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Modeling of prior exfoliation corrosion in aircraft wing skins
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Modeling of prior exfoliation corrosion in aircraft wing skins M. Liao, G. Renaud, D. Backman, D.S. Forsyth, N.C. Bellinger Structures, Materials, and Propulsion Laboratory, Institutefor Aerospace Research, National Research Council, Ottawa, Ontario K1A OR6, Canada
Abstract
This paper presents preliminary results on the development of an analytical framework for evaluating the effects of prior exfoliation corrosion on the residual strength and remaining life of wing skins. A number of pristine and corroded specimens were cut from naturally exfoliated wing panels fabricated from 7075-T6511 alloy, and an ultrasonic non-destructive inspection (NDI) was carried out on the exfoliated specimens. The maximum depth and three-dimensional (3D) profile of the exfoliation damage were determined from the NDI data analysis. These specimens were tested using static tension and compression loading and the strain distributions on the specimens were measured using strain gauges and photoelastic coating techniques. An analytical model was developed based on a "soft inclusion" technique, which was used to simulate the exfoliation damage. Using the PATRAN Command Language, automatic geometry and finite element (FE) model generation techniques were developed to create the 'soft inclusion' for any shape and depth of exfoliation including the detailed 3D profile determined by NDI data analysis. The 3D FE model was first verified with the pristine specimen test results. FE analyses were then carried out on the exfoliated specimens and compared with the tests results. From the comparisons, the parameters of the "soft inclusion" such as Young's modulus were calibrated. Finally the calibrated "soft inclusion" based FE model was applied to evaluate the effects of the exfoliation on the residual strength of the corroded specimens. The good agreements between the modeling and testing results indicated that the analytical models are capable of evaluating the effects of exfoliation on the residual strength. I. Introduction
In aircraft materials exfoliation is most common in the heat-treatable AI-Zn-Mg-Cu (7000 series), AI-Cu-Mg (2000 series), and A1-Mg alloys, but it has also been observed in AI-Mg-Si alloys [ 1]. Like pitting and stress corrosion cracking (SCC), exfoliation is commonly found in aircraft structures. It often occurs in the upper wing skins, either on the smooth surface where it starts from a corrosion pit, or around wing fastener holes where it originates at the exposed end grains in the countersink and hole bore surface [2].
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According to the current "f'md-it-fix-it" maintenance approach being implemented by many operators, whenever exfoliation is discovered, it must be repaired (ground out). With the aid of non-destructive inspection (NDI) techniques, or controlled search peening, very small levels of exfoliation can be identified, thus the current maintenance approach could result in more frequent corrosion maintenance actions. While this approach is costly, the effects of the grindout on the residual strength and remaining life of a structure are not well understood. Some studies have shown that a certain level of exfoliation may reduce the static strength and fatigue life less than the grindout repair [3]. A new corrosion management approach has been proposed with the intent of "anticipating, planning, and managing" corrosion, which stands in sharp contrast to the present "'find-it-fix-it" philosophy. To implement this new philosophy, analytical models need to be developed to quantify the effect of exfoliation on the structural integrity of wing skins. In the past years, considerable research has been carried out at the Institute for Aerospace Research of National Research Council Canada (IAR/NRC) on corrosion and fatigue [4]. An analytical framework was proposed to evaluate the effect of exfoliation on the residual strength and remaining life of aircraft structures and provide guidance for the repair actions [5,6]. The framework was designed to start the analysis from the NDI input and automatically generate a finite element (FE) model that contains the damage zone, simulated using a "soft inclusion" technique [7]. A commercial FE code is then used to carry out the stress analysis. Finally, the residual strength and remaining life of the structure are estimated. In this paper analytical models were developed in line with the proposed framework, and static tests were carried out to provide the data for calibrating and verifying the analytical models. 2. Static tests
The goals of the static tests were to (i) quantify the stiffness degradation of the exfoliated specimen and (ii) provide data to examine the FE modeling associated with the "soft inclusion" technique. To this end, the test program was designed as a set of comparative tests using both pristine and corroded specimens. The materials for the test specimens were cut from various upper wing skin panels that were naturally corroded in storage and therefore rendered unserviceable. These panels, fabricated from extruded 7075-T6511 alloy material, were originally supplied by Lockheed Martin Aeronautics for another project. Four wing skin panels were used for this study, and one of them is shown in Fig. 1. A specimen was designed for both tension and compression static tests with the maximum net stress lower than the yield strength. The specimen design was based on the ASTM E8-01 and ASTM E 111-97 standards, with some dimensions modified to be able to cover the exfoliation damage. Details of the specimen design are shown in Fig. 2. To date, six specimens, including three pristine and three corroded, have been tested. Fig. 3 shows the three corroded specimens, in which the indicated corrosion level, i.e. light, moderate, severe, was determined from visual inspection. To accurately measure the area and depth of the exfoliation, ultrasonic Time-of-Flight (ToF) inspections were carried out from the back surface (non-corroded side) of each corroded specimen. The resulting ultrasonic images, which were converted from the digitized images to display corrosion profiles, are also shown in Fig. 3.
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Fig. 1. C 141 aircraft wing panel.
Fig. 2. Static test specimen (in mm).
