Aircraft Wing Structures Research Articles

Multiscale Progressive Failure Analysis for Composite Stringers Subjected to Compressive Load.

The fiber-reinforced composite stringer is commonly used in large civil aircraft wing structures. Under compression loads, it exhibits complex failure modes, with matrix cracking being one of the most common. The quantitative analysis of matrix failure is important and difficult. To address this issue, a multiscale method combining the generalized method of cells (GMC) and macroscopic FEM models is employed to quantitatively predict matrix damage and failure. The extent of matrix damage in the composite structure is represented by the number of failed matrix subcells within the repeating unit cells. The 3D Tsai-Hill failure criterion is established for the matrix phase, and the maximum stress failure criterion is applied to the fiber subcell. Upon meeting the criterion, the stiffnesses of the failed subcells are immediately reduced to a nominal value. In the current study, the ultimate loads, failure modes and load-displacement curves of composite stringers subjected to compressive load are obtained by the experiment approach and the proposed multiscale model. The experimental and simulation results show good agreement, and the multiscale analysis method successfully predicts the extent of matrix damage in the composite stringer under compressive load. The number of failed matrix subcells quantitatively evaluates the damage extent within a 2 × 2 GMC model. The findings reveal that matrix subcell failures primarily occur in the 45° and -45° plies of the middle part of the stringer composite.

Active flutter suppression for an aircraft wing structure by utilizing FE analysis for Al7075+Sic alloy

An instable stimulated vibration resulting from the combination of aerodynamic, inertia, and aeroelastic forces is known as "active flutter" in the context of aircraft wing structures. The primary destabilizing element that the airplane structures display and which causes active flutter is limit cycle oscillation. Ultimately, it results in fatiguerelated harm and catastrophic aircraft mishaps. Two modern techniques, namely enhanced aero-elastic analysis and flutter control for wings and panel structures, are established by means of these developing technologies. comparable stiffness values and comparable stresses are assessed, together with theoretical and practical approaches, to enhance the features of aero-elastic analysis of aircraft structures. Additionally, it makes reducing flutter mandatory in order to enhance aero-elastic stability. Using Ansys software, a thorough FE analysis is carried out to improve and analyze the benefits of Al7075+ SiC alloy aircraft wing structure. With the aid of Ansys software, a thorough FE study is carried out to enhance their aeroelastic stability and to offer recommendations for the design of revised wings and panel structures. For the chosen NACA4412 aircraft wing, modal analysis is carried out, and frequencies are assessed. The advantages of AL7075+ SiC alloy material will be examined with equal stress, equivalent strain, and equivalent stiffness.

Effect of chemical milling and anodizing on the high cycle fatigue behaviour of 2024-T3 aluminium alloy

Aluminium 2024-T3 alloy, widely employed in transport aircraft wing and fuselage skin structures, necessitates thickness reduction for weight optimization. Chemical milling is utilized for this purpose; however, it removes Cladding (as a standard corrosion protection layer), prompting subsequent anodizing and sealing to provide corrosion protection. Literature indicates that anodizing or chemical milling can significantly reduce fatigue life, and limited data have been available on the life of chemically milled anodized materials. In this study, constant amplitude stress-life (S-N) testing was performed on two sets of fatigue samples: Cladded Aluminium Alloy (CAA) 2024-T3 sheet specimens (1.6 mm thick) and chemically milled, anodized and sealed (CMAS-AA) Aluminum Alloy sheet specimens (1 mm thick). S-N curves were generated at R=0.1 and room temperature under stress control. SEM micrographs showed micropores in CMAS-AA and a standard plain surface in CAA, consistent with prior findings. Both types exhibited 3-5 times shorter fatigue life than electropolished samples, as reported in MIL Handbook HDBK-5J, highlighting the importance of corrosion protection techniques for structural integrity. Residual stresses were observed, with cladded samples displaying compressive and CMAS-AA exhibiting tensile stresses. Fractography revealed no abnormal fatigue fracture features, and computed plane stress fracture toughness values estimated using measured crack size from the failed specimens matched standard values of aluminium alloys.

