This chapter is concerned with hybrid composite materials for very large lightweight wind turbine blades. It emphasizes the topic of structural and materials aspects of advanced composites in aerospace engineering applications. One of the perspectives of the development of high-energy wind turbines is dependent on the evolution of material concepts for wind turbine blades. Other prospective directions are reliant on the integrated structural health monitoring, active vibration reduction systems, advanced environmental durability, and novel maintenance issues. Upscaling of wind turbine blades requires specific solutions due to structural and material constraints. The chapter reviews several composite laminates including ply drop-off concepts, self-healing composites with nanostructures, and fracture mechanics of such material systems.

The incorporation of carbon nanotube-grafted carbon fibers (CNT-CF) into polymer matrices provides highly-enhanced physical properties to the materials. They could be ideal candidates to be integrated into multi-functional composites applications in several industries. However, in order to take advantage of carbon nanotubes (CNTs), it is essential to keep all the CNTs on the surface of fibers during their implementation phase in the composite manufacturing process. Besides, some unresolved safety issues remain during the integration phase of such hybrid forms into polymer matrices. This research is making an attempt to develop an effective plasma surface treatment method either to keep the maximum possible amount of CNTs on the fibers and to modify the surface properties of CNT-CF which is highly necessary prior to the composite fabrication. In order to provide better understanding of the modified structures in a domain ranging from microscopic to atomic scales, several characterization studies were realized by X-ray Photoelectron Spectroscopy (chemical structure) and Scanning Electron Microscopy (microstructure). This research is expected to provide valuable information for further studies to develop hybrid composites where multifunctionality is the main concern.

This paper deals with the dielectric barrier discharge assisted continuous plasma polypyrrole deposition on CNT-grafted carbon fibers for conductive composite applications. The simultaneous effects of three controllable factors have been studied on the electrical resistivity (ER) of these two material systems based on multivariate experimental design methodology. A posterior probability referring to Benjamini-Hochberg (BH) false discovery rate was explored as multiple testing corrections of the t-test p values. BH significance threshold of 0.05 was produced truly statistically significant coefficients to describe ER of two material systems. A group of plasma modified samples was chosen to be used for composite manufacturing to drive an assessment of interlaminar shear properties under static loading. Transversal and longitudinal electrical resistivity (DC, ω =0) of composite samples were studied to compare both the effects of CNT grafting and plasma modification on ER of resultant composites.

Among the principal load-carrying parts of an aircraft, the center wing box yields the main mechanical load-carrying area and is composed of many structural components and elements such as upper and lower skin panels, internal rods, ribs, stringers, front and rear spars. For the structural analysis of such real complex systems, simplified representations of the same sections would be used alternatively to determine whether components could withstand loadings without experiencing failure, thus stay in the margin of safety. Nevertheless, these components should contribute to weight efficiency while carrying in-plane loads and distribute out-of-plane loads to in-plane members safely. Materials selection plays a crucial role for the determination of such candidate materials along with their material properties in specific structural applications. It is usually performed by considering proper objectives, constraints and free variables with respect to functions of the components of a system. Since Composite Material Handbook (CMH-17), Metallic Materials Properties Development and Standardization (MMPDS), and Preliminary Material Properties Handbook (PMP-HDBK) databases are embedded in the materials selection software (CES Edupack), best candidate materials would be easily identified by using coupling constant(s) such as in multiple-constraints designs. In this study, Ashby’s methodology was applied to determine best candidate materials for constructing skin panels by considering them as bending plates. According to materials selection approach, continuous fiber reinforced epoxy composites was stipulated as one of the best candidate materials. Computational failure analysis was then carried out by referring the proper material properties from the materials selection software.

