Optimisation of fibre steering in composite laminates using a genetic algorithm

An optimization procedure using a genetic algorithm has been applied to define the optimum orientation of fibres in a uni-directional laminate in which the fibres were allowed to vary continuously across the domain. The domain was divided into two-dimensional finite elements and anisotropic properties corresponding to a carbon fibre laminate with all layers aligned in the zero element axis direction were applied to the laminate. The orientation of the material axis on each element was then prescribed as an independent variable for the genetic algorithm.     

Configuration of a genetic algorithm used to optimise fibre steering in composite laminates

In a previous study, a genetic algorithm (GA) was applied to define the optimum orientation of fibres in a unidirectional laminate in which the fibre orientations were allowed to vary continuously across the domain. The results were positive, but computation time was excessively long. The present study was undertaken to address this issue. The elements used by the GA (encoded representation, fitness criterion, operators affecting the population) were examined and optimised to reduce computation time. It was found that the use of a strain-based fitness criterion was better than use of a stress-based criterion, regardless of whether the load was above or below the elastic limit. A good balance of elitism and mixing was necessary in the GA parameters to obtain the fastest convergence. Automatic imaging of the GA output was achieved using the SAMCEF software package.     

Design of a cruciform specimen for biaxial testing of fibre reinforced composite laminates

To study the behaviour of fibre reinforced composite laminates under static and cyclic in-plane complex stress states, a biaxial loading frame has been developed. A cruciform type specimen is biaxially loaded in its plane using four independent servo-hydraulic actuators. An appropriate control unit keeps the centre of the specimen at the same position during the test. For obtaining reliable biaxial failure data the design of the cruciform specimen is of paramount importance. Finite element simulations of the cruciform specimen to study the influence of various geometrical parameters and experiments on nine selected cruciform specimen types using both strain gages and an optical-numerical full field method for strain determination over the entire biaxially loaded test zone have been carried out. In the present paper, the four most important geometries will be discussed. Since stresses in the biaxially loaded zone cannot be calculated from the experimentally applied loads due to the ill definition of the load bearing area, strain values – the only measurable quantity – are used for the comparison of the different geometry types. For the glass fibre reinforced composite material with a [(±45/0)₄/±45]_T lay-up this led to the proposal of an optimized geometry: a plane cruciform type specimen with a circular central region of reduced thickness and an adapted corner fillet at the intersection of the arms. Using this geometry failure will occur in the biaxially loaded test zone, rather than in the uniaxially loaded arms, at failure strains in the centre of the specimen comparable for the uniaxial load ratios to strains obtained on standard beam specimens. Tests at different load ratios were performed to obtain data as input for failure criteria. This paper focuses on the design of the specimen. In the near future, the biaxial test results will be compared with existing failure criteria.     

Impact damage to thick carbon fibre reinforced plastic composite laminates

Recently, very thick section laminates, up to 20 mm in thickness, have been proposed for the wing skins of large aircraft. Composite components in all aircraft have concerns relating to the presence of accidental damage, but there has been little work to investigate the mechanisms and effects of damage in such thick sections. In this work, carbon fibre composite laminates of up to 12 mm thickness have been subjected to dropped-weight impacts of at most 375 J. Two types of impacts were considered. The first is a central impact where the laminate is completely supported and the second a near edge impact where the laminate is partially supported so that one of its edges is free. The geometry of the damage has been studied using C-scan and deply techniques. The residual strengths of the impact-damaged laminates have been measured in tension and compression. The geometry of damage and level of strength reduction is different for central and edge impacts. Generally, an edge impact causes a greater reduction in compressive strength while a central impact causes more tensile strength reduction.     

单向炭布叠层C/C复合材料的热压法制备及其性能

以经表面处理的石墨、单向炭布、和沥青粉为原料,通过热压烧结制备炭布叠层C/C复合材料.考察了炭布含量对材料密度、孔隙率、弯曲强度以及摩擦磨损的影响,采用MM200摩擦磨损试验机进行了环-块摩擦磨损实验,并借助SEM表征了材料的弯曲断口和磨痕形貌.结果表明:当炭布质量分数为50％时,C/C复合材料的综合性能最好,抗弯强度为112.2MPa,密度为1.72 g/cm3,摩擦系数为0.28,磨损率为3.68×10-13 m3·N-1·m-1.弯曲实验中材料呈“假塑性”方式破坏,断口出现大量纤维的拔出.石墨相含量的增加有利于形成较好的摩擦膜,降低磨损率,保持摩擦系数稳定.     

