Carbon Fiber Composite Laminates Research Articles

New metal-epoxy-matrix carbon-fibre hybrids to tackle stress concentration

The paper explores the feasibility of improving the performance of composite structures with stress concentrators by local infusion of metals. The resulting material architecture presents Multi-Matrix Continuously-Reinforced Composites (MMCRC) – with shared domains of both the thermoset polymer and metal matrices with fibre-bridged interfaces between the domains. The study explores the manufacturing routes to creating such composites and assesses the resulting performance using the open-hole tensile test. It has been demonstrated that the presence of a rigid multi-matrix patch around the open hole, exhibiting minimal thickness variation, leads to a 15.0% improvement in failure load and a 16.6% improvement in failure strain for quasi-isotropic carbon fibre composite laminates. The strain history analysis of the MMCRC samples indicates the occurrence of plastic yielding within the metal matrix and strain redistribution mechanisms, leading to load sharing over the hybrid matrix area.

Large-ratio stiffness switching via harnessing the in-plane buckling and bi-stability of high-load capacity composite laminates

Stiffness tuning is crucial in developing many advanced engineering systems such as morphing aircraft, soft robotics, etc. Achieving a significant stiffness change while maintaining a high load-bearing capacity in compression is a challenging task. This study addressed this challenge by exploiting the in-plane buckling and bistability of asymmetric carbon fiber composite laminates. Due to the inverted curvatures of these laminates at their two stable states, they showed two distinct in-plane compression behaviors—stiff response where the laminate behaved like a thin curved column (with a stiffness coefficient near 68 N mm−1), and compliant response where the laminate acted like a soft non-linear spring (with stiffness near 0.7 N mm−1). As a result, a simple snap-though between these two stable states offered a significant stiffness change, with a high/low stiffness ratio of ≈97. By using extensive finite element simulations, experimentation, and a semi-empirical curved plate buckling model, this study examined the mechanics underpinning the laminate’s 3-step buckling and the influences of asymmetric boundary contacts with friction. Additionally, three strategies to enhance the laminate’s performance—adjusting the aspect ratio, applying lateral loads, and parallel stacking were examined. It also demonstrated cellular structure concepts using these bistable laminates. Remarkably, by simply adding a small lateral in-plane force (e.g., a rubber band), one could substantially increase the stiffness ratio of the composite laminates to ≈200 and load-bearing capacity at the stiff configuration, outperforming the state-of-art variable stiffness methods.

The decomposition of guided waves overlapped in the time–frequency domain in composite materials

Compared with traditional bulk wave detection methods, guided wave detection has higher efficiency and accuracy, and therefore has a wider range of applications. However, the guided waves have two inherent properties, namely dispersion and multimodality. In propagation, the various modes of guided wave signals are usually overlapped in the time–frequency domain, which greatly limits the applicability of ultrasonic detection. In this paper, the generalized dispersive mode decomposition method is introduced to separate the modes overlapped in the time–frequency domain in the composite materials. In the method, the overlapped modes in the time–frequency domain can be separated by estimating the group delay of corresponding dispersive modes without any prior knowledge. When the estimated group delay is equal to the true group delay, the corresponding mode is completely separated from the overlapped modes. The dispersion curves of the carbon fiber composite laminate with a determined propagation direction were reconstructed by employing a combination of the generalized dispersive mode decomposition and zero-crossing techniques. And the accuracy and effectiveness were further validated through comparison with the results obtained from empirical mode decomposition and independent component analysis. The proposed method presents a useful guideline for the mode decomposition in the carbon composites and helps detect the mechanical properties or defects of the composite materials.

Mechanical analysis of a carbon fibre composite woven composite laminate for ultra-light applications in aeronautics

Carbon fiber composites have emerged as a transformative technology, offering a fascinating alternative to traditional materials like aluminum and steel. Their unique combination of high strength, stiffness, and reduced density makes them an ideal choice for lightweight structural components, an attribute that aligns with the pursuit of fuel-efficient and eco-friendly aircraft designs. With the continuous race between countries and research organizations to find new materials that satisfies the above-mentioned characteristics, this article highlights the utilization of a new Ultra-Light Carbon-based Composite (ULCC) in the aeronautical sector developed within the industrial research project TERSA (Radar technologies for autonomus flying vehicles or TEcnologie Radar per Sistemi aerei a pilotaggio remoto (SAPR) Autonomi in italian). The composite material has been developed with the aim of achieving superior performance and efficiency compared to existing products on the market. To evaluate its effectiveness, first, the mechanical properties of the ULCC have been compared to T300/Epoxy and T1000/Epoxy, two of the materials commonly used in aeronautical industry and unmanned aerial vehicle (UAV). Second, finite element models were employed to verify and analyze the dynamic properties of aeronautical structural components made of ULCC. The results indicate that the new carbon-based composite exhibits remarkable strength-to-weight ratio, enhanced durability, and offering significant advantages in terms of weight reduction and overall performance. These findings validate its potential as a viable alternative in aeronautical industry.

