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Fatigue and damage tolerance issues of Glare in aircraft structures
International Journal of Fatigue 28 (2006) 1116–1123
International Journalof Fatigue www.elsevier.com/locate/ijfatigue
Fatigue and damage tolerance issues of Glare in aircraft structures R.C. Alderliesten *, J.J. Homan Delft University of Technology, Faculty of Aerospace Engineering, Structures and Materials Laboratory, P.O. Box 5058, 2600 GB Delft, The Netherlands Available online 20 March 2006
Abstract Airworthiness regulations require for most structural parts of the aircraft a damage tolerance design philosophy. The fatigue and damage tolerance issues related to the application of the fibre metal laminate Glare in the upper fuselage skin structure are discussed in this paper. As a result of the laminate lay-up (thin aluminium layers with in-between fibre/epoxy layers), the actual stresses in the aluminium layers are higher compared to the applied laminate stresses, resulting in a reduction of crack initiation life compared to monolithic aluminium. However, the crack propagation life is significantly longer than monolithic aluminium due to the bridging by the intact fibres over the fatigue crack. It is discussed that the prediction methods developed for monolithic aluminium can be modified for Glare. Durability issues with respect to moisture absorption and temperature effects are discussed as well as the residual strength of Glare. The through the thickness effects of fatigue in Glare joints are explained and compared to the monolithic aluminium behaviour. Because of the increased fatigue crack propagation life and the higher residual strength of Glare compared to aluminium, higher design allowable stresses are possible, increasing the inspection intervals. 2006 Elsevier Ltd. All rights reserved. Keywords: Fatigue initiation; Crack propagation; Residual strength; Glare; Joints
1. Introduction Airworthiness regulations (FAR/JAR 25) require for the design of most structural parts of the aircraft the use of a damage tolerant design philosophy. Many aircraft manufacturers comply with this philosophy for their aircraft fuselage skins by applying slow fatigue crack growth strategies. These strategies are intended to insure that the fatigue cracks can be discovered by inspection prior to failure. To achieve slow crack growth in the case of aluminium alloys, the allowable design stress levels are relatively low compared to the static strength values. However, in the case of Glare since the intrinsic fatigue crack growth resistance is high [1], much higher allowable design stress levels can be selected with respect to fatigue. For Airbus, this has been the major driver for the selection of Glare as a fuselage skin material [2], because weight savings can now be
*
Corresponding author. Tel.: +31 15 278 5492; fax: +31 15 278 1151. E-mail address: [email protected] (R.C. Alderliesten).
0142-1123/$ - see front matter 2006 Elsevier Ltd. All rights reserved. doi:10.1016/j.ijfatigue.2006.02.015
achieved by two contributions: higher operational stress levels and lower material density. This paper discusses the fatigue and damage tolerance issues related to the application of Glare in the fuselage skin structure with respect to material characteristics and airworthiness requirements in comparison to monolithic aluminium alloys. 2. Material definition The Fibre Metal Laminate Glare consists of thin layers of 2024-T3 aluminium alloy, bonded together with S2-glass fibres in a FM94 adhesive system. The material is cured in an autoclave cycle with a maximum pressure of six bar and a curing temperature of 120 C. Since several Glare grades with a large amount of layups are possible, a clear coding system is used to identify the Glare grade and lay-up. This coding system is important for design, production and material qualification. For instance a laminate with three aluminium layers of 0.3 mm thickness with two cross-ply fibre layers in-between
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
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Table 1 Overview of typical Glare grades Glare grade
Sub
Glare 1 Glare 2
– Glare Glare – Glare Glare – Glare Glare
Glare 3 Glare 4 Glare 5 Glare 6
2A 2B 4A 4B 6A 6B
Metal sheet thickness [mm] and alloy
Prepreg orientationa in each fibre layerb
Main beneficial characteristics
0.3–0.4 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5
0/0 0/0 90/90 0/90 0/90/0 90/0/90 0/90/90/0 +45/45 45/+45
fatigue, strength, yields stress fatigue, strength fatigue, strength fatigue, impact fatigue, strength in 0 direction fatigue, strength in 90 direction impact shear, off-axis properties shear, off-axis properties
7475-T761 2024-T3 2024-T3 2024-T3 2024-T3 2024-T3 2024-T3 2024-T3 2024-T3
a
All aluminium rolling directions in standard laminates are in the same orientation; the rolling direction is defined as 0, the transverse rolling direction is defined as 90. b The number of orientations in this column is equal to the number of prepreg layers (each nominally 0.133 mm thick) in each fibre layer.
Because of the higher stiffness of the aluminium layers as compared to the fibre layers as well as to the presence of residual tensile stresses after curing, the aluminium layers experience higher tensile stresses as compared to the applied laminate stresses on the laminate. This results in an increase of stress amplitude and a reduction of initiation life of the aluminium in Glare as compared to monolithic aluminium for equal applied stress. With respect to design, the allowable stresses for Glare following from the fatigue initiation requirements are lower compared to monolithic aluminium. However, fatigue initiation is in most cases not the limiting criterion determining the allowable.
is coded as Glare 3–3/2–0.3, which refers to the Glare grade, the lay-up and the aluminium layer thickness, respectively. The lay-up for this case is defined as ½2024 T 3=0 glass=90 glass=2024 T 3=90 glass=0 glass=2024 T 3 The Glare laminates have a symmetrical lay-up to avoid secondary bending effects due to unsymmetrical internal stresses. An overview of the typical Glare grades with the lay-up definition is given in Table 1. 3. Fatigue initiation Fatigue cracks initiate in the aluminium layers of Glare, as the fibres layers are insensitive to the cyclic loading. This means that the fatigue initiation is analogous to that of monolithic aluminium. Using a similarity approach - similar stress state in the aluminium layers in Glare and in monolithic aluminium yields the same initiation life - the fatigue initiation life of Glare can be determined by calculating the actual stresses in the aluminium layers. These stresses can be calculated using the classical laminate theory (CLT) and consist of a contribution due to the applied stresses and a contribution due to the presence of residual stresses after the curing process, see Fig. 1. The residual stress state after curing consists of tensile stresses in the aluminium layers and compressive stresses in the fibre layers as a result of the mismatch in coefficients of thermal expansion.