Non-probabilistic credible reliability analysis of the composite laminate

For the reliability analysis of composite laminate structures, there is a problem of difficult data acquisition. Traditional reliability analysis methods require a large amount of data to maintain calculation accuracy. Therefore, the traditional reliability analysis method is not applicable to composite laminates. For the above-mentioned problems, this paper presents a method to calculate the credible reliability of composite laminate structures using the Bayesian theorem. The method is distinguished from the traditional uncertainty analysis methods. The likelihood function and the newly introduced samples are first used to correct the prior distribution of the uncertain parameters. Furthermore, the posterior distribution, containing information from both the sample and prior distributions, can be obtained. Ultimately, based on the acquired posterior distribution, the probability density function can be updated with newly introduced samples. Using the Hashin failure criterion, the failure probabilities of the fibers and matrix of the composite laminate structure are quantified, respectively. Then, the credible reliability of the composite laminate is obtained. Numerical example results show the proposed method can maintain high accuracy with fewer samples and demonstrate the validity of the model. For the numerical calculations, the deterministic parameters of the distribution function and the functional distribution parameters were verified, composite laminates and aircraft wing structures were analyzed. The experimental results show that the error between the probability density function of the parameters and the probability density function predicted by the sample is less than 0.045. Compared with the traditional reliability analysis methods for composite structures, this method can effectively measure the credible reliability of composite laminate structures with fewer samples and similar accuracy. The method enables real-time updating of the parameter probability density function by newly introduced samples and improves the confidence level.

Multi-objective design and optimization of wings

Nowadays, with the fast development of technology, aircrafts have been successful advanced. As we all know, wings are one of the most significant parts of the aircrafts, therefore, multi-objective design and optimization of wings has become a critical issue among many researchers. This paper mainly introduces several approaches to improve the wing in different conditions, including optimization design of civil airplanes, design and optimization of aircraft wing structure at low Reynolds numbers, Wing Design and Aerodynamic Characteristics of Biomimetic Flapping Wing Micro Air Vehicles and so on, the purpose of this article is to give a brief introduction to these relatively new research methods and make the public understand more about this field. The primary approaches of this article are literature analysis and review. This paper finds that algorithm and artificial intelligence are fundamental to improving wings, and all the design needs to reduce the amount of energy as much as possible, which is also the most important requirement.

Modeling and analysis of multi-functional self-healing material using Runge-Kutta Method for investigation of aircraft wing structure

ABSTRACT This paper presents the deployment of Runge-Kutta method to overcome the main challenge in analysis of failure of aircraft wing structure subjected to wind pressure and point loading. Failure phenomenon of any structure is time dependents and is typically referred as dynamic in engineering mechanics and is fairly a complex to investigate. In this context, the dynamic analysis concept has successfully implemented by using computer program in SCILAB software. The numerical technique is adopted as Runge-Kutta fourth order (RK4) method for performing dynamic behaviour of wind structure. The demonstration of failure mode of wing structure is based on function of time. The numerical approach is deemed to provide a detailed description of these phenomena affecting the overall dynamic of failure envelope of wing structure. The parametric study is presented; likewise the effect on failure of wing structure by changing different wind pressure, length and moment, respectively. The wing structure is analytically validated against available literature. Finally, others important failure results obtained from this analysis has discussed in detail.

Non-probabilistic credible set model for structural uncertainty quantification

A non-probabilistic credible set model for uncertainty quantification structural parameters is proposed in this paper. The developed method surpasses the probabilistic uncertainty quantification method and traditional non-probabilistic uncertainty quantification method by providing both the distribution range of uncertain structural parameters and the corresponding credibility for scant samples, thereby reducing the computation time and improving the credibility of the uncertainty quantification process. The average inverse-distance information entropy, the expanded uncertainty and credibility are introduced to establish a one-dimensional non-probabilistic credible set uncertainty quantification procedure. The basic theory of non-probabilistic sets and a correlation analysis of multidimensional uncertain parameters are reviewed. This theory and analysis are used to develop a multidimensional non-probabilistic credible set uncertainty quantification method. The optimal non-probabilistic set is searched to determine the local coordinate system. The multidimensional uncertainty quantification is subsequently transformed to a one-dimensional uncertainty quantification. A civil aircraft wing structure case, a stiffened plate case and a dynamic case are utilized to verify the feasibility of applying the developed method to various engineering problems.

