Zoomed view of the finite element model around the hole

Results and discussions

In Table 4.2 we present the fault initiation load (obtained via FEA), where the fault starts (from any mode) in each of the PDM predicted elements in the panel. The final damage zone in an open cut panel predicted by PDM is found to be consistent with the experimental observations, as shown in the figure. Partial detachment of the patch occurs due to shear failures in the adhesive layer over the edge of the hole and due to the stress concentration. areas at the longitudinal overlap edge of the patch (see Figure 4.11(b)).

Once again it appears that the damage zone predicted by PDM fits well with the experimental observations, as shown in Figure 4.11(g)-(l). As the load continues to increase, damage propagates into the panel, with extensive matrix cracking.

Figure 4.5: Flowchart depicting PDM algorithm

Closure

The final failure of the panel occurs after the complete dissolution of the patch at load 50.26 kN. The ultimate strength and damage progression predicted by PDM are found to be in agreement with the experimental observations, thus confirming the accuracy of the developed PDM in conjunction with finite element method. Full-field surface deformation analysis of the composite laminates is performed using digital image correlation experiments.

The damage progression path predicted by PDM is consistent with experimental observations there confirming the accuracy of the developed PDM algorithm. Comparison of test techniques used to evaluate the unidirectional compressive strength of carbon fiber reinforced plastics.