SDHM. Structural Durability & Health Monitoring. Tech Science Press. Reprinted from. Editor-in-Chief: Ferri M.H.Aliabadi

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1 Reprinted from SDHM Structural Durability & Health Monitoring Editor-in-Chief: Ferri M.H.Aliabadi Honorary Editor: Satya N. Atluri ISSN: (print) ISSN: (on-line) Tech Science Press

2 Copyright 2011 Tech Science Press SDHM, vol.7, no.3, pp , 2011 Analytical and Experimental Investigations of Extending the Crack Growth Life of Integrally Stiffened Aluminum Panels by the Use of Composite Material Strips A. Brot 1 and I. Kressel 1 Abstract: Analysis and testing of integrally stiffened aluminum panels, reinforced by carbon-epoxy or boron-epoxy bonded strips, is described. Fatigue testing was performed at room-temperature and at -50 C. The test results show a very significant increase in the crack growth life of these panels after the reinforcement. The analytical results, based on finite-element models, correlated very well with the test results. 1 Introduction Israel Aerospace Industries (IAI) has investigated the damage-tolerance behavior of integrally stiffened metallic structures, reinforced by composite strips, as part of an international project called DaToN (Innovative Fatigue and Damage Tolerance Methods for the Application of New Structural Concepts). The DaToN project was partially funded by the European Commission (EC). IAI has performed both analytical and experimental studies of integrally stiffened metallic panels, in the framework of this EC project. This paper describes the analytical computations and fatigue testing that was performed in order to study the effect of adding composite material strips to integrally stiffened aluminum panels. 2 Testing of the Unreinforced Integrally Stiffened Panels A total of six integrally stiffened panels were crack-growth tested under constant amplitude fatigue loading. The panels were machined from 2024-T351 aluminum alloy. The overall dimensions of the panels were a 450 mm width and a 1000 mm length. Each panel was manufactured with two integral stringers. The first three panels were crack-growth tested, without any reinforcing strips, at several stress levels and stress-ratios. Figure 1 shows a two-stringer panel in the test fixture, 1 Engineering Division, Israel Aerospace Industries, Ben-Gurion Airport, Israel. abrot@iai.co.il

154 Copyright 2011 Tech Science Press SDHM, vol.7, no.3, pp.

An artificial crack of ±15mm length was inflicted at the panel

The panels had crack propagation gages bonded back-to-back to

a maximum stress level of 80 MPa, with R = 0.1.

growth of the front and back, right and left crack tips.

almost no resistance to the advance of the fatigue crack.

to evaluate the effect of the panel reinforcement using

Figure 1: A Two-Stringer Panel in the Test Rig and a Panel

to the Panels In recent years, there has been much discussion

composite materials bonded to the aluminum.

to bridge any cracks that may develop in the aluminum panel.

3 154 Copyright 2011 Tech Science Press SDHM, vol.7, no.3, pp , 2011 before and after failure. An artificial crack of ±15mm length was inflicted at the panel centerline. The panels had crack propagation gages bonded back-to-back to the panels, along the expected crack path, in order to monitor the crack growth. Figure 2 shows the measured results of an unreinforced panel at a maximum stress level of 80 MPa, with R = 0.1. The results shown in Figure 2 represent the mean value of the growth of the front and back, right and left crack tips. It is very clear from Figure 2 that the stringers offered almost no resistance to the advance of the fatigue crack. As such, their value as a damage-tolerance enhancer was found to be minimal. The results shown in Figure 2 were used as a baseline in order to evaluate the effect of the panel reinforcement using composite materials. Figure 1: A Two-Stringer Panel in the Test Rig and a Panel after Failure 3 Applying Composite Material Reinforcing Strips to the Panels In recent years, there has been much discussion of the advantages of a "hybrid" stiffened panel which has composite materials bonded to the aluminum. The composite material reinforces the aluminum panel and serves to bridge any cracks that may develop in the aluminum panel. This bridging effect was proven during the last 30 years in many composite bounded repairs of aging aircraft [Baker et al (2002)]. In order to improve the performance of the two-stringer integral panel, two 35mm wide strips, made from Hexcel Vicotex 913 unidirectional carbon-epoxy material were co-bonded to the panels at 120 C, using 3M AF adhesive, as is shown

