An experimental study on the fatigue performance of CFRP and BFRP repaired aluminium plates

Composite Structures 118 (2014) 589–599

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An experimental study on the fatigue performance of CFRP and BFRP repaired aluminium plates Jarkko Aakkula ⇑, Olli Saarela Aalto University, School of Engineering, Department of Applied Mechanics, P.O. Box 14300, FI-00076 Aalto, Finland

a r t i c l e

i n f o

Article history: Available online 19 August 2014 Keywords: Aluminium bonding Composite patch repair Crack growth Fatigue testing Thermal residual stresses

a b s t r a c t Cracked aluminium plates were repaired with multidirectional carbon/epoxy and boron/epoxy reinforcements. Aircraft repair conditions were simulated by using a steel bonding rig that constrained the effective thermal expansion to the level measured from real locally heated aircraft structures. Residual thermal stresses were varied using the rig and by varying the curing and testing temperatures. The repairs were tested using a variable amplitude loading (R = 0.14) with anti-buckling edge supports. The residual thermal stresses had a rectilinear and distinct effect on the crack growth rate. The structural support against bending had a significant effect on the fatigue life of the repairs. The fatigue life improvement factor with the centre-cracked single-sided repairs was greatest with the wet-laminated carbon/ epoxy repairs having a stiffness ratio of 1.04. With edge-cracked specimens the longest life was achieved with the double-sided boron/epoxy repairs having a stiffness ratio of 0.65. The cracked aluminium plate 2024-T3 clad had the longest fatigue life, followed by the S07-1020 and 7075-T6 clad aluminium grades. Ó 2014 Elsevier Ltd. All rights reserved.

1. Introduction The repair of cracked aircraft structures with adhesively bonded composite reinforcements has been proven to be effective by numerous studies and applications over decades [1–22]. The boltless repairs have several advantages and they can be used for temporary repairs and permanent crack or corrosion damage repairs. The efficiency of composite repairs of cracked aluminium structures depends on several factors. These factors include the repair materials, the adhesive curing temperatures, the effective thermal expansion during the curing, the type of fatigue loading, and the structural support against the secondary bending of single-sided repairs. Some of these factors may be difficult to simulate with plate specimens in laboratory testing. The main problems with the bonded repairs are the surface preparation of the metal and the residual thermal stresses with the composite reinforcements. Silane-based surface preparation methods are commonly applied. Their effectiveness has been previously reported by the authors [23]. The residual thermal stresses result from the mismatch of the coefficients of the thermal expansion (CTE) between the repaired aluminium plate and the composite reinforcement and from the temperature difference between the curing and operating ⇑ Corresponding author. Tel.: +358 505486842. E-mail address: Jarkko.aakkula@aalto.fi (J. Aakkula). http://dx.doi.org/10.1016/j.compstruct.2014.07.050 0263-8223/Ó 2014 Elsevier Ltd. All rights reserved.

temperatures (DT). The effect has been studied (or at least discussed) in almost all crack-patching investigations [1–22]. However, none of them have applied a bonding rig (as presented in this investigation) in order to constrain the thermal expansion of the repaired plate during the adhesive curing. Most authors have used unidirectional boron/epoxy (b/e) prepreg reinforcements. Only Sandow and Cannon [4] have tested multidirectional b/e patches. The unidirectional b/e plies have a high tensile stiffness (three times the aluminium stiffness), and their CTE in the fibre direction is one-fourth of the aluminium CTE. Carbon/epoxy (c/e) prepreg patches have been used in some repair studies [6–11]. The c/e patches can have a tensile stiffness ranging from the aluminium stiffness to the b/e patch stiffness. Usually it is in the middle of these two values. The CTE of the unidirectional c/e patches will cause more problems, since it is practically zero. The wet-laminated c/e repairs have been studied mainly for the repair of secondary structures [8]. GLARE material is made of aluminium and glass/epoxy plies. Consequently, the CTE of the GLARE is close to the CTE of aluminium but the stiffness is lower than the aluminium stiffness. The GLARE patches have been tested in Delft [18,19]. The wet laminated glass/epoxy (g/e) reinforcements have been used mostly in secondary structures or in non-structural repairs. The g/e-reinforcements have been tested by Hosseini-Toudeshky [20] and by Rao [21].

