Damage Tolerance Characteristics Research Articles

High Speed Machining of Titanium Alloys

Removal rates for machining titanium alloys are an order of magnitude slower than those for aluminum. The high strength and hardness coupled with the relatively low elastic modulus and poor thermal conductivity of titanium contribute to the slow speeds and feeds that are required to machine titanium with acceptable tool life. Titanium has extremely attractive properties for air vehicles ranging from excellent corrosion resistance to good compatibility with graphite reinforced composites and very good damage tolerance characteristics. At current Buy to Fly ratios, the F-35 Program will consume as much as seven million pounds of titanium a year at rate production. This figure is nearly double that of the F-22 Program, which has a much higher titanium content. As much as 50% of the final cost of titanium parts can be attributed to machining. Specifically, in this task, we are working to improve the material removal rate of titanium to reduce cost. Lockheed Martin is evaluating the potential to use lasers to heat the material ahead of the tool to reduce its strength. Coupled with other technologies that can improve the tool life and prevent the titanium material from welding to the tool, there is hope for a practical solution using similar milling machines to those which exist today, if not a simple retro-fit option. This presentation will present the current progress of this project and its potential impact to the Joint Strike Fighter.

Low-velocity impact damage on dispersed stacking sequence laminates. Part I: Experiments

The stacking sequence design of composite laminates is often limited to combinations of 0°, 90°, and ±45° fibre angle plies. Furthermore, in order to comply to certain stiffness requirements, clustering of plies becomes unavoidable. Although such laminates might have the desired stiffness properties, they may show poor impact and/or compression-after-impact behaviour. A method to redesign the traditional stacking sequences such that the alternative laminates have improved damage resistance whilst keeping similar in-plane and bending stiffness properties as their original traditional stacking sequences is proposed. This method makes use of optimisation tools based on genetic algorithms. In the alternative laminates, the difference between fibre angles of two consecutive plies is maximised and allowed to vary in the 0–90° fibre angle range at intervals of 5°. Manufacturing of such laminates is practical nowadays as the industry is changing its production techniques into accurate automated fibre-placement and tape-laying technologies. A two-step approach is proposed for the design of laminates. In the first step, the optimal laminate is designed in the traditional fashion to cope with the expected quasi-static loads on the structure. The second step consists of redesigning this laminate to better withstand impact loads by dispersing its stacking sequence while keeping similar stiffness properties as in the first step. A traditional laminate and two dispersed stacking sequence alternative layups were tested under low-velocity impact and compression-after-impact loads in order to compare their impact resistance and damage tolerance characteristics. The evaluation of these laminates will also be carried out by the innovative numerical tools proposed in the follow-up of the present paper.

Residual-Strength Assessment of Modified Transport Aircraft Fuselages

In this study, several combinations of fail-safe features were investigated to assess their effect on the damage-tolerance characteristics and residual strength of a typical military transport aircraft fuselage section composed of floating-frame construction. Detailed nonlinear finite element models were developed for a key area of a representative aft fuselage. These local fuselage models were refined to include longitudinal lead and multiple-site-damage cracks of various sizes. A global-to-local modeling procedure was used whereby kinematic boundary conditions applied to the edges of the local fuselage models were developed using internal field quantities from a finite element simulation of a structurally complete (global) fuselage for a given load case. A residual-strength assessment of the relevant aft section for load cases involving a combination of limit internal pressure and flight loads was conducted using highly refined finite element models for cases involving 1) no fail-safe features (i.e., the basic fuselage configuration), 2) inclusion of shear ties as part of a modified fuselage design, 3) inclusion of tear straps as part a modified fuselage design, and 4) inclusion of both shear ties and tear straps. The influence of discrete source damage and small-scale cracking at fastener holes (i.e., multiple-site damage) was considered in the analysis. The critical crack-tip opening-angle fracture criterion and plane-strain core concept were employed in performing large-deformation elastic-plastic stable-tearing analyses using three-dimensional fracture-analysis code and structural analysis of general shells code to predict the fuselage section's residual strength. These results suggest that shear ties and tear straps may be effectively used to increase the structural integrity of the aft fuselage. Key results from this study may provide useful information regarding the efficiency of incorporating fail-safe design modifications in the existing aircraft to enhance the residual strength of cracked sections of the fuselage.

