Aircraft Structural Components Research Articles

Fatigue and corrosion behavior of in-service AA7075 aircraft component after thermo-mechanical and retrogression and re-aging treatments

Fatigue crack growth and corrosion are the two important failure mechanisms for aircraft structural components and, therefore, various treatments have been developed to improve the fatigue and corrosion resistance of aircraft materials. In the present study, thermo-mechanical and retrogression and re-aging treatments were applied to AA7075T7352 specimens, which were extracted from a nearly 40 years in-service F-4 Phantom component. The in-service component was selected in order to observe the influence of thermo-mechanical and retrogression and re-aging treatments on the properties of a used aircraft material and it was expected that the service life of the material is extended in the maintenance stage. In the experimental work, electrical, mechanical, fatigue crack growth, and corrosion tests were carried out using the specimens with T7352 (as-received), thermo-mechanical and retrogression and re-aging conditions. Based on the results, fatigue crack growth resistance of the material benefited from the thermo-mechanical and retrogression and re-aging treatments; however, both treatments lowered the corrosion resistance of the material.

INVESTIGATION OF MECHANICAL AND CORROSION PROPERTIES OF AA2024–B4C–TiC HYBRID METAL MATRIX COMPOSITES

Metal matrix composites are widely applied in the automotive and aircraft structural components. This work investigates the influence of mechanical and corrosion properties hybrid aluminium 2024 metal matrix composites having 5%, 10%, 15% and 20% of TiC and B4C (equal amount) reinforcement compositions. The increase in volume fraction of B4C and TiC particle in matrix alloy increases the hardness, yield strength, ultimate strength and decreases the ductility of the composite. SEM images show microstructure of the composites with particles having a different volume percentage of reinforcement. The uniform distribution of B4C and TiC particles is most predominant in composite fabrication and the reinforcement particle additions have proved improved corrosion resistance in AA2024–B4C–TiC MMCs.

Effects of Vegetable Oil Based Cutting Fluid in Machining Kevlar Composite Material

Composite laminates are the first priority for many applications such as aerospace and aircraft structural components due to their superior properties. Composites offer a number of distinct advantages over more conventional engineering materials such as aluminium, steel etc. These benefits include higher specific strengths and stiffness’s, superior corrosion resistance as well as improved fatigue properties. However, the benefits of its application are limited by the difficulty of machining due to its poor thermal conductivity. Excess cutting temperature affects the dimensional accuracy of the drilled hole and deteriorates its surface finish. As a result, acquiring a good surface finish and dimensional accuracy greatly dependent upon the control of the cutting parameters and appropriate use of cutting environments. The objective of the article is to evaluate the performance of vegetable oil-based cutting fluid on machining Kevlar composite. Furthermore, machining responses are compared with dry and wet cutting conditions to have appropriate machining environment. Drilled holes were evaluated based on surface roughness and dimensional accuracy. Apart from machining parameters, tool wear and chips formations also have been observed on given different machining conditions. Experimental investigation revealed that Vegetable based cutting fluid is better for surface roughness where machining accuracy deteriorate. Furthermore, tool wear and chips formation also improved significantly on vegetable based cutting fluid.

The Effects of Sn Addition on the Microstructure and Surface Properties of Laser Deposited Al-Si-Sn Coatings on ASTM A29 Steel

Aluminium and its alloys have been successful metal materials used for many applications like commodity roles, automotive and vital structural components in aircrafts. A substantial portion of Al-Fe-Si alloy is also used for manufacturing the packaging foils and sheets for common heat exchanger applications. The present research was aimed at studying the morphology and surface analyses of laser deposited Al-Sn-Si coatings on ASTM A29 steel. These Fe-intermetallic compounds influence the material properties during rapid cooling by laser alloying technique and play a crucial role for the material quality. Thus, it is of considerable technological interest to control the morphology and distribution of these phases in order to eliminate the negative effects on microstructure. A 3 kW continuous wave ytterbium laser system (YLS) attached to a KUKA robot which controls the movement of the alloying process was utilized for the fabrication of the coatings at optimum laser parameters. The fabricated coatings were investigated for its hardness and wear resistance performance. The field emission scanning electron microscope equipped with energy dispersive spectroscopy (SEM/EDS) was used to study the morphology of the fabricated coatings and X-ray diffractometer (XRD) for the identification of the phases present in the coatings. The coatings were free of cracks and pores with homogeneous and refined microstructures. The enhanced hardness and wear resistance performance were attributed to metastable intermetallic compounds formed.