Multi-Objective Reliability-Based Partial Topology Optimization of a Composite Aircraft Wing

Reliability-based partial topology optimization (RBPTO) is a new approach for aircraft structural design; however, it brings computational complexity and makes aeroelasticity analysis quite challenging. Therefore, the present study proposed the multi-objective reliability-based partial topology optimization of a composite aircraft wing using a fuzzy-based metaheuristic (MRBPTOFBMH) approach. The objective is to obtain an optimal layout including partial topology and sizing of the aircraft wing structure. Here, an optimal aeroelastic structure is designed by taking into account the uncertain nature of material properties and the permitted transverse displacement. To increase computational efficiency in the design process, a non-probabilistic approach called a possibilistic safety index-based design optimization (PSIBDO) with fuzzy uncertainties is proposed to quantify the uncertainties due to aeroelastic and structural constraints. Various optimum partial topological shapes and sizing of aircraft wing structures with various PSI values in the range of [0.001, 1.0] have been obtained in a single optimization run. These outcomes, including deterministic and reliable optimal aircraft wing structures, demonstrate the high effectiveness of the proposed MRBPTOFBMH technique to alleviate the complexity of unconventional aircraft wing structure design. The findings also reveal the ease in cooperation of the suggested technique with a high-performance multi-objective evolutionary algorithm (MOEA) and its application in real-world multi-objective design optimization (MODO) problems with the least computational requirements against the traditional method’s multiple runs. Furthermore, the proposed methodology can generate potential aircraft wing structures in a range of m = [89.38–127.84] kg, and flutter speed = [285.61–632.78] m/s, that adhere to all the constraints requirements.

Thermal analysis of a high-altitude solar platform

High Altitude Long Endurance (HALE) aircraft operate under adverse thermal conditions, with ambient pressures and temperatures very low and at the same time high amounts of heat introduced by sun radiation. Thus, thermal management of the aircraft systems, such as electronics and batteries is a very challenging task. A first step in solving this is generating accurate models of the thermal dynamics of the HALE. This paper presents the thermal analysis of a solar-electric stratospheric HALE at the ground case. A thermal mathematical model based on first principles was developed for such an analysis. In a further step, an initial test campaign was performed. The test campaign included static and dynamic temperature measurements on an aircraft wing structure segment. The experiments throughout this test campaign showed good compliance of the results with the previously derived mathematical models, with differences of less than 5 ,^{circ }C between the measured and simulated temperature curves.

Laminar composite materials for unmanned aircraft wings

Unmanned Aerial Vehicles (UAVs) have high popularity, especially in the military field, but are now also being applied to the private and public sectors. One of the UAV components that require high material technology is the wing. The latest material technology developed as a material for unmanned aircraft wings is a composite material that has high strength and lightweight. This research aims to identify composite materials that can be used for unmanned aircraft wing structures. The method used in this research is a qualitative method with a literature study approach. The results of this theoretical study show that some of the latest composite materials that have been developed into materials for unmanned aircraft wings are Laminar Composites with a sandwich structure. Laminar and sandwich composites consist of various constituent materials such as Balsa wood fiber-glass and polyester resin, microparticles, Carbon Fibre Reinforced Polymer, polymer matrix composites reinforced with continuous fibers, Polymer matrix composites, E-glass/Epoxy, Kevlar/Epoxy, Carbon/Epoxy, woven fabrics, acrylonitrile butadiene styrene-carbon (ABS) laminated with carbon fiber reinforced polymer (CFRP) and uniaxial prepreg fabrics. Laminar and sandwich composite materials are a reference for developing unmanned aircraft wing structures that have resistant strength and lightweight.

