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Abstract The growing use of thermoplastic composites (TPCs) like low-melting polyaryletherketone (LM-PAEK) matrices reinforced with unidirectional carbon fiber (CF) in aircraft structures presents a significant challenge in terms of lightning strikes and electromagnetic interference shielding during aircraft operations. This is due to the weak electrical conductivity of TPC structures, which results in widespread damage when struck by lightning. The repair and maintenance of these extended damaged sites can increase operational costs and loss of flights. Several lightning strike protection (LSP) systems have been developed and implemented to address these concerns. This study evaluated a highly conductive coating with a low filler rate for its effectiveness as an LSP solution for TPCs on exterior aircraft surfaces. The TPC panel without any coatings was first studied. Subsequently, the level of conductivity was increased by incorporating the nanoscale conductive fillers, silver-coated copper (Ag/Cu) nanoflakes, with a silver content of 20 wt.% (Ag20/Cu) and 30 wt.% (Ag30/Cu), correspondingly, into the coating at two loadings of 55 wt.% and 70 wt.% in an epoxy carrier for the surface coatings. The behavior of electrical and surface conductivity was thoroughly examined to understand the impact of Ag/Cu with a high aspect ratio and the effectiveness of the LSP solution. In addition, the spray-coated TPC panels underwent rigorous Zone 2A lightning strike testing using simulated lightning current, in agreement with the industry standard of Society of Automotive Engineers (SAE) Aerospace Recommended Practice (ARP) 5412B. Despite the higher resistance due to the lower conductive coating weight, the TPC panels with Ag30/Cu at loading of 70 wt.% achieved better results than those with Ag30/Cu at loading of 55 wt.%. This is evidenced by the minor structural delamination and CF breakage on the front surface, which proposes a new economic route for a sustainable post-processed LSP system in the aviation industry.

The growing use of thermoplastic composites (TPCs) like low-melting polyaryletherketone (LM-PAEK) matrices reinforced with unidirectional carbon fiber (CF) in aircraft structures presents a significant challenge in terms of lightning strikes and electromagnetic interference shielding during aircraft operations. This is due to the weak electrical conductivity of TPC structures, which results in widespread damage when struck by lightning. The repair and maintenance of these extended damaged sites can increase operational costs and loss of flights. Several lightning strike protection (LSP) systems have been developed and implemented to address these concerns. This study evaluated a highly conductive coating with a low filler rate for its effectiveness as an LSP solution for TPCs on exterior aircraft surfaces. The TPC panel without any coatings was first studied. Subsequently, the level of conductivity was increased by incorporating the nanoscale conductive fillers, silver-coated copper (Ag/Cu) nanoflakes, with a silver content of 20 wt.% (Ag20/Cu) and 30 wt.% (Ag30/Cu), correspondingly, into the coating at two loadings of 55 wt.% and 70 wt.% in an epoxy carrier for the surface coatings. The behavior of electrical and surface conductivity was thoroughly examined to understand the impact of Ag/Cu with a high aspect ratio and the effectiveness of the LSP solution. In addition, the spray-coated TPC panels underwent rigorous Zone 2A lightning strike testing using simulated lightning current, in agreement with the industry standard of Society of Automotive Engineers (SAE) Aerospace Recommended Practice (ARP) 5412B. Despite the higher resistance due to the lower conductive coating weight, the TPC panels with Ag30/Cu at loading of 70 wt.% achieved better results than those with Ag30/Cu at loading of 55 wt.%. This is evidenced by the minor structural delamination and CF breakage on the front surface, which proposes a new economic route for a sustainable post-processed LSP system in the aviation industry. [ABSTRACT FROM AUTHOR]