Reliability-based design of blast-resistant composite laminates incorporating carbon nanotubes

Research efforts have reported on the ability of carbon nanotubes (CNTs) to enhance the stiffness and strength of carbon fiber composites. Systematic design of blast-resistant composites requires considering CNTs–carbon composites characteristics and the uncertainty of blast events. This article suggests a systematic reliability-based design approach where system reliability is used to compute the probability of failure of a composite laminate. A case study for the design of a five-layer composite laminate incorporating CNTs and subjected to uncertain blast event is demonstrated. 0.5–1% CNTs contents by weight seem capable of significantly reducing the composite probability of failure during blast.     

Yield and ultimate strengths of fibre composite laminates

A modified maximum strain failure criterion is proposed to predict the failure of the laminae of fibre-reinforced composites when the laminate is subjected to combined loading. The criterion includes two levels of failure. The lower level predicts the onset of stiffness degradation or ‘yield’ in the lamina. Beyond this level the lamina response is non-linear until the upper lever is reached, which corresponds to ‘ultimate’ failure. After the ultimate failure of a ply in a laminate in a certain mode it is assumed that the failed layer unloads gradually in the same mode only. The theoretical predictions of the present analysis are compared with predictions of other failure theories as well as with experimental data obtained from tests on composite tubular specimens.     

Notched strength of carbon fibre/epoxy composite laminates with a circular hole

This paper improves the two stress fracture criteria proposed by Whitney and Nuismer (known as the point stress criterion and the average stress criterion) to predict the strength of composite laminates with a circular hole. In the point stress criterion, it is assumed that the failure occurs when the stress over some distance (d ₀) away from the notch is equal to or greater than the un-notched laminate strength. In the average stress criterion it is assumed that failure occurs when the average stress over some distance (a ₀) ahead of the notch equals the unnotched laminate strength. Both stress fracture criteria are two parameter models based on the unnotched strength (σ ₀) and a characteristic dimension (d ₀ ora ₀). A simple relation is used for the characteristic length to improve the accuracy while evaluating the notched strength of carbon/epoxy composite laminates. The analytical results are compared well with the existing test results of AS4-carbon/948 Al epoxy [0/90]_4s and [0/ ± 45/90]_2S composite laminates with various hole diameters and specimen widths.     

Notched strength of carbon fibre/epoxy composite laminates with a circular hole

This paper improves the two stress fracture criteria proposed by Whitney and Nuismer (known as the point stress criterion and the average stress criterion) to predict the strength of composite laminates with a circular hole. In the point stress criterion, it is assumed that the failure occurs when the stress over some distance (d ₀) away from the notch is equal to or greater than the unnotched laminate strength. In the average stress criterion it is assumed that failure occurs when the average stress over some distance (a ₀) ahead of the notch equals the unnotched laminate strength. Both stress fracture criteria are two parameter models based on the unnotched strength (σ₀) and a characteristic dimension (d ₀ or a ₀). A simple relation is used for the characteristic length to improve the accuracy while evaluating the notched strength of carbon/epoxy composite laminates. The analytical results are compared well with the existing test results of AS4-carbon/948 A1 epoxy [0/90]_{4 s} and [0/±45/90]_{2 S} composite laminates with various hole diameters and specimen widths.     