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Most authors have used the FM73 or other adhesive films cured at 120 °C for bonding the patch [1–8,10–11,13–21]. Since the operating temperature is usually considered to be room temperature (RT) or lower, the residual stresses can be remarkable. In some studies, the authors have tested lower temperatures for curing the adhesive films [1,7,22]. Some authors have also used two-part paste adhesives cured at lower temperatures in order to reduce the residual thermal stresses [4,7,9,12]. Single-sided composite repairs have been found to be efficient for the repair of thin (less than 3.2 mm) structures [1–9,7–22]. Repairs of thicker aluminium structures from 6 to 10 mm have also been analysed and tested [10–16]. A limited number of tests have been done with medium (3.2 mm) thick plates [1–5,11,14,22]. Even fewer authors have compared single- and double-sided repairs [4–5,8,16–17]. The fatigue testing in most studies has been done with constant amplitude (CA) loading and a positive stress ratio R [1–10,12–22]. Vlot [19] has also used a filtered flight spectrum without any compressive loads. Only a few authors have used a simulated flight spectrum loading containing compressive loads [2–6]. In most investigations, the stress ratio R of the fatigue loading has been positive and no structural support has been used. However, with a single-sided repair, the neutral axis of the repaired structure will shift, and the secondary bending effects on the fatigue life of an unsupported structure can be significant. Some authors have used a sandwich support during the fatigue loading in order to eliminate the bending effects [1–3,5,7–8]. Sandow [4] and Kan [11] have used an anti-buckling support with edge rails. It has been standard practice since the beginning of the crackpatching technique to not use stop-drilling of the crack tip before the composite repairs [1–3]. Therefore, the effect of stop-drilling has been tested in only a few papers [7,21]. The reasons for omitting the stop-drilling are well known, but the simplified assumption based on the plastic zone in front of the crack tip is applicable only for a constant amplitude loading and does not take into account the compressive loads’ interaction effects [24]. The first objective of this study was to develop a representative laboratory simulation for bonded aluminium aircraft repairs with plate specimens. The surrounding aircraft structure will restrict the free displacement of the repair area in two ways. During the adhesive curing, the real aircraft structure will constrain thermal expansion of the heated area and thus affect the residual thermal stresses. After the repair, the aircraft structure will also support the area against secondary bending during the fatigue loading and thus affect the fatigue life. For the first issue we designed a bonding rig made of steel in order to constrain the effective thermal expansion during the adhesive curing to the same level as experienced in the real locally heated aircraft structures. For the second issue we used edge support rails during the fatigue loading. The second objective was to test the effect of residual thermal stresses using the designed rig and to test the efficiency of different repair materials with typical aircraft aluminium grades and with typical aircraft repair configurations. We used thin centre-cracked aluminium plates when we simulated the single-sided aircraft skin repairs. Typical stringer and frame structures were simulated with the narrower but thicker edge-cracked plates. Both single-sided and double-sided b/e repairs were tested with these edge-cracked specimens. In this investigation, all repaired cracks were stopdrilled before the repair because it was found to be beneficial in these cases [25]. The effects of stop-drilling the cracks will be investigated further in a separate study. The testing was started with clad aluminium 2024-T3 [26], since historically it has been used in skin plates prone to fatigue cracking. However, current aircraft operated by the Finnish Air Force (FINAF) use different aluminium grades. Therefore, it was

found necessary to compare the results achieved with 2024-T3 aluminium plates to the results achieved with these other grades. Typical prepreg and wet-laminated repair configurations were tested with the developed approach. Multidirectional patches were used in this investigation, because we wanted the patches to withstand the bi-directional loads that can occur in real-life applications. Also, as observed by Sandow and Cannon [4], multidirectional patches may withstand spectrum loading with compressive loads better than unidirectional patches. The repair materials included c/e prepreg, b/e prepreg and wet-laminated c/ e (wet c/e) reinforcements. We varied the residual thermal stresses in the repairs by using the steel bonding rig, by varying adhesive and resin curing temperatures and by varying the testing temperatures. A variable amplitude (VA) loading with tensile and compressive peaks was selected in order to realistically simulate the flight loading and to include the load interaction effects on fatigue crack growth. Because of the single-sided repairs and the compressive loads, a structural support against bending and buckling was used. The results achieved with the edge-supported specimens were compared to the results measured with the sandwich-supported and the unsupported specimens. 2. Materials and processes 2.1. Aluminium plates We did most of the testing with 1.6 mm thick 2024-T3 clad aluminium sheets. For medium thickness specimens we selected 3.2 mm thick French aluminium AU4G1 clad (equivalent to ASTM 2024-T3 [26]) due to its availability. Clad was not removed from the specimens prior to surface treatment. A comparison between different aircraft aluminium grades was accomplished with the 1.6 mm thick plates. The aluminium grades compared were BAE Military Systems S07-1020 (equivalent to ASTM 2014A-T4 [27]) used in the Hawk and ASTM 7075-T6 clad used in the F-18. The crack growth rate in the aluminium plates can be estimated using Paris’ law [28]:

da ¼ CðDKÞn dN

ð1Þ

where da/dN is the crack growth, DK is the variation of the stress intensity factor and C and n are empirically defined material constants. Young’s moduli, ultimate tensile strengths, Paris’ law material constants, fracture toughness and coefficients of thermal expansion of the used aluminium plates are given in Table 1. 2.2. Reinforcement materials Unidirectional c/e tape prepreg, balanced c/e fabric prepreg and unidirectional b/e tape prepreg were used in the pre-cured patches. Oriented and balanced carbon fabrics were used with epoxy resin in the wet-laminated patches. We used multidirectional patches in this study because we wanted the patch to withstand the occasional bi-axial loads, shear loads and compressive loads that can occur in real life structures. The 90° layers were placed in the middle of the laminate. The ±45° layers were placed on the top of the patch. This kind of unsymmetrical laminate has a coupling between tension and bending. This small bending effect will act against the secondary bending caused by the neutral axis offset in the repaired area. The c/e patches had g/e fabric layers under and over the carbon layers. The function of the bottom glass layer was to prevent the galvanic corrosion between the aluminium and the carbon

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J. Aakkula, O. Saarela / Composite Structures 118 (2014) 589–599 Table 1 Young’s moduli, ultimate tensile strengths, Paris’ law material constants, fracture toughness and coefficients of thermal expansion of the used aluminium plates [26–32].

a b c

Aluminium grade

EP [GPa]

rtu [MPa]

C  1012

n

pffiffiffiffiffi Kcr [MPa m]a

aP ½106 =  Cb

2024-T6 clad S07-1020c 7075-T6 clad

73.1 73.1 71.7

434 427 510

1.83 43.1 15.3

3.09 2.80 2.95

34 24 29

23.2 23.0 23.6

Average bare material fracture toughness. Average over 23–120 °C range. Equivalent to 2014A-T4.