Aircraft structures are exposed to a variety of operational and environmental loads that can cause structural deformation and fractures. Structural Health Monitoring (SHM) has emerged as a promising solution for in-situ monitoring of structural components. This article presents a state-of-the-art review of SHM in aviation, current regulations, data acquisition sensors and equipment, and damage detection and identification methods. The article discusses in detail the regulations SHM specific to both civil and military aviation. A comprehensive review of conventional electrical resistance sensors, fiber optic, piezoelectric sensors and smart materials used for SHM monitoring in aircraft structures is then presented. The pros and cons of each data acquisition approach were discussed individually. The damage detection and identification section begins by describing the traditional knowledge-based methods that are combined with expert knowledge and theory, then focuses on the applicability in aircraft SHM systems of spectral or frequency domain models. The last part investigates the new paradigm, machine learning and deep learning methods such as CNN and LSTM on different types of aircraft structures through the existing literature. Furthermore, it covers an emerging approach called physics-informed neural networks (PINN), which combines physics and machine learning, and explore its potential for SHM applications. [ABSTRACT FROM AUTHOR]

Cold expansion technology has been widely used in aviation industries as an effective method of improving the fatigue performance of fastener holes. It can improve the fatigue life several times over without adding weight, meeting the growing demand for lightweight and durable aircraft structures. In recent years, it has been extensively studied through extensive experiments and finite element simulations to analyze the residual stresses around the fastener hole. Appropriate process parameters lead to the generation of beneficial residual stresses that influence the material microstructure, thereby improving the fatigue life of the component. This paper summarized factors influencing residual stresses in cold expansion and their effects on fatigue life, and the strengthening mechanism, parameter optimization, and effect of anti-fatigue are discussed from the point of view of the residual stress and microstructure. The development of new cold expansion technologies and the research directions that can realize anti-fatigue technology efficiently are proposed. [ABSTRACT FROM AUTHOR]

The paper introduces an innovative aerospace component production approach employing Wire Arc Additive Manufacturing (WAAM) technology to fabricate near-finished preforms from Ti6Al4V titanium. Tensile tests on WAAM Ti6Al4V workpieces demonstrated reliable mechanical properties, albeit with identified anisotropic behavior in horizontal samples, underscoring the need for optimization. This alternative manufacturing strategy addresses the challenges associated with machining forged preforms, marked by a high Buy To Fly (BTF) ratio (>10), leading to material wastage, prolonged machining durations, elevated tool expenses, and heightened waste and energy consumption. Additionally, logistical and storage costs are increased due to extended delivery timelines, exacerbated by supply issues related to the current unstable situation. The utilization of WAAM significantly mitigates initial BTF, preform costs, waste production, machining durations, and associated expenditures, while notably reducing lead times from months to mere hours. The novelty in this study lies in the application of Wire Arc Additive Manufacturing (WAAM) technology for the fabrication of titanium aircraft components. This approach includes a unique height compensation strategy and the implementation of various deposition strategies, such as single-seam, overlapping, and oscillating. [ABSTRACT FROM AUTHOR]

The machine learning assisted structural optimization (MLASO) algorithm has recently been proposed to expedite topology optimization. In the MLASO algorithm, the machine learning model learns and predicts the update of the chosen optimization quantity in routine and prediction iterations. The routine and prediction iterations are activated with a predefined learning and predicting scheme; and in the prediction iterations, the design variable can be updated using the predicted quantity without running a finite element analysis and sensitivity analysis, and thus the computational time can be saved. Based on the MLASO algorithm, this work first proposes a novel generic criterion-driven learning and predicting (CDLP) scheme that allows the algorithm to autonomously activate prediction iterations in the solution procedure. Second, this work presents the convergence analysis and the computational efficiency analysis of the MLASO algorithm with the CDLP scheme. The MLASO algorithm is then embedded within the solid isotropic material with penalization topology optimization method to solve two-dimensional and three-dimensional problems. Numerical examples and results demonstrate the prediction accuracy and the computational efficiency of the MLASO algorithm, and that the CDLP scheme can remarkably improve the computational efficiency of the MLASO algorithm. [ABSTRACT FROM AUTHOR]