Analytical Modeling of Impact Resistance and Damage Tolerance of Laminated Composite Plates

The present study is concerned with a fully analytical model for predicting impact resistance and damage tolerance characteristics of laminated fiber reinforced composite plates. The energy balance for damage development during the impact process is established with the help of a localized deformation field, in which various effects such as the anisotropy of the delaminated sublaminates, elastic property degradation due to matrix cracking, multiple delaminations, and membrane deformation within the damage zone are taken into account. This results in upper and lower bounds of the delamination threshold load and a deflection-dependent critical load for stable delamination propagation. On the basis of the conservation of energy principle, simple relationships are found between the resulting damage area and the impact energy as well as the peak impact load. The global buckling load of damaged rectangular plates is derived from the energy-based stability criterion, and the residual ultimate strength is calculated using the von Karman postbuckling strength theory. The analytical solutions obtained turn out to be capable of providing highly accurate predictions compared with the standard impact and the compression-after-impact experiments of carbon-fiber-reinforced polyetherimide and aromatic polymer composites.

Numerical Estimates of the Compressive Strength of Impact-damaged Sandwich Composites

Improved numerical models have been developed for predicting the compressive strength of impact-damaged sandwich composites comprised of woven-fabric graphite-epoxy facesheets and Nomex honeycomb cores. The proposed methodology contains three key elements: (1) the use of nondestructive inspection estimates of damage, (2) the incorporation of nonlinear through-the-thickness core-crush constitutive response to account for core failure, and importantly (3) the simulation of progressive loss of facesheet structural integrity based upon userspecified ply failure criteria. The finite element predictions of residual strength for panels impacted with relatively blunt objects correlate well with the experimental observations; this is in contrast to a number of estimates from literature that tend to overestimate the strength. The proposed approach may potentially facilitate sandwich design by providing insight into the relationships between material configuration and damage that lead to improved damage tolerance characteristics.

Evaluation of hygrothermal effects on the shear properties of Carall composites

Fiber metal laminates are the frontline materials for aeronautical and space structures. These composites consists of layers of 2024-T3-aluminum alloy and composite prepreg layers. When the composite layer is a carbon fiber prepreg, the fiber metal laminate, named Carall, offers significant improvements over current available materials for aircraft structures. While weight reduction and improved damage tolerance characteristics were the prime drivers to develop this new family of materials, it turns out that they have additional benefits, which become more and more important for today's designers, such as cost reduction and improved safety. The degradation of composites is due to environmental effects mainly on the chemical and/or physical properties of the polymer matrix leading to loss of adhesion of fiber/resin interface. Also, the reduction of fiber strength and stiffness are expected due to environmental degradation. Changes in interface/interphase properties leads to more pronounced changes in shear properties than any other mechanical properties. In this work, the influence of moisture in shear properties of carbon fiber/epoxy composites and Carall have been investigated by using interlaminar shear (ILSS) and Iosipescu tests. It was observed that hygrothermal conditioning reduces the Iosipescu shear strength of CF/ E and Carall composites due to the moisture absorption in these materials.

A review on the development and properties of continuous fiber/epoxy/aluminum hybrid composites for aircraft structures

Weight reduction and improved damage tolerance characteristics were the prime drivers to develop new family of materials for the aerospace/aeronautical industry. Aiming this objective, a new lightweight Fiber/Metal Laminate (FML) has been developed. The combination of metal and polymer composite laminates can create a synergistic effect on many properties. The mechanical properties of FML shows improvements over the properties of both aluminum alloys and composite materials individually. Due to their excellent properties, FML are being used as fuselage skin structures of the next generation commercial aircrafts. One of the advantages of FML when compared with conventional carbon fiber/epoxy composites is the low moisture absorption. The moisture absorption in FML composites is slower when compared with polymer composites, even under the relatively harsh conditions, due to the barrier of the aluminum outer layers. Due to this favorable atmosphere, recently big companies such as EMBRAER, Aerospatiale, Boing, Airbus, and so one, starting to work with this kind of materials as an alternative to save money and to guarantee the security of their aircrafts.