Abstract Purpose Aircraft control cables are usually made of steel materials and subjected to deformation because of the motion of control surfaces such as aileron, rudder, elevator and trailing edge flaps. Investigation of the relationship between material properties and alloying elements would therefore be explored. Design/methodology/approach This study is focused on the modeling of mechanical properties of carbon steels concerning the content of alloying elements by using response surface methodology with false discovery rate (FDR) correction approach. SAS Institute JMP data analysis software was used to develop response and argument relationships in various carbon steels without including thermomechanical treatment effect. Mechanical properties were considered as tensile strength, yield strength, ductility, and Brinell hardness. Carbon (0.28 Wt.%-0.46 Wt.%) and manganese (0.7 Wt.%-0.9 Wt.%) proportions were gathered from ASM Handbook. Linear regression models were tested for the statistical adequacy by using analysis of variance and statistical significance analysis. A posterior probability, which refers to Benjamini–Hochberg FDR (BH-FDR), was embedded as multiple testing corrections of the t-test p-values. Findings BH-FDR significance threshold of 0.05 was produced statistically significant coefficients to describe mechanical properties of carbon steels. Originality/value The effect of statistically developed graphical interactions of alloying elements on the common mechanical properties of such steels would provide prompt comparison to material suppliers and part manufacturers except those subjected to thermomechanical treatment applications.

This study is based on the statistical analysis of interlaminar fracture toughness of various laminated polymer composites used in aerospace applications through parametric and non-parametric tests. Tukey’s, Dunkan’s, two-sample t-test, and Kolmogorov–Smirnov tests are used to analyze tensile mode interlaminar fracture toughness of various fiber-reinforced polymer composites obtained by beam theory, modified beam theory, and modified compliance calibration method. Among the studied composite samples, modified compliance calibration method provided the highest average interlaminar fracture toughness, whereas the modified beam theory showed the lowest one. Room temperature cured carbon fiber-reinforced composite samples exhibited higher interlaminar fracture toughness than the autoclave cured samples. Two-sample t-test show that all methods are found coherent with each other in terms of being significantly different. On the contrary, Kolmogorov–Smirnov test revealed no significant difference. Besides, Tukey’s and Duncan’s tests exhibited almost the same results in regard to significant differences except those obtained by the modified compliance calibration method. Two-sample t-test method should have to be performed in order to observe significant relations rather than Kolmogorov–Smirnov test, since the results of Tukey’s and Duncan’s tests are only consistent with each other for the beam theory and modified beam theory method.

Purpose Python codes are developed for the versatile structural analysis on a 3 spar multi-cell box beam by means of idealization approach. Design/methodology/approach Shear flow distribution, stiffener loads, location of shear center and location of geometric center are computed via numpy module. Data visualization is performed by using Matplotlib module. Findings Python scripts are developed for the structural analysis of multi-cell box beams in lieu of long hand solutions. In-house developed python codes are made available to be used with finite element analysis for verification purposes. Originality/value The use of python scripts for the structural analysis provides prompt visualization, especially once dimensional variations are concerned in the frame of aircraft structural design. The developed python scripts would serve as a practical tool that is widely applicable to various multi-cell wing boxes for stiffness purposes. This would be further extended to the structural integrity problems to cover the effect of gaps and/or cut-outs in shear flow distribution in box-beams.

In the typical structural design of an aircraft, the design criteria and sizing are considered within the scope of structural and aerodynamic data. The issues regarding the structural design and analysis are usually performed by considering the multiple constraints such as parts designed should be easily manufactured, strong enough for its intended purpose, and lightest possible. Within the framework of the structural design of an aircraft, tail, wing and control surfaces are used to be evaluated depending on the structural integrity concerns as well. For instance, one of the common structures used in aircraft substructures is in the form of a box beam. By using idealization approach, shear stress analysis and computation of stiffener loads could be performed quite easily. However, in order to overcome the conflicting objectives in an aircraft structural design, finite element models may undergo many updates. In most of these computational analyses, model verification could be achieved either by analytical or theoretical methods that provide approximate solutions. Besides, studies on the uncertainty analysis would be considered accordingly. In this study, subsequent to idealization, structural analysis of various aircraft box beams are realized using in-house developed python codes. Dimensional changes in spar and panel sections are studied in detail. The distribution of shear flows and stiffeners loads are obtained by numpy module. The visualization of the results is provided by matplotlib module. The developed python codes could easily serve as a practical tool for a wide variety of aerospace applications.