This paper presents a case study conducted in an assembly cell specifically designed for the automated drilling of an aeronautical structure. The study shows how techniques approached by the 4.0 industry have the potential to contribute to manufacturing, breaking the limits imposed by the previous state-of-the art systems. This paper proposes a method that utilizes a committee of neural networks to calculate an indicator for the final quality of drilled holes. The method analyzes data obtained by monitoring the electric current consumed by the drilling system drive. Considering the tests carried out on a real product, the method presents an accuracy of 95% and has the potential to increase the efficiency of the drilling process, reducing the cycle time by up to 25%, since it can avoid measurement steps and physical inspections which increase the cycle time of the drilling process. The proposal contributes to the literature by presenting an unprecedented application and to the praxis by solving a relevant problem of the aerospace industry. [ABSTRACT FROM AUTHOR]

Hidden corrosion remains a significant problem during aircraft service, primarily because of difficulties in its detection and assessment. The non-destructive D-Sight testing technique is characterized by high sensitivity to this type of damage and is an effective sensing tool for qualitative assessments of hidden corrosion in aircraft structures used by numerous ground service entities. In this paper, the authors demonstrated a new approach to the automatic quantification of hidden corrosion based on image processing D-Sight images during periodic inspections. The performance of the developed processing algorithm was demonstrated based on the results of the inspection of a Mi family military helicopter. The nondimensional quantitative measurement introduced in this study confirmed the effectiveness of this evaluation of corrosion progression, which was in agreement with the results of qualitative analysis of D-Sight images made by inspectors. This allows for the automation of the inspection process and supports inspectors in evaluating the extent and progression of hidden corrosion. [ABSTRACT FROM AUTHOR]

Structural health monitoring of riveted aircraft panels is a real challenge for maintenance engineers. Here, a diffused Lamb wave field is used for fatigue-crack detection in a multi-riveted strap-joint aircraft panel. The panel is instrumented with a network of low-profile surface-bonded piezoceramic transducers. Various amplitude characteristics of Lamb waves are used to extract information on fatigue damage. A statistical outlier analysis based on these characteristics is also performed to detect damage. The experimental work is supported by simplified modelling of wave scattering from crack tips to explain complex response features. The Local Interaction Simulation Approach (LISA) is used for this modelling task. The results demonstrate the potential and limitations of the method for reliable fatigue-crack detection in complex aircraft components. [ABSTRACT FROM AUTHOR]

Bonded joints used in wing sections and frames of aircraft structures are mostly exposed to cyclic loadings instead of static ones during their services. Bending types of dynamic loadings are mostly encountered. In this study, the fatigue response of a single lap joint (SLJ) exposed to bending loading was studied with the developed advanced finite-element (FE) model. The cohesive zone model describing the behaviour of the adhesive layer used the damage mechanism, where static and fatigue damages were linked to each other; i.e., the total damage was accumulated because of material deterioration and cyclic plastic separation. This enabled us to predict the fatigue characteristics including the finite fatigue life, crack propagation rate using Paris law. The model was implemented via a user-defined UMAT subroutine offered in ABAQUS-Standard. The numerical model was validated by experiments available in the literature. The fatigue performance of an SLJ subjected to bending loading was investigated for different lap joint configurations. A smaller bending load, a thicker adherend or a longer overlap length (OL) led to enhanced fatigue life. For instance, the fatigue life was observed to increase up to 50 times for a 66% increase in OL. [ABSTRACT FROM AUTHOR]

The paper introduces an innovative aerospace component production approach employing Wire Arc Additive Manufacturing (WAAM) technology to fabricate near-finished preforms from Ti6Al4V titanium. Tensile tests on WAAM Ti6Al4V workpieces demonstrated reliable mechanical properties, albeit with identified anisotropic behavior in horizontal samples, underscoring the need for optimization. This alternative manufacturing strategy addresses the challenges associated with machining forged preforms, marked by a high Buy To Fly (BTF) ratio (>10), leading to material wastage, prolonged machining durations, elevated tool expenses, and heightened waste and energy consumption. Additionally, logistical and storage costs are increased due to extended delivery timelines, exacerbated by supply issues related to the current unstable situation. The utilization of WAAM significantly mitigates initial BTF, preform costs, waste production, machining durations, and associated expenditures, while notably reducing lead times from months to mere hours. The novelty in this study lies in the application of Wire Arc Additive Manufacturing (WAAM) technology for the fabrication of titanium aircraft components. This approach includes a unique height compensation strategy and the implementation of various deposition strategies, such as single-seam, overlapping, and oscillating.