materials. The idea of the top glass layer was to protect the carbon layers against foreign object damage and abrasion. The c/e prepreg patch for the single-sided centre-crack repairs had five unidirectional layers of Hercules AS4/8552 3501-6 tape and one balanced layer of Hercules AW-193/8552 fabric. The fibre orientations of the carbon layers were 0°, 0°, 90°, 0°, 0°, ±45°. The orientations of the balanced g/e prepreg fabric layers were 0°/ 90°. The prepreg patches were cured two hours at 175 °C under 4 bar pressure. The thicknesses of the carbon patches were 1.2 mm. The patches are summarized in Table 2 and shown in Fig. 2. The given thicknesses and Young’s moduli are measured values. Thermal expansion coefficient values were calculated with the ESAComp software using the typical ply values [33]. The b/e patches for the single-sided centre-crack repairs and for the double-sided edge-crack repairs had six unidirectional layers of Textron 5521 prepreg. The fibre orientations of the boron layers were 0°, 0°, 90°, 0°, +45°, 45°. The b/e-patches for the single-sided edge-crack repairs had eight unidirectional layers of Textron 5521 prepreg. The fibre orientations of the boron layers were 0°, 0°, 0°, 90°, 0°, 0°, +45°, 45°. The prepreg patches were cured one hour at 120 °C under 4 bar pressure. The thicknesses of the cured

centre- and edge-crack patches were 0.78 and 1.04 mm, respectively. The patches are summarized in Table 2 and shown in Figs. 2 and 3. The co-bonded wet-laminated carbon patches had a balanced Interglass 92110 glass fabric layer between the aluminium plate and the carbon layer and also on top of the patch. The wet c/e patch had four layers of oriented Lyvertex G808 fabrics and two layers of balanced Interglass 98151 fabrics. The fibre orientations of the carbon layers were 0°, 0°, 0°/90°, 0°, 0°, ±45°. SP 690 laminating resin was used. Different curing temperatures were used in order to tailor the residual thermal stresses. The thicknesses of the wet-laminated carbon patches were 2.1 mm. The wet c/e-patch properties are summarized in Table 2. The stiffness ratios of the repairs were calculated with an equation given by Rose [1]:

S¼

tR ER tP EP

ð2Þ

where tR and ER are the thickness and Young’s modulus of the reinforcement and tP and EP are the thickness and Young’s modulus of the plate. The stiffness ratios for each repair are given in Table 2.

Table 2 Details of the composite reinforcements. Reinforcement

aR

0°/90° 0° 0° 90° 0° 0° ±45° 0°/90°

1.2

1.6

84.0

0.86

tape tape tape tape tape tape

0° 0° 90° 0° +45° 45°

0.78

6.0

94.6

0.65a

UD UD UD UD UD UD UD UD

tape tape tape tape tape tape tape tape

0° 0° 0° 90° 0° 0° +45° 45°

1.04

5.4

124.6

0.53

Interglass 92110 fabric G808 UD tape G808 UD tape Interglass 98151 G808 UD tape G808 UD tape Interglass 98151 Interglass 92110 fabric

glass

0°/90°

2.1

1.4

58.2

1.04

Material

c/e Prepreg for single-sided centre-crack repair

1 2 3 4 5 6 7 8

G120/8552 glass fabric AS4/8552 UD tape AS4/8552 UD tape AS4/8552 UD tape AS4/8552 UD tape AS4/8552 UD tape AW-193/8552 fabric G120/8552 glass fabric

b/e Prepreg for single-sided centre-crack and double-sided edge-crack repair

1 2 3 4 5 6

Textron Textron Textron Textron Textron Textron

5521 5521 5521 5521 5521 5521

UD UD UD UD UD UD

b/e Prepreg for single-sided edge-crack repair

1 2 3 4 5 6 7 8

Textron Textron Textron Textron Textron Textron Textron Textron

5521 5521 5521 5521 5521 5521 5521 5521

Wet c/e for single-sided centre-crack repair

1 2 3 4 5 6 7 8

a

tR [mm]

Ply number

For both single-sided and double-sided repairs.

Orientation

fabric

fabric glass

0° 0° 0°/90° 0° 0° ±45° 0°/90°

½106 =  C

ER [GPa]

S

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Fig. 1. Steel rig for controlling effective thermal expansion of the aluminium plates during the adhesive curing.

Fig. 3. Single-sided or double-sided edge-cracked aluminium plate repair.

3. Testing methods 3.1. Bonding rig A bonding rig was developed in order to control thermal expansion of the repaired plates during the elevated temperature adhesive curing. The purpose of the rig was to simulate a typical locally heated aircraft structure surrounded by cool structures. The rig restrains the expansion of the aluminium plate. The expansion coefficient was aimed at representing the typical effective thermal expansion attained during an on-the-aircraft repair. The bonding rig was constructed from steel having a thermal expansion coefficient of asteel ¼ 11:9  106 = C. The aluminium plates were tightened in-between the steel edges as shown in Fig. 1. The edges overlapped the aluminium plate 50 mm from the short edges and 30 mm from the long edges. The whole rig was vacuum-bagged and placed in an oven for curing. 3.2. Centre-crack specimens Fig. 2. Single-sided centre-cracked aluminium plate repair.

2.3. Surface treatment methods The aluminium surfaces were treated with the grit blast silane (GBS) process [34]. The repairs were bonded or wet- laminated in place within 60 min of the treatment. No primer was used. 2.4. Adhesives and bonding process Two layers of 290 g/m2 FM73M film were used as an adhesive with all prepreg repairs. The standard cure in this study for the adhesive film was 2 h at 120 °C in the bonding rig and in a vacuum bag. Alternative cure cycles were 8 h at 90 °C or at 70 °C in the bonding rig. 0.7 bar vacuum pressure was used. SP690 was used as an adhesive and as a laminating resin with the co-bonded wet c/e repairs. The resin was cured at 120 °C for 5 h in the bonding rig or alternatively pre-cured at RT for 16 h and post-cured at 120 °C in an oven for 5 h in the bonding rig. 0.5 bar vacuum pressure was used.

