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SOFT BODY IMPACT SIMULATION ON COMPOSITE STRUCTURES

M. Nejad Ensan\ D.G. Zimcik\ M, Lahoubi2, D. Andrieu2

lInstitute for Aerospace Research, National Research Council Canada,2Institut Catholique d'Arts et Metiers, Lille, France

Contact: [email protected]

Received November 2007, Accepted July 2008No. 07-CSME-66, E.I.C. Accession 3035

ABSTRACTThe paper describes recent progress on numerical simulation of soft body impact onto a fibre

reinforced composite wing leading edge structure. The work is based on the application of non-linearexplicit finite element analysis to simulate the response of composite wing structures under soft bodyimpact loads. Soft body impactors such as gelatine (substitute bird) or ice (hailstones) are highlydeformable on impact and flow over the structure spreading the impact load. Therefore, first benchmarksimulations were carried out for soft body impact onto a rigid target. Soft body impactor was modeledusing the Arbitrary Lagrangian-Eulerian (ALE) method. The results obtained using this impact model fordifferent velocity were compared to the experimental test results in terms of local pressure, includingHugoniot and stagnation pressures, and global load to validate the accuracy of the model. Then, theimpact of soft body onto a composite wing structures was described. A composite failure model whichincludes ply damage and interplay delamination model has been used to predict impact damage in thestructure modeled using shell elements. The simulation tool predicts the impact damage in leading edgestructure.

SIMULATION DE L'IMPACT DE CORPS DOUX SURDES STRUCTURES COMPOSITES

RESUMEL'article decrit Ie recent progres de la simulation numerique de l'impact de corps doux sur une structure

de bord d'aile en materiaux composites renforces. L'etude est basee sur l'utilisation de la Methode desElements Finis explicite non lineaire pour simuler la reponse des structures composites des ailes sous deschargements d'impact de corps mous. Les compacteurs de corps mous tels que la gelatine (pourrepresenter l'oiseau) ou la glace (pour representer les grelons) sont fortement deformables a l'impact etcirculent sur la structure ecartant Ie chargement d'impact. Par consequent, les premieres simulations dereference ont ete effectuees pour l'impact de corps mous sur une cible rigide. Le compacteur de corpsmou a ete modelise en utilisant la methode Lagrangienne-Eulerien Arbitraire (ALE). Les resultats obtenusen utilisant ce modele d'impact pour differentes vitesses ont ete compares aux resultats d'essaisexperimentaux en termes de pressions locales, y compris Hugoniot et pressions de stagnation, et auchargement global afm de valider la precision du modele. Puis, l'impact du corps mou sur des structuresd'aile composite a ete decrit. Un modele de rupture composite incluant les dommages de pli et Ie modeled'effet de decollement a ete employe pour prevoir les dommages d'impact dans la structure modelisee enutilisant des elements de plaque. L'outil de simulation prevoit les dommages d'impact sur la structureprincipale de bord d'aile.

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1. INTRODUCTION

Bird strikes are a significant threat both to civil and military aircraft and as such, take animportant place in the aircraft certification process. Also, fibre reinforced composite materialsare increasingly being used in primary aircraft structures such as wing leading edges, nacelles,propellers and aero-engine ducts and fan blades. These parts are susceptible to impacts from birdstrike or foreign objects that can pose a serious risk to aircraft safety. Before certification, anaircraft must demonstrate the ability to land safely after being struck by a two-kilogram birdanywhere on the structure, during flight at normal operating speeds. Impacted components mustmaintain structural integrity during the large transient loading resulting from bird strike loads.Past experience has been to demonstrate this compliance through full-scale test although there isa desire to improve modelling capabilities and enable verification by simulation. This will reducecertification and development costs.

Most reported research on impact in composite structure [1,2] concentrates on impact damageand modeling from rigid body impactors. However, for aircraft structures soft body impactorssuch as a bird or ice are highly deformable on impact and flow over the structure, spreading theimpact load. For reliable damage prediction in composite structures it is thus necessary todevelop modeling techniques and appropriate data for highly deformable impactors.