Response surface characterization of the mechanical behavior of impact-damaged sandwich composites

In this research, Response Surface Methodology (RSM) is employed to characterize the influence of material configuration on the damage tolerance and residual strength characteristics of sandwich composites. Test specimens used were comprised of carbon-epoxy woven fabric facesheets, and Nomex honeycomb cores. The ranges of the material configuration used are typical of those employed in aircraft applications. A series of carefully selected tests were used to isolate the coupled influence of various combinations of the number of facesheet plies, core density, and core thickness on the damage formation and residual strength degradation due to normal impact. Response surface estimates suggest that impact damage development and residual strength degradation are highly material and lay-up configuration dependent. Increasing the core thickness for a specific number of facesheet plies resulted in decreasing the impact damage, whereas increasing the number of facesheet plies for a given core thickness resulted in enhancing the residual strength. The derived damage tolerance and residual strength models can lead to a better understanding of the mechanical behavior of the impact-damaged sandwich composites, and hence improve their design and expand their applications.

Impact damage resistance of buckled carbon/epoxy panels

Results are presented for impact tests conducted on postbuckled, thin, orthotropic carbon/epoxy laminates manufactured by the RTM process from preforms of uniweave fabric, some of which were stitched with Kevlar sewing thread. They were impacted by a 13 mm diameter 8.9 g steel projectile at velocities ranging to 90 m/s on their convex and concave faces. For impact velocities above a critical value, either penetration occurred or the specimens failed catastrophically. Their corresponding critical velocity boundaries were established. Their projected impact damage area increased with increased impact velocity, however, it was independent of preload. Their postimpact properties remained largely unchanged. Stitching reduced significantly the projected damage area, however none of their other impact damage resistance and tolerance characteristics were effected, including their critical velocities for penetration and catastrophic damage. Impacting on the concave face, compared to impacting on the convex face, reduced the projected damage area and the penetration critical velocity boundary. No catastrophic damage was observed for these specimens.

Design case study for a comparative performance analysis of aerospace materials

Reduction of the life cycle cost of aircraft is the main focus of aerospace material development at the moment. By improving the performance of aluminium alloys and using other material systems such as carbon fibre reinforced plastics (CFRP), the life cycle cost can be reduced. This paper presents a design study of a pressurised fuselage section in aluminium and CFRP, focussing on the appropriate design criteria that are to be included for each material to get a sound comparison. Furthermore, a new reference for the aluminium fuselage will be presented applying new alloys and production techniques. Damage tolerance characteristics of both aluminium and CFRP are discussed, and appropriate design philosophies are presented. A correlation between residual strength and stiffening ratio is identified and used for the optimisation of the designs. It is concluded that limited fatigue damage growth and residual strength are the main design drivers for a pressurised fuselage, and the main weakness of CFRP is identified in this field, especially when looking at the compressive strength after impact and sensitivity to damages in tension. The new and improved aluminium alloys show an increase in yield stress and fracture toughness, and by laser-welding the stiffeners to the skin, the performance in compression and the fatigue life of the panel increases.

Impact response and damage tolerance characteristics of glass–carbon/epoxy hybrid composite plates

Impact behaviour and post impact compressive characteristics of glass–carbon/epoxy hybrid composites with alternate stacking sequences have been investigated. Plain weave E-glass and twill weave T-300 carbon have been used as reinforcing materials. For comparison, laminates containing only-carbon and only-glass reinforcements have also been studied. Experimental studies have been carried out on instrumented drop weight impact test apparatus. Post impact compressive strength has been obtained using NASA 1142 test fixture. It is observed that hybrid composites are less notch sensitive compared to only-carbon or only-glass composites. Further, carbon-outside/glass-inside clustered hybrid configuration gives lower notch sensitivity compared to the other hybrid configurations.

Fatigue behaviour of riveted Glare lap joints

Fibre metal laminates, such as Glare, are a family of materials with very interesting properties for fatigue‐critical applications. This article describes the results of a research programme carried out to evaluate the fatigue and damage tolerance characteristics of riveted Glare lap joints, representative of fuselage longitudinal joints. The comparison with the behaviour of metallic joints shows that different contributions are made to the total fatigue life, with the crack propagation life being by far the longer one for the Glare material whereas the crack nucleation life covers almost the entire fatigue life for metallic joints. Design rules should take this peculiar behaviour into consideration, to achieve the maximum benefit in the use of this class of materials in fatigue‐critical applications.

