Numerical and Experimental Study of Defects ...

NATO UNCLASSIFIED REALEASABE TO AUS AND SWE

Numerical and Experimental Study of Defects Evolution in a Composites Wing Box Under Compressive Loads R. Borrelli Aerospace Structures and Materials Laboratory – CIRA (Italian Aerospace Research Centre) ITALY [email protected]

U. Mercurio Aerospace Structures and Materials Laboratory – CIRA (Italian Aerospace Research Centre) ITALY [email protected]

D. Tescione Aerospace Structures and Materials Laboratory – CIRA (Italian Aerospace Research Centre) ITALY

[email protected]

B. Gambino Analysis Methods Development – Alenia Aeronautica S.p.A ITALY

[email protected]

A. Riccio Aerospace and Mechanical Engineering Department – Secondà Università di Napoli ITALY [email protected]

ABSTRACT Due to their high specific strength and stiffness, Carbon Fibres Reinforced Plastics (CFRP) are commonly considered suitable for aerospace structural applications. However, their failure mechanisms are not completely predictable and this is the main reason why the CFRP integration in the aerospace industry has been generally slow in the last twenty years. Indeed, the lack of robust numerical tools, able to take into account the damage tolerance of composite structures, has led to over-conservative designs, not fully realising the promised economic benefits of composites materials. The Project MACMES (Damage Management of Aircraft Composite Structures Monitored by Embedded Sensors), funded by the General Defence Secretariat/National Armaments Directorate, of the Italian Ministry of Defence, in the framework of the National Military Research Plan (PNRM), addresses this issue by suggesting an integrated approach consisting in a synergy between numerical predictive tools (FEM models) and experimental ones (embedded optic fibres) for the damage management of aircraft composite structure.Within the MACMES project, such integrated approach was applied to a composite wing-box in order to monitor the buckling and the internal damage evolution under compressive loads. The experimental compressive test was performed by ALENIA STO-MP-AVT-211

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Numerical and Experimental Study of Defects Evolution in a Composites Wing Box Under Compressive Loads

under displacement control. Two defects were artificially included into the wing box: a skin/stringer debonding and an embedded circular bay delamination. Position and size of the delamination and skinstringer debonding were chosen as a result of a sensitivity analysis performed during the preliminary design phase and aimed at determining the best configuration (in terms of defects size and position) which guarantees a satisfactory experimental measurement of the damage growth from initiation to the global buckling load and well before the failure load of the wing box. The non-linear post-buckling behaviour of the damaged composite structure was simulated by developing appropriate FE numerical models. The adopted numerical models use the Virtual Crack Closure Technique to simulate the inter-laminar damage evolution and the numerical analyses have been performed by means of the FEM code ABAQUS and B2000++. The obtained numerical results have been assessed and compared to each other in terms of delaminated area evolution, delamination growth initiation load and strain distributions in order to investigate the effectiveness of the adopted numerical platforms in predicting the evolution of inter-laminar damages. Comparisons with experimental data, in terms of load displacement curves and strains in the dedonding area, are presented to assess the accuracy of the numerical simulations.

1.0 INTRODUCTION Due to their high specific strength and stiffness, CFRP are commonly considered particularly suitable for aerospace structural applications. However, their failure mechanisms are complex and not completely predictable (especially the inter-laminar failure mechanism such as the delaminations) and this is the main reason why the CFRP integration in the aerospace Industry has been generally slow in the last twenty years. Indeed the lack of robust numerical tools, able to take into account the damage tolerance of composite structures, has led to over-conservative designs, not fully realizing the promised economic benefits of Composites. The Italian National MoD funded project MACMES, running under the PNRM (Programma Nazionale di Ricerca Militare) framework addressed this issue by suggesting an integrated approach for the inter-laminar damage management of aircraft composite structures monitored by embedded optic fibers. The mechanical behavior of composite structures with inter-laminar damages has been widely investigated in literature by experimental tests and by predictive numerical tools. Complex numerical models able to simulate the growth of delaminations are presented in [1]-[6] where the Virtual Crack Closure Technique (VCCT) [7], the Modified Virtual Crack Closure Technique (MVCCT) and the Cohesive Zone Models (CZM) are adopted for the calculation of the Energy Release Rate by means of proper interface elements and suitable criteria are used to identify the growth /no-growth status. In this paper, the structural behavior of a composite wing box under compressive load has been numerically analyzed taking into account the simultaneous presence of an artificial bay delamination and an artificial skin/stringer debonding. The numerical results obtained from two different advanced FEM models developed, respectively, in ABAQUSTM and B2000++, have been compared to each other and to the experimental data produced by the MACMES project in order to test the effectiveness of the numerical simulations and to provide an insight in the physical phenomena related to the inter-laminar failure mechanisms taking place in the damaged composite wing box. In section 2 the geometrical configuration of the wing box is presented and the Finite Element Models implemented in ABAQUSTM and B2000++ are described in detail. In section 3 the numerical results for the two models have been compared, in terms of delamination and skin-stringer debonding growth initiation loads and shapes. The strains at various locations of the wing-box have been analysed and compared to experimental data..