Enhanced damage tolerance and fracture toughness of lightweight carbon-Kevlar fiber hybrid laminate

Low damage tolerance and residual strength are the main drawbacks of carbon fiber composite laminates that limit their application in many lightweight structures. This study demonstrates the exceptional damage tolerance and high fracture toughness of carbon-Kevlar hybrid laminate, where Kevlar plies are placed between two carbon fiber face sheets. Flexural strength after damage and mode I translaminar fracture toughness of carbon and Kevlar and the hybrid laminates were evaluated using three-point bending and single-edge notched bending tests, respectively. The damage mechanisms in the three configurations were investigated using micro-computed tomography and correlated with their mechanical responses. The results showed that the hybrid laminate could sustain 70% of the laminate strength after fiber damage occurs and can sustain the same strength for large strains, unlike carbon and Kevlar fiber laminates, where they both lose their mechanical integrity after fiber breakage. Moreover, this laminate showed 200% and 170% larger specific absorbed energy than carbon and Kevlar laminates, respectively. The improvement can be justified by the propagation of fiber breakage at three different positions in the Kevlar core and the delamination at the carbon-Kevlar interface that allowed larger energy dissipation during fracture. Additionally, it showed 21% and 42.7% larger absolute and specific fracture toughness, respectively, than the carbon fiber laminate.

Research on low‐velocity impact resistance of carbon fiber composite laminates

AbstractThis research conducts a comprehensive experimental and numerical analysis of carbon fiber composite laminates under low‐velocity impact conditions. Split Hopkinson Pressure Bar (SHPB) was employed to conduct impact testing on laminates with two layup sequences ( and ) under impacts ranging from 2 to 6 J. For assessing damage in laminates, a comprehensive progressive damage model and a simplified model were developed. The primary distinction between these models lies in the fact that the former incorporates the strengthening effect of compression on the delamination behavior, while the latter does not. Both models produce a good agreement with experimental results in terms of force–time, force–displacement curves, and absorbed energies. However, notable differences were noted in the predicted internal damage; the simplified model was underpredicted in terms of damage size and distribution compared to the comprehensive progressive damage model.Highlights Impact tests were conducted on and composites. In order to simulate damage caused by impacts in laminates, an energy‐based damage model was created, and cohesive elements were utilized to model interlaminar delamination. The strengthening effect of compression was explored. Distinct damage criteria were established for two cases. Damage detection of composites by x‐rays.

Low‐velocity impact response and post‐impact assessment of self‐healing composites with different stacking configurations of core–shell nanofibers

AbstractThis work investigates the low‐velocity impact (LVI) and self‐healing behavior of carbon fiber composite laminates interlaced with electrospun polyacrylonitrile core–shell nanofibers (PAN@CFRP). The main goal is to investigate the influence of different stacking configurations of core–shell nanofibers on the impact resistance and self‐healing properties of composite laminates. To achieve the purpose, three different stacking configurations were designed and tested alongside virgin specimens. The damage mechanism and self‐healing of laminates following LVI were determined using ultrasonic C‐scan and cross‐sectional microscopy examinations. At an impact energy of 14.1 J, the flexural strength of the post‐healing PAN‐3@CFRP specimen reached 95.0% of the initial value, whereas at an impact energy of 21.8 J, severe impact damage was observed in the laminates with different configurations. Among the laminates, the PAN‐1@CFRP specimen had the highest flexural strength after healing, which reached 80.1% of its pre‐impact strength. As observed, the impact resistance and self‐healing capacity of composites could be significantly enhanced by optimizing the distribution configuration of core–shell nanofibers for various impact energies.Highlights The introduction of electrospun core–shell nanofibers has not significantly compromised the mechanical properties of the composites, but rather enhanced their toughness. The addition of core–shell nanofibers significantly enhances the impact resistance of the composites and imparts excellent self‐healing properties to the material. The flexural strength of impact‐damaged specimens with different stacking configurations can be restored up to 95% of the undamaged specimen after repair treatment. Significantly improved impact resistance and self‐healing ability of laminates by optimizing the distribution configuration of core–shell nanofibers between composite layers.

