Residual Strength and Fatigue Lifetime of ... - Solid Mechanics

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Hayman (2007) has described a damage assessment procedure for sandwich structures, which was originally developed for naval ships, but has potential for application to other structures with similar construction. The procedure exploits the fact that sandwich structures in ship’s hulls are to a large extent built up of a limited number of fairly large, flat panels that are supported at their edges. These panels are subjected primarily to local transverse pressure loadings and to in-plane loadings which are associated with global bending of the hull girder. Many cases of production defects and in-service damage initially involve a region of the panel that is small compared to the length and breadth of the panel. In such cases damage assessment requires use of a local strength reduction factor Rl, defined as the ratio between the far-field applied stress (or strain) causing failure in the presence of the damage to the corresponding value in the absence of the damage. The effect of the damage on the strength of both the panel and the structure as a whole may be determined from this local strength reduction factor Rl, which depends on the type and size of the damage, the material lay-up and the predominant stress state (tension, compression, shear) in the damage location. In this chapter, the above-mentioned damage assessment approach is adopted and strength reduction factors are determined for debond damaged sandwich panels under uniform compression loading, by a combination of finite element modelling and testing. Uniform compression tests were conducted on intact sandwich panels with three different types of core material (H130, H250 and PMI) and on similar panels with circular face/core debonds having three different diameters. The strains and out-of-plane displacements of the panel surface were monitored using the digital image correlation (DIC) technique. Mixed Mode Bending (MMB) tests were conducted to determine the fracture toughness of the interface of the panels for a full range of negative mode-mixities. Finite element analysis and linear elastic fracture mechanics were applied to determination of the critical buckling load and compression strength of the panels. Finally, the modelling approaches and failure criteria are discussed. Numerically determined crack propagation loads in most of the cases show fair agreement with experimental results, but in a few cases up to 45% deviation is seen between numerical and experimental results. This can mainly be ascribed to the large scatter in the measured interface fracture toughness and differing crack tip details. Tentative strength reduction curves are presented, but uncertainty concerning the intact strengths of the applied materials needs to be removed before these can be used with confidence. 3.2 Test Specimens Uniform compression loading tests were conducted on intact sandwich panels and panels with a predefined debond. A total of 30 panels were manufactured, each 460 mm long and 380 mm wide. The face sheets were of glass-reinforced plastics (GFRP) consisting of Devold AMT noncrimp fabrics and two different types of vinylester resin with the following lay-ups: 40
Type A: Three layers of DBLT-850 quadriaxial (0/90/+45/-45) glass with Dion 9102 vinylester Type B: As type A but with Dion 9500 rubber-modified vinylester Type C: Three layers of DBL-800 triaxial (0/0/+45/-45) glass with Dion 9102 vinylester In addition, a layer of chopped strand mat (CSM) was placed between each face sheet and the sandwich core. The core materials were PVC (H130 and H250) and PMI (51-IG) foams. The thicknesses of the core and the face sheets were 30 mm and approximately 2 mm, respectively, see Figure 3.1. All panels were reinforced with wooden inserts at the top and bottom edges of the panel to avoid crushing of the core at the loaded edges. The panels were resin injection molded and cured with vacuum consolidation. Furthermore, the top and bottom edges of the panels were machined straight and parallel following the specimen manufacturing. Figure 3.1: Geometry of panel specimens and an image of a manufactured panel with a debond diameter of 200 mm. Debond defects with three diameters (100, 200 and 300 mm) were introduced during the manufacturing process by laying a circular piece of 0.025 mm thick Airtech release film on the core and sealing the edges with resin before lamination. However, after manufacturing of the panels it was observed that the face sheet was partially bonded to the core in the debonded area, probably due to small perforations of the release film. To release the partial adhesion, a small hole with a 2 mm diameter was drilled into the backside through the thickness of the panels and the debonded face sheet was pushed, using a thin metallic bar, to a point where all partial cohesions were eliminated. The panel test specimens are listed in Table 3.1. Fracture mechanical characterisation tests were performed in connection with the debond studies. For these tests, special sandwich beam specimens were manufactured with face sheet lay-ups of type A, B and C and cores of a thickness of 10 and 20 mm. 41
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Residual Strength and Fatigue Lifet
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Published in Denmark by Technical U
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method to accelerate the simulation
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Synopsis Sandwich kompositter er i