This paper describes the finite element (FE) simulation efforts used to analyse impact of acomposite wing design to predict failure mode, large deformations and structural integrity inorder to get an insight into the behaviour of the wing component under soft body impact.

In order to accurately predict the behavior of the structure under soft body impact, it isnecessary to have a realistic soft body and material model. The development of the threedimensional soft body models based on the Arbitrary Lagrangian-Eulerian (ALE) [3] approachusing LS-DYNA [4] finite elements analysis software has been described. First the soft bodyimpact onto rigid panels was investigated to determine the properties of a soft body model underimpact. The results obtained using this model were compared to soft body test results to validatethe accuracy of the developed model [5]. Then a leading edge wing model has been developed tostudy the soft body impact onto a composite wing structure. The soft body is fired at thiscomposite (glass fabric/epoxy) wing model to predict failure mode, large deformations andstructural integrity.

2. SOFT BODY (BIRD) CHARACTERISTICS

In order to use the computational method for soft body impact simulation, a constitutive lawfor the soft body such as a bird with suitable materials parameters is required. Wilbeck [5]reported five regions of impact mechanics: elastic, plastic, hydrodynamic, sonic velocity andexplosive. The impact regime best suited to describe a bird strike is hydrodynamic. In thisregion, the yield stress of the projectile is greatly exceeded due to very rapid deceleration and thematerial may be treated as a fluid. The material model often used is referred to as "elastic-plastichydrodynamic" which was originally developed for ballistic impact in metals. This modeldescribes an isotropic elastic-plastic material at low pressures, with an equation of state (EOS)describing the "hydrodynamic" pressure-volume behavior at high pressure. The EOS is assumedto have the polynomial form ofdegree three:


where f1 is given by f1 = P/ Po -I. This is a dimensionless parameter defined in terms of the

ratio of current density P to initial density Po and represents the change in density during theimpact. Co, CI, C2, and C3 are material constants. This polynomial form is an establishedapproximation ofthe observed EOS for many materials [5].

Suitable values are required for the constants Ci in the EOS polynomial. Since theseconstants refer to the dynamic behavior of a bird at impact they are difficult to measure directlyand have to be determined indirectly. The approach generally used is to calibrate the materialparameters by comparing impact simulation results with test data on the behavior of impactors.Measured impact pressures of several materials; including gelatine show that pressure pulseshave a characteristic form [5]. This pressure pulse consists of a high peak pressure (Hugoniotshock pressure) caused by shock wave propagation in the impactor, followed by a lower constantpressure (stagnation pressure) due to steady flow of the impactor onto the target. Furthermore,the EOS for water could be used as a basis for predicting peak pressures behind a 1-D shockfront which were similar to those measured. For a material such as water which exhibits thelinear Hugoniot relation between shock velocity Vs and particle velocity vp :

vs= co+k vp (2)

where k is a material constant and Co the sound speed in the material, the pressure-densityrelation across a shock has the general form [2]:

It follows that when expanded for small and moderate values of J.1, equation (3) takes the formof the polynomial EOS (I) with:

Co initial equilibrium pressure, considered to be negligible

C\ = Poc~

Cz = (2k -I)c\

C3 = (k -IX3k -I)c\

where Po is the initial density, Co = 1482.9m/sis the speed of sound in water, k=2 is an arbitrary

parameter determined by Wilbeck's tests and C1 is the bulk modulus of the impactor material.It was shown that with values for the constant k and sound speed Co for water, the predicted

pressure peak across the shock front for gelatine impactors fitted well the test data [5]. However,for tests using real chickens as impactors, the pressure pulse had lower peak pressures and waswider in duration. Since the aim of the simulation tool is to simulate real bird impacts onstructures, it was recommended to use material constants in the EOS which represent a mixture


of water with about 10% air. The air content has the effect of reducing density, lowering the bulkmodulus and speed of sound. Thus it is possible to determine the EOS constants using a rule ofmixture model, as proposed in [5].

3. COMPOSITE MATERIAL FAILURE MODEL

An enhanced composite damage material model has been used in simulation efforts. Criteriaare defined for fiber direction tensile and compressive failure, and matrix direction tensile andcompressive failure. This material model is referred to as the Enhanced_Composite_Damage inLS-DYNA software.