Optimization of a structural Acoustic Black Hole embedded in a stringer-stiffened composite laminate panel

Many modern airframes are made of composite laminate materials due to their low mass but strong loadbearing capability. Carbon fiber composite laminates allow for a strong and lightweight aircraft, but create unhealthy levels of disruptive and uncomfortable noise within the fuselage. This work looks at the possibility of a structural Acoustic Black Hole (ABH) as a passive vibration damping solution for a common stringer-stiffened airframe panel. A 1-D symmetrically damped ABH was integrated vertically into the cross section of a stringer stiffener of a composite laminate plate. A multi-objective evolutionary algorithm was employed to search for the optimal ABH dimensions in terms of the tradeoff between weight, axial linear buckling load, and integrated vibration response. Computational results predicted the ability for damped ABH-stiffeners of a certain dimension to have less vibrational response than others with comparable linear density and axial strength. Mass normalized, non-tapered and ABH-tapered preliminary panels were manufactured and tested. These results showed a capability of ABH stiffeners to decrease vibrational response in the attached skin, more prominently above their cut-on frequency. Stiffened panels were manufactured to optimized specifications and their experimental results validate the presence of an optimal ABH dimension to achieve passive vibration damping in a system with mass and structural constraints.

A bistable pure carbon fiber composite symmetric laminate and its preparation, modeling, and experimental verification

A novel bistable symmetric laminate of pure carbon fiber composite is presented in this article. First, an analytical model was established based on the results of finite element analysis (FEA) using the segmented continuous displacement function. Then, the bistable pure carbon fiber composite symmetric laminate was prepared in the experiment. Both the analytical and FEA results are in considerable agreement with the experimental results. Finally, the effects of various laminate lengths, widths, ratios, thicknesses and materials were investigated. The results showed that the bistable pure carbon fiber composite symmetric laminate has larger deformation and lower snap-through load.

Study on Mechanical Properties of Carbon Nanotube Reinforced Composites.

In this study, carbon fiber composite laminates were modified by carbon nanotube films. In-plane and out-of-plane compression tests were carried out in a wide strain rate range (10-3-103/s). Results display that the out-of-plane compressive properties are improved by CNT interlaminar toughening because CNT can hinder the propagation of interlayer cracks; however, the dynamic in-plane compression performance is decreased due to the lack of resin in CNT film that leads to delamination inside of CNT film in advance. To optimize the material preparation process, two methods were used to prepare the mode I fracture test: (a) curing the prepreg by autoclave process; and (b) curing of resin preform by vacuum resin-transmitted molding (VARTM). Results showed that CNT prolonged the crack propagation path and improved the interlaminar fracture properties when the preform was infiltrated with resin and cured by VARTM. In addition, it was found that the interlaminar thickness was almost linear with the number of CNT layers.

High-velocity impact damage and compression after impact behavior of carbon fiber composite laminates: Experimental study

In this paper, the impact resistance of carbon fiber/resin composite laminates with a thickness of 2.5 mm at different impact velocities, angles and positions was investigated experimentally. In the process of high-velocity impact, fiber shear failure and tension failure are the main failure modes. When the impact velocity is 300 m/s, the fiber tension failure and matrix compression failure under oblique impact are more than that under normal impact, and the energy absorption of laminates is also lager than that under normal impact. The ballistic limits, energy absorption and delamination area of the center and edge impact of laminate are similar. In addition, the CAI tests are conducted to study the influence of different positions on the residual compressive strength of laminates after high-velocity impact, and the fracture loads and failure modes of two impact positions are quite different. The failure modes of the center impact laminate are main buckling failure, and the failure modes of the edge impact laminate are compression failure near the bullet hole side of the fixture and buckling failure near the center of the laminate. The research results can provide a reference for the residual strength of laminates after impact at different locations in practical situations.

Fatigue strength of self-piercing riveted joints of carbon fiber reinforced composite and Al5052 alloy plates

Self-piercing riveting (SPR) is an alternative to the conventional spot-welding method when joining sheets of dissimilar materials. In this study, the static and fatigue strengths of the SPR joints of carbon fiber composite laminate and an Al5052 aluminum alloy plate under various loading conditions were evaluated. Fatigue test results show that the load amplitudes, corresponding to the fatigue endurance limit, based on a lifetime of one million cycles, were 680 N, 323 N, and 305 N for loading angles of 0°, 45° and 90°, respectively, with corresponding fatigue ratios of the joints of 33%, 23% and 24%. The fatigue strengths at the loading angles of 45°, and 90° in terms of the equivalent stress intensity factor were found to be much higher than those at a loading angle of 0°. This is partially due to the difference in the fatigue failure modes. Regarding the fatigue strength designs of SPR joints consisting of composite laminate and aluminum alloy plates under various loads, it can be suggested that the tensile-shear strengths dominate the SPR fatigue strengths of the joints because they are highly vulnerable to tensile-shear loading.

