A finite element methodology for analysing degradation and collapse in postbuckling composite aerospace structures

A methodology for analysing the degradation and collapse in postbuckling composite structures is proposed. One aspect of the methodology predicts the initiation of interlaminar damage using a strength criterion applied with a global-local analysis technique. A separate approach represents the growth of a pre-existing interlaminar damage region with user-defined multi-point constraints that are controlled based on the Virtual Crack Closure Technique. Another aspect of the approach is a degradation model for in-plane ply damage mechanisms of fiber fracture, matrix cracking, and fiber-matrix shear. The complete analysis methodology was compared to experimental results for two fuselage-representative composite panels tested to collapse. For both panels, the behavior and structural collapse were accurately captured, and the analysis methodology provided detailed information on the development and interaction of the various damage mechanisms

Control of composite material crack branching for improved fracture toughness

An experimental investigation was conducted to develop techniques for controlling crack branching in composite laminates. Double cantilever beam specimens were tested to study mode I dominated crack growth. Embedded flaws were generated using ply gaps and strips of non-stick film, both individually and in combination as a "branch flaw". Crack branching in 0° plies was generated in an inconsistent manner using ply gaps, but in a consistent manner using branch flaws. Branching through 90° plies occurred automatically due to their orientation, and could be further controlled using embedded delamination flaws. Crack branching in 45° plies was more complex, but could be controlled using ply gaps as well as branch flaws. These discoveries were combined to demonstrate crack branch control through a quasi-isotropic laminate. The results have application to design of future high toughness and damage tolerant aerospace composites

Perturbation-based imperfection analysis for composite cylindrical shells buckling in compression

A numerical investigation was conducted into a perturbation-based analysis approach for assessing the imperfection sensitivity of composite cylindrical shells buckling under compression loading. The Single Perturbation Load Analysis (SPLA) approach was applied, which uses a single lateral load to introduce a realistic, worst-case and stimulating imperfection pattern. Finite element analysis was conducted for cylinders of both monolithic composite laminate and sandwich construction, with and without small and large cutouts. It was found that using a perturbation displacement equal to the shell thickness provides a suitable technique for estimating the reduction in buckling load caused by imperfections. Predictions of buckling knockdown factors using the SPLA approach showed advantages over the use of eigenmodes as the SPLA approach provides a clear design point and does not require experimental data for calibration. The effect of small and large cutouts was analogous to the effect of small and large perturbation loads. The location of the perturbation load influenced the buckling knockdown factors for both small and large cutouts, and worst-case locations were identified for both configurations

An experimental investigation into damage modes and scale effects in CFRP open hole tension coupons

An experimental investigation was conducted to study the effect of notch size and length scale on the damage of carbon fibre-reinforced composite specimens. Open hole tension specimens in a range of configurations were tested quasi-statically to ultimate failure. The load response, damage modes and strain field development were experimentally recorded. The results demonstrated that changing the ply thickness and specimen dimensions markedly affected the damage modes and specimen behaviour. This output provides key insights into the nature of composite behaviour, and is also critical for the development and validation of analysis methodologies capturing damage initiation and progression

Development of design guidelines for postbuckling composite aerospace structures

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Structural assessment of microvascular self-healing laminates using progressive damage finite element analysis

This paper presents a progressive damage analysis methodology to numerically analyse the effect of microvascular open channels on the structural properties of self-healing fibre-polymer laminates. The tensile and compression properties of self-healing carbon-epoxy laminates containing microvascular systems are analysed using finite element models which consider progressive in-plane ply damage and intra-ply damage (matrix and delamination cracking). The models predict with good accuracy (often within 5%) the stiffness and strength of laminates containing circular or elliptical microvascular channels of different sizes and orientations. The model calculates a progressive reduction in structural properties with increasing size of microvascular channels due to increased ply waviness, which was confirmed using experimental property data. The model also predicts the location and progression of damage under increasing tensile or compression loading to final failure. The model has application as a tool for the design of microvascular systems in self-healing composites used for structural applications

Benchmark assessment of automated delamination propagation capabilities in finite element codes for static loading

With the increasing implementation into commercial finite element (FE) codes of capabilities for simulating delamination propagation in composite materials, the need for benchmarking and assessing these capabilities is critical. In this study, the capabilities of the commercial FE code Marc (TM) 2008r1 with implementation of the Virtual Crack Closure Technique (VCCT) was assessed. Benchmark delamination propagation results for several specimen configurations were generated using a numerical approach. Specimens were analyzed with three-dimensional and two-dimensional models, and compared with previous analyses using Abaqus (R) with the VCCT implemented. The results demonstrated that the VCCT implementation in Marc (TM) was capable of accurately replicating the benchmark delamination growth results. The analyses in Marc (TM) were significantly more computationally efficient than previous analyses in Abaqus (R). This was due to a lack of convergence issues, and a solution process that maintained the use of large time increments. The results demonstrated the advantages of numerical over experimental and analytical benchmarks, particularly with regards to comparison of capabilities a cross codes. More broadly, the results illustrated key similarities and differences between two commercial FE codes implementing the same analysis technique, which reinforces the need for rigorous benchmarking and assessment

Experimental and numerical investigation of low velocity impacts of fibre-metal laminates

In this work, an experimental and numerical investigation into the impact behaviour of fibremetal laminates (FMLs) subject to low velocity impact is conducted. The study investigates the damage initiation and progression of FMLs in order to characterise and represent the complex damage responses and deformation that lead to strength and stiffness loss. Low velocity impact tests were experimentally conducted using a drop-test impact rig. Numerical analysis was conducted using Abaqus/Explicit, with damage models incorporated to represent in-plane composite damage, delamination, elastic-plastic behaviour of the metal layers at high strain rates, and failure in the adhesive layers. In addition to plastic deformation of the aluminium layers, matrix cracking, fibre failure and delamination of the composite glassepoxy prepreg were observed as critical damage modes. The numerical simulations exhibited good correlation with experimental impact tests in terms of kinematic response and damage development

Developments in structural proof testing for aircraft composite components

This paper evaluates structural proof testing methods for the detection of manufacturing defects in aircraft composite components. A static proof test method based on compliance and surface strain mapping and a dynamic (vibration) method based on mode shape curvature (MSC) analysis are evaluated for the detection of manufacturing defects. The evaluation of these structural proof methods was performed using finite element analysis to identify the test conditions best suited for damage/defect detection. The finite element modelling was validated by structural proof tests performed on T-joint composite specimens containing the manufacturing defects of large voids, porosity and delaminations. The coupled compliance and surface strain mapping technique was able to detect a delamination crack along the stiffener-skin bond-line or a void within the fillet region when the joint was elastically loaded, but failed to detect porosity at the concentrations which typically occur in joints with defects. The MSC technique successfully detected voids, porosity and delaminations in the T-joint excited by elastic stress waves induced by random frequency vibrations

Analysis of damage development in CFRP notched coupons with energy-based multi-axial failure data

A methodology that uses multi-axial testing and dissipated energy (DE) to characterise the mechanical behaviour of laminated composite materials has been implemented into an analysis approach. The goal of this approach is to quantify the damage development on a continuum basis and eventually relate this damage to the constitutive behaviour of the material. The approach was implemented into a commercial finite element package with a custom user subroutine. Doublenotch characterisation specimens and open hole tension coupons were assessed. DE was predicted well in tension and in-plane rotation cases apart from a pure shear case where DE was excessive. DE magnitude was satisfactory in the open hole case and damage propagation well represented. Future work will involve advancing the technique to include DEdependent constitutive modelling