The aircraft structures combat survivability is one of the critical factors,which affect the integration combat efficiency of aircraft.The efficient design of aircraft structures survivability is of great significance to improve the combat ability,decrease the expenses in service,and ensure the well combat readiness of aircraft.At first,the design factors of aircraft structures combat survivability,based on the connotation of survivability and the design requirement of structures survivability for aircraft are combed.Then the design guidelines of aircraft structures combat survivability are discussed.Finally,the design and analysis methods of aircraft structures survivability are presented,under the effects of explosions-overpressure field and gust field,which are the two different aspects of blast wave.The analysis example is presented.Results show that the design factors of aircraft structures combat survivability are similar to the damage tolerance design,and the design process of damage tolerance could be used in structures combat survivability design.The fuselage panel,leading edge,moveable surface and cover,which are sensitive to overpressure of blast wave,can be designed to improve structure survivability by configuration optimization.

Structural health monitoring represents an interesting enabling technology towards increasing aviation safety and reducing operating costs by unlocking novel maintenance approaches and procedures. However, the benefits of such a technology are limited to maintenance costs reductions by cutting or even eliminating some maintenance scheduled checks. The key limitation to move a step further in exploiting structural health monitoring technology is represented by the regulation imposed in sizing aircraft composite structures. A safety margin of 2.0 is usually applied to estimate the ultimate loading that composite structures must withstand. This limitation is imposed since physical nondestructive inspection of composite structures is really challenging or even impossible in some cases. However, a structural health monitoring system represents a viable way for a real time check for the health status of a composite structure. Thus, the introduction of structural health monitoring should help into reducing the stringent safety margin imposed by aviation regulation for a safe design of composite structures. By assuming a safety margin reduction from 2.0 to 1.75 thanks to the installation of permanently attached sensors for structural health diagnostics, this paper assesses the potential fuel savings and direct operating costs through a multidisciplinary analysis on a A220-like aircraft. According to the foreseen level of technology, addressed through the number of sensors per square meter, a DOC saving from 2% up to 5% is achievable preserving, at the same time, all the key aircraft performance. [ABSTRACT FROM AUTHOR]

Hidden corrosion remains a significant problem during aircraft service, primarily because of difficulties in its detection and assessment. The non-destructive D-Sight testing technique is characterized by high sensitivity to this type of damage and is an effective sensing tool for qualitative assessments of hidden corrosion in aircraft structures used by numerous ground service entities. In this paper, the authors demonstrated a new approach to the automatic quantification of hidden corrosion based on image processing D-Sight images during periodic inspections. The performance of the developed processing algorithm was demonstrated based on the results of the inspection of a Mi family military helicopter. The nondimensional quantitative measurement introduced in this study confirmed the effectiveness of this evaluation of corrosion progression, which was in agreement with the results of qualitative analysis of D-Sight images made by inspectors. This allows for the automation of the inspection process and supports inspectors in evaluating the extent and progression of hidden corrosion.