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Page 14 and 15: Contents Preface Executive Summary
Page 16 and 17: 5 Face/core Interface Fatigue Crack
Page 20 and 21: H11 bimaterial constant H22 bimater
Page 24 and 25: These peculiar damage modes often r
Page 26 and 27: 1.2 Overview of the Thesis In this
Page 28 and 29: Figure 1.3: Fracture modes, from Be
Page 30 and 31: In Equations (1.5) and (1.6) H11, H
Page 32 and 33: Figure 1.6: Schematic illustration
Page 34 and 35: The compact tension specimen (CT) i
Page 36 and 37: A Figure 1.10: CSB, DCB, TSD, DCB-U
Page 38 and 39: Chapter 2 Buckling Driven Face/Core
Page 40 and 41: (DIC) measurement system (ARAMIS 2M
Page 42 and 43: corresponds to the onset of debond
Page 44 and 45: Figure 2.7: Crack kinking into the
Page 46 and 47: (2.2) where and for plane str
Page 48 and 49: A modified version of the tilted sa
Page 50 and 51: previous observations of crack path
Page 52 and 53: of contact elements (CONTACT173 and
Page 54 and 55: the debond opening initially increa
Page 56 and 57: Table 2.4: Instability loads determ
Page 58 and 59: G (J/m2) 600 400 200 0 H100 IMP=0.1
Page 64 and 65: Table 3.1: Panel test specimens. La
Page 66 and 67: sheet and the core thickness, respe
Page 68 and 69: Load (N) 250 200 150 100 50 0 H130
Page 70 and 71: Table 3.4: Parameters in the face/c
Page 72 and 73: was used to monitor 3D surface disp
Page 74 and 75: arising due to the junction between
Page 76 and 77: (a) Debonded face sheet Figure 3.16
Page 78 and 79: Figure 3.19 shows the determined en
Page 80 and 81: H130 MMB H130 Panel H250 MMB Interf
Page 82 and 83: 3.6 Conclusion In this chapter the
Page 86 and 87: composite laminates under thermal c
Page 88 and 89: S S y( t ) y( t ) ( t ) 2 1 12 2
Page 90 and 91: listed in Table 4.1. The length and
Page 92 and 93: The strain energy release rate, G,
Page 94 and 95: high growth rate of G (see Figure 4
Page 96 and 97: 4.4 Face/Core Fatigue Crack Growth
Page 98 and 99: Debond 310 mm Symmetry B. C. a Figu
Page 100 and 101: B 2 1/ 2 ( S44 S55 S45) (4.17) wh
Page 102 and 103: 75), which illustrates possible ina
Page 104 and 105: Figure 4.18 (a) presents the deflec
Page 106 and 107: Debond radius (mm) 100 90 80 70 60
Page 108 and 109: using the cycle jump method, more t
Page 110 and 111: Chapter 5 Face/Core Interface Fatig
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of this chapter, sandwich panels wi
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H250 Specimen H100 Specimen H45 Spe
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Figure 5.5: Test setup. Initially,
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H100 Specimen Fibre bridging Figure
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propagation, the crack continues to
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Figure 5.14: Kinking of the crack i
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Crack length [mm] Crack length [mm]
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mode-mixity phase angle was chosen
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should be taken into account for a
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Figure 5.25: Kinking of the crack i
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log (da/dN) (mm/cycle) 10 log G (J/
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erroneous extrapolations in the tra
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compared to the 65% efficiency obta
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the case of uneven debond growth, t
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Load (kN) 2 1.5 1 0.5 0 Panel 1 Pan
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Figure 5.42: Zero and ninety degree
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G(J/m 2 ) 180 150 210 120 150 100 5
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110 q=qG=0.4 Test #1 100 Test #2 Te
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panels were determined for differen
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Chapter 6 Conclusion and Future Wor
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toughness using MMB fracture toughn
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For the specimens with H250 core th
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Development of testing methods for
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Bezazi A., Mahi A. E., Berthelot J.
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Kanny K. and Mahfuz H. (2005), Flex
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Ratcliffe J. and Cantwell W. J. (20
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Out-of-plane displacement (mm) 1 0.
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Out-of-plane displacement (mm) 3 2
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A.5 Out-of-plane deflection of Debo
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(a) Figure A.14: Out-of-plane defle
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(a) (b) Figure A.18: Out-of-plane d
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Load (kN) 250 200 150 100 50 250 30
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Displacement (mm) 4 3 2 1 0 8 9 (a)
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(a) (b) Figure B.11: Out-of-plane d
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(a) (b) Figure B.15: Out-of-plane d
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DTU Mechanical Engineering Section
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