Aluminum Alloy Skin Research Articles

Fracture Toughness and Shear Strength of the Bonded Interface Between an Aluminium Alloy Skin and a FRP Patch

The existing techniques to determine the fracture properties such as critical energy release rate in mode I (GIc) and mode II (GIIc) of an interface between two sheets of same material were modified to determine these properties between the sheets of dissimilar materials and thickness. In addition, the interface shear strength (ISS) was also determined. Experiments were carried out on the specimens made of a pre-cracked thin aluminium alloy skin and a Fiber reinforced polymer (FRP) patch. Two kinds of surface preparation of the aluminium skin were employed; (i) emery-paper roughened surface (ERS) and (ii) Sodium Hydroxide (NaOH) treated surface (NTS). GIc of ERS specimen was found to be 36.1 J/m2, while it was found to be much higher for NTS specimens, that is, 87.3 J/m2. GIIc was found to be 282.4 J/m2 for ERS specimens and much higher as 734.5 J/m2 for NTS specimens. ISS was determined as 32.6 MPa for ERS specimen and significantly higher for NTS specimen, that is, 44.5 MPa. The micrographs obtained from a field emission-scanning electron microscope (FE-SEM) and the surface roughness test showed that the NTS was significantly rougher than the ERS, explaining the higher values of all the three kinds of NTS specimens.

The deformation and residual stress simulation of dual laser-beam bilateral synchronous welding for Al-alloy aircraft panel structure

This paper reports a numerical investigation of dual laser-beam bilateral synchronous welding (DLBSW) for T-joint structure of aircraft fuselage. Finite element numerical simulation of DLBSW is carried out to obtain suitable matching of welding parameters for civil aircraft panels which composed of 6156 aluminium alloy skin and 6056 aluminium alloy stringer. The distribution of welding residual stress and welding distortion on the aircraft panels are predicted and discussed. Three-dimensional finite element model of the panel containing three stringers has been developed to simulate the temperature field, residual stress distribution and welded panel distortion. It is simulated that three stringers are welded to base plate of the specimen through different welding sequences and the welding sequence with the smallest distortion is acquired.

MODIFIED HAGG & SANKEY METHOD TO ESTIMATE THE BALLISTIC BEHAVIOUR OF LIGHTWEIGHT METAL/COMPOSITE/CERAMIC ARMOUR AND A FUSELAGE SKIN OF AN AIRCRAFT

Hagg & Sankey method assesses the functionality of containment rings, which prevent perforation of an aircraft's elements from turbine engine disc fragments after disc burst. However, the method assumption was based on studies of mechanism of destruction of ballistic shields made of among others ceramics, by small arms bullets. The hard ceramic facing of the ballistic shield blunts the projectile and breaks up the projectile's hard core used for armour piercing. As an impact result, a conoid of finely pulverized ceramic dust is formed, absorbing energy in the formation process. The dust, containing the remnants of the projectile's energy hits the backing, but is now spread over a larger area. With backings made of fibre, energy is absorbed in stretching, breaking and delamination. With backings made of ductile metal, energy is absorbed in elastic deformation. Modification of the method consisted of taking into account the stratification of the ceramic-metal composite and the occurrence of an aircraft stiffened skin, in order to better assess the effectiveness of ballistic shields. The concept of estimating the resistance was based on the described destruction mechanism, where the object of analysis is a metal containment ring with an additional ceramic protective ballistic shield. A comparison of two scenarios- with and without an additional 2 mm aluminum alloy skin was taken into account. For this particular scenario, results were sufficient for both of the two analyzed endurance criteria.

Experimental study on the fatigue behaviour of honeycomb sandwich panels with artificial defects

In order to characterise efficiently the honeycomb sandwich panels, the static behaviour does not seem to be sufficient and additional information about fatigue data properties is needed. This study presents an important experimental fatigue data for honeycomb sandwich panels with artificial defects and without defects. The results from the fatigue tests on honeycomb sandwich panels are compared to those of the aluminium alloy skin (the reference case). Experimental results showed that:-The presence of defect has no influence on the static behaviour of the material.-The lifetime of the L configuration is larger than in the W-direction.-The lifetime of sandwich panels is very sensitive to drilling hole type defect than Brinell one.