The strength in the fiber and matrix directions are different for tension and compression for acomposite material. The failure criteria are given in the following equations.

Tensile fiber mode (Fiber rupture) (TIl ~ 0 :

e 2 =(~J2 -1{~ Ofailed (4)f X t -< 0 elastic

Compressive fiber mode (Fiber buckling and kinking) (Til :s; 0 :

2 ((Til J2 {~Ofailede = - -1 (5)c Xc -< 0 elastic

Tensile matrix mode (Matrix cracking under transverse tension and shearing):

2_((T22 J2 ((T\2 J2 {~O failede - - + - -1m .r; Sc -< 0elastic

(6)

Compressive matrix mode (Matrix cracking under transverse compression and shearing):

Where:X t denotes longitudinal tensile strength

Xc denotes longitudinal compressive strength

.r; denotes transverse tensile strength

Yc denotes transverse compressive strength


Sc denotes shear strength

This composite material mode has the capability for impact simulation of shell typestructures.

4. FINITE ELEMENT SIMULATION

Three-dimensional models were developed to study the impact properties of a soft body ontoa target plate which was assumed to be rigid. This mass is frequently used for bird impactor testsfor which there are available impact pressure pulse test data. Since the body of a bird is mainlycomposed of water, the average density should be close to that of water. However, takingcavities, bones and different anatomical considerations into account, the averaged density islower than that of water. A typical value of density used by researchers is 934 kg/m3

, which wasthe value used for this study [5]. The impactor was given an initial impact velocity normal to thetarget plate.

4.1 Soft Body Impactor Geometry

The best impactor geometry was specified as a solid cylinder with hemispherical end caps andtotal length equal to twice the diameter [5] as shown in Fig. 1. It follows that the total volume isgiven by:

1rD37dJ2 D 51rD 3

Volume=-+--=-- (8)6 4 12

Accordingly, for the given density and weight of the bird, the diameter must be: D = 0.114 m.The duration of impact T is equal to the time it takes for the bird to travel through its length.

For this application knowing that the total length is twice the diameter, the duration of impact isgiven by:

where Do is the initial impact velocity of the projectile. This information will be useful todetermine the duration of the simulations.

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D

/

JI

II

---+---+---I,

\\

<:---D--->

Fig. 1: Bird model geometry.

4.2 Target PlateThe bird impact simulation was performed on a rigid plate for which comparative test data

were available from [5].The rigid panel was modelled using four node Belytschko-Tsai shell elements. Each element

had a uniform thickness of 10 mm. The material model used for the simulation wasMAT RIGID in LS-DYNA.

The material properties used for the rigid impactor are listed in Table 1. Elastic modulus andPoisson's ratio were used for determining sliding interface parameters such as interface stiffnessat the contact surface between the bird and the plate. The plate was constrained in all rotation andtranslation directions.

4.3 Benchmark Simulation of a Soft Body Impactor onto a Rigid PlateSimulations were performed using the LS-DYNA non-linear finite element code. The ALE

with multi-material method was applied to model bird impacts onto a rigid plate. Global load(applied to the rigid target) and deformation mode were used for comparison betweenexperimental data and fInite element results.

. I. 'dT bla e 1: Rigl . pate matena - Stee .Elastic modulus (MPa) 207000.

Poisson's ratio 0.33Density (kg/mJ

) 7800.

4.4 ALE with Multi-Material ModelThe soft body was modeled using multi-material ALE elements in combination with a multi

material group defInition and an initial volume fraction. This ALE mesh has to be fIne enough tocapture the material flow of the bird properly through the elements. The geometry of the initialvolume fraction is defIned by a part, which is composed of shell elements describing the outersurface of the original bird as shown in Fig. 2a and 2b. The initial volume fraction defInition

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leads to an initial density distribution of the bird material inside the shell elements and the ALEmesh. This was done using the INITIAL_VOLUME_FRACTION_GEOMETRY keyword inLS-DYNA. The volume fraction used was 10%.