A procedure was developed to detect defective areas in aircraft structure elements by electron shearography method. Nondestructive testing of the quality of full-scale elements of SC «Antonov» aircraft structures was performed: parts of helicopter blade and AN-74 aircraft elevator trimmer. Ref. 14, Fig. 8. [ABSTRACT FROM AUTHOR]

In this paper, a novel nonlinear aeroservoelastic modeling methodology for a morphing wing with bilinear stiffnesses in hinge joints is proposed to investigate the nonlinear aeroservoelastic behaviors during the morphing process efficiently. The notable advantage is that the bilinear nonlinearity in hinge joints and the variation of morphing configuration is modeled theoretically using a parameterized fictitious mode approach. The nonlinear, parameter-varying aeroservoelastic system is represented by piecewise-linear modal-based aeroservoelastic equations according to the different combinations of nonlinear parameters in hinge joints. The proposed nonlinear aeroservoelastic modeling methodology can efficiently investigate aeroservoelastic frequency responses, nonlinear aeroservoelastic vibration, and closed-loop stability because the nonlinear aeroservoelastic model is built via the parameterized fictitious modal coordinates, which can represent the bilinear structural stiffness and variation of morphing parameter simultaneously. To demonstrate the modeling accuracy and efficiency of the present method, a folding wing with two bilinear stiffnesses due to backlash in inboard and outboard hinge joints is selected as a test case. The numerical results show that the open-loop and closed-loop nonlinear behaviors of the morphing wing at different folding configurations can be efficiently investigated. Also, the chaotic phenomenon of the morphing wing subject to control surface excitation was firstly found. [ABSTRACT FROM AUTHOR]

Structural health monitoring of riveted aircraft panels is a real challenge for maintenance engineers. Here, a diffused Lamb wave field is used for fatigue-crack detection in a multi-riveted strap-joint aircraft panel. The panel is instrumented with a network of low-profile surface-bonded piezoceramic transducers. Various amplitude characteristics of Lamb waves are used to extract information on fatigue damage. A statistical outlier analysis based on these characteristics is also performed to detect damage. The experimental work is supported by simplified modelling of wave scattering from crack tips to explain complex response features. The Local Interaction Simulation Approach (LISA) is used for this modelling task. The results demonstrate the potential and limitations of the method for reliable fatigue-crack detection in complex aircraft components.

This paper presents an approach for optimizing practical commercial-scale aircraft wings using sandwich composites in a preliminary design stage. The approach uses lamination parameters as design variables in a continuous optimization step. Structural constraints for classic composite laminate design such as material failure and buckling, and for sandwich design such as crimping, wrinkling, dimpling, and core shear failure are accounted for using industrial-standard and empirical methods driven by finite element analyses. As an application case, optimization studies are performed at a skin panel level on the open-source Common Research Model wing. Optimization trends show areas of the wingbox where sandwich composites offer superior structural performance, as well as potential cost savings by requiring lesser number of stringers. The aim and novelty of this work is to present performance gains that can be achieved using sandwich composites in primary load-carrying aircraft structures when compared with monolithic composite designs and, through this, to provide a motivation for further research and development in sandwich composites and their applications. [ABSTRACT FROM AUTHOR]

Aircraft structures experience various kinds of loads over their entire lifetime, typically leading to fatigue, and ultimately structural failure. To avoid structural failure during operation, the maximum number of flight cycles and flight hours is regulated by laws ensuring continued airworthiness. However, because every flight impacts the aircraft differently, not all airframes are equally stressed at the time of decommissioning. This paper proposes a new retirement criterion based on the so-called fatigue damage index (FDI). The criterion takes into account that aircraft are differently operated and thus enables an individual decommissioning of aircraft without compromising its safety. Based on aircraft sample data covering 95% of the Airbus A320 fleet over two years, the enhanced decommissioning criterion is estimated to significantly extend the average aircraft service life. The impact varies within the fleet, depending on the experienced seat load factors, cruise altitudes, and taxi times considered for the individual aircraft during operation. Whereas seat load factors and flight altitudes significantly affect the defined FDI, the influence of taxi times is only minor. Based on the estimated increase in aircraft service life, the paper argues that for service life extensions, the FDI shall be considered as the limit of validity in the regulatory framework governing the decommissioning of aircraft. [ABSTRACT FROM AUTHOR]