The electrical resistance response of continuous carbon fibre composite laminates to mechanical strain

Measurements have been made of the effect of mechanical strain on potential distributions and resistance of unidirectional and multidirectional carbon fibre epoxy laminates. The effects of current flow direction and technique for current introduction on piezo-resistance have been studied. It was found that uniform current introduction at sample edges produced by sputtered Au–Cr contacts across the entire cross-section produced consistently low values of gauge factor of 1.75 for current flow parallel to the fibres and 2.7 for transverse current flow. Non-uniform current introduction, produced variously by local point introduction of current, or use of viscous adhesives producing intermittent contact, resulted in a wide range of apparent gauge factors ranging from 20.6 to −89. These anomalous values may be explained by a model in which the high anisotropy of resistance in unidirectional CFRP maintains initial non-uniform current throughout the sample. Under mechanical strain points of fibre contact will change, altering the distribution of current carrying fibres and leading to local changes in current. Thus changes in potential difference between two points produced by mechanical strain will not be exclusively caused by changes in local resistance. The presence of transverse plies in multidirectional laminates ensures that in plane non-uniform current distributions are largely eliminated, and the effect on piezo-resistance of non-uniform current introduction is minimised.     

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.     

A multi-objective optimization approach for design of blast-resistant composite laminates using carbon nanotubes

A reliable process for the design of blast-resistance composite laminates is needed. We consider here the use of carbon nanotubes (CNTs) to enhance the mechanical properties of composite interface layers. The use of CNTs not only enhances the strength of the interface but also significantly alters stress propagation in composite laminates. A simplified wave propagation simulation is developed and the optimal CNT content in the interface layer is determined using multi-objective optimization paradigms. The optimization process targets minimizing the ratio of the stress developed in the layers to the strength of that layer for all the composite laminate layers. Two optimization methods are employed to identify the optimal CNT content. A case study demonstrating the design of five-layer composite laminate subjected to a blast event is used to demonstrate the concept. It is shown that the addition of 2% and 4% CNTs by weight to the epoxy interfaces results in significant enhancement of the composite ability to resist blast.     

Comparative study of dynamic and static delamination behaviour of carbon fibre/epoxy composite laminates

Dynamic and static delamination characteristics of two unidirectional carbon fibre-reinforced epoxy composite laminates (Hercules MI 1610 and Torayca T300) have been studied under impact and low-speed (2 mm min¹) test conditions. The influence of interlaminar reinforcement with chopped Kevlar fibres on toughness has also been examined. The quasi-static or low-speed delamination tests were conducted with the usual double cantilever beam, end-notched flexure and mixed-mode flexure specimens. To determine the corresponding mode I, mode II and mixed-mode toughnesses under the impact condition, a special specimen design has been adopted and tests were performed with a Charpy impact machine. The novel aspect of the test scheme in the present study is that a single-plane delamination surface with a well-defined fracture mode has been obtained. The dynamic and static delamination characteristics of the same fracture mode were then studied by scanning electron microscopy, and special features were compared. While interlaminar reinforcement with a small amount of chopped Kevlar fibres resulted in an appreciable increase in the quasi-static delamination toughness, it was less effective under the impact condition.     

Improved fracture toughness of carbon fibre/epoxy composite laminates using dissolvable thermoplastic fibres

This paper reports on a novel toughening concept based on dissolvable phenoxy fibres, which are added at the interlaminar region in a carbon fibre/epoxy composite. The composites were prepared by resin infusion of carbon fibre fabric with the phenoxy introduced as a chopped fibre interleaf between the carbon fibre plies. The thermoplastic phenoxy fibre dissolved in the epoxy during curing at elevated temperatures and a phase separated morphology with phenoxy-rich secondary phase was formed upon curing. It was found that the average Mode-I fracture toughness value, G_1c increased tenfold with only 10 wt.% (with regard to the total matrix content) phenoxy fibre added. Other properties such as Young’s modulus, tensile strength and thermal stability were not adversely affected. The mechanical and thermal properties of the neat epoxy–phenoxy blends were also studied for comparison.     

Optimum design of composite laminates for frequency constraints

A method for minimum thickness multilayer rectangular composite laminates is presented which is based on a layerwise optimization procedure under natural frequency limitations. In this method, all types of boundary conditions and their combinations can be taken into account. Analysis is based on the Ritz approximations, and no limitations exist on the number of layers that can be considered or the plate’s aspect ratio. Coupling of flexural and torsional stiffness terms need not be zero in any of the cases. Optimization constraints can be either limits on the fundamental frequency or separation between two natural frequencies. Hybrid combination of layers is also made possible in order to enhance the laminate economy, and a short cut procedure is recommended to save total computational efforts.

Examples are offered that illustrate the method’s capabilities, and the results are verified by comparison to some published ones.     

