Aircraft Structural Components Research Articles

Exploring Microstructural, Interfacial, Mechanical, and Wear Properties of AlSi7Mg0.3 Composites with TiMOVWCr High-Entropy Alloy Powder.

This study explored the impact of varying weight percentages of TiMoVWCr high-entropy alloy (HEA) powder addition on A356 composites produced using friction stir processing (FSP). Unlike previous research that often focused on singular aspects, such as mechanical properties, or microstructural analysis, this investigation systematically examined the multifaceted performance of A356 composites by comprehensively assessing the microstructure, interfacial bonding strength, mechanical properties, and wear behavior. The study identified a uniform distribution of TiMoVWCr HEA powder in the composition A356/2%Ti2%Mo2%V2%W2%Cr, highlighting the effectiveness of the FSP technique in achieving homogeneous dispersion. Strong bonding between the reinforcement and matrix material was observed in the same composition, indicating favorable interfacial characteristics. Mechanical properties, including tensile strength and hardness, were evaluated for various compositions, demonstrating significant improvements across the board. The addition of 2%Ti2%Mo2%V2%W2%Cr powder enhanced the tensile strength by 36.39%, while hardness improved by 62.71%. Similarly, wear resistance showed notable enhancements ranging from 35.56 to 48.89% for different compositions. Microstructural analysis revealed approximately 1640.59 grains per square inch for the A356/2%Ti2%Mo2%V2%W2%Cr processed composite at 500 magnifications. In reinforcing Al composites with Ti, Mo, V, W, and Cr high-entropy alloy (HEA) particles, each element imparted distinct benefits. Titanium (Ti) enhanced strength and wear resistance, molybdenum (Mo) contributed to improved hardness, vanadium (V) promoted hardenability, tungsten (W) enhanced wear resistance, and chromium (Cr) provided wear resistance and hardness. Anticipating the potential applications of the developed composite, the study suggests its suitability for the aerospace sector, particularly in casting lightweight yet high-strength parts such as aircraft components, engine components, and structural components, underlining the significance of the investigated TiMoVWCr HEA powder-modified A356 composites.

Feed optimization based on force modelling and TLBO algorithm in milling Al 7075

Aluminium alloys find wide applications in structural industries. They are only second to steel when it comes to their applications. Their wide application is due to their numerous good properties. They have a high resistance to corrosion and high fatigue strength. Their specific density, electrical and thermal properties make them stand out amongst many structural metals. Al 7075 is an alloy of aluminium in which zinc is the dominant alloying element. This alloy is in the 7xxx series of aluminium alloys, a series considered to be the strongest with yield strengths greater or equal to 500 Mpa. It, therefore, finds a wide application in aircraft structural components manufacturing and other areas where high strength is a critical property. Therefore, this paper focuses on the CNC milling process of aluminium 7075 to establish force components models and the optimal values for the critical process parameters. The work reported in this paper involved designing experiments and the development of force components and resultant force models from the experimentally acquired data. The models were validated, and their prediction ability was within the accepted range. The validated force component models were used to investigate the interaction effects of the process parameters on the respective force components. The resultant force model was eventually used as the objective function in the TLBO algorithm to acquire the optimal force and control parameters. After that, the optimal resultant force was used as a constraint parameter to demonstrate feed optimization relative to the other parameters.

Simulation experiment on support strength of aircraft structural components

The design principles of strength test support methods for aircraft structural components were introduced in this paper. Aiming at the vertical tail static test of a certain type of aircraft, a supporting stiffness simulation method was invented, and a support fixture was designed. The supporting stiffness was accurately simulated by stiffness strength. The research results of this paper provide an important reference for the research of static test support methods for similar aircraft structural components.