Numerical Simulation of Bird Impact on Fibre Metal Laminates

A finite element model of bird impact on the Fibre Metal Laminates (FMLs) was established with the smoothed particle hydrodynamics (SPH) and finite element method in the software PAM-CRASH. The bird was modeled with the SPH elements in order to avoid the bird mesh distortion. It studied the sensitivities of some important parameters to the dynamic response of bird strike on FMLs, which included the configuration of the FMLs, the orientation of the composite layup, and the type and thickness of the metal. The results of the simulation show that the bird strike ability of the configuration (2/1) is better than the configuration (3/2). The bird strike ability of FMLs with the composite layup [45°/-45°] is better than the composite layup [0°/90°]. FMLs with aluminum alloy 2024T3 and metal thickness 0.4mm is the best for the anti-bird strike. Lastly, the bird strike simulation was carried out for the wing leading edge structures with three different skins, aluminum alloy skin and two FMLs skins. The analysis results show that the bird strike ability of the wing leading edge structures with FMLs skin is greatly improved.

Experimental and numerical investigations of low velocity impact on sandwich panels

Sandwich panels with honeycomb cores are used for aerospace frames where mass efficient structures are needed; however the structural integrity of these panels is reduced by low velocity impacts. Their behaviour on impact is investigated in this paper by means of refined numerical Finite Elements simulations. The sandwich panels considered are made of a Nomex™ honeycomb core and thin Al 2024-T3 aluminium alloy skins. A highly detailed FE model (micro mechanical level) has been built by using data obtained from flatwise compressive tests of honeycomb cores and a calibration of the constitutive law and ductile failure of aluminium skins. Numerical results are compared in terms of damage shapes and failure onset with an experimental campaign, carried out using a free-fall apparatus and a spherical steel impactor. The impact phenomena, including the failure of the skin, have been reproduced by means of numerical models with an elevated grade of similarity to a real impact scenario. Discussion and perspective are reported.

The influence of core density on the blast resistance of foam-based sandwich structures

This paper investigates the influence of varying core density on the blast resistance of sandwich panels based on crosslinked PVC cores and aluminium alloy skins. Initially, the findings of a series of compression and single edge notch bend and shear tests are presented in order to characterise the strength and toughness characteristics of the foams and generate data for input into the finite element models.Experimental blast tests employing a ballistic pendulum are then reported, where it is shown that damage within the sandwich panels becomes more severe as the density of the foam core is increased. Indeed, panels based on the lowest density foam (60 kg/m3) did not exhibit any fracture or debonding over the range of impulses considered, instead absorbing energy through plastic deformation in the metal skins and compression of the foam core. In contrast, significant damage, in the form of Mode I core cracking and debonding at the skin–core interface, was apparent in the higher density core systems. A finite element analysis has been employed to model the blast response of the sandwich structures. The FE models successfully predicted the post-test deformed shapes of the panels. These analyses were also used to determine the individual contributions of the skin and core to the energy-absorbing capacity of the various sandwich structures. Here, it was shown that the foam core absorbs more than fifty percent of the overall energy dissipated during the test.

Analysis of the Possibility to Assess the Occurrence of Hidden Corrosion in Lap Joints Using Active Thermography

Analysis of the Possibility to Assess the Occurrence of Hidden Corrosion in Lap Joints Using Active ThermographyThe article details the NDT technique of pulse thermography used for objective diagnosis of riveted lap joints construction. The degradation of materials manifesting in corrosion is inherent in the process of aircraft operation. One type of corrosion is galvanic corrosion occurring in the overlap joints (known as hidden corrosion). As a result of the potential difference between the two layers of the aluminum alloy skin, there occurs the phenomenon of oxidation of the material, producing corrosion products in the form of oxide compounds characterized by heat properties different than those of the base material. Active thermography techniques allow observing infrared energy, which changes due to the difference of thermal properties of the tested materials.

Perforation of sandwich plates with graded hollow sphere cores under impact loading

In this paper, sandwich plates made from 0.8 mm 2024 T3 aluminium alloy skin sheets and graded polymeric hollow sphere cores (having various density gradients) are studied. The experiments at 45 m/s were performed with an inversed perforation setup using SHPB system. Quasi-static tests using the same clamping system allow for the rate effect investigation. Numerical simulations are performed in order to get the indispensable local information (which is not experimentally available) to better understand the perforation process. With these experimental and numerical tools, it is found that the gradient profiles change the perforation process under studied impact loading, whereas they have no influence under quasi-static loading. Under impact loading, a competition of two deforming scenarios at the early stage governed the whole process afterwards. One is the global crush of the first layer in contact with the incident skin and the other is the piercing of the incident skin sheet. When the first layer is rather strong, the incident skin sheet breaks and the perforator makes a hole in the core afterwards. When the first layer is rather weak, the skin sheet folds into the core and develops a much more energy consuming process. The best gradient profile in terms of the energy absorption capacity as well as the non-sheet breaking criterion should be a weaker first layer and a progressively enhanced core.