Slam
4. Crack propagation During the crack propagation phase, the intact fibres carry a significant part of the load over the crack in the aluminium layers and restrain the crack opening. This socalled fibre bridging reduces the stress intensity at the crack tip in the aluminium layers and thus reduces the crack propagation rate considerably, although the far field stresses in the aluminium layers are higher than the far field stress in a monolithic aluminium sheet [3]. The approximately constant crack growth rate leads at the end to a much longer fatigue life in Glare compared to monolithic aluminium, see Fig. 2. Along with the crack growth in the aluminium layers, delamination growth occurs at the interfaces between the aluminium and fibre layers. This delamination growth is
SAl_tot
Slam
Slam
= SAl Sf,0 Sf,90 SAl
+ Slam
Glare
SAl_cure Sf,0_cure Sf,90_cure SAl_cure
SAl Sf,0 Sf,90 SAl
SAl_cure Sf,0_cure Sf,90_cure SAl_cure
SAl_tot
Fig. 1. Approach of calculating the stresses in the aluminium layers.
Aluminium
1118
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123 50 Smax = 120 MPa R = 0.05 W = 140 mm L = 580 mm f = 10 Hz 2a0 = 5 mm
45
Aluminium 2024 T3 - 2mm 40 35 30
GLARE 4B - 4/3 - 0.5 LT
a 25 (mm) 20
GLARE 3 - 3/2 - 0.3 L
15 10 5 0 0
20
40
60
80
100
120
140
160
180
200
N (kcycles)
Fig. 2. Comparison between crack growth curves for monolithic aluminium and Glare [3].
bridging stress
crack opening
delamination shape
Fig. 4. Schematic illustration of crack opening, delamination shape and bridging stress. Fig. 3. Illustration of fibre bridging in Glare.
induced by the shear stresses that occur as a result of the load transfer from the aluminium layer to the fibre layers, which is a result of the fibre bridging, see Fig. 3. The crack opening, the delamination length and the bridging stress are interrelated as shown in Fig. 4. An increase in the delamination length gives a reduction of the bridging stress and an increase of the crack opening. During crack growth, these mechanisms are in balance with each other, resulting in the approximately constant crack growth rates. The methodology for predicting fatigue cracks in flat Glare sheets is based on the assumption that all aluminium layers in the laminate contain cracks of equal lengths, see Fig. 5.
The crack growth behaviour in Glare can be described with the stress intensity factor following from Linear Elastic Fracture Mechanics by superposition. The contribution of the crack closing bridging fibres is subtracted from the crack opening contribution of the far field stresses present in the aluminium layers [4,5] K tip ¼ K farfield K bridging
ð1Þ
With a Paris relation the crack growth rate can then be determined. da ¼ CDK ntip dN
ð2Þ
where the constants are determined based on experimental results from fatigue tests on thin aluminium sheets. The effect of fibre bridging in the stress intensity factor Kbridging is
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
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acrack
taluminium tprepreg
starter notch Fig. 5. Through the thickness crack distribution in a flat Glare panel.
calculated by integrating the fibre bridging stress over the total crack length. The bridging stress at a certain location along the crack length follows from the delamination length, the crack opening and the shear deformation of the total prepreg layer at that location. Although the mechanisms discussed above become more complicated, when the loading direction is under an angle with respect to the fibre directions of the Glare panels, the approach remains the same. The laminate properties are determined for the given off-axis angle and are used in the calculation methods. A slightly different situation is the configuration with only a surface crack (i.e., a crack through the thickness of one single aluminium surface layer), which can be initiated as a scratch caused by a sharp object. Again, the stress intensity factor at the crack tip is the far field stress intensity factor subtracted with the factor caused by fibre bridging. However, in the case of a surface crack the intact aluminium layers give an additional contribution to crack bridging, thus giving an even lower stress intensity factor and crack growth rate than for the configuration with through cracks [6]. However, as a result of the surface crack, additional bending occurs as compared to the configuration with through cracks [3]. The resultant stress state depends on the crack length and laminate thickness. For small cracks in thick laminates, the effect can be ignored. Due to the differences in bridging mechanism and secondary bending, the applied calculation method is different [7,8]. However, analogous to the configuration with cracks through the thickness, the method is based on relations for the stress intensity factor based on the bridging. Part through cracks (e.g., corner cracks at holes) can be considered as a special form of surface cracks. Each cracked layer can be viewed as a surface crack with respect to the remaining (deeper) part of the laminate [9,10].
temperature and the other is the effect of moisture absorption. The effect of moisture on the material can be divided into three main effects: • corrosion and more rapid crack growth rates in aluminium • decrease of adhesive interface strength between fibre layer and aluminium • plasticising of the matrix. The chemical bond between the adhesive and the oxide layer of the aluminium and the oxide layer itself is degraded due to moisture absorption by the matrix, which results in a decrease of the adhesive interface strength. A key parameter is the glass transition temperature Tg of the epoxy, which is the temperature at which the characteristics of the epoxy change from a glassy to that of a rubbery state. This transition temperature decreases with the amount of moisture absorbed, as is illustrated with the stiffness drop in Fig. 6. Degradation of the mechanical properties and an increase of shear deformation in fatigue crack situations result in less crack opening restraint. This fact, together with larger delamination areas resulting from reduced delamination toughness, gives less fibre bridging and thus larger stress intensities at the crack tip in the aluminium layers.
E [MPa]
5. Temperature and moisture With respect to fatigue crack initiation and propagation, two aspects are especially important with respect to the durability characteristics of Glare: one is the influence of
Tg2
Tg1
T [˚C]
Fig. 6. Reduction of the glass transition temperature as result of moisture absorption.
