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Pull-off Tests and Analyses of Composite Skin and Frame T-Joint J. LI ABSTRACT Twenty pull-off specimens were tested to identify the mechanism and associated crucial factors that control the failure mode of frame to skin joint using angle clips stitched to the skin. The pull-off specimens were tested under both simply supported and clamped conditions. Four different support spans were used in either simply supported or clamped conditions. As expected, the support conditions and span lengths are greatly complicating failure analysis and pull-off allowable development based on pull-off load. As the test data indicate that the critical pull-off load reduces as the support span increases, and the critical pull-off loads are higher under clamped condition as compared with simply supported condition when the support span is the same. These behaviors greatly increase the cost of allowable development based on pull-off load and increase the likelihood of error in applying allowable defined in terms of pull-off load. On the other hand, the dependence of the critical bending moment on support span and support type is less significant. Which shows that the crack initiation is controlled by the maximum bending moment at the critical location. The mechanism behind the joint failure initiation is the maximum tension strain caused by bending. Allowable developed in terms of maximum critical bending moment is independent of the support condition and span length. The allowable bending moment developed from either clamped condition or simply supported condition becomes equivalent. Enormous savings of allowable development costs and greatly increased accuracy in analysis can be achieved by capturing the right crucial factor that governs the actual failure mode. Keywords: Composite angle clip joint, pull-off test, out-of-plane failure mode, delamination INTRODUCTION With the increased emphasis on reducing the cost of manufacturing composite structures, secondary bonding or co-curing is an attractive option to eliminate the need for mechanically fastening sub-assemblies. Many composite components in aerospace structures consist of flat or curved panels with co-cured frames and stiffeners. Out-of-plane loading such as internal pressure in a composite fuselage or out-of-plane deformations in a compression loaded post-buckled panel may cause the frame or stiffener to debond from the panel . The out-of-plane failure mode have been studied extensively using stiffener pull-off tests [1-12]. _____________ Jian Li, The Boeing Company, M530-B229, 5000 E. McDowell Road, Mesa, AZ 85215 Pull-off tests results of co-cured joints were analyzed to understand the failure mode and strength of attaching frames and bulkheads to the fuselage skin using angle clips . The angle clip joint is similar to blade stiffeners studied in Ref. 7. Ref. 12 indicated potential misrepresentation of pull-off strength when the clamps of the pull-off fixture get close to either the vertical flanges of the angle clip or the horizontal flange tip of the angle clip. The clamps pressed on the angle clip flanges introduced unrealistic constraints to the delamination failure mode in the noodle between the clip flanges and the skin that inflated the magnitude of the pull-off load at ultimate failure. Ref. 12 and 13 also indicated that the bending moment at the delamination front drives the failure initiation and growth from the corner region. To validate the analysis  and obtain the critical bending moment that characterizes the strength of the joint, twenty pull-off specimens were tested with four different support spans either under clamped conditions or simply supported conditions. In this paper, the results from these tests and their interrogations are presented. PULL-OFF TESTS The pull-off specimens were cut with diamond-tipped tool from a simple panel made of carbon fiber fabric and infused with the SI-ZG-5A resin system using the VARTM process . The fabrics were intermediate modulus AS4 with 3000 denier in the plain weave (PW) and 6000 denier in the 5- harness satin (5HS) plies. The symmetric skin lay-up was (±45PW, 0/905HS, ±45PW, 0/90PW)s. The angle clips were made of four plies of the same plain weave AS4 carbon fabric, all of which were laid-up in ±45° orientation. The angle clips were stitched to the skin before infusing. Pull-off Specimen Preparation and Test Setup Twenty 2-inch wide and 14-inch long specimens were cut from frame B (Figure 1) and milled to final finish. A schematic drawing of the top view of the simple panel and the nominal dimension of the