Rheology of Ti-6Al-4V alloy under non equilibrium conditions during β forging

During the “β-forging” of Ti-6Al-4V structural aircraft components that follows pre-heating above the β-transus, the skin zone of the blank actually cools down to below the β-transus whilst being deformed, as a result of heat losses towards the atmosphere or the dies. The core areas of the bulkiest sections of the blank, however, can be deformed above the β-transus during the whole processing time. This paper focuses on the development of an innovative rheological model to support the numerical modelling of the forging process, suitable for both the equilibrium conditions above the β-transus and the non-equilibrium conditions below. For that purpose, small-scale hot compression tests were performed with the thermomechanical simulator Gleeble 3500, after pre-heating in the β-field of Ti-6Al-4V at 1100 °C and rapid cooling to the compression temperature under the β-transus, which was varied between 750 °C and 950 °C. In some cases, a lag time -or stabilization time- at constant temperature was introduced between the cooling step and the deformation step. Effects of lag time and deformation temperature on microstructural evolutions and rheological properties of Ti-6Al-4V were investigated. Results show that the lack of stabilization time leads to a significant reduction of the flow stress at low strain. Such a behavior is associated with non-equilibrium thermodynamic conditions at the beginning of deformation and the progress of phase transformation during deformation. Based on that observation, an analytical formulation for the targeted rheological model is proposed. It includes a sub-model for the kinetics of phase transformation derived from the literature.

Damage monitoring of aircraft structural components based on large-area flexible graphene strain sensors

Conventional sensors face limitations when it comes to monitoring damage on intricately curved surfaces of aircraft structures. In this study, a flexible graphene/polydimethylsiloxane (GNP/PDMS) strain sensor was fabricated using a mechanical coating method. The sensor exhibits characteristics of large-area, high strain capacity, and excellent sensitivity, while maintaining low cost and simplicity of manufacture. It is a highly scalable technology that can monitor structural deterioration and deformation. After optimization, a flexible GNP/PDMS strain sensor exhibits a maximum sensitivity of 251 and an average response time of 0.23 s at 5% strain. Furthermore, it maintains steady performance during 10,000 fatigue cycles, while effectively detecting bending and torsion. The application of this flexible GNP/PDMS strain sensor in monitoring crack initiation and propagation in specimens with hole-edge cracks during fatigue processes, as well as strain measurements on aircraft wing box sections and composite material structure component subjected to various forms of impact loading was both demonstrated. Experiments verified the GNP/PDMS sensor’s capacity to locate fractures and track their progression in different specimens. The sensor has great potential for wide-ranging applications in important fields including aerospace, civil engineering, and robotic skins.

Research On MQL Flow Rate Based on CNC Milling of Aircraft Structural Parts

Micro lubrication technology (MQL) has been widely applied globally, mainly focusing on the design of lubrication mechanisms. There is a lack of research on the flow rate under actual processing conditions. The size of cooling capacity affects manufacturing costs, environmental governance, workpiece quality, and other factors. This article conducts research on the MQL flow rate of oil gas mixed lubrication, and experimentally obtains MQL flow parameters suitable for aircraft structural component processing.

Plastic flow and interfacial bonding behaviors of embedded linear friction welding process: Numerical simulation combined with thermo-physical experiment

In this study, a new linear friction welding (LFW) process, embedded LFW process, was put forward, which was mainly applied to combination manufacturing of long or overlong load-carrying titanium alloy structural components in aircraft. The interfacial plastic flow behavior and bonding mechanism of this process were investigated by a developed coupling Eulerian-Lagrangian numerical model using software ABAQUS and a novel thermo-physical simulation method with designed embedded hot compression specimen. In addition, the formation mechanism and control method of welding defects caused by uneven plastic flow were discussed. The results reveal that the plastic flow along oscillating direction of this process is even and sufficient. In the direction perpendicular to oscillation, thermo-plastic metals mainly flow downward along welding interface under coupling of shear stress and interfacial pressure, resulting in the interfacial plastic zone shown as an inverted “V” shape. The upward plastic flow in this direction is relatively weak, and only a small amount of flash is extruded from top of joint. Moreover, the wedge block and welding components at top of joint are always in un-steady friction stage, leading to non-uniform temperature field distribution and un-welded defects. According to the results of numerical simulation, high oscillating frequency combined with low pressure and small amplitude is considered as appropriate parameter selection scheme to improve the upward interfacial plastic flow at top of joint and suppress the un-welded defects. The results of thermo-physical simulation illustrate that continuous dynamic recrystallization (CDRX) induces the bonding of interface, accompanying by intense dislocation movement and creation of many low-angle grain boundaries. In the interfacial bonding area, grain orientation is random with relatively low texture density (5.0 mud) owing to CDRX.