A finite length cylindrical bar of an oxygen free high purity copper cylindrical bar impact on a rigid anvil is studied by using finite element method. Johnson-Cook plasticity is employed in terms of the material model. Simulation time is based on the time when the kinetic energy reaches zero. Mesh gradation is performed in order to represent the accuracy of the studied model. Computational cost in terms of CPU time and wall clock time with respect to corresponding time steps are evaluated accordingly. Comparison of the final profiles of cylindrical bar including mushrooming effect is realized by referring the previously published studies as well. The average error, which considers final length, final diameter, and final bulge width of the bar, is computed regarding the test values. Analytical deformation analysis is performed using Hawkyard’s energy approach rather than Taylor’s approach. Use of the global energy balance instead of the local momentum balance at the plastic front and use of true strain rather than engineering strain brought notable improvements in large deformation cases.

Ballistic impact is usually classied under military and general applications. Range of initial striking velocities, angle of incidence of impact, conguration and material properties of both target and projectile govern the ballistic impact. Among impact problems, projectile impact and bird strike in aircraft structures are usually classied under the same velocity regime called high velocity impact regime. ere are wide range of materials used to construct structural components against such ballistic impact threads. Among such candidate materials, ber reinforced laminated composites bring solutions where the weight gain is important without sacricing the overall performance. Failure modes in target due to impact loading would also be further identied in terms of delamination, penetration, perforation, etc. In this study, interlaminar and intralaminar failure analyses of composite plates due to impact of a mild steel projectile at a velocity of 54 m/s were performed in Abaqus Explicit v6.19. Johnson-Cook constitutive and damage model for the conical frontal portion with cylindrical shank steel projectile were adopted. For such conical projectiles attached shock waves in general considered unenviable. Fracture strain, stress triaxiality, strain rate, and displacement at failure data were introduced in nite element solver. General contact algorithm was employed to simulate the contact behavior between surfaces during penetration. Stress analysis and deformation of composite plates due to impact loading were studied accordingly. e maximum stress observed in bullet and composite plate were found around 250 MPa and 1480 MPa, respectively. Deformation was found almost localized in target material. Compared to data gathered from literature, implementation of cohesive zone modeling notably provided valuable contribution throughout the failure analysis of composite plates due to impact loading.

Computational impact analysis of an E-glass fiber reinforced laminated composite structure with a conical nose shaped projectile is performed including delamination analysis. In addition to intralaminar damage analysis, interlaminar damage analysis is considered due to the laminated conguration of the protective structure. reat is made of mild steel projectile and it is modeled using Johnson-Cook material model with ductile damage criterion. Intralaminar and interlaminar damage modeling of target material is realized in the frame of Hashin damage criteria and quadratic nominal stress criterion with Benzeggagh-Kenane fracture criterion, respectively. Stress and damage distribution both in target and threat materials are obtained by mesh gradation analysis via Abaqus/Explicit v6.19. Eect of delamination analysis in computational impact modeling is evaluated by considering the computational cost regarding CPU time and wall clock time. Larger von Mises stresses obtained when excluding interlaminar debonding in the analysis. It is found that delamination analysis signicantly improved the damage evaluation of composite laminates owing to impact loading with small computational eort.

Among the principal load-carrying parts of an aircraft, the center wing box yields the main mechanical load-carrying area and is composed of many structural components and elements such as upper and lower skin panels, internal rods, ribs, stringers, front and rear spars. For the structural analysis of such real complex systems, simplified representations of the same sections would be used alternatively to determine whether components could withstand loadings without experiencing failure, thus stay in the margin of safety. Nevertheless, these components should contribute to weight efficiency while carrying in-plane loads and distribute out-of-plane loads to in-plane members safely. Materials selection plays a crucial role for the determination of such candidate materials along with their material properties in specific structural applications. It is usually performed by considering proper objectives, constraints and free variables with respect to functions of the components of a system. Since Composite Material Handbook (CMH-17), Metallic Materials Properties Development and Standardization (MMPDS), and Preliminary Material Properties Handbook (PMP-HDBK) databases are embedded in the materials selection software (CES Edupack), best candidate materials would be easily identified by using coupling constant(s) such as in multiple-constraints designs. In this study, Ashby’s methodology was applied to determine best candidate materials for constructing skin panels by considering them as bending plates. According to materials selection approach, continuous fiber reinforced epoxy composites was stipulated as one of the best candidate materials. Computational failure analysis was then carried out by referring the proper material properties from the materials selection software.