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NATO UNCLASSIFIED REALEASABE TO AUS AND SWE CIRA-TR-13-0113 Rev. 0 P. 2/12



2.0 TEST AND NUMERICAL SIMULATION In this section, a description of the investigated wing box geometrical and material configuration is given together with the boundary conditions applied during the experimental test. The FEM models developed in B2000++ and ABAQUSTM are also illustrated providing detailed information on the simulation of interlaminar damages propagation.

2.1

Experimental Test

The geometrical features of the investigated wing-box configuration are given in figure 1 and table 1, while the material properties of the adopted material system are summarized in table 2. The stacking sequence of the wing box’s skin is [(0/+45/-45/90)2]S, while the stacking sequence of stringers foot and web are respectively [0/+45/-45/90]S and [0/+45/-45/90/90/-45/+45/0]S. In figure 1, the angle θ adopted to univocally identify a location along the delamination front, is also introduced. The artificial delamination is positioned between the 13th and the 14th ply while the artificial skin-stringer debonding is located in the first stringer as shown in figure 1. Figure 2 introduces the loading and boundary conditions applied to the structure. The wing box is considered compressed by means of applied displacements allowing the realistic simulation of the displacement controlled experiments. One edge of the structure is clamped while the opposite one is subjected to an applied uniform displacement uy keeping the rotations and the out of plane displacements blocked.

Figure 1: Geometrical description of the wing box.

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Table 1: Geometrical parameters of the wing box configurations.

SYMBOL

DESCRIPTION

VALUE

UNIT

Lx Ly Xc Yc Length Sd Sl Sb Wh Ft Wt Pt Dy Dx D Dt

WingBox Length WingBox Width Position of skin-stringer debond along Length Position of skin-stringer debond along Width Skin-Stringer Debond Length Distance between stringer feet Distance between the side edge and the stringer Width of stringer foot Height of the stringer Web Number of layers of stringer foot Number of layers of stringer web Number of layers of skin Position of delamination along Length Position of delamination along Width Delamination size Delamination depth

635 560 317.5 500.5 200 58 15 89 112.024 8 plies 16 plies 16 plies 133 127 50 13 plies

[mm] [mm] [mm] [mm] [mm] [mm] [mm] [mm] [mm] [#] [#] [#] [mm] [mm] [mm] [#]

Table 2: Material properties of the composite lamina used in wing box configuration.

Property

Average Value

Longitudinal Young’s modulus, E11 (GPa) - Tension

156

Transverse Young’s modulus, E22 = E33 (GPa) - Tension

8.35

In-plane shear modulus, G12 , G13 (GPa)

4.2

In-plane shear modulus, G23 (GPa)

2.52

Poisson’s ratio, n12 n13

0.33

Poisson’s ratio, n23

0.55

Critical ERR-Mode I, GIC [40mm] (Jm-2)

288

-2

Critical ERR-Mode II, GIIC [60mm] (Jm )

610

Critical ERR-Mode III, GIIIC (Jm-2)

610

Longitudinal tensile strenght Xt (MPa)

2500

Longitudinal compressive strenght Xc (MPa)

-1400

Transvese tensile strenght Yt = Zt (MPa)

75

Transverse compressive strenght Yc = Zc (MPa)

-250

Shear strength S12 = S13 (MPa)

95

Shear strength S23 (MPa)

108 -1

Thermal expansion coefficient α (°C )

-1.0e-6

Ply thickness , (mm)

0.186

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Clamped Edge

Z Y

Applied Displacement X

Ux=0 Uy=U0 Uz=0 ROTx=0 ROTy=0 ROTz=0

Ux=0 Uy=0 Uz=0 ROTx=0 ROTy=0 ROTz=0

Figure 2: Wing box FEM model - Applied load and boundary conditions.