Nanodiamond-structured zinc composite coatings with strong bonding and high load-bearing capacity.

The aerospace and automotive industries find that relying solely on the intrinsic resistance of alloys is inadequate to safeguard aircraft and automotive structural components from harsh environmental conditions. While it is difficult to attribute accidents exclusively to coating failure due to the involvement of multiple factors, there are instances where defects in the coating initiate a wear or degradation process, leading to premature and unplanned structural failures. Metallic coatings have been introduced to protect the aircraft mainly from wear due to the extreme temperatures and moisture exposure during their service life. Bare metallic coatings have a limited lifespan and need to be replaced frequently. Herein, the strength and wear resistance of zinc (Zn) coating is enhanced using varying concentrations of diamond particles as an additive in the Zn matrix (Zn-D). The dispersion strengthening mechanism is attributed to the high hardness (70 HRC), and reduced friction-of-coefficient (0.21) and dissipation energy (4.6 × 10-4 J) of electrodeposited Zn-D7.5 (7.5 g l-1 of diamond concentration) composite coating. Moreover, enhanced wear resistance with minimum wear volume (1.12 × 10-3 mm3) and wear rate (1.25 × 10-3 mm3 N-1 m-1) of the Zn-D7.5 composite coating resulted in perfect blending of diamond with Zn. The improved hardness and wear resistance for Zn-D7.5 (optimum 7.5 g l-1 diamond concentration) is due to the steadiness between well-dispersed diamonds in Zn and enrichment in load-bearing ability due to the incorporation of diamond particles. Electronic structure calculations on the zinc-diamond composite models (two configurations adopted) have been performed using the density functional theory (DFT) approach, and the in silico studies appeared to facilitate meaningful and evocative outcomes. Zn-doped diamond (C10@Zn) without hydrogen (H) atoms (binding energy: 418 kcal mol-1, i.e. showing an endothermic reaction and thermodynamically not favourable) was detected to be more stable than the Zn-doped diamond (C10H16@Zn) consisting of hydrogen (H) atoms (binding energy: -33.3 kcal mol-1, i.e. showing an exothermic reaction and thermodynamically preferable). Thus, a composite coating of zinc and diamond can be a suitable candidate for the aerospace and automotive industries.

Fatigue fracture mechanisms and life prediction of welded S310 - S321 joints at high temperature

Austenitic stainless steel is prevalent in the manufacturing of aircraft structural components and chemical pipelines, which are frequently subjected to high temperatures. To investigate the high cycle fatigue strength of welds involving two stainless steel materials S310 and S321, a series of high cycle fatigue tests were conducted at 550 °C, 650 °C, 750 °C, and 850 °C, respectively. Scanning electron microscopy (SEM) was used to observe the fracture surfaces in the base metal zone (BM), fusion zone (FZ), and heat affected zone (HAZ). The microstructure characteristics near the fusion line of the specimens were analyzed by the electron backscatter diffraction (EBSD). Results reveal a negative relavance between the presence of inclined coherent twin boundaries (CTBs) and material's fatigue performance. The initiation of cracks from the surface is caused by machining defects, welding voids, and high temperature oxidation of certain grains near the surface. For specimens that fracture in the FZ, the fatigue life show the divergence trend with temperature increases. Meanwhile, a fatigue life prediction model based on XGboost algorithm is developed with axial stress amplitude (σa), temperature (T), average grain size (Dmean), twin boundary density in the FZ (ρtb), and fracture location (Lf) as crucial analytical parameters. This model demonstrates better accuracy in predicting fatigue life with a mean absolute percentage error (MAPE) of 39.58 % compared with Basqin model's 46.81 %.