pull-off specimen is shown in Figure 1. Each specimen was measured at three locations for width, three locations for thickness and one location for length as shown in Figure 2. Where W1, W2 and W3 were width measurements taken from the skin and flange regions. Thickness T1 and T2 measured the combined thickness of the skin and flange, while thickness measurement T3 represented the skin thickness. The length of the specimen was also measured and denoted as L on Figure 2. The results of these measurements are given in Tables 1 and 2 for simply supported (roller) specimens and clamped specimens, respectively. Also provided in the Tables are the test rates associated with these specimens when tested. Frame B 14.0 2 Figure 1 Simple panel and pull-off specimen dimensions (in). W1 W2 W3 Adhesive Noodle T1 T2 T3 L Figure 2 Specimen dimension measurement locations. Table 1 Dimensions and test rates for simply supported pull-off specimens Roller Width (in) Thickness (in) Length (in) Test Rate Specimen ID Span (in) W1 W2 W3 Average T1 T2 T3 L (in./min) 101 4 2.003 2.002 2.005 2.003 0.147 0.136 0.090 14.1 0.1 102 4 1.999 2.000 2.000 2.000 0.145 0.139 0.087 14.1 0.1 121 6 2.000 2.002 2.000 2.001 0.145 0.138 0.087 14.1 0.1 122 6 2.000 2.000 2.001 2.000 0.139 0.142 0.086 14.1 0.1 111 8 2.000 1.999 2.000 2.000 0.150 0.141 0.089 14.1 0.1 112 8 2.000 1.999 2.001 2.000 0.149 0.139 0.090 14.1 0.4 131 10 2.000 2.001 2.001 2.000 0.134 0.134 0.087 14.1 0.8 132 10 2.000 1.998 1.998 1.999 0.135 0.134 0.090 14.1 0.7 Table 2 Dimensions and test rates for clamped pull-off specimens Clamp Width (in) Thickness (in) Length (in) Test Rate Specimen ID Span (in) W1 W2 W3 Average T1 T2 T3 L (in./min) 151 4 1.999 1.999 1.999 1.999 0.141 0.130 0.088 14.1 0.02 152 4 1.992 1.992 1.995 1.993 0.133 0.132 0.089 14.1 0.05 153 4 1.997 1.998 2.001 1.999 0.133 0.137 0.087 14.1 0.05 161 6 1.999 1.999 1.998 1.999 0.136 0.131 0.087 14.1 0.09 162 6 2.001 2.001 2.000 2.001 0.146 0.131 0.089 14.1 0.09 163 6 1.999 1.998 1.998 1.998 0.146 0.129 0.086 14.1 0.09 171 8 1.998 1.997 1.997 1.997 0.135 0.138 0.090 14.1 0.18 172 8 1.995 1.997 1.996 1.996 0.145 0.136 0.087 14.1 0.16 173 8 1.993 1.995 1.998 1.995 0.139 0.139 0.087 14.1 0.16 181 10 1.939 1.939 1.942 1.940 0.141 0.131 0.083 14.1 0.32 182 10 1.991 1.992 1.990 1.991 0.133 0.149 0.092 14.1 0.26 183 10 1.509 1.510 1.508 1.509 0.131 0.145 0.085 13.9 0.20 The simply supported and clamped pull-off test set-ups are schematically shown in Figures 3 and 4, respectively. Span P Figure 3 Simply support pull-off specimen test set-up. Span Base Plate P Figure 4 Clamped support pull-off specimen test set-up. Measured Test Results All specimens were tested at ambient conditions. The laboratory temperature during testing was between 76º-80º F. The relative humidity level in the laboratory was between 44-48%. The test was discontinued when the load dropped 20% or more from the peek load. The loading rate of the static test could vary from specimen to specimen so that failure would occur within 5 to 10 minutes. Simply Supported Test Results A total of eight specimens were tested under simply supported conditions where two specimens were tested for each of four different spans. Specimens 101 and 102 were tested with a four-inch span. The load versus displacement behavior for specimens 101 and 102 are shown in Figure 5. Test records indicate that audible cracking sounds were first heard at 100 lb and 145 lb for specimens 101 and 102, respectively. Notable load drops at these levels can be observed from the load-displacement curves shown in Figure 5. To be specific, the indicated first notable load drop shown in Figure 5 is 101 lb for specimen 101 and 148 lb for specimen 102. Visual observations of cracks were detected at higher load levels (170 lb for specimen 101 and 176 lb for specimen 102). 300 4-inch span simply supported 250 Specimen 102 first drop 200 Load (lb.) 150 100 Specimen 101 first drop 50 0 0 0.1 0.2 0.3 0.4 0.5 0.6 Displacement (in.) Figure 5 Load-displacement curves for 4-inch span simply supported. The pull-off loads associated with initial audible cracking, load drop and visual crack for all simply supported specimens are tabled in Table 3. The pull-off load at initial audible cracking was at the same load level as the initial load drop with the exceptions of specimen 122 and 112. This coincidence is significant in confirming the failure initiation by using the first load drop or deviation from linearity from the load-displacement curve. Also provided in Table 3 are the ultimate pull-off loads recorded from tests. Where the ultimate failure is defined as a 20% or