Specimen C 141-2-18,19A with light corrosion.
Specimen C141-2-21A with moderate corrosion.
Specimen C 141-2-20A with severe corrosion. Fig. 3. Corroded specimens and their corresponding ultrasonic images.
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Table 1 Summary of ultrasonic NDI data analysis results Specimen C 141-2-18, 19A (li~;ht) Nominal thickness (mm) 4.01 (0.1579 in.) 0.10 (0.0040 in.) Maximum thickness loss (mm) 10.2 (0.40 in.) Maximum damage width (mm) 2.5 Maximum thickness loss (%) Average thickness loss (%)
Fig. 4. Strain gauges on pristine specimen (SO, $4" uniaxial; S1-$2-$3" rosette).
C 141-2-21A (moderate) 3.88(0.1529 in.) 0.47(0.0185 in.) 74.4 (2.93 in.) 12.1 2.52
C 141-2-20A (severe) 3.99(0.1570 in.) 0.66 (0.0258 in.) 134.11 (5.28 in.) 16.4 6.93
Fig. 5. Photoelastic coating on pristine and corroded specimens.
The ToF data were analyzed using the in-house software package, NDIAnalysis, to determine the maximum exfoliation depth (maximum thickness loss) and threedimensional (3D) profile of each corroded specimen. Table 1 summarizes some of the results from the NDI data analysis. For the pristine specimens, strain gauges were placed on the front surface and a photoelastic coating (PEC) was attached to the back surface. The location and type of strain gauges are shown in Fig. 4. In the case of the corroded specimens, only a PEC was placed on the back surface (non-corroded side), as shown in Fig. 5. The test procedure incorporated interruptions at predetermined load levels to allow for the collection of strain gauge data and photoelastic images. The maximum tension load was set so as to create a net stress in the centre of the specimen equivalent to approximately 80% of the material yield strength. The maximum load used in the compression tests was set to a value just below the buckling load of each specimen. This value was determined using ASTM E9-89a. Each test was repeated three times. Of the six tested specimens, one pristine and two corroded specimens buckled during the compression tests and the buckling loads were recorded. The strain gauge results were then used to determine the Young's modulus for the tested material by acquiring the slope of the straight line fitted to the strain-stress data using the linear regression (least square) method, as recommended by ASTM E 111-97. The strain-stress data were taken from gauge $3 within the load range from 8.90 kN (2000 lb) to 22.24 kN (5000 lb). The Young's modulus (E), which was determined from the three pristine specimens, is given in Table 2. An automated photoelastic analysis [8], based on six PEC images taken at each load step, was used to determine the maximum shear strain (~'M~x)distribution on the back surface of the specimen. The YM~,results along both longitudinal and transverse centre lines (see Fig. 5) were used for quantitative study.
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Table 2 Young's modulus of tested 7075-T6511 AA from three pristine specimens Mean (GPa) Standard deviation (GPa)
Compression 71.2 (10.32 x 10 6 psi) 1.17 (0.17 x 10 6 psi)
Tension 69.6 (10.09 x 106 psi) 0.06 (0.01 x 106 psi)
3. Modeling A 3D FE-based analytical model was developed with the intention of quantifying the influence of exfoliation corrosion on the residual strength and life of a structure. The model used a "soft inclusion" technique to simulate the exfoliation damage present in the specimen [7]. This technique was previously developed and verified at IARfNRC to predict the compression strength after impact in composite structures [9]. The basic assumptions associated with the sott inclusion are: (i) the exfoliated damage zone, simulated by a sott inclusion with reduced stiffness, may have some load carrying capability depending on the corrosion level, and (ii) the soft inclusion, which has approximately the same volume as the exfoliated damage zone, is an isotropic material. In order to analyze various exfoliation corrosion damages with different depth and shape, automatic geometry and FE model generation techniques were developed based on the MSC PATRAN Command Language (PCL). The 3D profile of the damage, determined from the ultrasonic NDI and data analysis, was represented by the damage depth at each point of a rectangular grid covering the damage area. A PCL program was developed to automatically access the 3D profile data file, generate the FE models (geometry and mesh), launch the FE analysis using MSC NASTRAN, and retrieve the results. Fig. 6 illustrates this process schematically. An advantage of this approach is that complex damage topography can be modeled up to the same accuracy as provided by NDI. The damage is fully modeled by solids and appropriate mesh seeds are applied on their edges to automatically create the 3D mesh. This modeling strategy simplifies the meshing procedure and allows the use of
Fig. 6. FE modeling process based on NDI data for corroded specimen C141-2-21A (corrosion damage was removed for clarity).
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different mesh properties. With this modeling capability, multiple tasks were launched to automatically carry out the analyses on different geometries, material properties, loads, etc. Only the portion of the specimen between the grips was modeled. The tests were simulated by applying a total force on one end of the model while clamping the other end. Constant longitudinal displacement constraints and zero transverse and lateral displacements were imposed along the loaded edge to enforce symmetry in the boundary conditions. The FE model for a corroded specimen is shown in Fig. 6(c). 4. Results
4.1. Pristine specimens
To validate the FE modeling techniques, the pristine specimens were first analyzed and compared with the test results. Fig. 7 shows the strain results obtained from the FE analysis at 22.24 kN (5000 lb) on the back surface of specimen C141-2-15A along with the strain gauge and PEC measurements. This figure shows a very good agreement between the FE, strain gauge, and PEC results. Good agreement also occurred in the other two pristine tests and under different load levels, which indicates that the FE modeling techniques, including the mesh size, element type and boundary conditions, were appropriate.