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R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
As a result of the plasticising of the matrix, the glass transition will take place at lower temperatures than in a situation without moisture absorption [11]. For strength justification, this implies that if the glass transition temperature Tg is within the operational temperature range, the fatigue behaviour above Tg needs to be taken into account. For Glare, which is cured at a temperature of 120 C, the glass transition temperature Tg in dry condition is about 100 C [11], while after moisture absorption the Tg can reduce to about 80 C. Application of different adhesive systems in the prepreg of Glare with higher curing temperatures can increase the glass transition temperature both to levels above the operational temperature range. The environmental temperature has a significant effect on the fatigue behaviour of monolithic aluminium as well as Glare. An elevated temperature can have a negative effect on the fatigue crack growth behaviour of Glare, whereas a decrease in temperatures can result in a significant increase of fatigue life. The reduction of the fatigue life at elevated temperatures is mainly due to two factors. At elevated temperatures, the moisture absorption in the matrix increases. Further, elevated temperatures cause voids and microscopic defects to open and attract even more fluid. The second factor is the softening and weakening of the matrix due to the increasing temperatures. Weakening induces less crack opening restraint and less fibre bridging [12]. On the other hand, at low temperatures, the matrix becomes stiff and brittle, enhancing the crack opening restraining, resulting in a reduction in the stress intensity factor. In addition, the resistance of aluminium to fatigue crack growth increases at low temperatures, which adds
to the beneficial effect on the fatigue behaviour induced by the stiffer fibre layers. 6. Residual strength The airworthiness regulations require that the aircraft structure must be able to carry limited loads after the event of accidental damage, ignoring for this case the dynamic effects [14]. An example of a criterion to meet these requirements is the so-called two-bay crack criterion, which is related to the residual strength of a structure. The residual strength of a material is defined as the remaining static strength in case of any damage occurring during operation life. As part of a conservative design procedure, the Glare panel is assumed to contain sharp notches. Two types of damages are considered within the residual strength justification of Glare [15]: • Accidental damages • Fatigue damages The difference between the two types is related to the fibre layers in the wake of the crack. Whereas the fibre layers in a fatigue crack configuration are intact, all fibres are assumed to be broken in case of an accidental damage. Each type has its characteristic fracture mechanism. Glare panels with accidental damages loaded until fracture show stable crack extension prior to unstable rapid crack growth and failure, see Fig. 7. This means that the fracture behaviour is similar to the response of monolithic aluminium to static loading [16]. To predict the fracture behaviour and the residual strength of Glare the R-curve approach has been adopted,
Fig. 7. Crack extension versus applied load curve obtained from a residual strength test at room temperature in lab air on a Glare3 panel with an accidental damage case [15].
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
1121
Fig. 8. KR curves for Glare3-3/2-0.3 for different widths and starter notches, based on a compliance plastic zone correction. [15].
because the residual strength behaviour is similar to monolithic aluminium. The R-curve is treated as a material parameter valid between a given range of panel widths and starter notch lengths. The methods incorporate stable crack extension and limited amount of plasticity. An example of the R-curves for Glare 3 is given in Fig. 8. 7. Joints Major concerns in the design of pressurized aircraft fuselages are riveted longitudinal lap joints and circumferential butt joints. Each fastener hole acts as a stress raiser and there are literally thousands of them in an aircraft fuselage. Very often fatigue cracks start at about the same time at different holes, which will lead to a multiple site damage (MSD) situation. The presence of a large number of relatively small cracks will reduce the residual strength of a joint considerably, much more than a single small crack will do, and even more than one large crack with the same accumulated size of the small cracks. Furthermore, the
interaction between different cracks often leads to an increased crack growth rate. The consequence of this behaviour is either a low allowable design stress level or a small inspection interval. Glare however behaves in a completely different manner. Under the same stress level as monolithic aluminium joints, fatigue crack initiation in a Glare joint will, as mentioned before, start earlier. Fatigue crack initiation in Glare will occur in the layer with the highest stresses due to secondary bending and tensile loading. In a joint, this layer will be located at the mating surface, which is the surface in contact with the other sheet of the joint. Propagation of such cracks is a very slow process. The presence of the fibre/matrix layers prevents these cracks from propagating into the deeper aluminium layers and because the stresses in the deeper layers are lower, fatigue crack initiation in those layers will start at a (much) higher number of flights. This process leads to a crack distribution in thickness direction as shown in Fig. 9. Once a fatigue crack initiates at the mating surface of monolithic aluminium joints, a relatively
acrack
σnom+σbending
taluminium tprepreg
bore hole Fig. 9. Through the thickness crack distribution in a Glare joint.
1122
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
small part of the fatigue life is needed to propagate through the thickness and to become visible from the outside of the joint. However, in Glare the fatigue crack remains for a very long time in the layer at the mating surface, remaining invisible from the outside [13]. Because of the slow crack growth in the outer layer, the presence of intact fibres and the delay in crack initiation in the deeper layers, the reduction of residual strength during service develops at a much smaller rate than for comparable aluminium structures. The reduction of residual strength for a Glare joint as a result of slow crack propagation is schematically illustrated in Fig. 10. Certification of riveted joints in Glare requires a damage tolerance analysis of these joints. Methods for establishing
PUL Residual strength curve PLL
Crack propagation curve ac,LL ac,UL Ni
NUL
NLL
Flights
Fig. 10. Schematic crack initiation, crack propagation and residual strength behaviour of a joint in Glare [17].
the initiation life of a joint are based on available methods for joints in monolithic aluminium [18] and extended for fibre metal laminates [19]. Multiple Site Damage can be accounted for by applying stochastic approaches (Monte Carlo simulations) for crack initiation. Crack propagation in Glare such as shown that in Fig. 9 as well as the residual strength of joints can be modelled rather accurately [20]. As has been discussed, fatigue crack growth rates in Glare are relatively low. The difference in crack growth rates of aluminium and Glare is illustrated in Fig. 11 together with the corresponding decrease in residual strength. Although the first aluminium layer initiates significantly faster than monolithic aluminium, the residual strength of Glare remains higher than aluminium. This implies that when using Glare, it is even possible to select a relatively large allowable design stress level and still find an ultimate load life sufficiently large to avoid (small) inspection intervals. This ultimate load life can be well beyond the Design Service Goal of an aircraft. With this in mind, it becomes attractive to design a joint such that it will be capable to sustain ultimate load during its complete design life. In that case, minimum values (so-called B-values) instead of typical values must be used to prove ultimate strength capabilities. Fig. 12 gives an impression of initiation, crack propagation and residual strength for such a case. As long as ultimate loads can be sustained, no inspection is required. Since cracks are not visible from the outside of a joint, this can be a very beneficial side effect.