Multifunctional Carbon-Basalt hybrid composites against bird strike

A bird strike is a common occurrence during an aircraft's takeoff and landing and is classified as a high-velocity impact, which can cause severe damage, a decrease in aerodynamic efficiency, and may lead to catastrophic failure. High strength composite materials are used in the wing structure to resist the bird strike. Stealth is an important factor for military aircraft, as it reduces the detectability range of the target. The current aircraft wing structure incorporates different systems to perform various functions, so a composite wing capable of performing multiple functions of providing stealth while retaining its existing properties is required. The need for sustainable green materials has shifted the focus to natural materials, and hybridization is the way to improve the properties. The current study proposes a multifunctional hybrid carbon-basalt-polyaniline composite for aircraft wing skin that can absorb electromagnetic (EM) waves while maintaining its mechanical properties. A polyaniline with 5 % weight was found to be optimal, as the prepared sample has a dielectric loss tangent of 2.1396. The hybrid composite was simulated against a bird strike and a maximum deflection of −48.54 mm was observed for a thickness of 1.2 mm. The deflection behavior of the prepared composite was compared to several works, and a thickness of 3.6 mm was found to be optimal for a deflection of −28.97 mm. For a curved leading-edge profile, the observed deflection for 1.488 mm thickness was 580 mm. An all-composite wing model was designed and proposed having two ribs, one spar, and two stringers all modelled in carbon-basalt-PANI composite showing the maximum center point deflection of 450 mm. The prepared multifunctional hybrid composite has a high potential for use in aircraft wing skins due to its reasonable mechanical properties, good EM wave absorption, and ability to withstand bird strike.

Free Vibration Analysis of Tapered Composite Aircraft Wing via the Finite Element Method

The responses of the structures used in engineering applications under the effects of static and dynamic forces are significant in the design phase. Determination of the response of dynamic forces for a structure is initially performed by the evaluation of free vibration characteristics that are mode shape of the structure and vibration frequencies. This paper presents modal analyses of tapered aircraft wing structures that consist of NACA4415 design and different common materials used in the aviation industry. Furthermore, the effect of winglets on natural frequencies is examined. The main wing structures as ribs and shells are drawn using CATIA and imported to ANSYS Workbench. Analyses have been carried out considering the aircraft wing as a three-dimensional cantilever beam by fixing one end (root chord) of the aircraft wing while the other end (tip chord) is free. The first ten modes of free vibration with their respective natural frequencies and mode shapes of the wing structures of the aircrafts are obtained. The results show that the winglets decrease the natural frequency noticeably and the shell material as Carbon Epoxy UD has been observed to have higher natural frequency compared with Kevlar Epoxy.

Two-Step Multi-Objective Reliability-Based Design Optimization of Aircraft Wing Structures

The multi-objective reliability-based design optimization (MORBDO) of an aircraft structure employing a non-probabilistic model, at present, still has a high level of analysis complexity while solving the possibility safety index (PSI) as they are a triple-loop nested problem. Many techniques have been proposed to expedite the process of solving their inner loop with a single objective function; however, research on applying multi-objective optimization to complete this task is required. This research paper aims to reduce the solution complexity in the MORBDO of an aircraft wing structure, which is a symmetrical part of the aircraft structure. The present framework is comprised of a two-step technique that begins with the multi-objective optimization (MODO) of the wing structure, followed by its reliability analysis. A non-probabilistic model is adopted for uncertainty consideration, contrary to frequently used probabilistic models. The reliability design problem has aircraft wing mass, flutter speed, and the possibility safety index as objective functions. According to the results, the proposed MORBDO technique is highly effective in reducing the complexity of aircraft wing structural design and can generate more conservative and feasible design solutions with various PSI values. Such a design can be achieved within a single run, which has not been done in previous studies. The results show that the highest reliability aircraft wing structure mass is 104.8504 kg at a flutter speed of 584.5670 m/s. Additionally, the developed framework explicitly states the relationship between MODO and MORBDO.