2.2

B2000++ FEM model

B2000++ is a research oriented FEM code, written in C++, whose source files are available to users. Hence users are able to integrate in-house routines in the code at element, processor and solution algorithms levels. This is an unique feature which allows to easily implement new models in the frame of a numerical platform applicable to industrial case studies. B2000++ is owned by the major aerospace European research centers which collaborate, often in the frame of European financed projects, at the code development. In the B2000++ FE model the presence of both defects, skin/stringer debonding and embedded circular delamination, has been considered in the frame of a non- linear analysis aimed at verifying the influence of the two defects on their evolution. All the components of the wing box were modeled by using 20-noded layered solid elements as shown in figure 3. Contact elements were placed on the initial skin-stringer debonded zone and on the initial delaminated area in order to avoid penetrations, respectively, between the skin and stringer foot and the delaminated sub-laminates . Ad-hoc developed VCCT interface elements were placed in the skin-stringer debonding propagation zones and in the delamination propagation zone with the aim of allowing the propagation of both defects.

Figure 3: B2000++ FE model (with skin-stringer debonding and delamination)

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2.3

ABAQUS FEM model

In ABAQUSTM the wing box has been modeled with a combination of brick and shell elements. Eight nodes layered elements, SC8R (Continuum Shell element with 8 nodes), have been used to model the stringers and the adjacent portion of the skin while, for the rest of the skin, S4R elements (Shell elements with 8 nodes and reduced integration scheme) have been adopted. The stringers and the skin have been connected to each other with surface to surface tie constraints available in ABAQUSTM. The solid skin has been connected to the rest of the wing box with “solid-shell” coupling. To simulate the full debonding of the stringer, the two green surfaces surrounding the skin-stringer debonding, represented in Figure 4.a, have been connected by means of “node-to-surface” interface elements with the option “Virtual Crack Closure Technique”. The delaminated region has been modeled by two brick volumes representing the delaminated sub-laminates. In the surrounding area eventually involved by the delamination growth, the two volumes, representing the two sub-laminates, are connected by node to surface interface elements with the option “Virtual Crack Closure Technique” shown in Figure 4.b in green. The solid elements, used to represent the delamination area, have been connected to the shell model of the rest of the skin by the ABAQUSTM constraint type “Shell-to-solid Coupling”.

b)

a)

Figure 4: Node-to-Surface interaction zone modeled with VCCT interface elements: a) skinstringer debonding area; b) delamination area.

A representation of the ABAQUS Finite element discretization is shown in Figure 5.

Figure 5: FEM model of the wing box - ABAQUS.

3.0 RESULTS AND COMPARISONS In this section, B2000++ and ABAQUS numerical results are presented and compared to available experimental data. B2000++ non-linear analysis allowed to simulate the physical phenomena related to delamination buckling and growth, skin-stringer debonding buckling and growth and Global skin buckling. The delamination 1-6

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Numerical and Experimental Study of Defects Evolution in a Composites Wing Box Under Compressive Loads buckling, the skin-stringer debonding buckling and the global skin buckling modes are shown, as deformed shapes with out-of-plane displacements contour plots, in Figure 6.

Figure 6: Buckling Events at various strain level – B2000++.

The skin-stringer debonding status at the debonding propagation load and the delamination status at the delamination propagation load in B2000++ FE model are shown in Figure 7.