The Heat Treatment of Al-Cu-Li Alloys Ensured the Stability of Structure and Properties at Long Low Temperature Exposure

The structural transformations and properties changes of sheets from Al-Cu-Li alloys with two levels of Li content (alloy I with 1.7 wt. % and alloy II with 1.5 %) during long low temperature exposure (LLTE) at 85 °C, 1000 h and 90 °C, 4000 h have been investigated. Three-stage ageing ensured the strengthening of alloy II mainly at the expense of T 1 -phase precipitates with the favorable phases morphology on the grain boundaries and in their bodies. The alloy II hot rolled sheets have the high level of properties, including Impact Behavior Energy before and after LLTE The absence of structural changing during LLTE were confirmed by DSC and TEM data. The alloy I sheets, aged on offered regime, had some changes in properties, however they were no more than 5 %. 1. Introduction. The Al-Cu-Li alloys have a unique combination of high strength, low density, weldability, satisfactory damage tolerance characteristics in conditions of static and fatigue loading [1, 2] A lack of these alloys is the reduced plasticity and inherent to the majority of aluminium alloys with lithium insufficient stability of the structure and properties in conditions of long low temperature exposure (LLTE) [3]. The cause of this are the changes in quasiequilibrium solubility at ageing temperature and LLTE temperature and the fine dispersed δ'-precipitates formation from the nonequilibrium solid solution which has stayed supersaturated after ageing [4, 5], and also formation of GPI zones containing Al and Cu [6]. For increase of thermal stability in the previous work [5] the two-stage ageing 120°C, 20 h + 170°C, 12 h with slow cooling from temperature 170°C up to temperature LLTE was offered. This mode has ensured the sufficient stability for the alloy containing 1.75 weight % of lithium in conditions LLTE 85°C, 1000 hours. In the present work it was necessary to check up, as far as the structure and properties of the alloy will stable after the more rigid test - 90°C. 4000 hours. Assume that the reason of structure instability during LLTE is Li solid solution supersaturation [4, 5], for increase of stability it is reasonable to decrease Li content in alloy. Therefore in the given work the alloy with the content of 1.5 weight % lithium was investigated also. The optimum ageing regime was developed and the phase composition, precipitates morphology, complex of properties and their stability during LLTE were investigated. 2. Experimental. The study was carried out on the 10 and 7 mm hot rolled sheets of two Al-Li-Cu (type 1460) alloys with 1.75 wt. Li % (alloy 1) and 1.5 % (alloy II). The hot rolling temperature was 440-470°C. Sheets were solutionised at 530°C, 1.5 h and quenched in water. The time between quenching and stretching was less than 3 h. The residual deformation degree at the stretching was 3 %. After quenching sheets have unrecrystallized structure. The average fiber thickness was 15 μm. The tensile properties, KCU, K e0 , FCPR, Impact Behavior Energy were determined. The changes in the sheets structure were studied by TEM, X-Ray phase analysis, DSC techniques. The alloy thermal stability estimation during LLTE (85°C, 1000 h or 90°C, 4000 h) is difficult due to the limited researches time. There was a necessity to substitute these regimes for more brief high temperature imitative ones.

Development of fibre metal laminates for advanced aerospace structures

Fibre metal laminates offer significant improvements over current available materials for aircraft structures. While weight reduction and improved damage tolerance characteristics were the prime drivers to develop these new family of materials, it turns out that they have additional benefits which become more and more important for today’s designers, e.g. cost reduction and improved safety. The combination of these aspects in one material is an extraordinary achievement. It shows that GLARE is a unique material for aircraft applications. Moreover, it makes clear that, through its unique combination of properties, GLARE is a strong candidate material for fuselage skin structures of the new generation of aircraft.

Towards application of fibre metal laminates in large aircraft

Fibre metal laminates were developed at Delft University during the last two decades as a family of new hybrid materials consisting of bonded thin metal sheets and fibre/adhesive layers. This laminated structure provides the material with excellent fatigue, impact and damage tolerance characteristics and a low density. While the 20 per cent weight reduction was the prime driver behind the development of this new family of materials, it turns out that additional benefits like cost reduction and an improved safety level have become more and more important. The combination of these aspects in one material makes fibre metal laminates a strong candidate material for fuselage skin structures of the new generation of high capacity aircraft. The focus on this application currently leads to industrialization and qualification that makes this material available to the aircraft designer.