Effects of the building direction on fatigue crack growth behavior of Ti-6Al-4V manufactured by selective laser melting

Additive manufacturing selective laser melting (AM-SLM) are potential to be widely used in aircraft structural components. There are apparent layers along the building direction of titanium alloy samples manufactured by SLM. This will affect its fatigue properties. At present, few papers about effects of the building direction on fatigue crack growing behavior have been published. In this work, three kinds of titanium alloy Compact Tension (C(T)) specimens with different building directions were designed. Fatigue crack growth rate (FCGR) tests were performed. And fatigue crack growth rate versus stress intensity factor range (da/dN −ΔK) curves were obtained and then effects of the building direction on crack growth was analyzed. It is shown that FCGR of specimen with 90° building direction is the fastest, and then that of the 0° sample is slower, and 45° sample is the slowest. Fatigue life of 45° specimen is 1.49 times longer and 1.46 times longer than that of 90° and 0° specimen separately. Fracture morphology of samples was studied. It shows that the fracture surface with 0° building direction is smooth and fatigue striations are observed. The fracture surface with 45° building direction has a slope along the thickness. The fracture surface of the sample with 90°is cluttered, and its building layers can be observed. But its fatigue striations are not clear.

Residual life estimation of cracked aircraft structural components

Residual life estimation of cracked aircraft structural components

Development of the numerical model for evaluating the temperature field and thermal stresses in structural elements of aircrafts

An approximate method for estimating the thermal stresses of the aircraft key components of the simple geometric shape (the edges of the hull and wings, the nose fairing) has been developed. The mathematical model of such estimates is based on the solution of the quasi-static thermoelasticity problem. The solution is evaluated in the area with curvilinear boundaries, and the shape of these boundaries changes under the influence of thermal and mechanical loads. Thus the computational domain is transformed to an area where the regular Cartesian (structured) grid can be introduced. The initial validation and verification of the developed numerical methodology was carried out. Numerical modeling of temperature fields and thermal stresses in the simplest components of aircraft structures (cylinder blunted over the sphere and the shell) is performed.

Fatigue crack growth of multiple interacting cracks: Analytical models and experimental validation

AbstractThis article deals with the fatigue propagation of multiple cracks in finite width holed panels, which are typical of aircraft structural components. Theoretical studies in the literature have been considered and critically analyzed. Some of them have been translated into analytical models and implemented in a computer code. To check the effectiveness of the used models, a fatigue testing campaign has been conducted on six different configurations of notches and cracks. The comparison between experimental results and those obtained from the implemented models has shown a good agreement.

Ultrasonic Fatigue of CFRP - Experimental Principle, Damage Analysis and Very High Cycle Fatigue Properties

Structural aircraft components are often subjected to more than 108 loading cycles during their service life. Therefore the increasing use of carbon fiber reinforced polymers (CFRP) as primary lightweight structural materials leads to the demand of a precise knowledge of the fatigue behavior and the corresponding failure mechanisms in the very high cycle fatigue (VHCF) range. To realise fatigue investigations for more than 108 loading cycles in an economic reasonable time a novel ultrasonic fatigue testing facility (UTF) for cyclic three-point bending was developed and patented. To avoid critical internal heating due to viscoelastic damping and internal friction, the fatigue testing at 20 kHz is performed in resonance as well as in pulse-pause control resulting in an effective testing frequency of ~1 kHz and the capability of performing 109 loading cycles in less than twelve days. The fatigue behavior of carbon fiber twill 2/2 fabric reinforced polyphenylene sulfide (CF-PPS) and carbon fiber 4-H satin fabric reinforced epoxy resin (CF-EP) was investigated. To study the induced fatigue damage of CF-PPS and CF-EP in the VHCF regime in detail, the fatigue mechanisms and damage development were characterized by light optical and SEM investigations during interruptions of constant amplitude tests (CAT). Lifetime-oriented investigations showed a significant decrease of the bearable stress amplitudes of CF-PPS and CFEP in the range between 106 to 109 loading cycles. The ultrasonically fatigued thermoset matrix composite showed a significantly different VHCF behavior in comparison to the investigated thermoplastic matrix composite: No fiber-matrix debonding or transversal cracks were present on the specimen edges, but a sudden specimen failure along with carbon fiber breakage have been observed. The fatigue shear strength at 109 cycles for CF-PPS could be determined to τa, 13 = 4.2 MPa and to τa, 13 = 15.8 MPa for the thermoset material CF-EP.