Figure 7: Skin-stringer debonding status at 1274 με and delamination status at 2090 με - B2000++

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Numerical and Experimental Study of Defects Evolution in a Composites Wing Box Under Compressive Loads From figure 7, it can be noted that the skin-striger debonding growth initiation in B2000++ takes place at the center of the stringer in the upper debonding front at a load level of 1274 με. On the other hand, the delamination growth initiation takes place at an angle of 67.5° on the delamination front for a load level of 2090 με. In Figure 8, the non-linear deformed shapes with out-of-plane displacements contour plot obtained by means of a geometrically non-linear analysis ABAQUSTM , at delamination buckling, skin-stringer debonding buckling and global skin buckling load, are shown. DELAMINATION BUCKLING AT 433 με

DEBONDING BUCKLING AT 1064 με

GLOBAL BUCKLING AT 2388 με

Figure 8: Buckling Events at various strain level – ABAQUS.

Comparing figure 8 with figure 6, the excellent agreement, in terms of deformed shapes and relative load levels between B2000++ and ABAQUS results, can be noticed. The skin-stringer debonding status at the skin-stringer debonding propagation load and the delamination status at the delamination propagation load in ABAQUS FE model are shown in Figure 9. From figure 9 it is clear that the growth initiation in ABAQUS takes place at the center of the stringer in the upper debonding front at a load level of 1310 με. On the other hand, the delamination growth initiation, in ABAQUS, takes place at an angle of 72° on the delamination front for a load level of 2187 με. ABAQUS results are again in excellent agreement with the B2000++ results shown in figure 7.

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Loading Direction DEBONDING PROPAGATION LOAD at 1310 με

DELAMINATION PROPAGATION LOAD at 2187 με

Loading Direction

72°

Figure 9: Skin-stringer debonding status at 1310 με and delamination status at 2187 με ABAQUS.

A summary of the main non-linear analysis results obtained with B2000++ and ABAQUS FE models is reported in Table 3, in terms of applied strain, and in Table 4 in terms of applied load. Delamination buckling strain

Debonding buckling strain

Debonding propagation strain

Delamination growth strain

Global buckling strain

εdel [με]

εdeb [με]

εdebp [με]

εini [με]

εbuck [με]

B2000++ Results

390

1145

1274

2090

2381

ABAQUS Results

433

1064

1310

2187

2388

Table 3: B2000++ and ABAQUS results summary - applied strain

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Delamination buckling load

Debonding buckling load

Debonding propagation load

Delamination growth load

Global buckling load

Fdel [kN]

Fdeb [kN]

Fdebp [kN]

Fini [kN]

Fbuck [kN]

B2000++ results

130

390

430

693

777

ABAQUS Results

144

356

436

698

757

Table 4: B2000++ and ABAQUS results summary - applied load

From these last two tables, the good agreement between the two FE models can be established. In Figure 10 the comparison between the obtained numerical results and experimental data in terms of load vs. applied strain curves is presented.

Load vs Applied Strain

1400

Experimental Linear ABAQUS B2000++

1000 800 600 400

Applied Load [kN]

1200

200 0 ‐4000 ‐3500 ‐3000 ‐2500 ‐2000 ‐1500 ‐1000 ‐500 Applied Strain [με]

0

Figure 10: Load vs. applied strains – comparison between numerical results and experimental data

As it can be seen from figure 10, close to the global buckling event the wing box deviated from the linear trend and the global stiffness of the wing box starts to decrease. A good correlation has been found between experimental data and the numerical B2000++ and ABAQUS results. Indeed, the stiffness of the wing-box is well predicted in the pre-buckling and post-buckling phases, by both the numerical models. In Figure 11 the comparison between numerical results and experimental data in terms of strain at the center of the skin-stringer debonding vs. applied load curve is shown.

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Experimental

5000

B2000++

4000

ABAQUS

3000 2000 1000

Strain [με]

Applied Compressive Load vs Measured Strain ‐ Debonding Center

0 ‐1000 0

250

500 750 Applied Load [kN]