Airworthiness certification is a mandatory but expensive process in aircraft development. To reduce certification cost, it is desired to incorporate certification considerations into aircraft early design stages. The tradeoff between performance and certification constraints and the interactions between the behind-the-scene disciplines require a multidisciplinary design optimization (MDO) method implemented to incorporate certification considerations. Several MDO frameworks currently exist with mixing levels of fidelity and multidisciplinary coupling. However, few of them capture the impacts of certification constraints on the overall aircraft performance. Moreover, preceding MDO studies have mostly focused on optimizing a single design point, whereas little attention is paid to design space exploration. To fill these gaps, this paper proposes a certification-driven platform for airframe early preliminary design. With statistical methods applied, this platform allows efficient design space exploration and multi-objective optimization. To demonstrate the capabilities of this platform, a test case of preliminary horizontal tail design of a large twin-aisle aircraft is performed. The feasibility test and multi-objective optimization conducted in the test case prove that certification constraints play a critical role in the design space exploration at the early preliminary stage. [ABSTRACT FROM AUTHOR]

The problem of reducing the weight of aircraft structures, while maintaining the rest of the material properties, is urgent. Carbon fiber polymer matrix composite materials are used in such applications due to their light weight and high strength. Due to the need to develop radiation-resistant, heat-resistant, lightweight materials from carbon-carbon composite materials (CCCM), in this article, calculations of the stress-strain state (SSS) of a prototype frame by the finite element method were carried out. A detailed description of the initial data of the prototype of the frame was given and the loads acting on the structure were indicated. The physical and mechanical characteristics of the materials used in the construction were given, including the characteristics of carbon fiber based on filament, based on unidirectional carbon tape impregnated with a binder, and based on carbon fiber using epoxy prepreg. The authors created a loading scheme for a finite element model (FEM) of a prototype frame. The calculations allowed to determine the values of the structure rigidity, the prototype frame stability, critical load, stability margin, rigidity of the shell. Comparison of FEM in the initial state and after the loss of stability is given. Conclusions were made about the need to use beam finite elements that simulate a system of spiral, annular and longitudinal ribs of the mesh structure. [ABSTRACT FROM AUTHOR]

Bonded joints used in wing sections and frames of aircraft structures are mostly exposed to cyclic loadings instead of static ones during their services. Bending types of dynamic loadings are mostly encountered. In this study, the fatigue response of a single lap joint (SLJ) exposed to bending loading was studied with the developed advanced finite-element (FE) model. The cohesive zone model describing the behaviour of the adhesive layer used the damage mechanism, where static and fatigue damages were linked to each other; i.e., the total damage was accumulated because of material deterioration and cyclic plastic separation. This enabled us to predict the fatigue characteristics including the finite fatigue life, crack propagation rate using Paris law. The model was implemented via a user-defined UMAT subroutine offered in ABAQUS-Standard. The numerical model was validated by experiments available in the literature. The fatigue performance of an SLJ subjected to bending loading was investigated for different lap joint configurations. A smaller bending load, a thicker adherend or a longer overlap length (OL) led to enhanced fatigue life. For instance, the fatigue life was observed to increase up to 50 times for a 66% increase in OL.

Structural health monitoring represents an interesting enabling technology towards increasing aviation safety and reducing operating costs by unlocking novel maintenance approaches and procedures. However, the benefits of such a technology are limited to maintenance costs reductions by cutting or even eliminating some maintenance scheduled checks. The key limitation to move a step further in exploiting structural health monitoring technology is represented by the regulation imposed in sizing aircraft composite structures. A safety margin of 2.0 is usually applied to estimate the ultimate loading that composite structures must withstand. This limitation is imposed since physical nondestructive inspection of composite structures is really challenging or even impossible in some cases. However, a structural health monitoring system represents a viable way for a real time check for the health status of a composite structure. Thus, the introduction of structural health monitoring should help into reducing the stringent safety margin imposed by aviation regulation for a safe design of composite structures. By assuming a safety margin reduction from 2.0 to 1.75 thanks to the installation of permanently attached sensors for structural health diagnostics, this paper assesses the potential fuel savings and direct operating costs through a multidisciplinary analysis on a A220-like aircraft. According to the foreseen level of technology, addressed through the number of sensors per square meter, a DOC saving from 2% up to 5% is achievable preserving, at the same time, all the key aircraft performance.