The centre-crack specimens were manufactured from 1.6 mm thick clad aluminium to 280  200 mm plates (see Fig. 2). A 3.0 mm hole was drilled in the centre of the plate. Two 1.9 mm long notches on each side of the hole were then cut using an electric discharge machining method. The centre cracks were fatigued to the specimens from the cuts to the size a = 10 mm. After cycling the cracks, 3.0 mm diameter holes were drilled through the crack tips. Pre-cured c/e and b/e prepreg patches were bonded over the cracks after the surface treatment using the FM73M adhesive film. Wet-laminated c/e patches were laminated on the treated surface using the laminating resin. The bonding rig described in Section 3.1 was used in most cases in order to control the thermal expansion of the aluminium plates. Normally three plates were repaired and tested in each set. An exception was the testing at 10 °C, where only one specimen was tested. In some cases, one or two of the plates failed outside the repair. In some of those cases, we repeated the tests, and up to six plates could be repaired and tested. The numbers for the successful test results are given in the result tables for each set.

J. Aakkula, O. Saarela / Composite Structures 118 (2014) 589–599

One set of c/e prepreg repairs and one set of wet-laminated c/e repairs were environmentally conditioned. The repaired plates were immersed in 50 °C tap water for 30 days. The residual thermal stresses of the repaired aluminium plates were calculated using an equation given by Baker [1]:

rT ¼

t R ER EP DTðaP eff  aR Þ t P EP þ t R ER

ð3Þ

where tR and ER are the thickness and Young’s modulus of the reinforcement, tP and EP are the thickness and Young’s modulus of the plate, DT is the temperature difference between the curing and testing temperatures, aP eff is the effective thermal expansion coefficient of the aluminium plate during the cure, and aR is the thermal expansion coefficient of the reinforcement. 3.3. Edge-crack specimens The edge-crack specimens were manufactured from 3.2 mm thick clad aluminium to 280  100 mm plates (see Fig. 3). The edge-cracks were fatigued to the specimens as with the centrecrack specimens. Also, 3.0 mm holes were drilled through the crack tips. The pre-cured b/e prepreg patches were bonded over the cracks after the surface treatment using the FM73 adhesive film with the standard cure procedure in the bonding rig. Both single-sided and double-sided repairs were used.

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plate was a window for visual inspection of the crack length of the single-side repaired plates. The edge support rig is shown in Fig. 5. The sandwich support was provided by forming a sandwich structure from two repaired plates and honeycomb core. The plates were first repaired with the standard c/e prepreg repair procedure. It means that the plates had the same residual thermal stresses as the standard repairs. The repaired plates were bonded onto both sides of a 10 mm thick aluminium honeycomb core with the FM73 adhesive film at 120 °C. When the assembly cooled down, the bending normally caused by the residual stresses was eliminated. With the single-sided repairs, the crack length was measured visually from a polished unrepaired surface using a 10X magnifying dioptre. For inspection, the loading was paused after predetermined loading cycles to a tensile load peak. The accuracy of the visual crack length measurements was ±1 mm. With the double-sided repairs and with the sandwich-supported repairs, the crack length was measured with a Hocking eddy current tester. The accuracy of the eddy current measurements was at the same level as with the visual measurements. Normally the repairs were fatigue-tested at RT. Some c/e repairs were tested, however, at elevated or lower temperatures in order to change the magnitude of the residual thermal stresses. Two environmentally conditioned repairs were also tested in a hot/ wet environment. The testing temperature was 70 °C and the RH was between 70% and 100%.

3.4. Fatigue testing Variable amplitude (VA) loading was selected to simulate the actual loading spectrum of a typical aircraft repair. With the selected VA loading, the overload and underload interaction to the fatigue crack growth rate was included [24]. One flight from the Short Falstaff spectrum with 130 peaks was used [4]. The maximum peak was 140 MPa and the minimum peak was 20 MPa. A frequency of 10 Hz was used. The variable amplitude loading spectrum is shown in Fig. 4. Three different support cases were used during the fatigue testing. The standard case was testing with an edge support. For comparison, one set of specimens was tested without any support and one set with a sandwich structure support. The edge support against bending was provided on both sides of the repair and on both sides of the plate with four 40 mm wide PTFE rails which were bolted to 10 mm thick steel plates. The steel plates were also bolted together. In the middle of the front steel

Fig. 4. Variable amplitude loading spectrum used in the tests.

Fig. 5. Edge support rig consisting of a foot (1), filler plate (2), holding block (3), hinge band (4), anti-buckling plates (5), and PTFE rails (6).

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4. Results 4.1. Thermal expansion coefficients and residual thermal stresses The effective CTE of an aluminium plate repaired in the steel rig and placed in a vacuum bag was measured with strain gages. The measured value was aP eff ¼ 16:3  106 = C. We made a reference measurement from a locally heated Hawk aircraft rear fuselage structure. The result was aP eff ¼ 15:1 106 = C. The CTE of a locally heated structure was also approximated using FEM. Simple 800  800 mm size NISA [35] and IDEAS [36] aluminium plate models were constructed using two-dimensional shell elements. The models had a locally heated 300  300 mm central area. The models gave values between aP eff ¼ 11:2  14:8 106 = C. 4.2. Fatigue test results Results of the fatigue tests are given below. Tables 3–8 provide the number of test specimens that failed from the repair area and the average values and standard deviations of fatigue lives for these specimens. The life improvement factors (LIF) were calculated from the average values. Therefore, the fatigue lives given in the tables are slightly different from the number of cycles shown in Figs. 6–11. In general, the results shown in the figures correspond to the mean crack growth curve of three or more test specimens. In some cases, where some of the plates failed outside the repair area, the most representative fatigue crack growth curve is shown. The desired failure mode in the test specimens was the fatigue crack growth in the aluminium plate. After the crack reached both edges of the plate, the c/e patches usually failed cohesively between the bottom glass layer and the carbon layers. With the b/e prepreg repairs the adhesive layer or composite layers did not fail cohesively. Instead, the boron fibres of the patch snapped after the plate had cracked completely. 4.2.1. Crack growth with different patches The efficiency of different patch materials was tested with centre-cracked 1.6 mm thick 2024-T3 clad aluminium plates. The plates were repaired in the bonding rig with the single-sided c/e prepreg, b/e prepreg and wet-laminated c/e reinforcements. The average fatigue lives are given in Table 3 and the crack growth curves are shown in Fig. 6. The average fatigue life and the crack growth rate of the unrepaired plate are shown as a reference. The failure mode was in all cases the fatigue crack growth in the aluminium plate. 4.2.2. Crack growth with different supports The effect of different supports was tested with centre-cracked 1.6 mm thick 2024-T3 clad aluminium plates repaired in the bonding rig with the single-sided c/e prepreg reinforcements. The