Thermal joining of carbon fibre reinforced PEEK laminates

The heated press technique was used to thermally join carbon fibre reinforced PEEK. The resulting lap-shear specimens enabled the effect of varying certain geometric and processing parameters to be studied in detail. It was shown that the measured joint strength is dependent upon both the geometry of the lap-shear test specimen and the thickness of the joint itself. Tests at elevated temperatures indicated that these welded joints offer excellent strengths even when tested above the glass transition temperature of the PEEK matrix. It was also shown that the joint strength depends strongly upon the orientation of the outermost fibres in the composite adherends. Joints in which the surface fibres are oriented at 90° to the applied stress were shown to be 40% weaker than those in which the fibres are parallel to the loading direction. A series of tests were also undertaken in order to examine the possibility of joining this fibre reinforced thermoplastic at temperatures below the recommended processing temperature of 380°C. Here, it is shown that by employing a sufficiently long contact time, it is possible to obtain very high joint strengths at temperatures just above the melting temperature of the PEEK matrix.     

Characterization of melded carbon fibre/epoxy laminates

‘Melding’ is a novel in situ method for joining thermosetting composite structures, without the need of adhesives. Laminate joining is achieved using uncrosslinked resin matrix of the pre-preg. This study used Hexply914C pre-preg material to characterize melded CFRP structures produced using the melding method. A designated area of a laminate was maintained at temperatures below 40 °C retaining uncured (B-staged) material, while the remainder of the laminate was cured at 175 °C. After a 2.5 h cure cycle, the cured region showed a high degree of cure (0.88) and glass transition temperature (176 °C). The uncured area of the same laminate was cured in a second stage, simulating an in situ melded joint. By controlling the temperature and duration of the intermediate dwell and affecting minimum viscosity values prior to final cure, low values of porosity (<0.5%) were achieved. The mechanical properties of the resulting joint were consistent throughout the melded laminate. Flexural strength (1600 MPa), flexural modulus (100–105 MPa) and short beam strength (105–115 MPa) values observed where equivalent or greater than those found in the recommended autoclave cured control specimens. After the entire laminate was post cured, glass transition temperatures of 230 °C (peak tan δ) were observed in all areas of the laminate.     

Characterization of unidirectional carbon fibre reinforced plastic laminates

The purpose of the present paper is to outline a set of recommended tests in order to characterize the behavior of carbon fibre reinforced plastic laminates. The tests outlined in this contribution are classification of visual defects, dimension and density, fibre and resin content, tensible strength, apparent interlaminar shear strength, in-plane shear modulus, dynamic torsion property, static bending property, dynamic bending property, and thermal expansion coefficient. As a selected example, results of two tensile tests on CFRP laminates are reported.     

Automation of carbon fibre preform manufacture for affordable aerospace applications

At the present time, the weight and durability benefits of carbon fibre composites for aircraft structure components cannot be provided without very large manufacturing cost premiums. The new processing techniques of liquid and resin film moulding have to date only been proven to be justifiable for complex high value parts. This has primarily resulted from the high materials and labour costs required to manufacture the carbon fibre preform prior to moulding. A three year EPSRC (UK Research Council)/aerospace industry funded project, INFACS, addressed the automation of carbon fibre preform manufacture using low cost materials. The manufacture of a unidirectional tape using a binder coated thread was developed to form a low cost feed-stock for a preform laminating machine. A machine was conceived, designed and built to laminate single curvature parts to net shape with tailored thickness up to a size of 3 m by 1.5 m. It has achieved a net laminating rate of 45 kg/h. For the attachment of details such as stiffeners, the techniques of stitching and pinning were developed into effective processes. Components were processed by the Resin Transfer Moulding (RTM) method to a fibre volume fraction of 58%. Cost modelling of the manufacture of a range of sizes of skin components showed large reductions using the technology developed on the project compared to the current ‘state of the art’ manufacturing technology of automated pre-preg tape laying and autoclave curing. For a regional aircraft tail-plane skin the modelled cost reduction was 60%. The machine and preform development work is continuing within the EPSRC IMI Aerospace Sector project Automated Manufacture of Integrated Composite components (AMICC).