Fig. 7. Comparison of FE, strain gauge, and PEC results for pristine specimen C141-2-15A (center portion) at compression load 22.24 kN (5000 lb).
4.2. Corroded specimens
Fig. 8 presents the ~Maxresults measured from PEC during the compression tests for both the pristine and corroded specimens. This figure indicates that only the severely corroded specimen showed significantly different strain results as compared to the others. The exfoliation damage (depth and area) in the light and moderately corroded specimens was not severe enough to significantly affect the strain distribution on the non-corroded back surfaces. Therefore, the severely corroded specimen was used for calibrating the parameters of the soft inclusion.
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Fig. 8. Maximum shear strain (?M~x)results measured by PEC for pristine, light, moderate, and severely corroded specimens (center portions) at compression load 22.24 kN (5000 lb).
Fig. 9. Comparison of FE and PEC results for severely corroded specimen C141-2-20A (center portion) at compression load 8.90 kN (2000 lb).
Fig. 10. Comparison of FE and PEC results for severely and moderately corroded specimens loaded at 22.24 kN (5000 lb)compression.
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Since the soft inclusion was assumed to be an isotropic material, only two material properties, Poisson ratio (o) and Young's modulus (E), needed to be calibrated. FE analyses were carried out on the severely corroded specimen using different Poisson ratios and very similar strain results were obtained. Therefore, the actual Poisson ratio for AI (0.33) was used for the soft inclusion. The Young's modulus for the soft inclusion was varied to simulate the 'sotten' damage zone. Fig. 9 presents the YM~x results obtained from the FE analyses using different E degradations along with the PEC results. The comparison indicates that, - When the E degradation was changed from zero (i.e., no material loss) to 100% (i.e., complete material loss), the FE results were closer to the PEC results, and - The FE analyses using 50, 90 and 100% E degradations provided a good match with the PEC results at different positions along the longitudinal centre line, which suggests that different E values should be used for different damage areas. Optical microscopic observations on polished sections taken from specimen C 141-220A showed that deep intergranular (IG) corrosion was not present (deep IG corrosion is deemed to be able to maintain some load carrying capability in the exfoliated material). Overall, the sott inclusion/damage zone in C141-2-20A was calibrated to have 100% E (stiffness) degradation. FE analyses were carried out for C141-2-20A (severe) as well as C141-2-21A (moderate) at a compression load of 22.24 kN (5000 lb) using 100% E degradation. A comparison of the FE and PEC results is given in Fig. 10. Fig. 10 shows that the FE results compared well to the PEC results for the moderately corroded specimen, C141-2-21A, but not for the severely corroded specimen, C141-2-20A, which had a relative difference of 20% at the centre of the specimen. This discrepancy may have been caused by the bending that occurred during the test due to the severe corrosion on one side of the specimen. To verify this, an additional layer of elements were added to the FE model to simulate the photoelastic coating, which was 0.5 mm thick with a Young's modulus of 2.5 GPa and a Poisson ratio of 0.38. The FE results taken from the back surface of the coating, also presented in Fig. 10, are closer to the test results. In addition, Fig. 10 shows the FE results for the bottom surface of corrosion damage, which can not be measured in the tests. As expected these results show greater influence of the exfoliation on the front surfaces (corroded sides).
4.3. Effect of prior exfoliation corrosion on residual strength The residual strength analysis for the severely corroded specimen C141-2-20A was carried out using the developed FE model with a 100% E degradation soft inclusion. Using linear FE analysis, buckling loads were calculated for the pristine and corroded specimens. Table 3 presents the calculated and test results. Table 3 Buckling loads for pristine and corroded specimens Specimen Test (kN) FE analysis (kN)
C 141-2-22 (pristine) 43.77 (9840 lb) 58.29(13105 Ib)
C 141-2-20A (severe) 26.28 (5908 lb) 34.82(7828 lb)
Bucklingload ratio between corroded and pristine specimens 60.0% 59.7%
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Table 4 Compression yield FE analysis
Displacement (mm) Load (kN)
"C 141-2-20A" (assumed without corrosion) 0.0742(0.00292 in.) 61.28 (13776 lb)
C 141-2-20A (severe) 0.0742(0.00292 in.) 5.8.00 (13038 lb)
Ratio between corroded and pristine specimens 94.6%
Table 3 shows that the FE analyses overestimated the buckling loads for both the pristine and corroded specimens, which was expected from the linear buckling analysis. However, in terms of the buckling load ratio between the corroded and pristine specimens, the FE analyses gave the same results as the test. This indicates that the developed FE modeling is capable of estimating the relative impact of the exfoliation on the buckling load of the corroded specimen. The soft inclusion based FE model was used to evaluate the effect that the exfoliation had on the compression strength of the specimen. The compression yield load, which would cause global yielding of the specimen, was first calculated for a noncorroded specimen. The corresponding displacement was then applied to the same specimen with corrosion to calculate the compression yield load. The analytical results are presented in Table 4. Table 4 indicates that the compression yield load of the corroded specimen was reduced by less than 5% compared to the same specimen with no corrosion. The influence of the exfoliation on the compression strength is not as significant as that on the buckling load. To estimate the effect of the exfoliation on the remaining fatigue life of the corroded specimen, the sott inclusion based FE analyses were carried out for both corroded and pristine specimens under tension loading to examine the local stress concentration induced by the corrosion damage. Fig. 11 shows the maximum principal stress distribution for the centre portion of the corroded (C141-2-20A) and pristine (Cl41-222) specimens at a 22.24 kN (5000 lb) tension load. This figure reveals several stress concentrations on the bottom surface of corrosion damage with the maximum principal stress value of 244 MPa (35.4 ksi), which is about 40% higher than that of the pristine specimen (175 MPa (25.4 ksi)). This suggests that the effect of exfoliation on the remaining fatigue life of the corroded specimen could be significant. It should be noted that the local stress concentration is affected by the local damage 3D profile, which depends on the resolution and accuracy of an NDI technique. Since the existing NDI techniques have limited resolution for detecting small corrosion pits, destructive inspection (DI) measured 3D profile and detailed modeling, including pit modeling, may be needed for reliable fatigue analysis.