The ALE_MULTI-MATERIAL_GROUP command allows the appropriate ALE materialgrouping to be defined for interface reconstruction for two materials (soft body and air) in themodel. This card provides the interaction between different materials in the ALE mesh.

4.5 Simulation ResultsSimulation was performed for the impact velocities of 116, 197,253 mls. A typical simulation

sequence of the bird impacting a rigid plate at 116 mls is shown in Fig. 3. In this figure, the birdmaterial (red) is flowing out of the initial bird mesh. As expected, the mesh moves and expandsto track the bird material.

(a) 3D Bird Model (b)ALE Multi-material halfmeshFig. 2: ALE multi-material mesh


Oms 0.5 ms 1 ms 1.5 ms 2 ms

Fig. 3: Typical simulation sequence of the bird impacting rigid plate at 116 mls.

Test and simulation data were also presented as non-dimensional (nd) pressure-time curves,based on a non-dimensional time (tnd) and non-dimensional pressure (Pnd) given by:

tUot =

nd 2D'

where Do is the initial impact velocity of the projectile, Po the initial density and 2D the impactorlength. The time taken for the free end of the cylinder, with initial velocity Do and length 2D, toreach the target was tnd=1, which defined the impact pulse length. The pressure Pnd=1 was

equivalent to an impact pressure~tag = t PoU~, which was the stagnation pressure predicted

from Bernoulli's equation for a steady state fluid jet with impact velocity Do.As mentioned previously (see equation 9), the duration of impact was given by:

The time of interest was limited to the period for which the non-dimensional time tnd::;; 1.During this period, the pressure was determined over group of elements at the center of the targetplate.

Simulation was performed for the impact velocities of 116, 197, 253 mls with the impactduration corresponding to that velocity. Depending to the type and size of an aircraft, it mustdemonstrate the ability to land safely after being struck by a soft body (bird) anywhere on thestructure, at close to one of these operating speeds. The velocity, impact duration and stagnationpressure used in the simulation efforts are summarized in Table 2.


f£ rtI t'T bl 2 In da e put ata use III Slmu a Ion e 0 s.Velocity (mls) 116 197 253

Impact duration (s) 2.0xl0-J 1.2x 10-J 9xl0-4

Stagnation pressure (MPa) 6.3 18.1 29.9

9r------------------;:=============~

10.8

- Simulation: Multi-Mat

-Test: Wilbeck

0.60.40.2

III 8

~ 7II)

e 6Co

iij 5co'iii 4c~ 3:c§ 2

z 1 f+JYffi\JJ=rt~rs:;;::.2S..d~~~~~~~~~:::::::=:;;;;;::----10-1--'------''--------,------...,.-----...,.---------,-------1

oNondimensional time

Fig. 4: Pressure pulses at the centre ofthe plate during impact at 197 mls.

The comparison of the pressure history at the centre of the target plate at a typical impactvelocity of 197 mls from the simulation and test results is shown in Fig 4. Pressure and time arepresented in non-dimensional form. The results confirm that the pressure pulse predicted by theALE simulation had a short duration pressure peak at the Hugoniot shock, followed by the steadyflow region. The peak pressure is critically dependent on the fluid model used and the size of thecontact area. The shape of the simulated pressure pulse shown in Fig. 4 agreed with test datafrom [5]. Peak pressure occurred at approximately tnd= 0.06 with peak non-dimensional pressureof approximately 5.4 and steady non-dimensional flow pressure of approximately 1.0. At 197mls impact speed the predicted steady state pressure was 18.1 MPa and the peak shock pressurewas 97.7 MPa. Discrepancies between simulation and test results should be expected due to therounded geometry of the model and irregular surface of the bird, and the prior impact of suchthings as feathers, wings and legs. The good general agreement with measured pressure pulsedata from the literature gives confidence in the ALE model for the bird and the values of theconstants assumed in the equation of state.

5. SOFT BODY IMPACT SIMULATION ON A COMPOSITE WING STRUCTURE

5.1 Leading Edge Wing ModelThe analysis of the wing leading edge was performed, using the bird model developed in the

section 4.1, to assess the ability of the leading edge to absorb an impact and prevent detrimental

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damage to the front spar. The impact location was centered between rib#2 and #3 as shown inFigure 5.