4 Analytical and Experimental Investigations of Extending the Crack Growth Life half crack length a (mm) Stiffener Center Line Beginning of Reinforcment Life [cycles] Life (Cycles) Figure Figure 2: 2: Measured Measured Crack Growth Growth Curve Curve for for the the Two-Stringer Two-Stringer Unreinforced Unreinforced Panel Panel 3. Applying Composite Material Reinforcing Strips to the Panels in Figure 4. Each strip consisted of three plies of carbon-epoxy material. The In recent years, there has been much discussion of the advantages of a "hybrid" stiffened panel which has composite purpose materials ofbonded the strips to the was aluminum to reduce [1] the - [5]. stress-intensity The composite of material a crack that reinforces growsthe un-aluminuder it, to thereby bridge increasing any cracks the that crack may growth develop life in the of the aluminum panel. panel. On another This bridging identicaleffect was panel and serves proven during panel, the two last mm years wide in many stripscomposite of Textronbounded 5521 F/4 repairs unidirectional of aging aircraft boron-epoxy [4]. were bonded. Each strip consisted of two plies of boron-epoxy material. For both reinforcement improve the schemes, performance the composite of the two-stringer material strips integral were panel, bonded two 35mm only onwide the stringer strips, made from In order to Hexcel Vicotex side of 913 theunidirectional panels. carbon-epoxy material were co-bonded to the panels at 120 C, using 3M AF adhesive, as is shown in Figure 3. Each strip consisted of three plies of carbon-epoxy material. The purpose of the strips was to reduce the stress-intensity of a crack that grows under it, thereby increasing 4 Calculating the crack the growth Crack life Growth of the panel. Characteristics On another of identical a Reinforced panel, two Panel 35 mm wide strips of Textron 5521 F/4 unidirectional boron-epoxy were bonded. Each strip consisted of two plies of boronepoxy material. A NASTRAN For both finite-element reinforcement schemes, model (FEM) the composite was built to material study the strips effect were ofbonded the composite of the material panels. reinforcing strips on the stress-intensity factor. The model was only on the stringer side composed of CQUAD4 shell elements representing the skin and the reinforcement 4. Calculating strips. the 3D Crack HEXA Growth elements Characteristics were used for of a the Reinforced adhesive. Panel Due to symmetry, only a quarter-model was analyzed. A nonlinear analysis was performed for several crack A NASTRAN finite-element model (FEM) was built to study the effect of the composite material reinforcing lengths strips on in order the stress-intensity to examine the factor. contribution The model of the was composite composed strips of CQUAD4 on the stressintensity the skin values. and the The reinforcement first step was strips. to build 3D HEXA the finite-element elements were model used for for the unrein- adhesive. Due to shell elements representing symmetry, forced only a panel. quarter-model The nextwas stepanalyzed. was to add A the nonlinear composite analysis material was strips performed and adhesive for several crack lengths in to order the to model. examine The the final contribution step wasof tothe calculate composite the stress-intensity strips on the stress-intensity of the cracked values. The first step was panel, to build for a the range finite-element of crack lengths model from for 15mm the unreinforced to 100mm, panel. usingthe thenext displacement- step was to add the composite material strips and adhesive to the model. The final step was to calculate the stress-intensity of the cracked panel, for a range of crack lengths from 15mm to 100mm, using the displacementextrapolation method. The results of this stress-intensity analysis are shown in Figure 3 for the carbonepoxy and boron-epoxy reinforced panels. It should be noted that the stress-intensity of the cracked aluminum panel is much lower at the bonding surface interface than at the free surface of the aluminum panel, as is shown in Figure 3. This means that the effect of the reinforcements is to introduce both tensile and bending effects on the aluminum panel. Figure 3 also showed a convergence of the mean stress-

156 Copyright 2011 Tech Science Press SDHM, vol.7, no.3, pp.

The results of this stress-intensity analysis are shown in Figure 3 for

It should be noted that the stress-intensity of the cracked aluminum

surface of the aluminum panel, as is shown in Figure 3.

tensile and bending effects on the aluminum panel.

to a nearly constant value beneath the strip, verifying the good

constant stress-intensity factor under a bonded composite patch.

of crack length, for the unreinforced panel, and for the panel with the

reinforcement, and at the free surface of the panel).