50
400
45 350 40 RS Glare3-5/4-0.3
300
250 a [mm]
30
25
200 RS 2024-T3
20 150
Residual Strength [MPa]
35
Limit load
15
CG Glare3-5/4-0.3 layer 1
CG Glare3-5/4-0.3 layer 2
100
10 CG 2024-T3
50
5
0 0
20000
40000
60000
80000
100000
120000
0 140000
N [cycles]
Fig. 11. Comparison between the crack growth (CG) of monolithic aluminium and the first two layers of Glare with the corresponding decrease in residual strength (RS).
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
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MSD onset Typical residual strength
PUL
PLL Minimum residual strength ac,LL Propagation curve of lead crack Average crack
ac,UL Ni
DSG
NUL
NLL
Flights
Fig. 12. Initiation, crack propagation and residual strength behaviour of a joint with MSD in Glare.
8. Conclusions To reach the technology readiness of Glare for application in primary skin structures of aircraft fuselages, all the issues with respect to fatigue and damage tolerance have been investigated and covered by new developed calculation methods. The different approaches compared to monolithic aluminium have been validated and justified based on large amount of experimental results. References [1] Roebroeks G. Towards GLARE - The Development of a fatigue insensitive and damage tolerant aircraft material, PhD Thesis, Delft University of Technology, 1991. [2] Pora J, Hinrichsen J. Material and Technology Developments for the Airbus A380, in: Proceedings of the 22nd International SAMPE Europe Conference of the society for the Advancement of Materials and Process Engineering, Paris, 2001. [3] Alderliesten RC. In: Vlot A, Gunnink JW, editors. Fatigue, Fibre Metal Laminates - An introduction. Kluwer Academic Publishers; 2001. p. 155–71. [4] Alderliesten RC. Fatigue crack propagation and delamination growth in Glare, PhD Thesis, Delft University of Technology, 2005. [5] Alderliesten RC. Energy Release Rate approach for delamination in a fatigue crack configuration in Glare, Proceedings of the 21st International Congress of Theoretical and Applied Mechanics, Warsaw, 2004. [6] Alderliesten RC, Homan JJ. In: Varvani-Farahani A, Brebbia CA, editors. Fatigue Crack Growth Behaviour of Surface Cracks in Glare, Fatigue Damage of Materials - Experiments and Analysis. Southampton, Boston: WIT press; 2003. p. 213–22. [7] Gonesh KAM. Development of a fatigue crack growth model for surface cracks in GLARE, Master Thesis, Delft University of Technology, 2000.
[8] Gonesh KAM. Crack growth prediction of surface cracks in Glare, Preliminary Thesis, Delft University of Technology, 1999. [9] Randell C, van der Zwaag S. On subsurface crack growth in fibre metal laminate materials, Proceedings of the 35th INternational SAMPE Technical Conference, Dayton OH, USA, 2003. [10] Randell C. Subsurface Fatigue Crack Growth in Fibre Metal Laminates, PhD Thesis, Delft University of Technology, 2005 (title/ date may be subject to changes). [11] Borgonje B, van der Hoeven W. In: Vlot A, Gunnink JW, editors. Long-term behaviour, Fibre Metal Laminates - an introduction. Kluwer Academic Publishers; 2001. p. 155–71. [12] Beumler Th. Flying GLARE, A contribution to aircraft certification issues on strength properties of non-damaged and fatigue damaged GLARE structures, PhD Thesis, Delft University of Technology, 2004. [13] Homan JJ, Mu¨ller RPG, Pellenkoft F, de Rijck JJM. In: Vlot A, Gunnink JW, editors. Fatigue of riveted joints, Fibre Metal Laminates - an introduction. Kluwer Academic Publishers; 2001. p. 173–95. [14] Damage tolerance and fatigue evaluation of structure, Federal Aviation Regulations Part 25 – Airworthiness Standards: Transport Category airplanes, Section 25.571, Federal Aviation Administration, Department of Transportation, Washington D.C., 1997. [15] de Vries TJ. Blunt and sharp notch behaviour of Glare laminates, PhD Thesis, Delft University of Technology, 2001. [16] de Vries TJ. In: Vlot A, Gunnink JW, editors. Residual strength, Fibre Metal Laminates - an introduction. Kluwer Academic Publishers; 2001. p. 197–217. [17] Beumler Th. In: Vlot A, Gunnink JW, editors. Damage tolerance aspects, Fibre Metal Laminates - an introduction. Kluwer Academic Publishers; 2001. p. 219–33. [18] Homan JJ, Jongebreur AA. Calculation method for predicting the fatigue life of riveted joints, ICAF Symposium Proceedings (ed. Blom), 1993;175–191. [19] Homan JJ. Procedure for estimating the fatigue life in mechanically fastened joints in Glare, ICAF Symposium Proceedings (ed. J. Rouchon), 2001;941–946. [20] Alderliesten RC, Hagenbeek M, Homan JJ, Hooijmeijer PA, de Vries TJ, Vermeeren CAJR. Fatigue and Damage Tolerance of Glare. Applied Composite Materials 2003;10:223–42.
