Failure mechanisms on composite specimens subjected to compression after impact

Composite panels are widely used in aeronautic and aerospace structures due to their high strength/weight ratio. The stiffness and the strength in the thickness direction of laminated composite panels is poor since no fibres are present in that direction and out-of-plane impact loading is considered potentially dangerous, mainly because the damage may be left undetected. Impact loading in composite panels leads to damage with matrix cracking, inter-laminar failure and eventually fibre breakage for higher impact energies. Even when no visible impact damage is observed at the surface on the point of impact, matrix cracking and inter-laminar failure can occur, and the carrying load of the composite laminates is considerably reduced. The greatest reduction in loading is observed in compression due to laminae buckling in the delaminated areas. The objective of this study is to determine the mechanisms of the damage growth of impacted composite laminates when subjected to compression after impact loading. For this purpose a series of impact and compression after impact tests were carried out on composite laminates made of carbon fibre reinforced epoxy resin matrix. An instrumented drop-weight-testing machine and modified compression after impact testing equipment were used together with a C-scan ultrasonic device for the damage identification. Four stacking sequences of two different epoxy resins in carbon fibres representative of four different elastic behaviours and with a different number of interfaces were used. Results showed that the delaminated area due to impact loading depends on the number of interfaces between plies. Two buckling failure mechanisms were identified during compression after impact, which are influenced more by the delamination area than by the stacking sequence.     

Impact damage characterisation of carbon fibre/epoxy composites with multi-layer reinforcement

Low-velocity impact tests were performed to investigate the impact behaviour of carbon fibre/epoxy composite laminates reinforced by short fibres and other interleaving materials. Characterisation techniques, such as cross-sectional fractography and scanning acoustic microscopy, were employed quantitatively to assess the internal damage of some composite laminates at the sub-surface under impact. Scanning electron microscopy was used to observe impact fractures and damage modes at the fracture surfaces of the laminate specimens. The results show that composite laminates experience various types of fracture; delamination, intra-ply cracking, matrix cracking, fibre breakage and damage depending on the interlayer materials. The trade-off between impact resistance and residual strength is minimised for composites reinforced by Zylon fibres, while that for composites interleaved by poly(ethylene-co-acrylic acid) (PEEA) film is substantial because of deteriorating residual strength, even though the damaged area is significantly reduced. Damages produced on the front and back surfaces of impact were also observed and compared for some laminates.     

Damage growth analysis of low velocity impacted composite panels

Low velocity impact loading in aircraft composite panels is a matter of concern in modern aircraft and can be caused either by maintenance accidents with tools or by in-flight impacts with debris. The consequences of impact loading in composite panels are matrix cracking, inter laminar failure and, eventually, fiber breakage for higher impact energies. Even when no visible impact damage is observed on the surface at the point of impact, matrix cracking and inter laminar failure can occur, and the carrying load of the composite laminates is considerably reduced. The greatest reduction in loading is observed in compression due to laminae buckling in the delaminated areas.

The objective of this study is to determine the limit loading capacity and the damage growth mechanisms of impacted composite laminates when subjected to compression after impact loading. For this purpose a series of impact and compression after impact tests were carried out on composite laminates made of carbon fiber reinforced epoxy resin matrix. Four stacking sequences representative of four different elastic behaviours were used. Results show that the compressive, after impact, failure stress is influenced by the stacking sequence but a relatively independent strain to failure is observed.     

The effect of stacking sequence on impact damage in a carbon fibre/epoxy composite

The effect of stacking sequence on impact damage in a carbon fibre/toughened epoxy composite was studied. The major form of damage was delamination, which initiated at almost every interface through the panel. During the impact event the force-time response was monitored and the energy absorbed analysed in terms of an initiation and a propagation energy. The energy absorbed in delamination initiation was influenced by the stacking sequence, being increased by placing 45° fibres in the surface plies and by increasing the number of dissimilar interfaces. The residual energy absorbed in delamination propagation was found to increase linearly with increasing total delamination area. The compression-after-impact strength was related to the maximum delamination area.     