support conditions during the testing were no support, edge support with PTFE rails, and sandwich support. The average fatigue lives are given in Table 4 and the mean crack growth curves are shown in Fig. 7. The failure mode was in all cases the fatigue crack growth in the aluminium plate. With the sandwich support specimens, testing was stopped when the crack on either side of the sandwich panel reached both edges of the plate. 4.2.3. Crack growth with c/e prepreg patch and different residual thermal stresses The effect of residual thermal stresses on c/e prepreg repairs was tested with centre-cracked 1.6 mm thick 2024-T3 clad aluminium plates repaired with the single-sided reinforcements. The reinforcements were bonded at 120 °C, at 90 °C and at 70 °C in the steel rig. One set was bonded at 120 °C without the steel rig for comparison. One test series was conditioned by immersion, as described in Section 3.2. The repairs were tested at different temperatures. The conditions during the curing and during the testing resulted in different residual thermal stresses. The average fatigue lives and residual thermal stresses are given in Table 5 and the mean crack growth curves are shown in Fig. 8. The failure mode was in most cases the fatigue crack growth in the plate, but with two exceptions. The environmentally conditioned wet-laminated repairs had corrosion pits in the aluminium plate outside of the repair. New cracks initiated from those pits and they grew faster than the centre-crack under the patch. With the repair cured at 70 °C, the crack growth was slow until the crack reached the edges of the repair. The final failure of the repair after that was partly cohesive in the bonding adhesive and partly adhesive between the adhesive film and the patch. 4.2.4. Crack growth with wet-laminated c/e patch and different residual thermal stresses The effect of residual thermal stresses on wet-laminated c/e repairs was tested with centre-cracked 1.6 mm thick 2024-T3 clad aluminium plates repaired with the single-sided reinforcements. The repairs were cured at 120 °C in the steel rig. The steel rig was also used during the RT cure followed by post-curing at 120 °C. One test series was conditioned by immersion, as described in Section 3.2. The repaired plates were tested at RT or at elevated temperatures. These conditions during the curing and during the testing resulted in different residual thermal stresses. The average fatigue lives and residual thermal stresses are given in Table 6 and the mean crack growth curves are shown in Fig. 9. The failure mode of two standard cure and RT tested repairs was the fatigue crack growth in the plate. With other repairs the centre-crack growth was very slow and the specimens failed from new cracks outside the repair area. With the environmentally conditioned repairs tested at 70 °C, new cracks initiated from corrosion pits outside of the repair. Also in the RT-cured repairs tested

Table 3 Number of tested plates, average fatigue lives, standard deviations of the fatigue lives, life improvement factors and residual thermal stresses of the centre-cracked 2024-T3 clad aluminium plates repaired with c/e prepreg, b/e prepreg and wet-laminated c/e single-sided reinforcements in the bonding rig and tested with edge supports. Series

Unrepaired c/e Prepreg b/e Prepreg c/e Wet lam. a b

No. of plates

3 5 2b 2b

Average fatigue life in cycles

153,759 483,520 795,176 984,348

Referenced to the unrepaired plate. Third plate failed from a crack initiated outside the repair area.

Standard deviation of fatigue lives In cycles

In%

37,433 31,501 244,803 22,725

24 6.5 31 2.3

Life improvement factora

rT in MPa (Eq. (3))

n/a 3.3 5.2 6.1

n/a 48 27 53

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Table 4 Number of tested plates, average fatigue lives, standard deviations of the fatigue lives, life improvement factors and the residual thermal stresses of the centre-cracked 2024-T3 clad aluminium plates repaired with the single-sided c/e prepreg reinforcements in the bonding rig and tested with no support, edge support and sandwich support. Series

Unrepaired No support Edge support Sandwich support a

No. of plates

3 3 5 3

Average fatigue life in cycles

153,759 322,099 48,3520 1,123,481

Standard deviation of fatigue lives In cycles

In%

37,433 20,978 31,501 73059

24 6.5 6.5 6.5

Life improvement factora

rT in MPa (Eq. (3))

n/a 2.1 3.3 7.3

n/a 48 48 48

Referenced to the unrepaired plate.

Table 5 Number of tested plates, average fatigue lives, standard deviations of the fatigue lives, life improvement factors and residual thermal stresses of the centre-cracked 2024-T3 clad aluminium plates repaired with the single-sided c/e prepreg reinforcements bonded at different curing conditions and tested at different temperatures with the edge support. Series

Unrepaired No bonding rig, tested at RT Std. cure, tested at -10 °C Std. cure, tested at RT 90 °C cure, tested at RT 70 °C cure, tested at RT Std. cure, tested at 70 °C Std. cure, tested at 70 °C/wet a b c

No. of plates

3 6 1 5 3 2b 3 c

Average fatigue life in cycles

153,759 476,571 365,756 483,520 589,801 528,292 603,395 434,461

Standard deviation of fatigue lives In cycles

In%

37,433 51,010 n/a 31,501 85,484 53,476 24,188 51,661

24 10 n/a 6.5 14 10 4.1 12

Life improvement factora

rT in MPa (Eq. (3))

n/a 3.1 2.4 3.3 3.8 3.4 3.9 n/a

n/a 73 64 48 33 23 25 25

Referenced to centre-crack failure of unrepaired plate. Third plate failed due poor adhesion. Repairs immersed in 50 °C water for 30 days before testing; all three plates failed from new cracks initiated from corrosion pits outside repair area.