5. Concluding remarks In conjunction with static tests using specimens taken from upper wing skin panels, a "soft inclusion" based FE model was developed to evaluate the residual strength of a corroded specimen. The 3D FE model was first verified with the test results from pristine specimens. For the corroded specimens, an approach to automatically generate the 3D geometry of the damage zone (soft inclusion) using NDI input was developed
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(a) C141-2-20A (severe), bottom surface of corrosion damage (fringe unit: psi).
(b) C141-2-22 (pristine), same fringe definition as that used in (a). Fig. 11. Maximum principal stress distribution under 22.24 kN (5000 lb) tension load.
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with the aid of a PCL program. The parameters of a "soft inclusion" were calibrated using the tests results and the calibrated FE model was used to evaluate the buckling load, compression yield load, and stress concentration for the corroded specimens. The analysis results indicate that the effects of the exfoliation on the buckling and stress concentration of the specimen are more significant than that on the compression strength. Future work will validate the analytical model ft~her with additional compression and bearing tests using fastener holes containing exfoliation, and integrate the analytical model with corrosion pit/IG modeling to evaluate the remaining fatigue life.
Acknowledgements This work has been performed with the support of Defence Research and Development Canada (DRDC) and National Research Council Canada, project on Exfoliation Analysis and Repair Tool. Thanks to Mike Brothers for carrying out the ultrasonic inspections, Tom Benak, and Tony Marincak for conducting the static tests. Special thanks to Graeme Eastaugh for his efforts to obtain permission from Lockheed Martin Aeronautics for using the wing skin panels.
References [1] M.O. Speidel, M.V. Hyatt, Stress-corrosion cracking of high-strength aluminum alloys, in: M.G. Fontana, R.W. Staehle (Eds.), Advances in Corrosion Science and Technology, vol. 2, Plenum Press, New York, 1972, pp. 115-335. [2] W. Wallace, D.W. Hoeppner, Aircraft Corrosion: Causes and Case Histories, AGARD Corrosion Handbook, vol. 1, AGARD-AG-278, Neuilly-Sur-Seine, 1985, p. 93. [3] M. Worsfold, The effect of corrosion on the structural integrity of commercial aircraft structure, Presented at the Workshop on Fatigue in the Presence of Corrosion organized by the NATO Applied Vehicle Technology Panel, held October 7-8, 1999, Corfu, Greece. [4] J.P. Komorowski, N.C. Bellinger, R.W. Gould, D.S. Forsyth, G. Eastaugh, Can. Aeron. Space J. 47 (2001) 275-288. [5] M. Liao, J.P. Komorowski, Effect of prior exfoliation corrosion on fatigue and fracture behavior of aging aircraft structures and materials, LTR-SMPL-2002-0136, Institute for Aerospace Research, National Research Council, Ottawa, 2002. [6] N.C. Bellinger, J.P. Komorowski, M. Liao, D. Carmody, T. Foland, D. Peeler, Preliminary study into the effect of exfoliation corrosion on aircraft structural integrity, in: Proc. 6th Joint FAA/DoD/NASA Conference on Aging Aircraft, held 2002, San Francisco, USA, (CD-ROM). [7] M. Liao, N.C. Bellinger, J.P. Komorowski, Int. J. Fatigue 25 (2003) 1059-1067. [8] M. Liao, T. Benak, D. Backman, M. Brother, P. Vesely, T. Marincak, Static testing of C141 upper wing skin specimens containing exfoliation corrosion - preliminary results, LTRSMPL-2004-0016, Institute for Aerospace Research, National Research Council, Ottawa, 2004. [9] Y. Xiong, C. Poon, P.V. Straznicky, H. Vietinghoff, Comp. Struct. 30 (1995) 357-367. [10] N.C. Bellinger, Y. Xiong, C. Poon, Determination of finite width correction factors for composite laminates containing impact damage, LTR-ST-1899, Institute for Aerospace Research, National Research Council, Ottawa, 1992.