On the impact performance of carbon fibre laminates with epoxy and PEEK matrices

A study was made of the mechanical properties and impact performance of carbon fibre/PEEK (0,90), ($\text{±45}$) and ($\text{±45,0}$) laminates and comparisons were made with similar carbon fibre/epoxy laminatesFibre dominated properties such as plain tensile strength were similar to those of epoxy laminates with similar fibres and volume fractions. Because of the increased toughness of PEEK there was less extensive matrix cracking, even though there was fibre debonding, and this gave increased transverse and shear cracking strains, increased shear strengths but decreased notched tensile strengths. The lower modulus and yield stress of PEEK caused lower compressive strengths, but PEEK absorbed little moisture and at 120°C moisture had little effect on mechanical properties.Dropweight impact produced less extensive damage in carbon fibre/PEEK laminates. Residual tensile strengths were similar but, because of the less extensive damage and greater delamination fracture energy, the residual compressive strengths were significantly greater with a PEEK matrix.Microscopic examination showed less matrix cracking and more fibre buckling in the carbon fibre/PEEK and this is discussed in relation to mechanical properties.     

Effects of fibre/matrix adhesion on carbon-fibre-reinforced metal laminates—II. impact behaviour

Quasi-static, low-hanging and high-velocity impact tests have been conducted in order to study the effect of fibre/matrix adhesion on the impact properties of fibre-reinforced metal laminates. Differences in fibre/matrix adhesion were achieved by using treated or untreated carbon fibres in an epoxy resin system. Chemical removal of the aluminium layers and a sectioning technique were applied to examine and characterize the impact damage in the laminates. The results show that the laminates with the weaker fibre/matrix adhesion exhibit larger damage zones, although the back face crack length and permanent indentation after impact are smaller for a given impact energy. Residual tensile strength after impact is also higher for the untreated fibre laminates due to increased fibre/matrix splitting in the composite layer.     

碳纤维复合材料假脚工艺参数对其冲击后疲劳性能的影响

基于三维逐渐损伤理论和有限元法,对碳纤维复合材料假脚的冲击及冲击后疲劳破坏过程进行分析,研究了不同的复合材料体系、几何尺寸、纤维铺设方式等工艺参数对碳纤维假脚的冲击损伤及疲劳性能的影响规律。结果表明,在冲击载荷作用下,碳纤维复合材料假脚的损伤模式主要为基体开裂、纤维压缩和分层;复合材料体系的横向和法向拉伸强度以及剪切强度等参数越小,假脚的冲击损伤面积越大,所能承受的疲劳循环次数越低;随着后龙骨厚度的增加,基体开裂损伤面积越来越大,分层损伤面积略有减小,而纤维压缩损伤几乎没有变化。尽管随着后龙骨厚度的增加,假脚的疲劳循环次数逐渐增大,但是相对于厚度的增加量,疲劳循环次数的增加量相对较小;不同铺层参数对碳纤维复合材料假脚的冲击损伤模式几乎没有影响。适度增加0°铺层的含量,可有效提高碳纤维复合材料假脚的疲劳性能。     

Post-impact compressive behaviour of low temperature curing woven CFRP laminates

The polymeric matrix in a fibre-reinforced composite serves to bind the fibres together, transfer load to the fibres and protect them against environmental attack and damage due to handling. The matrix has a strong influence on several mechanical properties of the composite such as transverse modulus and strength, impact resistance, shear properties and properties in compression. This paper describes the results of an experimental study to determine the effect of resin (matrix) on the post-impact compressive behaviour of carbon fibre woven laminates. Three new low temperature cure (50–125°C) epoxy resins are examined: an unmodified (LTM12), a rubber-modified (LTM25) and a thermoplastic toughened epoxy resin (MT8E). Note, however, that the first two are post-cured at 190°C. Velocities and impact energies were used to simulate momenta typical of low velocity impact hazards associated with aircraft in-service. Measurements of impact damage and damage growth during compression are made using ultrasonic C-scanning and penetrant-enhanced X-ray radiography techniques. For low impact energies the superior performance of the thermoplastic toughened epoxy is confirmed. Its residual compressive strength compares favourably with that obtained for high strength carbon fibre/epoxy laminates manufactured from unidirectional sheets cured at 190°C.     