Prognostic analysis of fastener joints in straight attachment lugs

PurposeMany failures of aircraft structural components in the past were attributed to cracks emanating from joints, which are identified as the most critical locations. In cases using the recently emerging structural health monitoring (SHM) systems, continuous monitoring needs be carried out at many major joint locations. The purpose of this paper is to develop computational techniques for fastener joints, including the possible change in contact conditions and change in boundary values at the pin-hole interface. These techniques are used for the prognostic analysis of pin-loaded lug joints with rigid/elastic pin subjected to fatigue loading by estimating the residual life of the component at any given instance to assist the SHM systems.Design/methodology/approachStraight attachment lug joints with rigid/elastic push-fit pin and smooth pin-hole interface are modelled in commercial software MSC PATRAN. In each case, the joint is subjected to various types of fatigue load cycles, and for each type of cycles, the critical locations and the stress concentrations are identified from the stress analysis. Later, for each type of fatigue cycle, the number of cycles required for crack initiation is estimated. A small crack is located at these points, and the number of cycles required to reach the critical length when unstable crack growth occurs is also computed. The novelty in the analysis of life estimations is that it takes into account possible changes in contact conditions at the pin-hole interface during load reversals in fatigue loading.FindingsThe current work on fastener joints brings out the way the load reversals leading to change in contact conditions (consequently changing boundary conditions) are handled during fatigue loading on a push-fit joint. The novel findings are the effect of the size of the hole/lug width, elasticity of the material and the type of load cycles on the fatigue crack initiation and crack growth life. Given other parameters constant, bigger size hole and stiffer pin lead to lesser life. Under load controlled fatigue cycles, pull load contributes to significant part of fatigue life.Originality/valueThe analysis considers the changing contact conditions at the pin-hole interface during fatigue cycles with positive and negative stress ratios. The results presented in this paper are of value to the life prediction of structural joints for various load cycles (for both pull-pull cases, in which the load ratios are positive, and pull-push cycles, where the load ratios are negative). The prognostic data can be used to effectively monitor the critical locations with joints for SHM applications.

Reliability analysis of the PZL-130 Orlik TC-II aircraft structural component under real operating conditions

Reliability analysis of the PZL-130 Orlik TC-II aircraft structural component under real operating conditions

Simulation of plastic deformation in Ti-5553 alloy using a self-consistent viscoplastic model

Titanium alloy Ti-5Al-5Mo-5V-3Cr (Ti-5553) is a near beta alloy used in structural aircraft components because of its excellent mechanical properties. Simulating the mechanical response of this material using constitutive models is an important step in understanding the relationship between its microstructure and properties. Uniaxial compression tests were conducted at temperatures both below and above the β transus and at different loading rates. A viscoplastic self-consistent (VPSC) model was used to match the stress-strain response of the Ti-5553 alloy based on uniaxial compression tests across a range of temperatures and strain rates. Sets of parameter values were determined for two different hardening models, namely the modified Voce model, which is empirical, and the Mechanical Threshold Stress (MTS) model, which is based on dislocation theory. No consistent trends in the Voce parameter values were found as a function of temperature or strain rate. By contrast, the physically-based MTS model, which is explicitly designed to cover wide ranges of deformation conditions, was able to fit a wide range of temperatures and strain rates. It was further validated by comparison with experimental measurements with other β titanium alloys that had similar compositions.