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. [ABSTRACT FROM AUTHOR]

The paper presents a summary of non-destructive testing and evaluation (NDT&E) methods applied in inspections of structural aircraft components. This brief review covers the most commonly applied methods such as visual and penetrant inspections, tap-testing, eddy current inspections, shearography, thermography, acoustic emission testing, radiographic and tomographic inspections, and ultrasonic inspections. The general operating principles of these methods as well as their main advantages, limitations and application areas are described below.

This article provides an overview of the evolution of the in-situ consolidation (ISC) process over time. This evolution is intimately linked with the advancements in each of the steps of the ISC manufacturing process, is additive in nature, and is limited by the orthotropic nature of composite materials and the physicochemical behavior of the thermoplastic matrix. This review covers four key topics: (a) Thermal models-simulation tools are critical to understand a process with such large spatial gradients and fast changes. Heating systems once marked a turning point in the development of industrial ISC systems. Today, lasers are the most recent trend, and there are three key issues being studied: The absorption of energy of light by the material, the laser profile, and the laser focusing. Several approaches have been proposed for the distributed temperature measurements, given the strong temperature gradients. (b) Adhesion-this refers to two subsequent mechanisms. In the first place, the process of intimate contact is one by which two surfaces of thermoplastic pre-impregnated composite materials are brought into contact under pressure and temperature. This enables closure of the existing gaps between the two microscopic irregular surfaces. This process is then followed by the healing or diffusion of polymer molecules across the interface. (c) Crystallinity-mostly influenced by the cooling rate, and strongly affects the mechanical properties. (d) Degradation-this refers to the potential irreversible changes in the polymer structure caused by the high temperatures required for the process. Degradation can be avoided through adequate control of the process parameters. The end goal of the ISC manufacturing process is to achieve a high product quality with a high deposition rate through an industrial process competitive with the current manufacturing process for thermoset composites [ABSTRACT FROM AUTHOR]

Structural health monitoring is recognized as a viable solution to increase aviation safety and decrease operating costs enabling a novel maintenance approach based on the actual condition of the airframe, mitigating operating costs induced by scheduled inspections. However, the net benefit is hardly demonstrated, and it is still unclear how the implementation of such an autonomic system can affect performance at aircraft level. To close this gap, this paper presents a systematic analysis where the impact of cost and weight of integrating permanently attached sensors—used for diagnostics- affect the main performance of the aircraft. Through a multidisciplinary aircraft analysis framework, the increment of aircraft operating empty weight is compared with the possible benefits in terms of direct operating costs to identify a breakeven point. Furthermore, the analysis allows to establish a design guideline for structural health monitoring systems returning a safer aircraft without any economic penalties. The results show that the operating costs are lower than those of the reference aircraft up to 4% increase in maximum take-off weight. Paper findings suggest to considering a condition monitoring strategy from the conceptual design stage, since it could maximize the impact of such innovative technology. However, it involves in a design of a brand-new aircraft instead of a modification of an existing one.

Purpose: The main objective of this paper is to develop a model that will combine economic and environmental assessment tools to support the composite material selection of aircraft structures in the early phases of design and application of the tool for an aircraft elevator. Methods: An integrated life cycle cost (LCC) and life cycle assessment (LCA) methodology was used as part of the sustainable design approach for the laminate stacking sequence design. The model considered is the aircraft structure made of carbon fiber reinforce plastic prepreg and processed via hand layup-autoclave process which is the preferred method for the aircraft industry. The model was applied to a cargo aircraft elevator case study by comparing six different laminate configurations and two different carbon fiber prepreg materials across aircraft's entire life cycle. Results and discussion: The results show, in line with other studies using different methodologies (e.g., life cycle engineering, or LCE), that the combination of LCA with LCC is a worthwhile approach for comparing the different laminate configurations in terms of cost and environmental impact to support composite laminate stacking design by providing the best trade-off between cost and environment. Elevator LCC reduces 19% by changing the material type and applying different ply orientations. Elevator LCA score reduces 53% by selecting the optimum instead of best technical solution that minimizes the displacement. Improving the structural performance does not always lead to an increase in the cost. [ABSTRACT FROM AUTHOR]