Modeling of prior exfoliation corrosion in aircraft wing skins M. Liao, G. Renaud, D. Backman, D.S. Forsyth, N.C. Bellinger Structures, Materials, and Propulsion Laboratory, Institutefor Aerospace Research, National Research Council, Ottawa, Ontario K1A OR6, Canada
Abstract
This paper presents preliminary results on the development of an analytical framework for evaluating the effects of prior exfoliation corrosion on the residual strength and remaining life of wing skins. A number of pristine and corroded specimens were cut from naturally exfoliated wing panels fabricated from 7075-T6511 alloy, and an ultrasonic non-destructive inspection (NDI) was carried out on the exfoliated specimens. The maximum depth and three-dimensional (3D) profile of the exfoliation damage were determined from the NDI data analysis. These specimens were tested using static tension and compression loading and the strain distributions on the specimens were measured using strain gauges and photoelastic coating techniques. An analytical model was developed based on a "soft inclusion" technique, which was used to simulate the exfoliation damage. Using the PATRAN Command Language, automatic geometry and finite element (FE) model generation techniques were developed to create the 'soft inclusion' for any shape and depth of exfoliation including the detailed 3D profile determined by NDI data analysis. The 3D FE model was first verified with the pristine specimen test results. FE analyses were then carried out on the exfoliated specimens and compared with the tests results. From the comparisons, the parameters of the "soft inclusion" such as Young's modulus were calibrated. Finally the calibrated "soft inclusion" based FE model was applied to evaluate the effects of the exfoliation on the residual strength of the corroded specimens. The good agreements between the modeling and testing results indicated that the analytical models are capable of evaluating the effects of exfoliation on the residual strength. I. Introduction
In aircraft materials exfoliation is most common in the heat-treatable AI-Zn-Mg-Cu (7000 series), AI-Cu-Mg (2000 series), and A1-Mg alloys, but it has also been observed in AI-Mg-Si alloys [ 1]. Like pitting and stress corrosion cracking (SCC), exfoliation is commonly found in aircraft structures. It often occurs in the upper wing skins, either on the smooth surface where it starts from a corrosion pit, or around wing fastener holes where it originates at the exposed end grains in the countersink and hole bore surface [2].
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According to the current "f'md-it-fix-it" maintenance approach being implemented by many operators, whenever exfoliation is discovered, it must be repaired (ground out). With the aid of non-destructive inspection (NDI) techniques, or controlled search peening, very small levels of exfoliation can be identified, thus the current maintenance approach could result in more frequent corrosion maintenance actions. While this approach is costly, the effects of the grindout on the residual strength and remaining life of a structure are not well understood. Some studies have shown that a certain level of exfoliation may reduce the static strength and fatigue life less than the grindout repair [3]. A new corrosion management approach has been proposed with the intent of "anticipating, planning, and managing" corrosion, which stands in sharp contrast to the present "'find-it-fix-it" philosophy. To implement this new philosophy, analytical models need to be developed to quantify the effect of exfoliation on the structural integrity of wing skins. In the past years, considerable research has been carried out at the Institute for Aerospace Research of National Research Council Canada (IAR/NRC) on corrosion and fatigue [4]. An analytical framework was proposed to evaluate the effect of exfoliation on the residual strength and remaining life of aircraft structures and provide guidance for the repair actions [5,6]. The framework was designed to start the analysis from the NDI input and automatically generate a finite element (FE) model that contains the damage zone, simulated using a "soft inclusion" technique [7]. A commercial FE code is then used to carry out the stress analysis. Finally, the residual strength and remaining life of the structure are estimated. In this paper analytical models were developed in line with the proposed framework, and static tests were carried out to provide the data for calibrating and verifying the analytical models. 2. Static tests
The goals of the static tests were to (i) quantify the stiffness degradation of the exfoliated specimen and (ii) provide data to examine the FE modeling associated with the "soft inclusion" technique. To this end, the test program was designed as a set of comparative tests using both pristine and corroded specimens. The materials for the test specimens were cut from various upper wing skin panels that were naturally corroded in storage and therefore rendered unserviceable. These panels, fabricated from extruded 7075-T6511 alloy material, were originally supplied by Lockheed Martin Aeronautics for another project. Four wing skin panels were used for this study, and one of them is shown in Fig. 1. A specimen was designed for both tension and compression static tests with the maximum net stress lower than the yield strength. The specimen design was based on the ASTM E8-01 and ASTM E 111-97 standards, with some dimensions modified to be able to cover the exfoliation damage. Details of the specimen design are shown in Fig. 2. To date, six specimens, including three pristine and three corroded, have been tested. Fig. 3 shows the three corroded specimens, in which the indicated corrosion level, i.e. light, moderate, severe, was determined from visual inspection. To accurately measure the area and depth of the exfoliation, ultrasonic Time-of-Flight (ToF) inspections were carried out from the back surface (non-corroded side) of each corroded specimen. The resulting ultrasonic images, which were converted from the digitized images to display corrosion profiles, are also shown in Fig. 3.
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Fig. 1. C 141 aircraft wing panel.
Fig. 2. Static test specimen (in mm).
Specimen C 141-2-18,19A with light corrosion.