the bonding surface while "outer" refers to the free surface) The

5 156 Copyright 2011 Tech Science Press SDHM, vol.7, no.3, pp , 2011 extrapolation method. The results of this stress-intensity analysis are shown in Figure 3 for the carbon-epoxy and boron-epoxy reinforced panels. It should be noted that the stress-intensity of the cracked aluminum panel is much lower at the bonding surface interface than at the free surface of the aluminum panel, as is shown in Figure 3. This means that the effect of the reinforcements is to introduce both tensile and bending effects on the aluminum panel. Figure 3 also showed a convergence of the mean stress-intensity factors to a nearly constant value beneath the strip, verifying the good agreement with the Rose Model [Baker et al (2002)] that predicts a constant stress-intensity factor under a bonded composite patch. The reduction of the stress-intensity due to the reinforcement was taken into account by the NASTRAN analysis. The stress-intensity factors were extracted from the FEM, as a function of crack length, for the unreinforced panel, and for the panel with the reinforcing strips (at the interface between the panel and the reinforcement, and at the free surface of the panel). Figure 3: Stress-Intensity Results for Cracked Panels ("inner" refers to the bonding surface while "outer" refers to the free surface) The stress-intensity results, as obtained from the FEM, were input into NASGRO ver. 5 (crack growth software) as a data table, in order to compute the predicted crack growth characteristics. The effects of the stress-intensity variation (between the free edge and at the interface) were also accounted for by this analysis. The results of the crack growth analysis, compared to experimental results, are shown in Section 7.

Analytical and Experimental Investigations of Extending the Crack

Panels The first two reinforced panels were fatigue tested at

for the unreinforced panel (80 MPa at R = 0.1).

additional EA cross-section contribution of the reinforcing strip.

significantly slower crack growth rate than the unreinforced

Figure 5 also shows that the crack growth life of the three-layer

and the metal substrate, or delaminations between the layers,

Figure 4: Panel with Carbon-Epoxy Reinforcing Strips, Before and

panels seems to be constant, almost up to failure.

6 Analytical and Experimental Investigations of Extending the Crack Growth Life Room Temperature Testing of the Reinforced Panels The first two reinforced panels were fatigue tested at room temperature, under a 7% higher loading than what was used for the unreinforced panel (80 MPa at R = 0.1). The purpose of the 7% increase was to compensate for the additional EA cross-section contribution of the reinforcing strip. Figure 4 shows the reinforced panels installed in the testing fixture, before and after failure. Figure 5 shows the crack growth test results of both reinforced panels at roomtemperature. The results clearly show that both reinforced panels had a significantly slower crack growth rate than the unreinforced panel. Figure 5 also shows that the crack growth life of the three-layer carbon-epoxy strips gave somewhat better results than the two-layer boron-epoxy strips. It should be noted that no debonds between the composite strips and the metal substrate, or delaminations between the layers, were observed up to failure for all the tested panels. Figure 4: Panel with Carbon-Epoxy Reinforcing Strips, Before and after Failure The crack propagation rate of all the reinforced panels seems to be constant, almost up to failure. This phenomenon is in good agreement with the Rose Model [Baker et al (2002)] that predicts a constant stress-intensity factor under a bonded composite patch. Detrimental residual thermal stresses exist in aluminum panels reinforced by composite material patches, induced by the thermal expansion coefficient mismatch

158 Copyright 2011 Tech Science Press SDHM, vol.7, no.3, pp.

) between the carbon-epoxy or boron-epoxy materials and the aluminum substrate.

temperature 120 C and the operating temperature.

residual stress in the aluminum panel was approximately 8 MPa for both the three-layer

It should be noted that the compressive residual stress in the composite reinforcement strips

6 Testing a Reinforced Panel at -50 C The residual stress phenomenon, as described above, was

Finite-element studies showed that the residual stress in the aluminum panel will reach

On the other hand, the inherent crack growth rate in the 2024-T351 aluminum panel is much

7 158 Copyright 2011 Tech Science Press SDHM, vol.7, no.3, pp , 2011 Figure 5: Comparison of the Crack Growth of Both Reinforced Panels to the Unreinforced Two-Stringer Panel. (The loading on the reinforced panels were increased by 7% relative to the unreinforced panel.) between the carbon-epoxy or boron-epoxy materials and the aluminum substrate. These residual stresses may be significant because of the difference between the curing temperature 120 C and the operating temperature. When tested at roomtemperature (approximately 25 C), finite-element studies showed that the residual stress in the aluminum panel was approximately 8 MPa for both the three-layer carbon-epoxy strips and the two-layer boron-epoxy strips, a relatively insignificant value. It should be noted that the compressive residual stress in the composite reinforcement strips was significantly higher than that of the aluminum substrate. 6 Testing a Reinforced Panel at -50 C The residual stress phenomenon, as described above, was shown to be more pronounced at the reduced temperatures that occur at higher altitudes. Finite-element studies showed that the residual stress in the aluminum panel will reach approximately 14 MPa for the carbon-epoxy strips at -50 C. On the other hand, the inherent crack growth rate in the 2024-T351 aluminum panel is much slower at -50 C than at room temperature. Therefore, an additional test was performed on a carbonepoxy reinforced panel at an ambient temperature of -50 C. Figure 6 shows the test setup and the refrigeration unit that was used to cool the test chamber to -50 C. Also for this test, 7% higher loading was used, compared to what was used for the