International Journalof Fatigue www.elsevier.com/locate/ijfatigue
Fatigue and damage tolerance issues of Glare in aircraft structures R.C. Alderliesten *, J.J. Homan Delft University of Technology, Faculty of Aerospace Engineering, Structures and Materials Laboratory, P.O. Box 5058, 2600 GB Delft, The Netherlands Available online 20 March 2006
Abstract Airworthiness regulations require for most structural parts of the aircraft a damage tolerance design philosophy. The fatigue and damage tolerance issues related to the application of the fibre metal laminate Glare in the upper fuselage skin structure are discussed in this paper. As a result of the laminate lay-up (thin aluminium layers with in-between fibre/epoxy layers), the actual stresses in the aluminium layers are higher compared to the applied laminate stresses, resulting in a reduction of crack initiation life compared to monolithic aluminium. However, the crack propagation life is significantly longer than monolithic aluminium due to the bridging by the intact fibres over the fatigue crack. It is discussed that the prediction methods developed for monolithic aluminium can be modified for Glare. Durability issues with respect to moisture absorption and temperature effects are discussed as well as the residual strength of Glare. The through the thickness effects of fatigue in Glare joints are explained and compared to the monolithic aluminium behaviour. Because of the increased fatigue crack propagation life and the higher residual strength of Glare compared to aluminium, higher design allowable stresses are possible, increasing the inspection intervals. 2006 Elsevier Ltd. All rights reserved. Keywords: Fatigue initiation; Crack propagation; Residual strength; Glare; Joints
1. Introduction Airworthiness regulations (FAR/JAR 25) require for the design of most structural parts of the aircraft the use of a damage tolerant design philosophy. Many aircraft manufacturers comply with this philosophy for their aircraft fuselage skins by applying slow fatigue crack growth strategies. These strategies are intended to insure that the fatigue cracks can be discovered by inspection prior to failure. To achieve slow crack growth in the case of aluminium alloys, the allowable design stress levels are relatively low compared to the static strength values. However, in the case of Glare since the intrinsic fatigue crack growth resistance is high [1], much higher allowable design stress levels can be selected with respect to fatigue. For Airbus, this has been the major driver for the selection of Glare as a fuselage skin material [2], because weight savings can now be
*
Corresponding author. Tel.: +31 15 278 5492; fax: +31 15 278 1151. E-mail address: [email protected] (R.C. Alderliesten).
0142-1123/$ - see front matter 2006 Elsevier Ltd. All rights reserved. doi:10.1016/j.ijfatigue.2006.02.015
achieved by two contributions: higher operational stress levels and lower material density. This paper discusses the fatigue and damage tolerance issues related to the application of Glare in the fuselage skin structure with respect to material characteristics and airworthiness requirements in comparison to monolithic aluminium alloys. 2. Material definition The Fibre Metal Laminate Glare consists of thin layers of 2024-T3 aluminium alloy, bonded together with S2-glass fibres in a FM94 adhesive system. The material is cured in an autoclave cycle with a maximum pressure of six bar and a curing temperature of 120 C. Since several Glare grades with a large amount of layups are possible, a clear coding system is used to identify the Glare grade and lay-up. This coding system is important for design, production and material qualification. For instance a laminate with three aluminium layers of 0.3 mm thickness with two cross-ply fibre layers in-between
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
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Table 1 Overview of typical Glare grades Glare grade
Sub
Glare 1 Glare 2
– Glare Glare – Glare Glare – Glare Glare
Glare 3 Glare 4 Glare 5 Glare 6
2A 2B 4A 4B 6A 6B
Metal sheet thickness [mm] and alloy
Prepreg orientationa in each fibre layerb
Main beneficial characteristics
0.3–0.4 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5 0.2–0.5
0/0 0/0 90/90 0/90 0/90/0 90/0/90 0/90/90/0 +45/45 45/+45
fatigue, strength, yields stress fatigue, strength fatigue, strength fatigue, impact fatigue, strength in 0 direction fatigue, strength in 90 direction impact shear, off-axis properties shear, off-axis properties
7475-T761 2024-T3 2024-T3 2024-T3 2024-T3 2024-T3 2024-T3 2024-T3 2024-T3
a
All aluminium rolling directions in standard laminates are in the same orientation; the rolling direction is defined as 0, the transverse rolling direction is defined as 90. b The number of orientations in this column is equal to the number of prepreg layers (each nominally 0.133 mm thick) in each fibre layer.
Because of the higher stiffness of the aluminium layers as compared to the fibre layers as well as to the presence of residual tensile stresses after curing, the aluminium layers experience higher tensile stresses as compared to the applied laminate stresses on the laminate. This results in an increase of stress amplitude and a reduction of initiation life of the aluminium in Glare as compared to monolithic aluminium for equal applied stress. With respect to design, the allowable stresses for Glare following from the fatigue initiation requirements are lower compared to monolithic aluminium. However, fatigue initiation is in most cases not the limiting criterion determining the allowable.
is coded as Glare 3–3/2–0.3, which refers to the Glare grade, the lay-up and the aluminium layer thickness, respectively. The lay-up for this case is defined as ½2024 T 3=0 glass=90 glass=2024 T 3=90 glass=0 glass=2024 T 3 The Glare laminates have a symmetrical lay-up to avoid secondary bending effects due to unsymmetrical internal stresses. An overview of the typical Glare grades with the lay-up definition is given in Table 1. 3. Fatigue initiation Fatigue cracks initiate in the aluminium layers of Glare, as the fibres layers are insensitive to the cyclic loading. This means that the fatigue initiation is analogous to that of monolithic aluminium. Using a similarity approach - similar stress state in the aluminium layers in Glare and in monolithic aluminium yields the same initiation life - the fatigue initiation life of Glare can be determined by calculating the actual stresses in the aluminium layers. These stresses can be calculated using the classical laminate theory (CLT) and consist of a contribution due to the applied stresses and a contribution due to the presence of residual stresses after the curing process, see Fig. 1. The residual stress state after curing consists of tensile stresses in the aluminium layers and compressive stresses in the fibre layers as a result of the mismatch in coefficients of thermal expansion.