1000

‐2000 1250

Figure 11: Strain at center of the skin-stringer debonding vs. applied load

The experimental strain curve was measured by a strain gage located on the upper surface of the upper skin at the center of the skin-stringer debonding region. The strain is initially negative due to the compressive load. At the skin-stringer debonding buckling load the slope of the curve suddenly changes. At this point the upper skin starts to detach from the stringer foot and the stress state becomes a tensile stress state. The numerical curves were able to reproduce these main events with a good accuracy. The numerical B2000++ FE model slightly overestimate the skin-stringer debonding buckling load while ABAQUS is able to perfectly replicate the experimental behavior. This difference is probably due to the use of the continuum shell formulation in ABAQUS which is characterized by a better bending behavior with respect to the solid layered formulation adopted in B2000++. Concerning the post-buckling regime, the B2000++ model behaves substantially better than the ABAQUS model. The discrepancy between experimental data and numerical results in ABAQUS can be strictly related to the mesh and time step dependency of the VCCT interface elements. These elements needs to be tuned on experimental data to determine the right combination of time step and element size in order to provide reliable results in terms of delamination size as a function of the applied load. ABAQUS clearly overestimates the skin-stringer debonding propagation with respect to experimental data. Indeed, the stringer is almost completely separated from the skin a few load increments after the buckling load leading to a general underestimation of the tensile strains reported in figure 10 with respect to experimental data. The experimental observations showed that the skin-stringer debonding growth in the post buckling regime is limited. B2000++ provides numerical results which are in excellent agreement with experimental data in the post buckling regime, probably because the skin-stringer debonding is not allowed to grow behind the propagation area shown in figure 4. The mesh and time step dependency issue can be avoided by implementing the alternative and time costly VCCT based delamination growth model proposed in [8] and [9]. Both the FE models were found able to provide information on the compressive behavior of the investigated composite wing box as well as on the phenomena governing the inter-laminar failure mechanisms.

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4.0 CONCLUSIONS In the present paper, the compressive behavior of a composite wing box, with artificial inter-laminar damages, has been numerically investigated. Two FE models have been implemented in ABAQUS and B2000++ . These models have been demonstrated to be effective and accurate by comparing the obtained numerical results to experimental data. Both the models have been demonstrated to provide reliable results in terms of skin-stringer debonding buckling, skin-stringer debonding growth, delamination buckling, delamination growth and skin global buckling. However, the behavior of the skin-stringer debonding in the post-buckling regime has not been predicted correctly by the ABAQUS model due to a relevant overestimation of the skin-stringer debonding growth phenomenon. This overestimation was not allowed in the B2000++ model as the propagation area surrounding the skin-stringer debonding was restrained. Finally the proposed FE models have contributed to investigate and understand the phenomena related to the inter-laminar damage growth in the composite wing box. Some considerations have been outlined in the paper on the best FE formulation and delamination growth tool to be adopted to investigate the compressive behavior of composite wing boxes .

5.0 REFERENCES [1] Perugini P. , Riccio A. and Scaramuzzino F. Influence of Delamination Growth and Contact Phenomena on the Compressive Behaviour of Composite Panels. Int. Journal of Composite Materials 1999; 33(15):1433-1456. [2] Riccio A. , Perugini P. and Scaramuzzino F. Modelling Compression Behaviour of Delaminated Composite Panels. Computers & Structures 2000; 78:73-81. [3] Nilsson K. -F., Thesken J.C. , Sindelar P. , Giannakopoulos A. E. and Storakers B. A Theoretical and Experimental Investigation of Buckling Induced Delamination Growth. Journal of Mech. Phys. Solids 1993; 41(4):749-782. [4] Gaudenzi P. , Perugini P. and Riccio A. Post-buckling behaviour of Composite Panels in the presence of Unstable Delaminations. Composite Structures 2001; 51(3):301-309. [5] Riccio A. , Perugini P. and Scaramuzzino F. Embedded Delamination Growth In Composite Panels Under Compressive Load. Composites part B: Engineering 2001; 32(3):209-218. [6] Gudmundson, P. Micromechanically based constitutive models for damage evolution in composite laminates. International Journal of Damage Mechanics 2000; 9(1):29-39. [7] Krueger, R., The virtual crack closure technique: history, approach and applications, NASA/CR2002-211628 [8] A. Riccio, A. Raimondo and F. Scaramuzzino. A study on skin delaminations growth in stiffened composite panels by a novel numerical approach. Applied Composite Materials. DOI: 10.1007/s10443-0129282-7, 20 July 2012. [9] Elisa Pietropaoli, Aniello Riccio. On the robustness of finite element procedures based on Virtual Crack Closure Technique and fail release approach for delamination growth phenomena. Definition and assessment of a novel methodology. Composites Science and Technology 2010;70:1288–13003.

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