8 Analytical and Experimental Investigations of Extending the Crack Growth Life 159 unreinforced panel (80 MPa at R = 0.1). Figure 6: Test set-up for the Carbon-Epoxy Reinforced Panel Tested at -50 C Figure 7 shows that the crack grew significantly slower at -50 C than at room temperature, showing that the reduced crack growth rate of aluminum at -50 C was more decisive than the presence of tensile residual stresses. It should be noted that, as in the previous tests performed at room temperature, no debonds between the composite strips and the metal substrate, or delaminations between the layers, were observed up to the failure.

9 160 Copyright 2011 Tech Science Press SDHM, vol.7, no.3, pp , 2011 half crack length a (mm) Stiffener Center Line Unreinforced Specimen Carbon-Epoxy RT Specimen Carbon-Epoxy -50ºC Specimen Beginning of Reinforcment Life (cycles) Figure 7: Comparison of Carbon-Epoxy Reinforced Panels at Room-Temperature and at -50 C (results are compared to an unreinforced panel) 7 Comparison between the Analytical and Experimental Results Figure 8 compares the measured crack growth characteristics for the carbon-epoxy reinforced panel to the analytical results computed by the methods described in Section 4. Similarly, Figure 9 compares the measured crack growth characteristics for the boron-epoxy reinforced panel to the analytical results. The results shown in both Figure 8 and Figure 9 indicate very good agreement between the test and analytical results. 8 Summary and Conclusions This analytical and experimental study demonstrated that the use of carbon-epoxy or boron-epoxy strips will significantly increase the crack growth life of integrally stiffened aluminum panels. Further testing and analysis is needed to quantitatively confirm these results. The analytical results, derived from finite-element models, correlate very well with the test results. The effect of tensile residual stresses in the aluminum panels at -50 C, introduced by the coefficient of thermal expansion mismatch between the aluminum and composite materials, is not detrimental to the crack growth rate since the reduced crack growth rate of aluminum at -50 C more than offsets the effect of the tensile residual stresses.

10 Analytical and Experimental Investigations of Extending the Crack Growth Life 161 Figure 8: Crack Growth of the Carbon-Epoxy Reinforced Panel Figure 9: Crack Growth of the Boron-Epoxy Reinforced Panel

11 162 Copyright 2011 Tech Science Press SDHM, vol.7, no.3, pp , 2011 No debonds between the composite strips and the metal substrate, or delaminations between the layers, were observed up to failure for all the panels tested. References Baker, A. (1984): Repair of Cracked or Defective Metallic Aircraft Components with Advanced Fibre Composites an Overview of Australian Work, J. of Composite Structures, vol. 2. Baker, A. (1999): Bonded Composite Repair of Fatigue-Cracked Primary Aircraft Structure, Composite Structures, 47: Baker, A., Rose, F. and Jones, R. (2002): Advances in the Bonded Composite Repair of Metallic Aircraft Structure, Elsevier. Heinimann, M. (2007): Validation of Advanced Metallic Hybrid Concept with Improved Damage Tolerance Capabilities for Next Generation Lower Wing and Fuselage Applications, Proceedings of the 24th ICAF Symposium, Naples. Zhang, X. et al (2007): Improving Fail-Safety of Aircraft Integral Structures through the Use of Bonded Crack Retarders, Proceedings of the 24th ICAF Symposium, Naples.

12 SDHM: Structural Durability & Health Monitoring ISSN: (Print); (Online) Journal website: Manuscript submission Published by Tech Science Press 5805 State Bridge Rd, Suite G108 Duluth, GA , USA Phone (+1) Fax (+1) Website: Subscription: SDHM is Indexed & Abstracted in Cambridge Scientific Abstracts (Aerospace and High Technology; Materials Sciences & Engineering; and Computer & Information Systems Abstracts Database); Engineering Index; INSPEC Databases; Mechanics; Science Navigator.

THE DAMAGE-TOLERANCE BEHAVIOR OF INTEGRALLY STIFFENED METALLIC STRUCTURES

THE DAMAGE-TOLERANCE BEHAVIOR OF INTEGRALLY STIFFENED METALLIC STRUCTURES A. Brot, Y. Peleg-Wolfin, I. Kressel and Z. Yosef Engineering Division Israel Aerospace Industries e-mail: abrot@iai.co.il ABSTRACT