Specimen C141-2-21A with moderate corrosion.
Specimen C 141-2-20A with severe corrosion. Fig. 3. Corroded specimens and their corresponding ultrasonic images.
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Table 1 Summary of ultrasonic NDI data analysis results Specimen C 141-2-18, 19A (li~;ht) Nominal thickness (mm) 4.01 (0.1579 in.) 0.10 (0.0040 in.) Maximum thickness loss (mm) 10.2 (0.40 in.) Maximum damage width (mm) 2.5 Maximum thickness loss (%) Average thickness loss (%)
Fig. 4. Strain gauges on pristine specimen (SO, $4" uniaxial; S1-$2-$3" rosette).
C 141-2-21A (moderate) 3.88(0.1529 in.) 0.47(0.0185 in.) 74.4 (2.93 in.) 12.1 2.52
C 141-2-20A (severe) 3.99(0.1570 in.) 0.66 (0.0258 in.) 134.11 (5.28 in.) 16.4 6.93
Fig. 5. Photoelastic coating on pristine and corroded specimens.
The ToF data were analyzed using the in-house software package, NDIAnalysis, to determine the maximum exfoliation depth (maximum thickness loss) and threedimensional (3D) profile of each corroded specimen. Table 1 summarizes some of the results from the NDI data analysis. For the pristine specimens, strain gauges were placed on the front surface and a photoelastic coating (PEC) was attached to the back surface. The location and type of strain gauges are shown in Fig. 4. In the case of the corroded specimens, only a PEC was placed on the back surface (non-corroded side), as shown in Fig. 5. The test procedure incorporated interruptions at predetermined load levels to allow for the collection of strain gauge data and photoelastic images. The maximum tension load was set so as to create a net stress in the centre of the specimen equivalent to approximately 80% of the material yield strength. The maximum load used in the compression tests was set to a value just below the buckling load of each specimen. This value was determined using ASTM E9-89a. Each test was repeated three times. Of the six tested specimens, one pristine and two corroded specimens buckled during the compression tests and the buckling loads were recorded. The strain gauge results were then used to determine the Young's modulus for the tested material by acquiring the slope of the straight line fitted to the strain-stress data using the linear regression (least square) method, as recommended by ASTM E 111-97. The strain-stress data were taken from gauge $3 within the load range from 8.90 kN (2000 lb) to 22.24 kN (5000 lb). The Young's modulus (E), which was determined from the three pristine specimens, is given in Table 2. An automated photoelastic analysis [8], based on six PEC images taken at each load step, was used to determine the maximum shear strain (~'M~x)distribution on the back surface of the specimen. The YM~,results along both longitudinal and transverse centre lines (see Fig. 5) were used for quantitative study.
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Table 2 Young's modulus of tested 7075-T6511 AA from three pristine specimens Mean (GPa) Standard deviation (GPa)
Compression 71.2 (10.32 x 10 6 psi) 1.17 (0.17 x 10 6 psi)
Tension 69.6 (10.09 x 106 psi) 0.06 (0.01 x 106 psi)
3. Modeling A 3D FE-based analytical model was developed with the intention of quantifying the influence of exfoliation corrosion on the residual strength and life of a structure. The model used a "soft inclusion" technique to simulate the exfoliation damage present in the specimen [7]. This technique was previously developed and verified at IARfNRC to predict the compression strength after impact in composite structures [9]. The basic assumptions associated with the sott inclusion are: (i) the exfoliated damage zone, simulated by a sott inclusion with reduced stiffness, may have some load carrying capability depending on the corrosion level, and (ii) the soft inclusion, which has approximately the same volume as the exfoliated damage zone, is an isotropic material. In order to analyze various exfoliation corrosion damages with different depth and shape, automatic geometry and FE model generation techniques were developed based on the MSC PATRAN Command Language (PCL). The 3D profile of the damage, determined from the ultrasonic NDI and data analysis, was represented by the damage depth at each point of a rectangular grid covering the damage area. A PCL program was developed to automatically access the 3D profile data file, generate the FE models (geometry and mesh), launch the FE analysis using MSC NASTRAN, and retrieve the results. Fig. 6 illustrates this process schematically. An advantage of this approach is that complex damage topography can be modeled up to the same accuracy as provided by NDI. The damage is fully modeled by solids and appropriate mesh seeds are applied on their edges to automatically create the 3D mesh. This modeling strategy simplifies the meshing procedure and allows the use of
Fig. 6. FE modeling process based on NDI data for corroded specimen C141-2-21A (corrosion damage was removed for clarity).
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different mesh properties. With this modeling capability, multiple tasks were launched to automatically carry out the analyses on different geometries, material properties, loads, etc. Only the portion of the specimen between the grips was modeled. The tests were simulated by applying a total force on one end of the model while clamping the other end. Constant longitudinal displacement constraints and zero transverse and lateral displacements were imposed along the loaded edge to enforce symmetry in the boundary conditions. The FE model for a corroded specimen is shown in Fig. 6(c). 4. Results
4.1. Pristine specimens
To validate the FE modeling techniques, the pristine specimens were first analyzed and compared with the test results. Fig. 7 shows the strain results obtained from the FE analysis at 22.24 kN (5000 lb) on the back surface of specimen C141-2-15A along with the strain gauge and PEC measurements. This figure shows a very good agreement between the FE, strain gauge, and PEC results. Good agreement also occurred in the other two pristine tests and under different load levels, which indicates that the FE modeling techniques, including the mesh size, element type and boundary conditions, were appropriate.