Various levels of armed conflicts continue in many parts of the world. Many of these conflicts occur in regions close to o ur geography and are closely related to the security of Turkey. Multinational forces such as NATO are taking m easures to prevent these conflicts and are often involved in these conflicts through their military forces. Conflicts naturally force countries to need security measures. Therefore, the need for lig ht weapons and light weapons ammunition is at th e center of these increasing needs. In this study, evalu ations are driven by using Machine and Chemical Industries Corpo ration, NATO Army Armaments Group (NAAG), US Army Research Development Engineerin g Command (RDECOM), US Army Armament Research Development and Engineering Center (ARDEC) and Alliant Techsystems Inc. (ATK) data for the light weapons and the production of light weapon ammunition.

Analyzing the failure behavior and damage characteristics of polymer composites at different scales have become a rising challenge to many researchers. From the microscale aspect, a number of early studies states that most interfaces are buried inside the composite material, thus some of the material combinations or substitutions for existent reinforcing fibers, adherents or epoxy matrices are usually difficult to achieve due to lack of understanding of structure and composition of fiber-matrix interface. However, with proper characterization of fibers and composite interfaces via various recent characterization techniques, in-situ analysis of cohesive zone that physically formed in composite materials would be satisfactorily established. This would further be used to obtain better interpretation on the failure behavior of such composites, since most of computational failure analysis apparently weighs on defect free reinforcement phase and perfectly uniform interfacial boundary region. Moreover, interfacial adhesion promotion in polymer composites via various approaches reveals the fact that interfacial region would not be simply assumed as a weak uniform boundary layer so that uniform cohesive zone modeling could not be fully capable of characterizing fiber-matrix debonding in such materials. In order to perform more realistic micromechanical damage analysis by finite element method, the statistical models generated by in-situ obtained reinforcement and interface/interphase geometry via various microscopic and spectroscopic analyses were used in this study. Finite element model was generated at a certain level of statistical significance. Smoothing process was performed in order to avoid the formation of saw-like apperance of interphase-matrix region and/or reinforcement phase. The mechanical properties of interface/interphase can be determined by interfacial shear strength tests; the interphase material data can be defined in the finite element program accordingly. Since the interfacial debonding was considered the propagating mode of failure, the use of cohesive zone model was revealed the differences in failure behavior of polymer composites in Abaqus/Explicit 6.14 with the developed Python code. Instead of using the idealized geometry of reinforcement phase and interface/interphase region of the composite material, the finite element model was generated using statistically resembling in-situ characteristics of polymer composites in order to improve current micromechanical damage analysis that may further be helpful to perform more accurate multiscale computational failure analysis.

Fiber reinforced composites have already been integrating in various structural applications for several decades. The determination of their failure behavior under the applied loads are being extensively researched with the applicable damage models which mainly depend on the continuum mechanics. Even though the mechanical properties of carbon fibers are vulnerable to several micro structural defects such as surface flaws, micro crazes and cracks, integration of the cohesive zone in order to represent the fiber-matrix interface/interphase remain uniform in the structural finite element analyses. This research has been introducing an attempt to develop a numerical model in order to implement an arbitrary cohesive zone for the determination of failure behavior of balanced symmetric composite laminates under the applied in-plane mechanical loadings. The thickness of the cohesive zone is generated randomly by a way of fitted statistical distributions/empirical distributions, obtained by statistical data analysis, resembling the real sample behavior along the entire length of continuous fiber reinforced laminates for the damage modeling.

The aircraft structural components should withstand the major stresses especially in take-off, flight, and landing phases. Throughout the design process of such structural components, engineers are usually used to modify their design to meet several constraints. Since the conservative design methodologies were time-consuming, Ashby's methodology was introduced by CES EduPack in order to indicate the best candidate materials for the construction of light and safe components in an aircraft prior to finite element analysis. Depending on the chose constraints and objectives in the models, the best candidate materials were mainly attributed to polymer matrix composites. The influence of the curing method on the mechanical properties of the candidate materials were also examined in this paper. Based on the Hashin's damage criteria, the effect of variations in elastic and ultimate properties have been studied via structural finite element analysis.