The fatigue damage mechanics of notched carbon fibre/PEEK laminates

A model is presented for the strength, post-fatigue residual strength and damage propagation in notched, cross-ply carbon fibre/polyetheretherketone (PEEK) laminates. Fracture mechanics principles are used to predict quasi-static damage growth, and the application of a Paris law permits extension to fatigue damage. Strength is predicted by applying a failure criterion based on the tensile stress distribution in the 0° plies, as modified by damage (either quasi-static or fatigue). The volume dependence of strength is included by using a simple Weibull distribution. The parameters of the model are determined from independent experiments. Good agreement with experimental results is obtained. Comparisons are made with previous results from carbon fibre/epoxy laminates. The behaviour of the carbon fibre/PEEK is similar, although the extent of delamination and matrix cracking is reduced owing to the higher inherent toughness of the matrix.     

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.     

The damage and failure of GRP laminates by underwater explosion shock loading

This paper examines the development of microstructural damage in a glass-reinforced polymer (grp) laminate subjected to explosive shock loading in water. GRP is commonly used in small naval vessels, and may be subjected to underwater explosions. In the experiments, the laminates were exposed to increasing amounts of shock loading produced by underwater explosions. The laminates were backed with either water or air to modify the amount of bending experienced under loading, with the air-backed laminates having the higher amount of bending. Examination of the grp microstructure by optical and scanning electron microscopy after shock testing failed to reveal any damage to either the polymer matrix or glass fibres when the laminate was backed with water. In contrast, when the laminate was backed with air, small cracks were produced in the polymer matrix at low shock pressures. Raising the shock pressure above a threshold limit caused complete failure of the laminate by cracking in the polymer matrix, cracking of the glass fibres, and delamination of the glass fibres from the polymer. The differences in the shock resistance of the water- and air-backed grp are discussed. Measurements of the residual tensile fracture strength of the laminates after shock loading are also presented. The fracture strength of the water-backed laminate was not affected by shock, but the fracture strength of the air-backed laminate deteriorated with the onset of glass fibre breakage and delamination in the grp microstructure.     

Damage Tolerance of Pre-Stressed Composite Panels Under Impact Loads

An experimental test campaign studied the structural integrity of carbon fibre/epoxy panels preloaded in tension or compression then subjected to gas gun impact tests causing significant damage. The test programme used representative composite aircraft fuselage panels composed of aerospace carbon fibre toughened epoxy prepreg laminates. Preload levels in tension were representative of design limit loads for fuselage panels of this size, and maximum compression preloads were in the post-buckle region. Two main impact scenarios were considered: notch damage from a 12 mm steel cube projectile, at velocities in the range 93–136 m/s; blunt impact damage from 25 mm diameter glass balls, at velocities 64–86 m/s. The combined influence of preload and impact damage on panel residual strengths was measured and results analysed in the context of damage tolerance requirements for composite aircraft panels. The tests showed structural integrity well above design limit loads for composite panels preloaded in tension and compression with visible notch impact damage from hard body impact tests. However, blunt impact tests on buckled compression loaded panels caused large delamination damage regions which lowered plate bending stiffness and reduced significantly compression strengths in buckling.     

Effect of micro-randomness on macroscopic properties and fracture of laminates

Composite materials demonstrate a considerable extent of heterogeneity. A non-uniform spatial distribution of reinforcement results in variations of local properties of fibrous laminates. This non-uniformity not only affects effective properties of composite materials but is also a crucial factor in initiation and development of damage and fracture processes that are also spatially non-uniform. Such randomness in microstructure and in failure evolution is responsible for non-uniform distributions of stresses in composite specimens even under externally uniform loading, resulting, for instance, in a random distribution of matrix cracks in cross-ply laminates. The paper deals with statistical features of a distribution of carbon fibres in a transversal cross-sectional area in a unidirectional composite with epoxy matrix, based on various approaches used to quantify its microscopic randomness. A random character of the fibres’ distribution results in fluctuations of local elastic moduli in composites, the bounds of which depend on the characteristic length scale. A lattice model to study damage and fracture evolution in laminates, linking randomness of microstructure with macroscopic properties, is discussed. An example of simulations of matrix cracking in a carbon fibre/epoxy cross-ply laminate is given.     