more load drop from the peak load. The sequence of events is first audible sound, initial load drop, visual crack and ultimate failure. Table 3 Initial audible, load drop, visual crack and ultimate for simply supported specimens Span (in.) 4 6 8 10 Specimen ID 101 102 121 122 111 112 131 132 Audible (lb.) 100 145 95 78 65.7 57 42 42 Load drop (lb.) 100.6 148.2 95.7 93.5 65.7 64.9 42.4 42.5 Visual (lb.) 170 176 123 99 81 81 50 43 Utimate (lb.) 214 260 149 143 91 92 60 64 500 4-inch span clamped 450 152 400 350 300 Load (lb.) 151 250 200 150 153 100 50 0 0 0.05 0.1 0.15 0.2 0.25 0.3 0.35 Displacement (in.) Figure 6 Load-displacement curves for 4-inch span clamped. Clamped Test Results There were twelve specimens tested under clamped conditions with three specimens tested at each span of 4, 6, 8 and 10 inches. The first audible, load drop, visual crack and ultimate failure loads are summarized in Table 4 for the clamped specimens. Reasonable agreement can be seen between the first audible crack and first load drop from the load-displacement curve for most of the specimens. Typical load-displacement curve can be seen from Figures 6 for clamped span of 4 inches. The first load drops are more identifiable in clamping spans of 4, 6 and 8 inches. It is much harder to identify the initial load drops in those 10-inch span specimens due to their strong nonlinear load-displacement behaviors. Table 4 First audible, load drop, visual crack and ultimate failure for clamped specimens Span (in.) 4 6 8 10 Specimen ID 151 152 153 161 162 163 171 172 173 181 182 183 Audible (lb.) 198 189 165 137 143 147 109 119 118 60 72 67 Load drop (lb.) 206.5 206.8 170.2 137.7 159.8 147.8 130.3 120.7 118.4 87.7 130.4 71.7 Visual (lb.) 297 291 310 200 280 243 164 170 170 140 160 136 Utimate (lb.) 460 433 406 307 322 329 235 314 276 210 276 184 Trends in Pull-off Test Results The simply supported critical pull-off loads associated with each of the four failure definitions (first audible cracking sound, initial load drop, first visual crack observed and ultimate failure) are shown in Figure 7. The critical pull-off load follows the trend of decreasing magnitude with increasing support span irrespective with the type of failure definition. The first audible crack and initial load drop are close to one another and representative of crack initiation. The pull-off load for initial visual crack is less important as the crack initiation is likely to occur at the center away from both edges of the pull-off specimen . The stitching seems to increase the ultimate failure load and maximum deflection after crack initiation. If the joint is designed by crack initiation, the stitching is a welcome feature for damage tolerance reserves. Whether the joint can be designed based on the ultimate pull-off load should depend on the application and be carefully reviewed. Similar to Figure 7, critical pull-off loads for clamped specimens are plotted in Figure 8. The trends are similar with the results of simply supported specimens. The critical pull-off loads are higher for specimens under clamped conditions than for those under simply supported conditions for the same span as seen from Figure 9. Since the difference in pull-off load varies significantly with the support conditions, allowable strength of a joint should not be based on pull-off load. A parameter controlling the failure mode that is independent of the span length and support condition is crucial to the design of this type of joints. If such a parameter can be identified, the joint strength allowable could be efficiently developed based on this parameter for design and analysis. Critical Pull-off Loads 300 Audible (lb.) 250 Load drop (lb.) Visual (lb.) Pull-off Loads (lb.) 200 Utimate (lb.) 150 100 50 0 3 4 5 6 7 8 9 10 Simply Supported Span (in.) Figure 7 Critical pull-off loads by four failure definitions for simply supported specimens. Critical Pull-off Loads 500 Audible (lb.) 450 Load drop (lb.) 400 Visual (lb.) Utimate (lb.) Pull-off Loads (lb.) 350 300 250 200 150 100 50 0 3 4 5 6 7 8 9 10 Clamped Span (in.) Figure 8 Critical pull-off loads by four failure definitions for clamped specimens. Crack Initiation Pull-ff Loads 250 Clamp supports 200 Roller supports Pull-off Loads (lb.) Clamped = -15.95x + 252.32 150 100 Simply = -13.758x + 177.99 50 0 3 4 5 6 7 8 9 10 Support Span (in.) Figure 9 Crack initiation loads for clamped and simply supported conditions. Maximum Bending Moments at Crack Initiation 100 90 Clamped = 1.9049x + 54.226 80 Moment/Width (lb-in./in.) 70 60 50 Simply = -1.6454x + 74.393 40 30 Clamp supports 20 Roller supports 10 0 3 4 5 6 7 