Aircraft duraluminum wing spar fatigue testing

Experimental fatigue crack growth analysis of the aircraft structural component subjected to different cycling loading has been presented. The purpose of the experimental test was to evaluates the fatigue behavior of the real aircraft spar, made in aluminum 2024-T3. In-house laboratory fatigue test installation is described in the paper. The spar-test specimen has been manufactured specifically for the cycling damage testing. Test results were obtained with most modern equipment for data acquisition. Applied loads and obtained spar fatigue results are described in detail. Conclusions have been made with reference to the previous work on the duraluminum spar fatigue testing. As it is presented in the previous work with advanced numerical methods for fatigue analysis, in this paper is given an opinion on the same. At the end, possible directions of the further experimental spar fatigue testing is given and shows us that it makes sense to make the comparison between real and numerical results only in situations where cracks behave similarly and the numbers of required cycles are about the same.

Computational Modeling of Multi-Purpose Amphibious Aircraft Be-103

The paper is dedicated to the aspects of 3D computational modeling of Be-103 amphibious aircraft. The paper contains initial Figures and drawings; the computational modeling is performed in 3 ds Max 3D graphic modeling system. Models of amphibious aircraft structural components are produced by polygonal extrusion process. Shading is performed at the sub-objects level as well as assignment of materials. Figures illustrating realistic rendering of amphibious aircraft 3D-model are presented as well.

Machining of large scaled CFRP-Parts with mobile CNC-based robotic system in aerospace industry

Industrial robots have already demonstrated their advantages in smart and efficient production in a wide field of applications and industries. However, their use for machining of structural aircraft components is still impeded by the disadvantage of low absolute accuracy, sensitivity to process loads and limited workspace compared to large machining centers.A mobile robotic system is presented as a new approach for machining applications of large aircraft components. The system presented in this paper consists of a CNC-based serial robot kinematic with additional secondary encoder systems on every axis. The entire system is based on a Siemens CNC control, which evolves the robot to a full-featured machine tool. This setup enables additional possibilities for the implementation of extended control strategies and advanced calibration routines. Thus, high demands and challenging tolerances in aircraft manufacturing will be fulfilled, so that process times and investment cost can be reduced significantly.

The algorithm of layering calculation for corner plunge milling tool path

Cavity corner is one of the most common machining features in the complex structural components of aircraft, and processing pocket corner by plunge milling has the advantages of higher metal removal rate, smaller machining distortion, and higher machining precision, and especially is suitable for cavity corner machining with large depth. However, there are some problems in current corner plunge milling approach, such as sidewall over-cut, invalid tool location points, and uncontrollable residual height. Therefore, a universal layering algorithm for plunge milling tool path generation is projected and established. First, define and represent two-dimensional corner model. Next, the terminologies about machining layer circle, cutter circle, and cuspidal point circle are introduced and layering model is built. Then, construct cutter circle arrangement program on the basis of the intersection between layer circle and corner model. Meanwhile, the relationship between residual height and layering model is determined. Finally, the layering tool path is recursively generated in line with radial cutting width. Two groups of contrast experiments prove effectively that the proposed method excels the conventional methods in guaranteeing residual height and avoiding invalid tool location points. In addition, the layering method can maintain an efficient processing tool path ratio of more than 80%. Furthermore, the application of a CAD solid model of an aircraft panel demonstrates the feasibility and effectiveness of the layering approach.

Effects of Cutting Parameters on Quality of Surface Produced by Machining of Titanium Alloy and Their Optimization

Abstract Titanium alloy (Ti-6Al-4V) has been extensively used in aircraft turbine-engine components, aircraft structural components, aerospace fasteners, high performance automotive parts, marine applications, medical devices and sports equipment. However, wide-spread use of this alloy has limits because of difficulty to machine it. One of the major difficulties found during machining is development of poor quality of surface in the form of higher surface roughness. The present investigation has been concentrated on studying the effects of cutting parameters of cutting speed, feed rate and depth of cut on surface roughness of the product during turning of titanium alloy. Box-Behnken experimental design was used to collect data for surface roughness. ANOVA was used to determine the significance of the cutting parameters. The model equation is also formulated to predict surface roughness. Optimal values of cutting parameters were determined through response surface methodology. A 100% desirability level in the turning process for economy was indicated by the optimized model. Also, the predicted values that were obtained through regression equation were found to be in close agreement to the experimental values.