Table 6 Number of tested plates, average fatigue lives, standard deviations of the fatigue lives, life improvement factors, and residual thermal stresses of the centre-cracked 2024-T3 clad aluminium plates repaired with the single-sided wet-laminated c/e reinforcements at different curing temperatures and tested at different temperatures with edge supports. Series

Unrepaired Std. cure, tested at RT Std. cure, tested at 70 °C/wet RT cure, tested at RT RT cure, tested at 70 °C

No. of plates

3 2b c d e

Average fatigue life in cycles

153,759 968,279 n/a n/a n/a

Standard deviation of fatigue life In cycles

In%

37,433 22,725 n/a n/a n/a

24 2.3 n/a n/a n/a

Life improvement factora

rT in MPa (Eq. (3))

n/a 6.1 n/a n/a 1

n/a 53 27 0 26

a

Referenced to unrepaired plate. Third plate failed from a new crack initiated outside the repair area. c Repair immersed in 50 °C water for 30 days before testing, centre-crack grew in two plates, but all three plates failed from the cracks initiated from corrosion pits outside repair area. d Centre-crack grew in two plates, but all three plates failed from the cracks initiated outside the repair area. e No crack growth in the centre-crack, all three plates failed from the cracks initiated outside the repair area. b

Table 7 Number of tested plates, average fatigue lives, standard deviations of the fatigue lives, life improvement factors and residual thermal stresses of the edge-cracked AU4G1 clad aluminium plates repaired with the single-sided and with the double-sided b/e prepreg reinforcements in the bonding rig and tested at RT with no supports. Series

Unrepaired Single-sided b/e patch Double-sided b/e patch a

No. of plates

3 3 1a

Average life

56,859 121,225 4 352 466

Standard deviation of fatigue lives In cycles

In%

20,447 9966 n/a

36 8.2 n/a

Life improvement factor

rT in MPa (Eq. (3))

n/a 1.9 71

n/a 26.7 27.2

Two other plates failed from new cracks initiated outside the repair area.

at RT new cracks initiated from the outer edges of the reinforcements or next to the end tabs. The new cracks grew faster than the centre-crack and resulted in the final failure of the plate. With RT-cured repairs tested at 70 °C, there was no growth in the centrecrack, and all specimens failed from new cracks initiated on the edges of the reinforcements.

4.2.5. Crack growth of edge-cracked specimens The edge-cracked specimens consisted of 3.2 mm thick AU4G clad aluminium plates repaired in the bonding rig with the single-sided or the double-sided b/e prepreg reinforcements. The plates were tested without the edge support rails. The average fatigue lives are given in Table 7 and the representative crack growth

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Table 8 Number of tested plates, life improvement factors, standard deviations of the fatigue lives and residual thermal stresses of the centre-cracked 2024-T3 clad, S07-1020 and 7075T6 clad aluminium plates repaired with the single-sided c/e prepreg reinforcements in the bonding rig and tested at RT with edge supports. Series

Unrepaired 2024-T3 c/e Prepreg on 2024-T3 Unrepaired S07-1020 c/e Prepreg on S07-1020 Unrepaired 7075-T6 c/e Prepreg on 7075-T6 a

No. of plates

3 5 3 4 2a 3

Average life in cycles

153,759 483,520 150,676 475,086 41,399 245,008

Standard deviation of fatigue lives In cycles

In%

37,433 31,501 25,715 136,482 3096 24,457

24 6.5 17 29 7.5 10

Life improvement factor

rT in MPa (Eq. (3))

n/a 3.3 n/a 3.2 n/a 5.9

n/a 48.2 n/a 48.1 n/a 47.7

One plate failed from the end tab area.

Fig. 6. Crack growth in the centre-cracked 2024-T3 clad aluminium plates repaired with c/e prepreg, b/e prepreg and wet-laminated c/e single-sided reinforcements in the bonding rig and tested with edge supports.

Fig. 7. Crack growth in the centre-cracked 2024-T3 clad aluminium plates repaired with the single-sided c/e prepreg reinforcements in the bonding rig and tested with no support, edge support, and sandwich support.

curves are shown in Fig. 10. The crack growth rate of the unrepaired plate is shown as a reference. In single-sided repairs the failure mode of the repairs was the fatigue crack growth of the plate. Two double-sided repairs failed from new cracks initiated next to the end tabs. With the remaining double-sided repair, the crack growth in the beginning of the testing was minimal and was not detected. After 4.17 million cycles, a

Fig. 8. Crack growth in the centre-cracked 2024-T3 clad aluminium plates repaired with the single-sided c/e prepreg reinforcements bonded at different curing conditions and tested at different temperatures with the edge support.

Fig. 9. Crack growth in the centre-cracked 2024-T3 clad aluminium plates repaired with the single-sided wet-laminated c/e reinforcements at different curing temperatures and tested at different temperatures with edge supports.

50 mm crack length was measured and the final crack growth outside the patch still took 180 000 cycles. 4.2.6. Crack growth of different aluminium grades The crack growths in different aluminium grades were tested with 1.6 mm thick 2024-T3 clad, S07-1020 (2014) and 7075-T6 clad aluminium plates. The centre-cracked specimens were repaired with single-sided c/e prepreg reinforcements. The average

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The CTE and residual thermal stresses of the c/e prepreg repairs used in this investigation were also calculated using an analytical model and FEM by Varis [39]. For the repair cured in the rig, the analytical model gave the value aP eff ¼ 13:1  106 = C, which is close to the values achieved in this study (see Section 4.1). For the repair cured at 120 °C without the rig the FEM calculations gave the value rT = 75 MPa [39] while Eq. (3) gave the value rT = 73 MPa. 5.2. Fatigue testing and the relevance of the results

Fig. 10. Crack growth in the edge-cracked AU4G1 clad aluminium plates repaired with the single-sided and with the double-sided b/e prepreg reinforcements in the bonding rig and tested at RT with no supports.