Fig. 7. Comparison of FE, strain gauge, and PEC results for pristine specimen C141-2-15A (center portion) at compression load 22.24 kN (5000 lb).
4.2. Corroded specimens
Fig. 8 presents the ~Maxresults measured from PEC during the compression tests for both the pristine and corroded specimens. This figure indicates that only the severely corroded specimen showed significantly different strain results as compared to the others. The exfoliation damage (depth and area) in the light and moderately corroded specimens was not severe enough to significantly affect the strain distribution on the non-corroded back surfaces. Therefore, the severely corroded specimen was used for calibrating the parameters of the soft inclusion.
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Fig. 8. Maximum shear strain (?M~x)results measured by PEC for pristine, light, moderate, and severely corroded specimens (center portions) at compression load 22.24 kN (5000 lb).
Fig. 9. Comparison of FE and PEC results for severely corroded specimen C141-2-20A (center portion) at compression load 8.90 kN (2000 lb).
Fig. 10. Comparison of FE and PEC results for severely and moderately corroded specimens loaded at 22.24 kN (5000 lb)compression.
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Since the soft inclusion was assumed to be an isotropic material, only two material properties, Poisson ratio (o) and Young's modulus (E), needed to be calibrated. FE analyses were carried out on the severely corroded specimen using different Poisson ratios and very similar strain results were obtained. Therefore, the actual Poisson ratio for AI (0.33) was used for the soft inclusion. The Young's modulus for the soft inclusion was varied to simulate the 'sotten' damage zone. Fig. 9 presents the YM~x results obtained from the FE analyses using different E degradations along with the PEC results. The comparison indicates that, - When the E degradation was changed from zero (i.e., no material loss) to 100% (i.e., complete material loss), the FE results were closer to the PEC results, and - The FE analyses using 50, 90 and 100% E degradations provided a good match with the PEC results at different positions along the longitudinal centre line, which suggests that different E values should be used for different damage areas. Optical microscopic observations on polished sections taken from specimen C 141-220A showed that deep intergranular (IG) corrosion was not present (deep IG corrosion is deemed to be able to maintain some load carrying capability in the exfoliated material). Overall, the sott inclusion/damage zone in C141-2-20A was calibrated to have 100% E (stiffness) degradation. FE analyses were carried out for C141-2-20A (severe) as well as C141-2-21A (moderate) at a compression load of 22.24 kN (5000 lb) using 100% E degradation. A comparison of the FE and PEC results is given in Fig. 10. Fig. 10 shows that the FE results compared well to the PEC results for the moderately corroded specimen, C141-2-21A, but not for the severely corroded specimen, C141-2-20A, which had a relative difference of 20% at the centre of the specimen. This discrepancy may have been caused by the bending that occurred during the test due to the severe corrosion on one side of the specimen. To verify this, an additional layer of elements were added to the FE model to simulate the photoelastic coating, which was 0.5 mm thick with a Young's modulus of 2.5 GPa and a Poisson ratio of 0.38. The FE results taken from the back surface of the coating, also presented in Fig. 10, are closer to the test results. In addition, Fig. 10 shows the FE results for the bottom surface of corrosion damage, which can not be measured in the tests. As expected these results show greater influence of the exfoliation on the front surfaces (corroded sides).
4.3. Effect of prior exfoliation corrosion on residual strength The residual strength analysis for the severely corroded specimen C141-2-20A was carried out using the developed FE model with a 100% E degradation soft inclusion. Using linear FE analysis, buckling loads were calculated for the pristine and corroded specimens. Table 3 presents the calculated and test results. Table 3 Buckling loads for pristine and corroded specimens Specimen Test (kN) FE analysis (kN)
C 141-2-22 (pristine) 43.77 (9840 lb) 58.29(13105 Ib)
C 141-2-20A (severe) 26.28 (5908 lb) 34.82(7828 lb)
Bucklingload ratio between corroded and pristine specimens 60.0% 59.7%
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Table 4 Compression yield FE analysis
Displacement (mm) Load (kN)
"C 141-2-20A" (assumed without corrosion) 0.0742(0.00292 in.) 61.28 (13776 lb)
C 141-2-20A (severe) 0.0742(0.00292 in.) 5.8.00 (13038 lb)
Ratio between corroded and pristine specimens 94.6%
Table 3 shows that the FE analyses overestimated the buckling loads for both the pristine and corroded specimens, which was expected from the linear buckling analysis. However, in terms of the buckling load ratio between the corroded and pristine specimens, the FE analyses gave the same results as the test. This indicates that the developed FE modeling is capable of estimating the relative impact of the exfoliation on the buckling load of the corroded specimen. The soft inclusion based FE model was used to evaluate the effect that the exfoliation had on the compression strength of the specimen. The compression yield load, which would cause global yielding of the specimen, was first calculated for a noncorroded specimen. The corresponding displacement was then applied to the same specimen with corrosion to calculate the compression yield load. The analytical results are presented in Table 4. Table 4 indicates that the compression yield load of the corroded specimen was reduced by less than 5% compared to the same specimen with no corrosion. The influence of the exfoliation on the compression strength is not as significant as that on the buckling load. To estimate the effect of the exfoliation on the remaining fatigue life of the corroded specimen, the sott inclusion based FE analyses were carried out for both corroded and pristine specimens under tension loading to examine the local stress concentration induced by the corrosion damage. Fig. 11 shows the maximum principal stress distribution for the centre portion of the corroded (C141-2-20A) and pristine (Cl41-222) specimens at a 22.24 kN (5000 lb) tension load. This figure reveals several stress concentrations on the bottom surface of corrosion damage with the maximum principal stress value of 244 MPa (35.4 ksi), which is about 40% higher than that of the pristine specimen (175 MPa (25.4 ksi)). This suggests that the effect of exfoliation on the remaining fatigue life of the corroded specimen could be significant. It should be noted that the local stress concentration is affected by the local damage 3D profile, which depends on the resolution and accuracy of an NDI technique. Since the existing NDI techniques have limited resolution for detecting small corrosion pits, destructive inspection (DI) measured 3D profile and detailed modeling, including pit modeling, may be needed for reliable fatigue analysis.