Experimental characterization and validation of skin components in aircraft entails multiple evaluations (structural, aerodynamic, acoustic, etc.) and expensive campaigns. They require different rigs and equipment to perform the necessary tests. Two of the main dynamic characterizations include the energy absorption under impact forcing and the identification of modal parameters through the vibration response under any broadband excitation, which also includes impacts. This work exploits the response of a stiffened aircraft composite panel submitted to a multi-impact excitation, which is intended for impact and energy absorption analysis. Based on the high stiffness of composite materials, the study worked under the assumption that the global response to the multi-impact excitation is linear with small strains, neglecting the nonlinear behavior produced by local damage generation. Then, modal identification could be performed. The vibration after the impact was measured by high-speed 3D digital image correlation and employed for full-field operational modal analysis. Multiple modes were characterized in a wide spectrum, exploiting the advantages of the full-field noninvasive techniques. These results described a consistent modal behavior of the panel along with good indicators of mode separation given by the auto modal assurance criterion (Auto-MAC). Hence, it illustrates the possibility of performing these dynamic characterizations in a single test, offering additional information while reducing time and investment during the validation of these structures.

The use of integrally stiffened panels (ISP's) in wings of small and medium-sized aircraft is frequent in aviation. These thin-walled structures are often subjected to compressive loads during their life-time and, although the finite element method has become an important tool for engineers to obtain the buckling load of structural members, some results based in charts published by NACA are still used in the preliminary stages of the wing design in many aircraft companies. However, these charts used for calculation of the critical compressive stress consider only idealized stiffened panels and neglect geometric details as the fillet radius used in the current design of ISP's. The objectives of this paper are twofold: (i) to show that the charts published by NACA provide good results for the critical buckling stresses for several geometries of ISP's, when compared to finite element results with proper boundary conditions, providing fillet radii are also neglected in the finite element models and (ii) to show that the values of the critical buckling stresses for local instability of ISP's may be significantly increased when one considers the effect of the fillet radius, meaning that this parameter should also be considered in the optimal design of such structures. Several numerical results obtained with finite element simulations based on different geometrical parameters of ISP's are presented in this study. [ABSTRACT FROM AUTHOR]

The concept of a curvilinearly stiffened panel is promising for aerospace and aircraft structures with cutouts, since the stiffness distribution and loading path can be flexibly tailored in terms of cutouts. However, uncertainties from the manufacturing process can significantly reduce the practical load-carrying capacity of a curvilinearly stiffened panel. Traditional nondeterministic design optimization methods would suffer from a high computational burden because numerous discrete-continuous variables, in which thousands of finite element analyses are usually required, are involved. Therefore, a hierarchical nondeterministic optimization framework of a curvilinearly stiffened panel is established, containing two steps (global search and local search) and four main parts (layout design, nondeterministic design optimization, reshape layout design, and stress validation). In the first-step deterministic optimization, all variables are considered, and a rough result set will be obtained by a global search. In the second-step nondeterministic optimization, all discrete-continuous layout parameters are fixed in the initial stage, and the efficient adaptive-loop method is applied to significantly increase the computational efficiency, in which the variations of material properties and geometric dimensions are considered as uncertainties. The layout also will be reshaped to explore the potential of the load-carrying capacity. Finally, the stress constraint and crippling constraint are verified for the optimum design. An illustrative example indicates the advantage of the proposed framework not merely provides a competitive efficiency and robustness over other nondeterministic approaches but also improves the rationality of deterministic design optimization. [ABSTRACT FROM AUTHOR]