The effect of adhesively bonded external hybrid patches on the residual strength of repaired glass/epoxy‐curved laminates

AbstractComposite materials are increasingly used in aircraft structural components with complex structures like L, C, or T shapes, which can cause delamination failure in the thickness direction under complex loading. This research focuses on the experimental investigation of the effect of different adhesive reinforcement phases with an adhesively bonded external patch to examine the residual strength and delamination resistance of repaired glass/epoxy curved samples. The damaged region in curved samples was repaired using adhesive reinforcement phases like neat epoxy, chopped glass fiber, and particulate glass fiber, along with adhesively bonded external patches like hybrid patches of glass/Kevlar (HP G/K), glass/carbon (HP G/C), and intraply patch carbon/Kevlar (IP C/K). The findings show that chopped glass fiber HP G/C patch‐repaired curved samples had a curved beam strength (CBS) and interlaminar tensile strength (ILTS) recovery of 222.87% and 149.54%, respectively, as compared to damage area removed samples. In addition, it exhibited higher peak force, residual strength, and less elongation when compared to HP G/K and IP C/K patch‐repaired curved samples. According to the findings of this research, composite angle brackets in aircraft structural components can be repaired utilizing an adhesively bonded HP G/C patch filled with chopped glass fiber at the dressed site.Highlights The effect of patch hybridization on repaired curved laminates was studied. Four‐point bending test was conducted on indented and repaired samples. Residual strength and delamination behavior were studied. Chopped glass fiber HP G/C patch repaired samples had higher CBS and ILTS. Delamination onset and propagation were assessed using Photographic images.

Predicting the Remaining Useful Life of Light Aircraft Structural Parts: An Expert System Approach

This paper introduces an expert system approach for predicting the remaining useful life (RUL) of light aircraft structural components by analyzing operational and maintenance records. The expert system consists of four modules: knowledge acquisition, knowledge base, inference, and explanation. The knowledge acquisition module retrieves data from mandatory records, such as aircraft logbooks and mass and balance sheets. The knowledge base stores specific remaining useful lives (SRULs) for different load profiles that are determined using numerical strength analysis. The inference module utilizes the Palmgren-Miner rule to estimate the accumulated fatigue damage of the structural component based on the input data and the knowledge base. Lastly, the explanation module links the accumulated damage to the maintenance program and suggests the appropriate maintenance action. The Cessna 172R main landing gear leg is utilized as a case study, demonstrating the variance of RUL depending on the operating conditions. The objective of this approach is to enhance light aircraft maintenance decision making and advance operational safety.

Role of intra-ply hybridization and nano reinforcement on crack growth of epoxy

Purpose Epoxy resins are widely used in a variety of engineering applications, including composite wind turbine blades used in the renewable energy industry, highly complex structural components for aircraft, paints, coatings, industrial tooling, biomedical systems, adhesives, electronics and automotive. Epoxies' low fracture toughness is one of the key obstacles preventing its adoption in a wider range of applications. To address epoxy's low fracture toughness, this paper aims to examine the roles of intra-ply hybridization and nano reinforcing. Design/methodology/approach This paper investigates the role of intra-ply hybridization of glass-carbon woven fibers and adding 0.8 wt.% of multiwall carbon nanotube (MWCNT) nano reinforcement to overcome the low fracture toughness of epoxy. A bending test is used to calculate the composites elastic parameters, and a notched sample three-point bending test is used to show crack behavior in addition to using materials characterization methods to reveal the effect of the MWCNT on structure, bonding, glass transition temperature (Tg) and dispersion of MWCNT in the matrix. Furthermore, this paper suggests using the finite element method to overcome the difficulty in calculating the crack extension. Findings Intra-ply hybridization and MWCNT reinforcement decrease the crack extension of epoxy with time. The inclusion of high-strength carbon fiber increased the fracture toughness of glass composite. Furthermore, the existence of MWCNT in the surrounding area of the notch in epoxy composites hinders crack propagation and provides stiffness at the interface by bridging the crack and eventually enhancing its fracture toughness. Originality/value Studying the role of intra-ply hybridization of glass-carbon woven fibers and adding 0.8 wt.% of MWCNT nano reinforcement to overcome the low fracture toughness of epoxy. Additionally, this research recommends using the finite element method to overcome the challenge of computing the crack extension.