8 9 10 Support Span (in.) Figure 10 Trend of maximum bending moments at crack initiation. Critical Bending Moment The crack initiation pull-off load decreases linearly as the support span of the pull-off specimen increases as shown in Figure 9 for either clamped or simply supported conditions. The bending moments at mid span at crack initiation defined by the first load drop are shown in Figure 10 for clamped and simply supported conditions. The calculation of the mid-span bending moment for simply supported pull-off specimen is M simply P L = (1) W W 4 Where P is the pull-off load and L is the span length. The bending moment is normalized by the specimen width, W. The bending moment at the mid-span for clamped condition is evaluated from M clamped P L f s 1 − D11 ≈ 1 + (2) W W 8 L sf D11 Where f is the length between the tips of the horizontal angle clip flanges (or, approximately from the S middles of the taper regions if the flanges taper down at the tips) D11 is the skin bending stiffness and sf D11 is the bending stiffness of the composite skin and flange. The dependence of the critical bending moment on support span and support type is insignificant as seen from Figure 10. The results shown in Figure 10 indicate that the initiation of crack is controlled by the maximum bending moment at the critical location. Figure 10 also shows the advantage of developing pull-off strength in terms of critical bending moment by eliminating the span length effect. The allowable bending moment developed from either clamped condition or simply support condition becomes less important. There are eight simply supported data points with an average of 62.9 in.-lb/in and a standard deviation of 9.4 in.-lb/in. There are twelve clamped data points with an average of 67.6 in.-lb/in and a standard deviation of 8.8 in.-lb/in. The average of the twenty data points is 65.7 in.-lb/in with a standard deviation of 9.1 in.-lb./in. The averages and standard deviations are very close among the three different sample methods. If the numbers of data points are equal, equivalent allowable could be developed from any of the sample methods. The test data validated that the crucial factor controlling the joint corner failure initiation is the bending moment and showed that the pull-off test under simply supported condition and clamped condition are similar in failure mechanism and equivalent in allowable development. The allowable bending moment is potentially significant for unitized composite structures since it serves as a crucial link between the joint failure model and global nonlinear analysis that constitutes the high fidelity failure analysis methodology outlined in Ref. 13. CONCLUSION Tests and analyses of stitched pull-off specimens show that crack initiation is controlled by the maximum bending moment at the critical location. The mechanism behind the joint failure initiation is the maximum tension strain caused by bending. Allowable developed in terms of maximum critical bending moment is independent of the support condition and span length. The allowable bending moment developed from either clamped condition or simply supported condition becomes equivalent. Enormous savings of allowable development costs and greatly increased accuracy in analysis can be achieved by capturing the right crucial factor governing the actual failure model. On the other hand, the support conditions and span lengths greatly complicate failure analysis and allowable based on pull-off load. As the test data indicate the critical pull-off load reduces as the support span increases, and the critical pull-off loads are higher under clamped condition than under simply supported condition for the same support spans. The dependence of pull-off load on support span and support conditions greatly increases the cost of allowable development based on pull-off load and the likelihood of error in design when this allowable is applied. ACKNOWLEDGEMENTS This work was supported with shared funding by the U. S. Rotorcraft industry and government under the RITA/NASA Cooperative Agreement Contract No. NCC2-9019, under WBS No. 01-B-03-7.2-P2 for Structural Joining Technologies. The contributions of Integrated Technologies Corporation in specimen preparations and testing are greatly appreciated. REFERENCES 1. Minguet, P. J., Fedro, M. J., O’Brien, T. 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And Chen, T. “High Fidelity Failure Analysis for A Composite Fuselage Section,” Proceedings of the American Helicopter Society 57th Annual Forum, Washington, DC, May 9-11, 2001.
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