Slam
4. Crack propagation During the crack propagation phase, the intact fibres carry a significant part of the load over the crack in the aluminium layers and restrain the crack opening. This socalled fibre bridging reduces the stress intensity at the crack tip in the aluminium layers and thus reduces the crack propagation rate considerably, although the far field stresses in the aluminium layers are higher than the far field stress in a monolithic aluminium sheet [3]. The approximately constant crack growth rate leads at the end to a much longer fatigue life in Glare compared to monolithic aluminium, see Fig. 2. Along with the crack growth in the aluminium layers, delamination growth occurs at the interfaces between the aluminium and fibre layers. This delamination growth is
SAl_tot
Slam
Slam
= SAl Sf,0 Sf,90 SAl
+ Slam
Glare
SAl_cure Sf,0_cure Sf,90_cure SAl_cure
SAl Sf,0 Sf,90 SAl
SAl_cure Sf,0_cure Sf,90_cure SAl_cure
SAl_tot
Fig. 1. Approach of calculating the stresses in the aluminium layers.
Aluminium
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R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123 50 Smax = 120 MPa R = 0.05 W = 140 mm L = 580 mm f = 10 Hz 2a0 = 5 mm
45
Aluminium 2024 T3 - 2mm 40 35 30
GLARE 4B - 4/3 - 0.5 LT
a 25 (mm) 20
GLARE 3 - 3/2 - 0.3 L
15 10 5 0 0
20
40
60
80
100
120
140
160
180
200
N (kcycles)
Fig. 2. Comparison between crack growth curves for monolithic aluminium and Glare [3].
bridging stress
crack opening
delamination shape
Fig. 4. Schematic illustration of crack opening, delamination shape and bridging stress. Fig. 3. Illustration of fibre bridging in Glare.
induced by the shear stresses that occur as a result of the load transfer from the aluminium layer to the fibre layers, which is a result of the fibre bridging, see Fig. 3. The crack opening, the delamination length and the bridging stress are interrelated as shown in Fig. 4. An increase in the delamination length gives a reduction of the bridging stress and an increase of the crack opening. During crack growth, these mechanisms are in balance with each other, resulting in the approximately constant crack growth rates. The methodology for predicting fatigue cracks in flat Glare sheets is based on the assumption that all aluminium layers in the laminate contain cracks of equal lengths, see Fig. 5.
The crack growth behaviour in Glare can be described with the stress intensity factor following from Linear Elastic Fracture Mechanics by superposition. The contribution of the crack closing bridging fibres is subtracted from the crack opening contribution of the far field stresses present in the aluminium layers [4,5] K tip ¼ K farfield K bridging
ð1Þ
With a Paris relation the crack growth rate can then be determined. da ¼ CDK ntip dN
ð2Þ
where the constants are determined based on experimental results from fatigue tests on thin aluminium sheets. The effect of fibre bridging in the stress intensity factor Kbridging is
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
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acrack
taluminium tprepreg
starter notch Fig. 5. Through the thickness crack distribution in a flat Glare panel.
calculated by integrating the fibre bridging stress over the total crack length. The bridging stress at a certain location along the crack length follows from the delamination length, the crack opening and the shear deformation of the total prepreg layer at that location. Although the mechanisms discussed above become more complicated, when the loading direction is under an angle with respect to the fibre directions of the Glare panels, the approach remains the same. The laminate properties are determined for the given off-axis angle and are used in the calculation methods. A slightly different situation is the configuration with only a surface crack (i.e., a crack through the thickness of one single aluminium surface layer), which can be initiated as a scratch caused by a sharp object. Again, the stress intensity factor at the crack tip is the far field stress intensity factor subtracted with the factor caused by fibre bridging. However, in the case of a surface crack the intact aluminium layers give an additional contribution to crack bridging, thus giving an even lower stress intensity factor and crack growth rate than for the configuration with through cracks [6]. However, as a result of the surface crack, additional bending occurs as compared to the configuration with through cracks [3]. The resultant stress state depends on the crack length and laminate thickness. For small cracks in thick laminates, the effect can be ignored. Due to the differences in bridging mechanism and secondary bending, the applied calculation method is different [7,8]. However, analogous to the configuration with cracks through the thickness, the method is based on relations for the stress intensity factor based on the bridging. Part through cracks (e.g., corner cracks at holes) can be considered as a special form of surface cracks. Each cracked layer can be viewed as a surface crack with respect to the remaining (deeper) part of the laminate [9,10].
temperature and the other is the effect of moisture absorption. The effect of moisture on the material can be divided into three main effects: • corrosion and more rapid crack growth rates in aluminium • decrease of adhesive interface strength between fibre layer and aluminium • plasticising of the matrix. The chemical bond between the adhesive and the oxide layer of the aluminium and the oxide layer itself is degraded due to moisture absorption by the matrix, which results in a decrease of the adhesive interface strength. A key parameter is the glass transition temperature Tg of the epoxy, which is the temperature at which the characteristics of the epoxy change from a glassy to that of a rubbery state. This transition temperature decreases with the amount of moisture absorbed, as is illustrated with the stiffness drop in Fig. 6. Degradation of the mechanical properties and an increase of shear deformation in fatigue crack situations result in less crack opening restraint. This fact, together with larger delamination areas resulting from reduced delamination toughness, gives less fibre bridging and thus larger stress intensities at the crack tip in the aluminium layers.
E [MPa]
5. Temperature and moisture With respect to fatigue crack initiation and propagation, two aspects are especially important with respect to the durability characteristics of Glare: one is the influence of
Tg2
Tg1
T [˚C]
Fig. 6. Reduction of the glass transition temperature as result of moisture absorption.