Fig. 11. Crack growth in the centre-cracked 2024-T3 clad, S07-1020 and 7075-T6 clad aluminium plates repaired with the single-sided c/e prepreg reinforcements in the bonding rig and tested at RT with edge supports.

In many investigations it has been suggested that the stiffness ratio of composite repairs, calculated with Eq. (2), should be between 1 and 1.2, or even 1.6 [40–42]. In this study, the prepreg reinforcements used had lower stiffness than the repaired plate. Stiffness ratios of the single-sided c/e and the b/e prepreg reinforcements were 0.86 and 0.65, respectively. This was considered acceptable since we wanted the crack to grow during the fatigue testing. The wet-laminated reinforcement had the tensional stiffness slightly above the plate stiffness, the stiffness ratio being 1.04. No adhesive failures were observed either with the film adhesives or the resins from the aluminium surface. The silane surface treatment without a primer was thus satisfactory, as also confirmed in a separate investigation [23]. The repairs were not tested with NDI for local disbonds between the adhesive and the patch or between the adhesive and the plate. In any case, the patches stayed bonded to the final fracture of the cracked aluminium plate. After the crack had reached both edges of the plate, the c/e patches disbonded between the bottom glass fabric ply and the first c/e ply. With the b/e reinforcements the final failure mode was the breaking of the boron fibres. The crack lengths were measured after predetermined cycles. With the first specimens, we sometimes defined too long interval and consequently had only a few crack length measurements before the final failure. It was impossible to make direct comparisons of the test results against the results published by other authors, since the repair and loading configurations were always more or less different. For example, no-one has used stop-drilling combined with edge support and VA loading with compressive peaks. Usually life improvement factors measured in VA loading were lower than those measured in CA loading [11]. This was to be expected because the compressive loads in the used VA loading will accelerate the fatigue crack growth and shorten the life [24]. 5.3. Efficiency of the patches

fatigue lives are given in Table 8, and the mean crack growth curves are shown in Fig. 11. The crack growth rates of the unrepaired plates are shown as a reference. 5. Discussion 5.1. Bonding rig and residual thermal stresses The steel bonding rig provided a suitable constraint to the thermal expansion of the aluminium plate during the elevated temperature curing. The measured effective CTE was in good agreement with the value measured from the locally heated aluminium aircraft structure. Some amount of residual bending was observed with the repairs cured over 100 °C. Several authors have used Eq. (3) in estimating residual thermal stresses [1–3,8,37–38]. They have also shown that the values calculated with Eq. (3) are in good agreement with the measured and numerically computed values. In this investigation, the values used in Eq. (3) were based on several measurements (tR, ER, tP, DT, aP eff ), on one calculated value (aR) and on one literature value (EP). Therefore, we assume that the rT values calculated here are relevant.

All repairs remarkably increased the fatigue life. With the standard cured c/e prepreg patches tested with edge supports, the fatigue life improvement factor (LIF) was 3.3. The single-sided b/e prepreg patches provided a LIF of 5.2. The wet-laminated cobonded c/e patch with higher stiffness ratio resulted in a LIF of 6.1. The c/e repairs had low standard deviations in the fatigue lives while the b/e repairs had a high deviation. The closest reference we found to our c/e repairs was for a single-sided unidirectional c/e repair of a centre-cracked 3.1 mm thick aluminium plated under VA loading and with edge supports [11]. Kan and Ratwani [11] measured the life improvement factor of 3.7. The stiffness ratio was not given, but we assumed it was close to 1, while we had 0.86. The LIF from reference [11] was, anyhow, close to the LIF achieved in our study. In the referenced papers, only Sandow and Cannon [4] used single-sided multidirectional b/e patches, but they used them on edge-crack repairs. They measured the LIF of 5-22 under VA loading. The stiffness ratios were varied from 0.5 to 1.9 and had a significant effect on the results. In that investigation, the multidirectional b/e patches provided better performance than unidirectional b/e patches under VA loading.

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In our tests with the co-bonded wet-laminated c/e patch, the brittle resin was also used as an adhesive. The shear modulus of the resin is higher than the shear modulus of the ductile FM73 adhesive film. As analysed by Gu [17], the thin and stiff adhesive layer will transmit more load to the patch. In contrast, the ultimate shear strain of the resin is much smaller than the ultimate shear strain of the film adhesive, and it will be prone to delamination and breakage of the joint. In this study, all results achieved with the co-bonded wet-laminated c/e repairs were better than the results achieved with the c/ e prepreg and b/e prepreg reinforcements. One reason for that is the higher stiffness ratio of the patch. However, the hot/wet and cold/dry conditions, which were not tested here, can cause limitations with the wet lay-up materials. We could not find any test results of similar wet-laminated repairs and testing from the literature.

With the wet-laminated c/e repairs cured at RT, it was possible to omit the tensile residual thermal stresses completely or even to change them to compressive. However, since the repairs were post-cured at 120 °C, some amount of additional curing of the composite patch occurred. This resulted in small residual bending observed from the repaired specimens. The result of lower residual thermal stresses was that the crack growth slowed down remarkably or stopped. The final failure location shifted from the centrecrack to the unrepaired area of the plate – next to the patch edges or next to the end tab edges. The only case where we could properly measure the LIF of the wet-laminated repairs was the standard cured repair tested at RT, where the LIF was 6.1. With all other specimens, the thermal residual stresses were lower, the crack growth was slower and they failed outside the repair area.