Impact resistance and tolerance of interleaved tape laminates

This paper presents and discusses the results of low-velocity impact and compression-after-impact (CAI) tests conducted on interleaved and non-interleaved carbon/epoxy tape laminates. Olefin film interleaves provided a strong interface bond, resulting in a reduction in projected damage area. These interleaves changed the stress distribution under impact and restricted delamination formation at the ply interface. An investigation into the compression behaviour of these laminates revealed a reduction in undamaged strength using olefin interleaves. This was attributed to the lack of lateral support for fibres at the fibre/interleaf interface, allowing fibre microbuckling to occur at a low load. Low modulus copolyamide web interleaves resulted in an increase in damage area and minor changes to CAI strength. Examination of laminate cross-sections revealed that this was due to both the open structure of the interleaf and poor resin/interleaf adhesion. High shear modulus polyethylene interleaves resulted in a significant decrease in damage area at various impact energies, with CAI strength improved compared to the non-interleaved laminates.     

A single-lap shear specimen for determining the effect of strain rate on the interlaminar shear strength of carbon fibre-reinforced laminates

A new design of single-lap shear specimen for determining the effect of loading rate on the interlaminar shear strength of laminated composites is described. Finite element analyses are used to optimize the specimen geometry and minimize the variation in the shear stress and the magnitude of the normal stress along the interlaminar failure plane. Experimental results are obtained at a quasi-static and an impact rate of loading for the interlaminar shear strength parallel to the fibres in both unidirectional carbon/epoxy and unidirectional carbon/polyetheretherketone (peek) laminates and at interfaces across which the fibre orientation is 0°/90° and ±45°. Results for carbon/epoxy laminates are compared with those from an earlier investigation using a double-lap specimen geometry and show a similar small dependence on loading rate. No significant effect of loading rate was observed for the carbon/peek laminates.     

Modelling off-axis ply matrix cracking in continuous fibre-reinforced polymer matrix composite laminates

The fracture process of composite laminates subjected to static or fatigue tensile loading involves sequential accumulation of intra- and interlaminar damage, in the form of transverse cracking, splitting and delamination, prior to catastrophic failure. Matrix cracking parallel to the fibres in the off-axis plies is the first damage mode observed. Since a damaged lamina within the laminate retains certain amount of its load-carrying capacity, it is important to predict accurately the stiffness properties of the laminate as a function of damage as well as progression of damage with the strain state. In this paper, theoretical modelling of matrix cracking in the off-axis plies of unbalanced symmetric composite laminates subjected to in-plane tensile loading is presented and discussed. A 2-D shear-lag analysis is used to determine ply stresses in a representative segment and the equivalent laminate concept is applied to derive expressions for Mode I, Mode II and the total strain energy release rate associated with off-axis ply cracking. Dependence of the degraded stiffness properties and strain energy release rates on the crack density and ply orientation angle is examined for glass/epoxy laminates. Suitability of a mixed mode fracture criterion to predict the cracking onset strain is also discussed.     