This study shows the selection of materials for light and safe internal rod for an aerospace application which is assembled to a selected strut fitting in the center wing-box of an aircraft. Particularly the minimization of mass is considered as an objective while the safety performance of the material is preserved. In order to do that, the strength and the stiffness limited design is examined separately at first. After that, the selection of material is developed by considering multiple constraints as strength and stiffness. To identify the material index for each case, the relevant equations are introduced to observe the effect of selected constraints. CES EduPack 2016 is used while selecting the material since it has already included the steps for the selection process. For the material selection process in CES EduPack, aerospace materials are chosen from Level 3 Database. Either in strength or stiffness-limited design at minimum weight with specified rod length, the best candidate materials are found as PEEK/carbon fiber, Cyanate ester/carbon fiber, and Epoxy/carbon fiber composites. When the selection of material is developed by considering multiple constraints as strength and stiffness at minimum mass, both stiffness and strength leads two independent equations for the mass. It is obvious that, the mass of the rod must be same for both limitation cases (m1= m2) and it gives the coupling constant Cc as L/δ. By applying the rules of multiple constraints design, the best choice is found as Cyanate ester/carbon fiber composite. As the selection of engineering materials become very crucial due to severe application conditions, the best material for the preferred design should be identified carefully. CES EduPack provides highly effective and flexible route in selecting materials for a particular application as examined in the case of internal rod design for an aerospace application.

Due to the discovery of their incredible functional properties, carbon nanotubes (CNTs) have drawn a great deal of interest from both academic and industrial research teams in the past few years. Since novel materials are to be integrated in structural and functional applications in several fields, inclusion of CNTs as a reinforcement component in polymer matrix composites (PMC) could bring new solutions. However, in order to obtain more advanced CNTs composites, the amount of strong bonding between CNTs and matrix must be realized to ensure the effective stress transfer in a PMC. This research aims to establish an efficient dielectric barrier discharge technique for the surface modification of CNTs grafted carbon fibers (CNTs-CF) with plasma polypyrrole (PPPy) in order to be used in PMC. It is found that response surface methodology can be applicable in modeling to evaluate the effects of important process variables on electrical resistivity of CNTs-CF. From low to high plasma powers, X-ray Photoelectron Spectroscopy studies revealed the loss of α- and β-carbons in pyrrole ring. The higher the plasma power the lower the electrical conductivity and the higher the mechanical properties.

The radiofrequency plasma (13.56 MHz) was employed to polymerize trimethyl borate (TMB) monomer/N 2 gas mixture on the surface of biaxially oriented polypropylene (BOPP) films. Plasma polymer coated poly-propylene films were examined by flame retardancy test (limiting oxygen index, LOI). The highest LOI value calculated for the untreated BOPP sample was 18.4 (v/v O 2 %) and 24.2 (v/v O 2 %) for the 55 W 30 minutes treated sample. The plasma polymers were characterized by FTIR spectroscopy and AFM. Accord-ing to the FTIR results, the -OH, B-CH 3 , B-O, and BH 2 functional groups were detected. It was found that the highest surface roughness belonged to 40 W 30 min treated BOPP sample which was calculated as 9.78 nm (10 μm × 10 μm). Moreover, the wettability of the modified BOPP film surfaces was characterized via contact angle measurements. The water contact angle values have decreased from 109.6o the lowest value of 68.2åfter the plasma treatment. The results showed that TMB/N 2 plasma modification could be used as an alternative method for the enhancement of flame retardancy and hydrophilicity of BOPP film.

Flame retardancy of polypropylene films was studied by plasma polymerisation technique. The surface of BOPP film was modified by boron containing plasma polymers at different plasma conditions. Plasma polymer coated polypropylene films were examined by flame retardancy test (limiting oxygen index, LOI) and TGA. Boron containing plasma polymer deposition on the film surface showed an improvement of flame retardancy. Furthermore, TGA thermograms pointed out that the first degradation temperature of treated polypropylene film was increased from 331 to 396°C, and the second degradation temperature was shifted from 401 to 455°C. Also, the plasma polymers were characterised by FTIR spectroscopy and XPS. According to XPS results, the BOPP surfaces treated with TMB showed significant difference in the composition with respect to the untreated sample. The FTIR spectra of plasma polymers obtained indicated that when the treatment time was increased to 60 min with a constant discharge power at 80 W, the absorption intensities of all the functional groups increased. As a result, boron containing plasma polymer treatment was found to be an effective method in enhancing the flame retardancy of BOPP film.