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As a result of the plasticising of the matrix, the glass transition will take place at lower temperatures than in a situation without moisture absorption [11]. For strength justification, this implies that if the glass transition temperature Tg is within the operational temperature range, the fatigue behaviour above Tg needs to be taken into account. For Glare, which is cured at a temperature of 120 C, the glass transition temperature Tg in dry condition is about 100 C [11], while after moisture absorption the Tg can reduce to about 80 C. Application of different adhesive systems in the prepreg of Glare with higher curing temperatures can increase the glass transition temperature both to levels above the operational temperature range. The environmental temperature has a significant effect on the fatigue behaviour of monolithic aluminium as well as Glare. An elevated temperature can have a negative effect on the fatigue crack growth behaviour of Glare, whereas a decrease in temperatures can result in a significant increase of fatigue life. The reduction of the fatigue life at elevated temperatures is mainly due to two factors. At elevated temperatures, the moisture absorption in the matrix increases. Further, elevated temperatures cause voids and microscopic defects to open and attract even more fluid. The second factor is the softening and weakening of the matrix due to the increasing temperatures. Weakening induces less crack opening restraint and less fibre bridging [12]. On the other hand, at low temperatures, the matrix becomes stiff and brittle, enhancing the crack opening restraining, resulting in a reduction in the stress intensity factor. In addition, the resistance of aluminium to fatigue crack growth increases at low temperatures, which adds
to the beneficial effect on the fatigue behaviour induced by the stiffer fibre layers. 6. Residual strength The airworthiness regulations require that the aircraft structure must be able to carry limited loads after the event of accidental damage, ignoring for this case the dynamic effects [14]. An example of a criterion to meet these requirements is the so-called two-bay crack criterion, which is related to the residual strength of a structure. The residual strength of a material is defined as the remaining static strength in case of any damage occurring during operation life. As part of a conservative design procedure, the Glare panel is assumed to contain sharp notches. Two types of damages are considered within the residual strength justification of Glare [15]: • Accidental damages • Fatigue damages The difference between the two types is related to the fibre layers in the wake of the crack. Whereas the fibre layers in a fatigue crack configuration are intact, all fibres are assumed to be broken in case of an accidental damage. Each type has its characteristic fracture mechanism. Glare panels with accidental damages loaded until fracture show stable crack extension prior to unstable rapid crack growth and failure, see Fig. 7. This means that the fracture behaviour is similar to the response of monolithic aluminium to static loading [16]. To predict the fracture behaviour and the residual strength of Glare the R-curve approach has been adopted,
Fig. 7. Crack extension versus applied load curve obtained from a residual strength test at room temperature in lab air on a Glare3 panel with an accidental damage case [15].
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
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Fig. 8. KR curves for Glare3-3/2-0.3 for different widths and starter notches, based on a compliance plastic zone correction. [15].
because the residual strength behaviour is similar to monolithic aluminium. The R-curve is treated as a material parameter valid between a given range of panel widths and starter notch lengths. The methods incorporate stable crack extension and limited amount of plasticity. An example of the R-curves for Glare 3 is given in Fig. 8. 7. Joints Major concerns in the design of pressurized aircraft fuselages are riveted longitudinal lap joints and circumferential butt joints. Each fastener hole acts as a stress raiser and there are literally thousands of them in an aircraft fuselage. Very often fatigue cracks start at about the same time at different holes, which will lead to a multiple site damage (MSD) situation. The presence of a large number of relatively small cracks will reduce the residual strength of a joint considerably, much more than a single small crack will do, and even more than one large crack with the same accumulated size of the small cracks. Furthermore, the
interaction between different cracks often leads to an increased crack growth rate. The consequence of this behaviour is either a low allowable design stress level or a small inspection interval. Glare however behaves in a completely different manner. Under the same stress level as monolithic aluminium joints, fatigue crack initiation in a Glare joint will, as mentioned before, start earlier. Fatigue crack initiation in Glare will occur in the layer with the highest stresses due to secondary bending and tensile loading. In a joint, this layer will be located at the mating surface, which is the surface in contact with the other sheet of the joint. Propagation of such cracks is a very slow process. The presence of the fibre/matrix layers prevents these cracks from propagating into the deeper aluminium layers and because the stresses in the deeper layers are lower, fatigue crack initiation in those layers will start at a (much) higher number of flights. This process leads to a crack distribution in thickness direction as shown in Fig. 9. Once a fatigue crack initiates at the mating surface of monolithic aluminium joints, a relatively
acrack
σnom+σbending
taluminium tprepreg
bore hole Fig. 9. Through the thickness crack distribution in a Glare joint.
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small part of the fatigue life is needed to propagate through the thickness and to become visible from the outside of the joint. However, in Glare the fatigue crack remains for a very long time in the layer at the mating surface, remaining invisible from the outside [13]. Because of the slow crack growth in the outer layer, the presence of intact fibres and the delay in crack initiation in the deeper layers, the reduction of residual strength during service develops at a much smaller rate than for comparable aluminium structures. The reduction of residual strength for a Glare joint as a result of slow crack propagation is schematically illustrated in Fig. 10. Certification of riveted joints in Glare requires a damage tolerance analysis of these joints. Methods for establishing
PUL Residual strength curve PLL
Crack propagation curve ac,LL ac,UL Ni
NUL
NLL
Flights
Fig. 10. Schematic crack initiation, crack propagation and residual strength behaviour of a joint in Glare [17].
the initiation life of a joint are based on available methods for joints in monolithic aluminium [18] and extended for fibre metal laminates [19]. Multiple Site Damage can be accounted for by applying stochastic approaches (Monte Carlo simulations) for crack initiation. Crack propagation in Glare such as shown that in Fig. 9 as well as the residual strength of joints can be modelled rather accurately [20]. As has been discussed, fatigue crack growth rates in Glare are relatively low. The difference in crack growth rates of aluminium and Glare is illustrated in Fig. 11 together with the corresponding decrease in residual strength. Although the first aluminium layer initiates significantly faster than monolithic aluminium, the residual strength of Glare remains higher than aluminium. This implies that when using Glare, it is even possible to select a relatively large allowable design stress level and still find an ultimate load life sufficiently large to avoid (small) inspection intervals. This ultimate load life can be well beyond the Design Service Goal of an aircraft. With this in mind, it becomes attractive to design a joint such that it will be capable to sustain ultimate load during its complete design life. In that case, minimum values (so-called B-values) instead of typical values must be used to prove ultimate strength capabilities. Fig. 12 gives an impression of initiation, crack propagation and residual strength for such a case. As long as ultimate loads can be sustained, no inspection is required. Since cracks are not visible from the outside of a joint, this can be a very beneficial side effect.