Scalable High Tensile Modulus Composite Laminates Using Continuous Carbon Nanotube Yarns for Aerospace Applications.

An approach is established for fabricating high-strength and high-stiffness composite laminates with continuous carbon nanotube (CNT) yarns for scaled-up mechanical tests and potential aerospace structure applications. Continuous CNT yarns with up to 80% degree of nanotube alignment and a unique self-assembled graphitic CNT packing result in their specific tensile strengths of 1.77 ± 0.07 N/tex and an apparent specific modulus of 92.6 ± 3.2 N/tex. Unidirectional CNT yarn reinforced composite laminates with a CNT concentration of greater than 80 wt % and minimal microscale voids are fabricated using filament winding and aerospace-grade resin matrices. A specific tensile strength of up to 1.71 GPa/(g cm-3) and specific modulus of 256 GPa/(g cm-3) are realized; the specific modulus exceeds current state-of-the-art unidirectional carbon fiber composite laminates. The specific modulus of the laminates is 2.76 times greater than the specific modulus of the constituent CNT yarns, a phenomenon not observed in carbon fiber reinforced composites. The results demonstrate an effective approach for fabricating high-strength CNT yarns into composites for applications that require specific tensile modulus properties that are significantly beyond state-of-the-art carbon fiber composites and potentially open an unexplored performance region in the Ashby chart for composite material applications.

Low-velocity impact and compression-after-impact behaviors of twill woven carbon fiber/glass fiber hybrid composite laminates with flame retardant epoxy resin

Composite structures suffer from significant residual strength reduction as a result of invisible damage caused by impacts, which can cause severe harm without any prior warning. Fibers play a crucial role in improving the impact resistance of composite materials by providing strength, toughness, and energy absorption. This work explores the dynamic response and residual compressive strength of composite laminates reinforced with twill woven carbon fiber, glass fiber, and carbon fiber /glass fiber hybrid, subjected to low-velocity impact (LVI) and compression-after-impact (CAI) testing. First, the performance of woven carbon fiber composite laminates infused with flame retardant epoxy resin and general epoxy resin was studied. The results indicated a lower load-bearing capacity and elongation of the former. To enhance the impact resistance of the laminates, we explored the effects of incorporating glass fibers into the woven carbon fiber composite laminates. Our findings demonstrated that the introduction of glass fibers alteres the impact damage mode of the laminates and effectively improved their impact resistance and CAI strength. Further, several techniques were adopted to characterize the corresponding damage morphologies and failure mechanisms of the woven carbon fiber composite laminates, which provides valuable insights for their structural design and performance improvement.

A scientific view on composite inlays in running shoes: The influence of epoxy crosslink density on bending and fatigue properties

This study highlights the need for a deeper understanding of the structure-property relationships of carbon fiber reinforced thermosets applications, whenever considered as a functional part for running footwear. This is done, investigating the static and dynamic flexural properties of three different epoxy resins in prepreg-based carbon fiber composite laminates, varying the glass transition temperature (Tg). In this way, the inherent morphology of the resin was correlated with the fatigue properties of the laminate, which is crucial for the design of composite stiffening inlays for running shoes. First, it was shown that the crosslink density and aromatic content of the resin, which correlates with Tg, drastically affect the compressive strength of the matrix and are therefore crucial for the static and dynamic flexural properties of CFRP. Increasing the Tg from 147 °C to 269 °C leads to stiffening of the resin and thus increased resistance to fiber buckling, resulting in improved bending stiffness, flexural strength, bending angle, overall fatigue strength and thus durability of CFRP plates for running shoes. Moreover, the percentage of fibers in the loading direction can be considered as the main factor affecting the flexural properties of the laminate. Here, a quasi-isotropic stacking order leads to an increased bending angle but to a decrease in bending stiffness, flexural strength and fatigue strength compared to a unidirectionally reinforced composite due to a high load bearing capacity of the 0° fibers. This knowledge creates a basis to bridging the gap between composite material and biomechanical sports science for future work.