Failure analysis of a landing gear nose wheel fork produced in Ti6Al4V(ELI) through selective laser melting

This paper presents the failure analysis of a topologically optimised landing gear nose wheel fork of a light aircraft that was produced in Ti6Al4V (ELI) through selective laser melting (SLM). The study was motivated by the intention of the South African producer of the Advanced High-Performance Reconnaissance Light Aircraft (AHRLAC) to consider replacing conventional manufacturing of some structural components with additive manufacturing (AM). However, the defects typically associated with AM, such as inherent surface roughness and porosity, still complicate the certification of AM parts due to their detrimental effects on fatigue performance. An experimentally validated numerical simulation, fractographic analysis and microscopic investigation were carried out on a scaled-down Ti6Al4V(ELI) component to devise preventative actions to avoid failure of the full-scale fork. The mass saving, strength and fatigue limitations of the actual nose wheel fork of the AHRLAC are discussed. This study also outlines an effective failure analysis and data collection method necessary to qualify SLM of Ti6Al4V(ELI) to produce other airworthy mission-critical structural aircraft components.

Nanocomposites for Aircraft Applications

The steady increase in jet fuel prices in recent years has accelerated the quest for greater economy in airline operations. While improvements in aerodynamic efficiency, engine fuel efficiency, etc., have been sought, lighter and stronger materials for aircraft structural components remain a very important goal. The percentage of advanced composites, utilizing carbon fibers, in modern aircraft has reached 50 percent. The trend now is to use nano reinforcements such as graphene and carbon nanotubes to further reduce the structural weight. Due to the difficulties in fabricating nanocomposites with a high enough volume fraction of nano reinforcement, the nano reinforcements are either added to the polymer matrix or grown on the carbon fibers. The nano reinforcements are multi-functional; thus, in addition to contributing strength, they contribute to other properties such as electrical conductivity which in turn enhances the lightning strike protection, electrostatic discharge, electromagnetic shielding, etc. This paper summarizes these aspects and also describes the applications of nanocomposites in aircraft interiors, brake components, aircraft repair, etc.

Study on Non-linear Contact Stress and Prognosis Analysis of Lug Joints With Misfits

Components of aircraft and other large scale structures are generally made in parts in order to facilitate easy manufacturing, assembly and dis-assembly. Lug joints with fasteners are preferred in several aerospace applications where their purpose is to connect primary load carrying members. Study on structural integrity of these types of joints becomes mandatory from the safety point of view. This requires in-depth understanding of science and technology of Structural Health Monitoring (SHM) at these joints. Analysis presented here deals with the critical locations of interference and clearance fit lug joints which continuously operate under fatigue loading. Both interference and clearance fits exhibit non-linear contact-stress variation above certain load level and this effect needs to be addressed for both fatigue and monotonic loading. The maximum Pull or Push loads causes high stress concentrations at critical locations along the pin-hole interface. With continuous usage of the component, crack like damages initiate at these locations and they grow till failure of the component. The non-linear prognosis analysis is carried out to generate the data required for SHM at any stage which can be used to estimate remaining life of the component.

Fatigue life estimation of aircraft structural component using FE approach and validation through analytical and experimental methods

Aircraft structural design requirements encompass safe-life and damage-tolerant design, analysis, and verification through testing to arrive at the overall service life. The fatigue phenomenon is most critical and local; hence fatigue studies are necessary to conduct first at the coupon level before attempting at the component or full-scale level. The present work demonstrates the development of fatigue life estimation procedures such as high cycle fatigue (HCF) life and fatigue crack growth (FCG) life on specimens as per ASTM standards.The HCF life is predicted for the Aluminum (Al) 2024-T3 ASTM E466 specimen subjected to constant amplitude loading with a stress ratio (R) of 0.1 based on Basquin's method. The computational 2D finite element (FE) model is developed to predict HCF life using a safe-life approach through Nastran Embedded Fatigue (NEF). The effect of fatigue strength modifying factors affecting fatigue life is also explored. Similarly, the FCG life is predicted for the Al 2024-T3 ASTM E647 specimen subjected to constant amplitude loading with a stress ratio of 0.1 using Walker's equation. The computational 3D FE models with a single edge crack in each are developed to predict FCG life using FRANC3D. The FE-based HCF and FCG life prediction procedures demonstrated in this work are verified by comparing FE results with analytical and experimental ones. Therefore, these FE-based methodologies for HCF and FCG life prediction can be adopted at the feature and structural component levels. The developed computational methods reduce the experimental effort, cost, and time involved in the overall fatigue design of the aircraft structures.