50
400
45 350 40 RS Glare3-5/4-0.3
300
250 a [mm]
30
25
200 RS 2024-T3
20 150
Residual Strength [MPa]
35
Limit load
15
CG Glare3-5/4-0.3 layer 1
CG Glare3-5/4-0.3 layer 2
100
10 CG 2024-T3
50
5
0 0
20000
40000
60000
80000
100000
120000
0 140000
N [cycles]
Fig. 11. Comparison between the crack growth (CG) of monolithic aluminium and the first two layers of Glare with the corresponding decrease in residual strength (RS).
R.C. Alderliesten, J.J. Homan / International Journal of Fatigue 28 (2006) 1116–1123
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MSD onset Typical residual strength
PUL
PLL Minimum residual strength ac,LL Propagation curve of lead crack Average crack
ac,UL Ni
DSG
NUL
NLL
Flights
Fig. 12. Initiation, crack propagation and residual strength behaviour of a joint with MSD in Glare.
8. Conclusions To reach the technology readiness of Glare for application in primary skin structures of aircraft fuselages, all the issues with respect to fatigue and damage tolerance have been investigated and covered by new developed calculation methods. The different approaches compared to monolithic aluminium have been validated and justified based on large amount of experimental results. References [1] Roebroeks G. Towards GLARE - The Development of a fatigue insensitive and damage tolerant aircraft material, PhD Thesis, Delft University of Technology, 1991. [2] Pora J, Hinrichsen J. Material and Technology Developments for the Airbus A380, in: Proceedings of the 22nd International SAMPE Europe Conference of the society for the Advancement of Materials and Process Engineering, Paris, 2001. [3] Alderliesten RC. In: Vlot A, Gunnink JW, editors. Fatigue, Fibre Metal Laminates - An introduction. Kluwer Academic Publishers; 2001. p. 155–71. [4] Alderliesten RC. Fatigue crack propagation and delamination growth in Glare, PhD Thesis, Delft University of Technology, 2005. [5] Alderliesten RC. Energy Release Rate approach for delamination in a fatigue crack configuration in Glare, Proceedings of the 21st International Congress of Theoretical and Applied Mechanics, Warsaw, 2004. [6] Alderliesten RC, Homan JJ. In: Varvani-Farahani A, Brebbia CA, editors. Fatigue Crack Growth Behaviour of Surface Cracks in Glare, Fatigue Damage of Materials - Experiments and Analysis. Southampton, Boston: WIT press; 2003. p. 213–22. [7] Gonesh KAM. Development of a fatigue crack growth model for surface cracks in GLARE, Master Thesis, Delft University of Technology, 2000.
[8] Gonesh KAM. Crack growth prediction of surface cracks in Glare, Preliminary Thesis, Delft University of Technology, 1999. [9] Randell C, van der Zwaag S. On subsurface crack growth in fibre metal laminate materials, Proceedings of the 35th INternational SAMPE Technical Conference, Dayton OH, USA, 2003. [10] Randell C. Subsurface Fatigue Crack Growth in Fibre Metal Laminates, PhD Thesis, Delft University of Technology, 2005 (title/ date may be subject to changes). [11] Borgonje B, van der Hoeven W. In: Vlot A, Gunnink JW, editors. Long-term behaviour, Fibre Metal Laminates - an introduction. Kluwer Academic Publishers; 2001. p. 155–71. [12] Beumler Th. Flying GLARE, A contribution to aircraft certification issues on strength properties of non-damaged and fatigue damaged GLARE structures, PhD Thesis, Delft University of Technology, 2004. [13] Homan JJ, Mu¨ller RPG, Pellenkoft F, de Rijck JJM. In: Vlot A, Gunnink JW, editors. Fatigue of riveted joints, Fibre Metal Laminates - an introduction. Kluwer Academic Publishers; 2001. p. 173–95. [14] Damage tolerance and fatigue evaluation of structure, Federal Aviation Regulations Part 25 – Airworthiness Standards: Transport Category airplanes, Section 25.571, Federal Aviation Administration, Department of Transportation, Washington D.C., 1997. [15] de Vries TJ. Blunt and sharp notch behaviour of Glare laminates, PhD Thesis, Delft University of Technology, 2001. [16] de Vries TJ. In: Vlot A, Gunnink JW, editors. Residual strength, Fibre Metal Laminates - an introduction. Kluwer Academic Publishers; 2001. p. 197–217. [17] Beumler Th. In: Vlot A, Gunnink JW, editors. Damage tolerance aspects, Fibre Metal Laminates - an introduction. Kluwer Academic Publishers; 2001. p. 219–33. [18] Homan JJ, Jongebreur AA. Calculation method for predicting the fatigue life of riveted joints, ICAF Symposium Proceedings (ed. Blom), 1993;175–191. [19] Homan JJ. Procedure for estimating the fatigue life in mechanically fastened joints in Glare, ICAF Symposium Proceedings (ed. J. Rouchon), 2001;941–946. [20] Alderliesten RC, Hagenbeek M, Homan JJ, Hooijmeijer PA, de Vries TJ, Vermeeren CAJR. Fatigue and Damage Tolerance of Glare. Applied Composite Materials 2003;10:223–42.