5.2 Geometry and Material PropertiesThe wing design used in the numerical simulations is shown in Figure 5. This is

representative of outer wing section of a Piper Aztec aircraft. The leading edge structure consistsof four main parts: the outer skin which is made from composite material, the main spar, the noseribs and the main ribs which are made from aluminum. Thickness of the components rangedfrom 0.03 in to 0.09 in. The model simplified some ofthe details ofthe actual wing fabrication.

The main spar, ribs, were modeled using a Belytschko-Tsay shell element formulation.The material definition used in the main spar and rib models were

MAT_PLASTIC_KINEMATIC, which was suited to model isotropic and kinematic workhardening. This material selection was consistent with the standard properties stated forAluminum-Alloy 6061-T6 standard structural profiles [6]. The material properties used for themain spar and ribs are listed in Table 3.

An eight integration point shell finite element model was developed for the skin of the leadingedge structure, based on eight integration point for the laminate which corresponds to the eightply lay-up (0/45/90/135).

/C~m.tact <ilSned bctwlca thie t)~c.n

andbit4 $: ~:bs1i,tu"

TrM,sl*<e co,t$1tiint (ntXI. ud Z) ,attheback <>f the wmg

Figure 5: Wing Model

The skin is defined using the material model MAT_ENHANCED_COMPOSITE_DAMAGEin LS-DYNA. The composite material used in this model was made of Kevlar fibres and anepoxy resin matrix. Some of the main material properties were listed up in Tables 4 and 5.


M . IPT bl 3 Ala e ununum atena ropertIes.Mass Density (kg/m j

) 2700Young's Modulus (Pa) 6.89SxlO lU

Poisson's Ratio 0.33Yield Stress 2.62x10lS

Tangent Modulus 6.7Sx10 lS

. M . IPT bl 4 Ca e omposlte atena ropertIes.Density (kg/mj

) 1 SOOEa- Young's Modulus (Longitudinal

6.2x 1010

Direction) (Pa)Eb -Young's Modulus (Transverse

6.1x1010

Direction) (Pa)Poisson's Ratio 0.02

Gab - Shear Modulus (pa) 2.6x101U

~c - Shear Modulus (Pa) 1.6x101u

Gca - Shear Modulus (pa) 1.6x101U

The spot welds were used to link the main ribs, the main spar and the nose ribs together. Theywere also used to attach the skin to the ribs. Failure of the spot welds occurred when thefollowing criterion was reached:

where fn and fs are the normal and the shear interface force, Sn and Ss denotes normal and shear

strength, respectively. Spot weld failure due to plastic straining occurred when the effectivenodal strain exceeded the input value 8max.

The spot welds used in this model had the same properties than aluminum rivets, described inTable 6. This option has been selected because it can model the tearing out of a spot weld fromthe outer skin [4].

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ldfST hI 6 Pa e ropertIes 0 )pot we s.Normal Force at Spot Weld Failure Sn (N) 5 100Shear Force at Spot Weld Failure Ss (N) 3200Exponent for Normal Spot Weld Force n 1.5Exponent for Shear spot Weld Force m 2.1Effective Plastic Strain at Failure 8max 15%

5.3 Contact DefinitionThe model used five contacts: the contact between the skin and the nose rib, the contact

between the nose rib and the main spar, the contact linking the skin and the main rib, the contactbetween the main spar and the main rib, and finally the contact between the skin and the bird.The first four contacts used the keyword AUTOMATIC_SURFACE_TO_SURFACE in LSDYNA. The last one, between the skin and the bird, used a constrainedLAGRANGE_IN_SOLID contact, which rules the contact between fluid (bird) and solid (skin).The coefficient of friction used during the simulation efforts was 0.2.

5.4. Boundary ConditionsThe back of the wing must defined with translational constraint (in X, Y and Z direction)

using 84 Single Point Constraint (SPC) nodes. The bird model struck the wing at impact speed of225 knots (116 mls).