Evaluation on Piezoaeroelastic Energy Harvesting Potential of A Jet Transport Aircraft Wing with Multiphase Composite by means of Iterative Finite Element Method

This paper presents new insight on the potential of piezoaeroelastic energy harvesting on the transport aircraft wing structure. A novel numerical investigation is conducted in the present study. An advanced iterative finite element method (FEM) is applied to estimate the amount of harvested energy. Currently, FEM-based commercial software has a limited application on piezoelectric structures, i.e., actuator and sensor. The iterative FEM algorithm extends the commercial software implementation for the energy harvesting analysis. The multidisciplinary issue of the aeroelastic phenomenon and piezoelectric energy harvesting is evaluated in the present case. Likewise, interestingly, stress and failure analysis of a lifting surface with an active energy harvesting component could be enabled. Implementation of a wing with an embedded piezoelectric layer is concerned. A cruise flight condition and gust disturbance are evaluated. The analysis concerning the occurrence of gust also provides a more realistic insight into the power harvested during a flight operation.

Natural convection cooling of aircraft wingbox structures during turnaround period

Modern commercial aircraft can experience significant solar loading during turnaround when the aircraft is stationary on the tarmac. With the increased usage of composite materials, a change in the aircraft thermal environment results due to the substantial difference in metal and composite thermal properties. In this paper, an experimental study is undertaken to compare the thermal environment established in a CFRP and aluminium wingbox compartment due to solar loading for levels of 81 W/m2, 396 W/m2, 700 W/m2 and two further cases with a constant wing skin temperature of 100 °C. It was found that conduction through the vertical front and rear spars established a complex but stable flow environment comprising of four counter-rotating circulations which interact strongly with the spar walls, a criterion using scale analysis was also established to determine whether the induced flow would persist. Nusselt number measurements were carried out for each circulation and compared to existing correlations from the literature for differentially heated and top wall heated cavities. The measurements provide aircraft thermal designers with representative values for the heat transfer coefficient which can be used in the thermal modeling of aircraft wing structures.

Optimum Design of UAV Wing Skin Structure with a High Aspect Ratio Using Variable Laminate Stiffness

This paper quickly calculated the design variables to satisfy the strength and stability conditions, the dominant design verification conditions of the composite wing structure, through a genetic algorithm. It developed a variable stiffness stacking optimization process applicable to aircraft wing structure design. We proposed a laminate parameter-based optimization strategy that considers the stiffness characteristics of the two-dimensional elements used as design variables. Compared with the optimization results obtained using continuous stiffness optimization as an optimization process function of laminate sequences with genetic algorithms, we obtained a reasonable stiffness distribution while complying with critical guidelines related to individual composite layup designs. The results of the stiffness optimization design were implemented as a finite element model, and the results were verified through NASTRAN. We demonstrated the functionality of the stiffness optimization process by obtaining results satisfying the set response conditions, i.e., strength and stability, in many design areas of the wing.

Performance analysis of composite laminates wing skin with the aid of fluid structure interaction of aerodynamic loading-structural analysis

Advances in today’s technology have led to the use of composite materials with their advantages of significant high strength with lightweight structure in aircraft design worldwide. With the aid of simulation, ANSYS, two main common problems in designing the aircraft composite wing, which is the aerodynamic loading on the wing and composite modeling; whereby both can be coupled to solve the aforementioned problems. The paper addresses with the use of fluid Structure Interaction (FSI) simulation and developed a standard numerical approach by integrating the aerodynamic loading from fluid domain with Finite Element (FE) analysis from structural domain in the application of hybrid composite laminates of aircraft wing structure. In this study, the Computational Fluid Dynamics (CFD) is conducted on the selected wing model of Selig 1223 in order to obtain the pressure loading from fluid domain before transferring it as an input to FE composite of structural domain. The FSI simulation was performed for two parametric conditions, namely Composite Wing Skin without Honeycomb (CWS) and Composite Wing skin with Honeycomb (CWSwH). From the result, CWSwH has lower total deformation with 3.06% percentage difference compared to CWS. However, the shear stress and strain values of CWSwH showed higher value than those of CWS due to the effect of the elasticity of honeycomb core. Therefore, the numerical approach of FSI with the application composite material has proven to be feasible and significant of current research on the behavior of fluid flow and structural for the wing skin, which replicates the real conditions of the aircraft wing during take-off or cruise.