5. Concluding remarks In conjunction with static tests using specimens taken from upper wing skin panels, a "soft inclusion" based FE model was developed to evaluate the residual strength of a corroded specimen. The 3D FE model was first verified with the test results from pristine specimens. For the corroded specimens, an approach to automatically generate the 3D geometry of the damage zone (soft inclusion) using NDI input was developed
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(a) C141-2-20A (severe), bottom surface of corrosion damage (fringe unit: psi).
(b) C141-2-22 (pristine), same fringe definition as that used in (a). Fig. 11. Maximum principal stress distribution under 22.24 kN (5000 lb) tension load.
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with the aid of a PCL program. The parameters of a "soft inclusion" were calibrated using the tests results and the calibrated FE model was used to evaluate the buckling load, compression yield load, and stress concentration for the corroded specimens. The analysis results indicate that the effects of the exfoliation on the buckling and stress concentration of the specimen are more significant than that on the compression strength. Future work will validate the analytical model ft~her with additional compression and bearing tests using fastener holes containing exfoliation, and integrate the analytical model with corrosion pit/IG modeling to evaluate the remaining fatigue life.
Acknowledgements This work has been performed with the support of Defence Research and Development Canada (DRDC) and National Research Council Canada, project on Exfoliation Analysis and Repair Tool. Thanks to Mike Brothers for carrying out the ultrasonic inspections, Tom Benak, and Tony Marincak for conducting the static tests. Special thanks to Graeme Eastaugh for his efforts to obtain permission from Lockheed Martin Aeronautics for using the wing skin panels.
References [1] M.O. Speidel, M.V. Hyatt, Stress-corrosion cracking of high-strength aluminum alloys, in: M.G. Fontana, R.W. Staehle (Eds.), Advances in Corrosion Science and Technology, vol. 2, Plenum Press, New York, 1972, pp. 115-335. [2] W. Wallace, D.W. Hoeppner, Aircraft Corrosion: Causes and Case Histories, AGARD Corrosion Handbook, vol. 1, AGARD-AG-278, Neuilly-Sur-Seine, 1985, p. 93. [3] M. Worsfold, The effect of corrosion on the structural integrity of commercial aircraft structure, Presented at the Workshop on Fatigue in the Presence of Corrosion organized by the NATO Applied Vehicle Technology Panel, held October 7-8, 1999, Corfu, Greece. [4] J.P. Komorowski, N.C. Bellinger, R.W. Gould, D.S. Forsyth, G. Eastaugh, Can. Aeron. Space J. 47 (2001) 275-288. [5] M. Liao, J.P. Komorowski, Effect of prior exfoliation corrosion on fatigue and fracture behavior of aging aircraft structures and materials, LTR-SMPL-2002-0136, Institute for Aerospace Research, National Research Council, Ottawa, 2002. [6] N.C. Bellinger, J.P. Komorowski, M. Liao, D. Carmody, T. Foland, D. Peeler, Preliminary study into the effect of exfoliation corrosion on aircraft structural integrity, in: Proc. 6th Joint FAA/DoD/NASA Conference on Aging Aircraft, held 2002, San Francisco, USA, (CD-ROM). [7] M. Liao, N.C. Bellinger, J.P. Komorowski, Int. J. Fatigue 25 (2003) 1059-1067. [8] M. Liao, T. Benak, D. Backman, M. Brother, P. Vesely, T. Marincak, Static testing of C141 upper wing skin specimens containing exfoliation corrosion - preliminary results, LTRSMPL-2004-0016, Institute for Aerospace Research, National Research Council, Ottawa, 2004. [9] Y. Xiong, C. Poon, P.V. Straznicky, H. Vietinghoff, Comp. Struct. 30 (1995) 357-367. [10] N.C. Bellinger, Y. Xiong, C. Poon, Determination of finite width correction factors for composite laminates containing impact damage, LTR-ST-1899, Institute for Aerospace Research, National Research Council, Ottawa, 1992.