Analysis of adhesively bonded joints of laser surface treated composite primary components of aircraft structures

The performance of the adhesively bonded aerospace structures highly depends on the adhesion strength between the adhesive and adherents, which is affected by, in particular, the condition of the bonding surface. Among the various surface treatment methods, as state of the art, laser surface treatment is a suitable option for the CFRP composite structures to enhance the adhesion performance, adjusting the roughness and surface free energy with relatively minimizing the damage to the fibers. The aim of this study is the validation and evaluation of the adhesive bonding behavior of the laser surface-treated CFRP composite structures, using the finite element technique to perform a conservative prediction of the failure load and damage growth. Such objectives were achieved by executing both experimental and numerical analyses of the secondary bonded CFRP parts using a structural adhesive. In this regard, to complement physical experiments by means of numerical simulation, macro-scale 3D FEA of adhesively bonded Single Lap Joint and Skin-Spar Joint specimens has been developed employing the Cohesive Zone Model (CZM) technique in order to simulate bonding behavior in composite structures especially skin-spar relation in the aircraft wing-box.

Non-destructive evaluation of GTAWA, GTAWF and LBWA welded Ti6Al4V alloy

Ti6Al4V is the most commonly used titanium alloy which features high strength with low density, high fracture toughness, excellent resistance against corrosion and good weldability. Ti6Al4V finds wide application in variety of sectors such as aircraft structure, jet engines, gas turbines, defence armour and IC engine components. The quality of the titanium weldments is significant in producing titanium structures and components. This requires critical Non-Destructive Testing (NDT) procedures to be followed. There is very scarce literature reported on NDT of titanium weldments. In this investigation the Non-Destructive Testing of Ti6Al4V weldments welded using Gas Tungsten Arc Welding – Autogenous (GTAWA), Gas Tungsten Arc Welding – with filler (GTAWF) and Laser Beam Welding – Autogenous (LBWA). The weldments were subjected NDT procedures such as visual and X-ray radiography. The uniqueness of this study is that the Image Quality Indicator (IQI) of aluminium was used as the IQI for titanium is not available in the industries. Generally, the titanium gets contaminated by gases in the atmosphere, leading to loss of mechanical strength and corrosion resistance of the weldment. Hence in this investigation a novel approach for ruling out the contamination of the weld pool and the Heat Affected Zone (HAZ) has been proposed viz. Weld Quality Indicator (WQI), i.e. the discolouration on the surface of the weldment is taken as an indicator of the contamination.

Analysis of low velocity impact response of glass fiber reinforced epoxy resin composite with shape memory alloy incorporation

The current article investigates about degradation of lightweight materials in aircraft structural components resulting from low-velocity impacts, a common cause of failure. Although subsonic collisions do not cause immediate external damage to the structure, they significantly reduce the material's stiffness and impact strength. The study explores the performance of intelligent hybrid thermoplastic composite plates subjected to low-speed impacts, both numerically and experimentally. Particularly, the behavior of laminated composites with integrated shape memory alloy (SMA) fibers is discussed in this study. SMA wires are inserted into the composite laminate fabric layer located on the intermediate planes. The study revealed that incorporating SMA fibers increases the damage tolerance and stiffness of the laminated composites compared to standard composite structures. The compounds exhibit a noticeable increase in hardness and energy levels before failure.

Investigation on fatigue failure of split-sleeve cold expansion holes of 7075-T651 aluminum alloy

7075-T6 is widely used in the manufacture of aircraft structural components, the majority of which are joined and assembled through hole structures. However, the geometric discontinuity near the fastener holes facilitates stress concentration and makes the joints susceptible to crack nucleation and fatigue failure. It is essential to improve the fatigue resistance of the hole structure. In this paper, the factors affecting the split-sleeve cold expansion on the fatigue behavior of fastener holes of 7075-T651 are investigated based on the compressive residual stress, microstructure, surface roughness, and fatigue fracture. Results revealed that cold expansion causes sub-grain refinement near the hole wall, which increases the surface micro-hardness and favors inhibition of crack growth. Grain refinement, plastic deformation on the surface of the holes, and residual compressive stress produced by the cold expansion process all work together to extend the fatigue life of the hole structure. The constituent particles and microstructure after expansion are important factors affecting fatigue performance, higher level of Fe content could result in the formation of a large number of iron-bearing particles leading to the acceleration of the failure of the workpiece.