5.5 Simulation ResultsThe results obtained from the wing with composite skin has been compared to the traditional

wing made of aluminum skin.The propagation of the crack can be observed in Figure 6. In the aluminum case (Figure 6.a),

the crack propagated mainly through the z-direction (see Figure 5), and the propagation causedsignificant damage to the wing. However, for the composite wing (Figure 6.b), there was no realcrack propagation. The damage was caused by the high value of stress where the bird hit thewing, but the fibers in all the directions (0°, 45°, 90°, 135°, 0°, 45°, 90°, 135°) acted as crackstoppers.

(a) (b)Figure 6: (a) Global shape of the impact zone for aluminum wing and (b) composite wing


The global shape of the impact zone looked quite different when comparing the aluminumand the composite wing (see Figure 6.a and 6.b). The aluminum impact zone was wider than thecomposite impact area; the elements deformed more in the aluminum case, instead of failingfaster as occurred in the composite structure. Furthermore, there were large deformations in thearea around the crack for the aluminum case whereas, for the composite wing, the damageseemed to be limited to the diameter ofthe bird itself.

It is worth to mention that the mesh size of the soft body (bird) must be at the same order ofmagnitude of the target (wing section) to obtain consistent results for different mesh size duringthe simulation efforts.

Figure 7: (a) Distribution of the Von Mises stress within the aluminum wing and (b) within theComposite wings, at t = 2 ms.

The distributions of the stress in the skin looked quite different for the two wing structure (seeFigure 7.a and Figure 7.b). Indeed, for the aluminum wing, the stress seemed to propagatethrough the whole skin, whereas, for the composite wing, the stress was much more localizedaround the impact zone. This is explained by the nature of the materials; aluminum is anisotropic material, therefore, the stress can easily propagate in all directions where as thecomposite material is anisotropic. Hence the stress in the structure is "stopped" due to theplacement of the fibers in each layer which act as stopper.

6. CONCLUSION

Penetration of thin composites plates by a soft body impactor has been investigated. Thesimulations were performed to assist in the development of the modeling requirements forsimulating bird impact on composite wing skin.

A three dimensional bird model, based on the Arbitrary Lagrangian-Eulerian (ALE) multimaterial approach has been analysed using finite element formulations in LS-DYNA. Simulationresults have been compared to experimental data in terms of shock and steady state pressure forthree impact velocities. Analyses have shown that simulation results using ALE formulationswith a hydrodynamic fluid law and an equation of state (EOS) were close to experimental data interm of shock and steady state pressures. Results from the ALE method showed good agreementcompared with available experimental results in terms of local pressure, including Hugoniot andstagnation pressures, and global load.


The impact of soft bodies on composite wing structures was subsequently described. Acomposite failure model which included ply damage and an interplay delamination model hasbeen used to predict impact damage in the shell structures. Modeling the penetration processusing LS-DYNA keywords provides a capability to tear nodes and elements during the impactevent. The results obtained demonstrate that it is possible to simulate impact failure progressionduring soft body impact loading in composite wing structure. The developed capability could beapplied to reduce the number ofcertification test and costly redesign process.

REFERENCES

[1] Abrate S. Impact on composite structures. Cambridge, UK, Cambridge University Press,1998.[2] Iannucci L, Dechaene R, Willows M, Degriek J., A failure model for the analysis of thinwoven glass composite structures under impact loadings, Computational Structure, pp. 79-99,2001[3] Olovsson L., Souli M., Do 1., "LS-DYNA ALE (Arbitrary Lagrangian-Eulerian) Capabilities,Fluid-Structure Interaction Modeling", Livermore Software Technology Corporation, 2001.[4] Hallquist lO., "LS-DYNA Theoretical Manual", Version 970, Livermore Software

Technology Corporation, 2001.[5] Wilbeck J.S., "Impact Behavior of Low Strength Projectiles", Air Force MaterialsLaboratory, Technical Report AFML-TR-77-134, 1977.[6] American Society of Testing and Materials (ASTM) Handbook, Designation: B 3081B 308M- 00, May 10, 2000.


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