The impact resistance of polypropylene-based fibre–metal laminates

The high velocity impact response of a range of polypropylene-based fibre–metal laminate (FML) structures has been investigated. Initial tests were conducted on simple FML sandwich structures based on 2024-O and 2024-T3 aluminium alloy skins and a polypropylene fibre reinforced polypropylene (PP/PP) composite core. Here, it was shown that laminates based on the stronger 2024-T3 alloy offered a superior perforation resistance to those based on the 2024-O system. Tests were also conducted on multi-layered materials in which the composite plies were dispersed between more than two aluminium sheets. For a given target thickness, the multi-layered laminates offered a superior perforation resistance to the sandwich laminates. The perforation resistances of the various laminates investigated here were compared by determining the specific perforation energy (s.p.e.) of each system. Here, the sandwich FMLs based on the low density PP/PP core out-performed the multi-layer systems, offering s.p.e.’s roughly double that exhibited by a similar Kevlar-based laminate. A closer examination of the panels highlighted a number of failure mechanisms such as ductile tearing, delamination and fibre failure in the composite plies as well as permanent plastic deformation, thinning and shear fracture in the metal layers. Finally, the perforation threshold of all of the FML structures was predicted using the Reid–Wen perforation model. Here, it was found that the predictions offered by this simple model were in good agreement with the experimental data.

Formability of AA5182/polypropylene/AA5182 sandwich sheets

A sandwich sheet, AA5182/polypropylene/AA5182 (AA/PP/AA), has been developed by the roll bonding of two AA5182 aluminum alloy skin sheets with a pre-rolled polypropylene core sheet at 140 °C. The strain-hardening exponent of the aluminum skin was higher than that of the sandwich sheet, while the strain rate sensitivity of the sandwich sheet was higher than that of the aluminum skin. In order to optimize the sandwich sheet, the formability has been analyzed by constructing the forming limit diagram (FLD) based on the modified Marciniak–Kuczynski (M–K) theory. To account for the planar anisotropic property of the aluminum alloy skin sheet, Hill’s and Barlat’s new yield functions, published in 1948 and 2000, respectively, were employed. When compared with experimental results, the calculated FLD showed reasonably good qualitative agreement. Further analysis showed that the positive effect of the higher strain rate sensitivity of the sandwich sheet is compensated for by the negative effect of its lower strain-hardening exponent so that the greater thickness was the main cause of the higher formability of the sandwich sheet compared to that of the skin sheet.

Mechanical Design for JET-X Aperture Door

The joint European X-ray Telescope, JET-X, is one of the core instruments in the scientific payload of the Russian Spectrum Roentgen-Gamma-RG which is a high energy astrophysics mission. The JET-X structure is constructed from three carbon fibre tubular sections, and contains the instruments and acts as an optical bench. The forward section of the structure has a deployable door, which remains closed until the satellite has reached orbit. The function of the door is to protect sensitive optical surfaces from contamination and moisture from ground operations, during launch and early orbits operations. The door is fabricated from a honeycomb core faced with aluminium alloy skins. The door is approximately 1 m in diameter and weighs about 6 Kg only. The door body, the cutter assembly and the hinge assembly were designed to withstand the mechanical stresses arising during the launch environment. The door assembly was integrated to the main structure and vibration tests which simulate the conditions during the launch were conducted. These tests were: Sine, Random and Shock. This paper presents the door assembly design, the finite element modelling and the dynamic responses during the tests.

Aircraft floor panel developments at British Airways (1967–1973)

In 1967 testing started on balsa, pvc and aluminium alloy cores with aluminium alloy skins and a cost-effectiveness formula was developed to provide a basis for comparison. A specification for improved aluminium/balsa floors was produced and flight trials began with aluminium/aluminium honeycomb floors. Carbon fibre and later glass fibre came on the scene and a new specification was raised, based on more fundamental criteria. As a result over 500 cfrp/Nomex panels have been fitted in 747 aircraft and about 70 grp/Nomex panels. These are much lighter than earlier types of flooring and more cost-effective.