Post-impact compressive strength of curved GFRP laminates

Low velocity impacts to fibre reinforced plastic composites cause a pattern of damage consisting in general of delamination, fibre breakage and matrix cracking. Such damage is accidental and may go unnoticed; therefore composite structures must be designed assuming impact damage exists. Previous work on flat composite laminates has resulted in a reasonable understanding of the mechanisms of compressive strength reduction. There are, however, many instances where curved laminates are used in structures where impact is likely. Furthermore, due to the mechanisms of strength reduction, it may be expected that curvature would have a significant effect on the behaviour of the laminates.The work described here consists of experimental measurement of the post-impact compressive strength of curved GFRP laminates. The laminates were of 8 plies of 0.3 mm thick pre-impregnated glass fibre/epoxy tape in a (0, ±45, 0°)_s lay-up. Each laminate was 200 mm in length by 50 mm wide with the plane of curvature normal to the length. Laminates were impacted on the convex surface of the laminate by dropping a steel mass from 1 m vertically above it.Impacted laminates were loaded in compression and the out-of-plane displacements of the top and bottom surfaces were recorded. Final failure was typically due to fibre breakage occurring through the centre of the impacted area of the laminate. Possible differences in the impact response, and measurable differences in the sizes of the impact damage area, were found to arise from these curvatures, and differences were observed in their post-impact buckling behaviour. However, perhaps unexpectedly, the post-impact compressive strength for a curved laminate was found to be similar to that for a flat laminate. The failure loads for the impact damage laminates are shown to be comparable with those for laminates containing artificial delaminations.     

An experimental study of low velocity impact response in 2/2 twill weave composite laminates manufactured by a novel fabrication process

A novel fabrication process for advanced composite components—the Quicktep^TM process was described. 2/2 twill weave MTM56/CF0300 carbon epoxy composite laminates were manufactured by the Quickstep and the autoclave processes. The response of these laminates to drop-weight low velocity impact at energy levels ranging from 5 to 30 J was investigated. It was found that the laminates fabricated by the Quickstep had better impact damage tolerance than those fabricated by the autoclave. Optical microscopy revealed extensive matrix fracture in the center of the backside of the autoclave laminates indicating the more brittle property of the epoxy matrix cured by the autoclave process. Interfacial shear strength (IFSS) for two composite systems were measured by micro–debond experiments. The MTM56/CF0300 material cured by the Quickstep showed stronger fibre matrix adhesion. Since the thickness and density of the impact targets produced by two processes were different, finite element analysis (FEA) was performed to study the effect of these factors on the impact response. The simulation results showed that the difference in thickness and density affects the stress distribution under impact loading. Higher thickness and lower density caused by processing lead to less endurance to drop weight impact loading. Therefore the better performance of Quickstep laminates under impact loading was not due to the thickness and density change, but resulted from stronger mechanical properties.     

Stiffness and fracture analysis of laminated composites with off-axis ply matrix cracking

Matrix cracks parallel to the fibres in the off-axis plies is the first intralaminar damage mode observed in laminated composites subjected to static or fatigue in-plane tensile loading. They reduce laminate stiffness and strength and trigger development of other damage modes, such as delaminations. This paper is concerned with theoretical modelling of unbalanced symmetric laminates with off-axis ply cracks. Closed-form analytical expressions are derived for Mode I, Mode II and the total strain energy release rates associated with off-axis ply cracking in [0/θ]_s laminates. Stiffness reduction due to matrix cracking is also predicted analytically using the Equivalent Constraint Model (ECM) of the damaged laminate. Dependence of the degraded stiffness properties and strain energy release rates on the crack density and ply orientation angle is examined for glass/epoxy and carbon/epoxy laminates. Suitability of a mixed mode fracture criterion to predict the cracking onset strain is also discussed.     

Development of novel cost effective hybrid ply carbon-fibre composites

A carbon-fibre/epoxy hybrid material has been developed which comprises ultra-high-performance fibres in the principal load-bearing direction and standard carbon fibres in the secondary orientations. The effect of stacking sequence on mechanical properties, including impact damage tolerance, has been studied and the results compared with non-hybrid data. By optimising the stacking sequence the hybrid material exhibited a mechanical performance similar to that of the ultra-high-performance material and a performance superior to the standard material. The energy absorbed during low velocity impact was analysed in terms of an initiation and propagation energy. The energy absorbed through delamination initiation was increased by placing 45 ° fibres in the surface plies and by placing ultra-high-performance fibres in the 0 ° plies. The energy absorbed during delamination growth was independent of the fibre type and determined solely by the matrix material. On the basis of current pre-preg prices the hybrid material corresponds to a significant cost saving of 12% through the use of lower cost standard fibres in the secondary stressed layers.