Study on the Stability of a Repair Process for Carbon Fiber Laminates

Carbon fiber-reinforced resin matrix composites are often used in the surface structure of helicopters. In the process of flight, the structure often suffers from different degrees of damage due to the influence of the harsh environment. To explore the superiority of the repair method and obtain a good repair process, the intact and damaged carbon fiber composite laminates were prepared in this paper. The damaged specimens were repaired by single-sided double-layer patch bonding. The allowable value of tensile strength of carbon fiber laminates was tested, and the mechanical properties of carbon fiber composite laminates after repair were studied. Combined with the failure mode of the specimen after failure, the results show that the mechanical properties of the repaired specimen can be restored to about 81.9 % of the non-destructive specimen. The single-sided double-layer patch can effectively improve the mechanical properties of the middle loading area. The repair process has high reliability and good stability, which provides a new method for the repair of the composite laminate structure.

Compression-After-Impact analysis of carbon/epoxy and glass/epoxy hybrid composite laminate with different ply orientation sequences

This article is a continuation study of the LVI and CAI response of the plain carbon fiber composite laminates and explores the experimental & numerical study of the post-impact damage propagation in carbon/epoxy and glass/epoxy hybrid laminate under compression. LVI (Low-Velocity Impact) tests are first carried out at three different energy levels, and the same specimens are put under a displacement-controlled compression test. The damage initiation and propagation in both impact tests and compression tests are studied using X-ray computed tomography. The results matched well with the results obtained from the finite element model, simulated using ABAQUS/explicit. The effect of the changing impact energy and the influence of the stacking orientation on post-impact damage propagation is studied. The strain maps during the compression tests are studied through Digital Image Correlation (DIC) technique based on which the stress–strain histories are plotted by using well-calibrated strain values. Upon completing the study, it is observed that the CAI (Compression-After-Impact) strength depends on the factors like properties of the constituent material, ply orientations and stacking sequence of the laminate. It is observed that even when the damage is barely visible on the surface of the specimen, compressive strength has been reduced drastically. The observed mode of damage propagation in CAI testing is buckling of the sub-laminates in the vicinity of the damage zone formed by the impact event. The size of the impact damage zone depends on the impact energy, which actually governs the CAI strength.

Optimum Design of Ply-up of T800 Carbon Fiber Composite Laminates with Holes

Based on the HyperWorks software, this paper takes the layup sequence of carbon fiber composite laminates as the design variables, the main strain of the laminates and the continuous angle layup not exceeding four layers as constraints, and the maximum bearing capacity as the design goal. The optimized design of the laminate and the corresponding experimental verification were carried out. The experimental results are in good agreement with the simulation analysis results, indicating that the optimization design method proposed in this paper is reasonable.

Impactor Diameter and Ply Number Effects on the Impact Behavior of Carbon Fiber Composite Laminates

As it is known, impact damage is a major mechanical phenomena for composite materials especially used in the aerospace structures. The factors affecting the impact behaviour of the composites depend on the impactor systems as well as the target material. In this study ply number and impactor geometry effects of carbon fiber reinforced epoxy composites were investigated by impact tests. In this context, drop weight impact tests were carried out at 6J, 12J and 24J energy levels by using hemispherical impactors with 10 mm and 20 mm diameters. Laminated composites were manufactured in 6, 10 and 14 plies with vacuum infusion method. The effects of laminate thickness, impactor diameter and impact energy effects on the force, velocity, absorbed energy and damage surfaces were investigated. It is observed that impactor geometries and velocities caused the different damage mechanisms in composites and impactors played an important role in determining the penetration and perforation behaviours of composites.

An experimental and numerical study of industrially representative wrinkles in carbon fibre composite laminates

Wrinkles (out-of-plane waviness) are a common carbon fibre composite manufacturing defect that causes significant reductions in the components’ mechanical properties. The substantial cost implications of empirical testing has led to industrial demands for reliable and representative wrinkle models. This paper presents a novel modelling method for laminates with wrinkle geometries measured from non-destructive testing and/or sample cross-sections. The resulting finite element modelling pre-processor offers a self-contained system of mesh generation, local material orientation calculation, material assignment, boundary condition selection and solver specification. The approach was validated by comparison with pre-existing analytical models and experimental data. Through the resulting rapid model generation, the influence of various physical wrinkle parameters were assessed at an unprecedented rate and sensitivity. For example, this highlighted the dominant dependence of wrinkle angle on the laminate’s compressive strength. Additionally, increases in wrinkle extent and reductions in amplitude decay rate were shown to be detrimental to mechanical properties.The model was then validated against an experimental study on industrially relevant carbon-epoxy laminate samples to understand the effect out-of-plane waviness on compressive properties. The model results showed good agreement to the experimental values (<15% error) and at least a 32% reduction in compressive strength.