An analytical sensitivity theory-based optimization method for composite structures with large-scale variables

This research focusses on large-scale variable optimization techniques for commercial composite wing structures. To achieve a more complete set of results, the layer thickness for composite layers with the same material and ply direction, are set as a design variable in this study. An analytical approach was utilized to deal with the sensitivity analysis so as to study the relationship between displacement, stress and strain constraints. In order to improve the applicability and user-friendliness, the procedure of sensitivity analysis has been programmed within Fortran and integrated into a Computer-Aided Engineering (CAE) software named HAJIF. To compare the consistency and efficiency, the analytical sensitivity calculation method, along with the finite difference method and a semi-analytical method, were employed to optimise a composite aircraft wing structure. The analytical method was found to have the best consistency and efficiency amongst the three optimization methods. With the aid of HAJIF, the wing was optimized with a weight loss ratio of 8.9%. The total optimization time was is 56 times quicker than the finite difference method and semi-analytical method.

The Need for Multi-Sensor Data Fusion in Structural Health Monitoring of Composite Aircraft Structures

With the increased use of composites in aircraft, many new successful contributions to the advancement of the structural health monitoring (SHM) field for composite aerospace structures have been achieved. Yet its application is still not often seen in operational conditions in the aircraft industry, mostly due to a gap between research focus and application, which constraints the shift towards improved aircraft maintenance strategies such as condition-based maintenance (CBM). In this work, we identify and highlight two key facets involved in the maturing of the SHM field for composite aircraft structures: (1) the aircraft maintenance engineer who requires a holistic damage assessment for the aircraft’s structural health management, and (2) the upscaling of the SHM application to realistic composite aircraft structures under in-service conditions. Multi-sensor data fusion concepts can aid in addressing these aspects and we formulate its benefits, opportunities, and challenges. Additionally, for demonstration purposes, we show a conceptual design study for a fusion-based SHM system for multi-level damage monitoring of a representative composite aircraft wing structure. In this manner, we present how multi-sensor data fusion concepts can be of benefit to the community in advancing the field of SHM for composite aircraft structures towards an operational CBM application in the aircraft industry.

Retracted] Application and Optimization of Wing Structure Design of DF‐2 Light Sports Aircraft Based on Composite Material Characteristics

Compared with ordinary metal structures, advanced composite materials have the characteristics of high strength, high rigidity, and light weight. The use of composite materials in aircraft structures is currently a hot research topic. This research mainly discusses the optimization design of the composite wing structure of the DF‐2 light sports aircraft. This article takes the DF‐2 light sports aircraft planned to be produced by the company as the source. Based on its overall design basis, aerodynamic requirements, and the original wing structure design, according to the composite material aircraft structure design theory and method, the aircraft wing structure is carried out. Composite materials are materials with new properties that are composed of two or more materials with different properties at the macroscale by physical and chemical methods. Composite materials can be divided into functional composite materials and structural composite materials according to the nature of the application. Functional composites are materials with special functions, such as conductive composites, ablative materials, and frictional composites. At present, the main research is on structural composite materials, which are composed of two components: matrix material and reinforcing material. The new structural scheme design and structural strength analysis are designed to meet the structural strength requirements of the wing and the lightest weight. In this paper, according to the force transmission characteristics of different structural types of the wing, the characteristics of the load transmission are analyzed, and the shape parameters and load parameters of the wing structure design are used as initial conditions, and the quantitative analysis model of the wing structure is constructed according to the requirements of strength, stiffness, and stability. Through rapid mathematical modeling and analysis of the wing structure, the weight and efficiency of different configurations can be evaluated. Through the quantitative analysis model of the wing, the wing structure type can be quickly determined according to the wing parameters in the preliminary design, which makes the basis for the selection of the wing structure type. After optimization, the weight of the wing structure decreased from 0.966 kg to 0.803 kg, a decrease of 16.87%. The designability of composite materials is one of its major characteristics. By optimizing the layup angle, layup sequence, and dropout area, the performance indicators of the structure are finally improved. This research will promote the further development of the aerospace field.