항공기 낙뢰 전산 시뮬레이션 및 보호시스템 설계

항공기는 결빙, 난기류, 낙뢰 등 환경적 요인에 의해 안전성을 위협받는다. 그 중 낙뢰는 항공기에 강한 전류와 전압을 유입시켜 고온의 열과 자기장을 발생시키므로 항공기 안정성에 큰 위협이 된다. 본 연구에서는 낙뢰의 발생원리 및 항공기 낙뢰 피격 메카니즘을 설명하고 항공기에 대한 낙뢰의 직접, 간접 영향성을 다룬다. 먼저 항공기 낙뢰 지침서인 ARP를 분석하여 항공기 낙뢰모사와 부착위치에 대한 분석을 진행하였으며, 이를 전산 시뮬레이션에 적용하였다. 또한 ABAQUS 소프트웨어를 활용한 열적, 전기적 시뮬레이션을 통해 연료탱크의 낙뢰 영향성 분석 및 낙뢰보호시스템 설계를 진행하였다. 그리고 Maxwell 방정식에 기초한 EMA3D 소프트웨어를 활용하여 항공기 전기체의 전류 흐름에 대한 전산수치해석을 진행하였다. The safety of aircraft can be threatened by environmental factors, such as icing, turbulence, and lightning strike. Due to its adverse effects on aircraft structure and electronic components of aircraft, lightning strike is one of the biggest hazards on aircraft safety. Lightning strike can inject high voltage electric current to the aircraft, which may generate strong magnetic field and extreme hot spots, leading to severe damage of structure or other equipment in aircraft. In this work, mechanism of lightning strike and associated direct and indirect effects of lightning on aircraft were studied. First, on the basis of aircraft lightning regulations provided by Aerospace Recommended Practice (ARP), we considered different lightning waveform and zones of an aircraft. A coupled thermal-electrical computational model of ABAQUS was then used for simulating flow of heat and electric current caused by a lightning strike. A study on fuel tank, with and without lightning protection system, was also conducted using the computational model. Finally, electric current flow on two full scale airframes was analyzed using the EMA3D code.

Dynamic creation and evolution of gradient nanostructure in single-crystal metallic microcubes.

We demonstrate the dynamic creation and subsequent static evolution of extreme gradient nanograined structures in initially near-defect-free single-crystal silver microcubes. Extreme nanostructural transformations are imposed by high strain rates, strain gradients, and recrystallization in high-velocity impacts of the microcubes against an impenetrable substrate. We synthesized the silver microcubes in a bottom-up seed-growth process and use an advanced laser-induced projectile impact testing apparatus to selectively launch them at supersonic velocities (~400 meters per second). Our study provides new insights into the fundamental deformation mechanisms and the effects of crystal and sample-shape symmetries resulting from high-velocity impacts. The nanostructural transformations produced in our experiments show promising pathways to developing gradient nanograined metals for engineering applications requiring both high strength and high toughness-for example, in structural components of aircraft and spacecraft.

Forming limit in thermal cruciform biaxial tensile testing of titanium alloy

In this study, a new cruciform biaxial tensile testing technology was employed to evaluate the thermal limit strains of a TA1 titanium alloy, which is widely used as an aircraft structural component, so as to prevent fracture defects during thermoplastic processing. The effect of specimen shape on the forming limit and its measurement range is discussed. The results of this study show that a design that reduces the semispherical thickness results in localized necking and fracture at the center of the specimen when normal stress is applied to it. The forming limit diagrams obtained at different angles with respect to the material rolling direction suggested that the limit strain was anisotropic in nature. The experimental results were in good agreement with the substituted values obtained from the finite element analysis and Oyane ductile fracture criterion. Finally, the efficacy of the related test and prediction methods was corroborated.