5.4. Effect of the structural support

The 3.2 mm thick edge-cracked plates were repaired only with the b/e prepreg reinforcements. The efficiency of the single-sided repair was low. The stiffness ratio was 0.53, and the life improvement factor was 1.9. The life improvement factor achieved here was less than the value 5 reported by Sandow and Cannon [4] for a rather similar case with a stiffness ratio of 0.63. With the double-sided repairs, the stiffness ratio was 0.65. The secondary bending effects were eliminated and the crack almost stopped. The life improvement factor was 71. This result is based on one test specimen. The other two test specimens failed outside the repair area without any growth of the repaired edge-crack. For double-sided repairs, Sandow and Cannon [4] used CA loading and a stiffness ratio of 1.2. The LIF for a double-sided patch was 150. Their specimens also failed outside the repair area. The difference measured here between the single-sided and the double-sided repairs was greater than the difference measured by Sandow and Cannon [4] and Poole [8]. On the other hand, Gu [17] has made a similar estimate for double-sided repairs using FEM analysis. According to his studies, the crack growth will practically stop with double-sided b/e repairs of a 1.6 mm aluminium plate with a reinforcement that has the stiffness ratio S = 1 [17].

We varied the structural support during the fatigue loading by using or omitting edge support rails and by using sandwich structure specimens. As expected, with single-sided repairs and VA fatigue loading including compressive loads, the support increased the fatigue lives remarkably. The life improvement factor of centrecracked specimens reinforced with single-sided c/e prepreg patches grew from 2.1 with unsupported repairs to 3.3 with edge-supported repairs and to 7.3 with sandwich-supported specimens. Surprisingly, the standard deviation was 6.5% for all cases. The deviation was considered to be very small for fatigue test specimens. In the referenced papers, there was usually no support or in some cases a sandwich structure support. The results with sandwich support are always very good [1–3,7–8]. When edge support was used, the referenced results were much closer to our results; the LIF of c/e repairs 3.3 vs. 3.7 [11] and of b/e repairs 5.2 vs. 522 [4]. The obtained results underline the need to have a realistic structural support when testing the efficiency of the single-sided repairs with plate specimens. 5.5. Effect of residual thermal stresses We varied the residual thermal stresses by using different curing and different testing temperatures and by using or omitting the bonding rig during adhesive curing. The effect of residual stresses was rectilinear. Higher residual stresses resulted in higher crack growth rates. The results are in line with the measurements made by Baker [1,2]. We found two exceptions to that order. At low temperatures (10 °C), the crack growth was faster than expected. This was assumed to come from the more brittle behaviour of the adhesive, as also reported by Vlot [19] and Butkus [43]. The repairs immersed in hot water for 30 days also had higher crack growth rate, as expected. The wet repair materials tested at elevated temperature were assumed to have lower tensile stiffness and lower shear stiffness than the dry materials. The c/e repair bonded at 70 °C had the lowest crack growth rate due to the lowest residual thermal stresses. However, when the crack reached the edges of the patch, the adhesive could not anymore carry the loads as well as with the other repairs. Cohesive failure modes were found from the adhesive film, and adhesive failure modes were found from the interfaces. It was assumed, that the adhesive film was not completely cured. Rider [44] has reported that the FM73 adhesive film cured at lower temperature (80 °C) had a lower performance in the wedge tests than the FM73 after the standard cure. On the other hand, Ong [7] measured the best results when he lowered the curing temperatures of the FM73 from 120 to 70 °C.

5.6. Efficiency of the edge-crack repairs

5.7. Crack growth in different aluminium grades Aluminium 2024-T3 has the highest fracture toughness Kcr (see Table 1) and also the highest fatigue life of the tested aluminium grades. 7075-T6 aluminium has a higher static tensile strength, but the fracture toughness is lower and also the fatigue life of the cracked plate was shorter. The fracture toughness value for S07-1020 aluminium is lower than for the other aluminium grades. The exponent n in the Paris’ equation (1) is, however, the lowest for S07-1020 and the crack growth before fracture was slower. The average fatigue life of the unrepaired S07-1020 plates was close to the fatigue life of 2024-T3 aluminium plates. The life improvement factors of the c/e prepreg repaired 2024 and S07-1020 plates were also close to each other being 3.0 and 3.2, respectively. The repaired S07-1020 plates had a higher standard deviation than the 2024-T3 plates. Repaired 7075-T6 clad plates had a life improvement factor of 5.9, but the fatigue life was still shorter than with the other aluminium grades. Since the crack growth rate in 7075-T6 clad aluminium is high, composite repair of the cracks is highly beneficial for this material. 6. Conclusions An experimental investigation was conducted to characterize the effect of several parameters on fatigue crack growth of precracked aluminium plates repaired with multidirectional c/e and

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b/e composite patches. We draw the following conclusions from the study:  The steel bonding rig we developed in order to control the thermal expansion of the repaired plates restricted the expansion to the level measured from locally heated real aircraft structures.  The residual thermal stresses we varied using the rig and different curing temperatures had a rectilinear effect on the crack growth rate of the patched plates.  Multidirectional b/e prepreg, c/e prepreg and wet-laminated c/e reinforcements were all efficient in the repairs of centre-cracked aluminium plates. Best performance was achieved with the wetlaminated patches, which had the highest stiffness ratio.  A support structure against secondary bending of the singlesided repairs provided a remarkable difference in the test results when compared to unsupported or sandwich-supported repairs. The support was emphasized, since we used a loading spectrum that included compressive peaks.  The double-sided b/e repairs stopped the edge-crack growth almost completely, while the single-sided repairs were not as efficient.  The 2024-T3 clad aluminium and the S07-1070 aluminium plates had a longer fatigue life when compared to the 7075T6 clad aluminium plates. The life improvement factor due to the bonded composite repairs, on the other hand, was highest with 7075-T6 clad aluminium.

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