Damage Progression in Bolted Composites

Structural durability, damage tolerance, and progressive fracture characteristics of bolted graphite/epoxy composite laminates are evaluated via computational simulation. Constituent material properties and stress and strain limits are scaled up to the structure level to evaluate the overall damage and fracture propagation for bolted composites. Single and double bolted composite specimens with various widths and bolt spacings are evaluated. The effect of bolt spacing is investigated with regard to the structural durability of a bolted joint. Damage initiation, growth, accumulation, and propagation to fracture are included in the simulations. Results show the damage progression sequence and structural fracture resistance during different degradation stages. A procedure is outlined for the use of computational simulation data in the assessment of damage tolerance, determination of sensitive parameters affecting fracture, and interpretation of experimental results with insight for design decisions.

Processing and Response of Aluminum-Lithium Alloy Composites Reinforced with Copper-Coated Silicon Carbide Particulates

Lithium-containing aluminum alloys have shown promise for demanding aerospace applications because of their light weight, high strength, and good damage tolerance characteristics. Additions of ceramic reinforcements to an aluminum-lithium alloy can significantly enhance specific strength, and specific modulus while concurrently offering acceptable performance at elevated temperatures. The processing and fabrication of aluminum-lithium alloy-based composites are hampered by particulate agglomeration or clustering and the existence of poor interfacial relationships between the reinforcing phase and the matrix. The problem of distribution of the reinforcing phase in the metal matrix can be alleviated by mechanical alloying. This article presents the results of a study aimed at addressing and improving the interfacial relationship between the host matrix and the reinforcing phase. Copper-coated silicon carbide particulates are introduced as the particulate reinforcing phase, and the resultant composite mixture is processed by conventional milling followed by hot pressing and hot extrusion. The influence of extrusion ratio and extrusion temperature on microstructure and mechanical properties was established. Post extrusion processing by hot isostatic pressing was also examined. Results reveal the increase in elastic modulus of the aluminum-lithium alloy matrix reinforced with copper-coated SiC to be significantly more than the mechanically alloyed Al-Li/SiC counterpart. This suggests the possible contributions of interfacial strengthening on mechanical response in direct comparison with a uniform distribution of the reinforcing ceramic particulates.

Progressive Damage and Fracture of Unstiffened and Stiffened Composite Pressure Vessels

Structural durability and damage tolerance characteristics of pressurized graphite/epoxy laminated thin composite cylinders are investigated via computational simulation. Both unstiffened and integral hoop stiffened cylinders are considered. A computer code is utilized for the simulation of composite structural degradation under loading. Damage initiation, growth, accumulation, and propagation to structural fracture are included in the simulation. The increase of burst pressure due to hoop stiffening is quantified. Results demonstrate the significance of the type and size of local defects on the structural durability of pressurized composite cylindrical shells.

Structural performance of discontinuous metal matrix composites

Discontinuous metal matrix composites (i.e. short fibre and particle reinforced materials) have attained a significant degree of scientific and technological maturity as advanced structural materials. Initial commercialisation has been achieved, with the unique combinations of mechanical and physical properties afforded by metal-ceramic systems proving appropriate for a variety of structural and semistructural applications. In recent years there has been important consolidation in the understanding of basic structural properties in such composites, which are addressed in the present review. The outstanding requirement for an improved understanding of damage tolerance characteristics in these materials is particularly noted. ‘Mesoscopic’ materials architectures (e.g. laminated and functionally graded materials) are also discussed, and the associated potential for development offracture resistant discontinuous metal composite materials highlighted.

The fatigue resistance of TiAl-based alloys

TiAl-based alloys are a class of intermetallic alloys that show potential as a high-temperature structural material. Many applications considered for TiAl-based alloys require certain damage-tolerance characteristics. In this article, the fatigue resistance of two-phase TiAl alloys is reviewed to assess the current understanding of the fatiguedamage processes in this class of emerging materials. Salient features of fatigue-crack growth and stress-life curves are summarized to identify the challenges these fatigue characteristics present in the life assessment of TiAl components.