Failure analysis of single-bolted joint for lightweight composite laminates and metal plate

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1 IOP Coferece Series: Materials Sciece ad Egieerig PAPER OPEN ACCESS Failure aalysis of sigle-bolted joit for lightweight composite lamiates ad metal plate To cite this article: Lijie Li et al 018 IOP Cof. Ser.: Mater. Sci. Eg View the article olie for updates ad ehacemets. This cotet was dowloaded from IP address o 05/11/018 at 11:47

2 ICIEM 017 IOP Cof. Series: Materials Sciece ad Egieerig (017) 0107 doi: / x/84/1/0107 Failure aalysis of sigle-bolted joit for lightweight composite lamiates ad metal plate Lijie Li 1, Juli Qu 1 ad Xiagdog Liu 1 1 Departmet of Mechaics, Xi a Uiversity of Sciece ad Techology, 58 Yata Road, Beili District, Xi a , Chia lijie_happy@163.com Abstract. A three-dimesioal progressive damage model was developed i ANSYS to predict the damage accumulatio of sigle bolted joit i composite lamiates uder i-plae tesile loadig. First, we describe the formulatio ad algorithm of this model. Secod, we calculate the failure loads of the joit i fibre reiforced epoxy lamiated composite plates ad compare it with the experimet results, which validates that our model ca appropriately simulate the ultimate tesile stregth of the joits ad the whole process of failure of structure. Fially, this model is applied to study the failure process of the light-weight composite material (USN15). The study also has a great potetial to provide a strog basis for bolted joits desig i composite Lamiates as well as a simple tool for comparig differet lamiate geometries ad bolt arragemets. 1. Itroductio Composite is a state-of-art material i aircraft structure. To have a safer structural desigs of aircraft with low weight, more ad more advaced composites have bee developed, which are ot oly light weight but also have higher stregth to weight ratio [1]. Oe of the most useful applicatios of lamiated composites as weight reducig elemets is i aircraft wig where composite strigers has huge potetial of beig attached with the wig ski by meas of a bolted joit []. Although composite materials are ow beig widely used as primary aircraft structures, the use is rather limited due to lack of research i composite mechaics ad complicated failure mechaics of the composite joits, especially for the advaced composites [3].. Numerical simulatios usig ANSYS.1 The geometry of the compoet The test specime cofiguratio is show i Figure 1, ad the specific values are show i Table 1. Figure 1. Geometries of the specimes Cotet from this work may be used uder the terms of the Creative Commos Attributio 3.0 licece. Ay further distributio of this work must maitai attributio to the author(s) ad the title of the work, joural citatio ad DOI. Published uder licece by Ltd 1

3 ICIEM 017 IOP Cof. Series: Materials Sciece ad Egieerig (017) 0107 doi: / x/84/1/0107 Table 1. Geometrical model of the composite joit. Plate legth Plate thickess Hole diameter Plate width Distace betwee the ceter of the hole (mm) (mm) (mm) (mm) ad the right margi of plate(mm) Fiite elemet modelig As show i Figure, bolts, uts ad gaskets are simplified as a whole i this 3D model. The metal plate ad bolt are modelled usig Solid185 elemets ad the composite plate is modelled usig Shell99 i asys software. I order to make the simulatio closer to the real situatio, the impact of the cotact model, frictioal ad preload were also cosiderate. The cotact areas betwee bolt ad plates are modelled usig CONTA173 cotact elemet. The magitude of the frictioal force betwee the cotact surfaces is cotrolled by coulomb frictio, the frictio coefficiet was 0.. The preload applied o the bolt is 3 Nm. The boudary coditios are show i the Figure 3. The right edge of the model is fixed ( U U U 0 ) ad a uiformly icreasig tesile-load is applied o the left edge of x y z x y z the composite plate alog the x-axial directio of the model ( U U 0 ). The mesh y z x y z covergece aalysis is performed o a various umber of elemets to esure the accuracy of calculatios ad freedom of the model to mesh size. The fial mesh distributios through thickess ad aroud the bolt hole are show i Figure 3. Figure. Simplified model of bolt 3. Aalysis of failure Figure 3. FEM mesh of composite joit 3.1 Lamiated plate cumulative damage aalysis processig Dao et al. [4] developed a progressive damage model to ivestigate the effect of failure criteria ad the material property degradatio rules o the behaviour of joits i a graphite/epoxy composite lamiate. A flow chart for the program is show i Figure 4.

4 ICIEM 017 IOP Cof. Series: Materials Sciece ad Egieerig (017) 0107 doi: / x/84/1/0107 Figure 4. Progressive damage model algorithm 3. Stress aalysis Apply a costrait to oe ed of the composite lamiates ad apply load to aother ed. The load actig o the ier surface of the lamiate icreases gradually, whe the load icreases to a step, i.e. P =P -1 +P. Because there is o volume force ad accordig to icremetal form of virtual displacemet priciple ad force boudary coditios, we have balace equatio as follows. dv S Ti ui ds 0 i, j 1,,3, 1 where ad ui represet strai icremet ad displacemet icremet from the load step 1 to -1 the load step,respectively. represet the lamiates stress uder load P,, where is the icremetal load from load step 1 to the load step. Ti (1) represet the per uit area force actig o the boudary S uder load P. V 1 is the volume deformatio at step -1. Due to the geometric large deformatio, strai icremet is show below. where e is the strai icremet ad aisotropic elastomers is show below [5]. 6 e, () is the quadratic term. The stress-strai relatioship of C ( i1,,,6). (3) i j j1 3

5 ICIEM 017 IOP Cof. Series: Materials Sciece ad Egieerig (017) 0107 doi: / x/84/1/ Whe the load from the step 1 to step, the icreased load variable is i Ckl kl 1 kl, where C is the composite material modulus uder the load step 1, which ca be obtaied by the followig equatio. where m is the umber of composite layers. Q m k 1 1 Qrs C C, (4) kl rs k 1 rs k1 m 1 is the composite material modulus of the k-th layer after coverted by coordiates uder the 1 load step. ad are liberalized: e C e,the the equatio (1) ca be coverted ito equatio (5): i 1 kl kl, 1 e dv dv S Ti ui ds 1 e. 1 1Ckl ekl dv 1 The result of the variatio equatio o the above equatio will give a liear system of equatios for displacemet icremets, which ca be further solved by fiite elemet method. Whe the load icreases to P, if the material damage occurs, it will lead to damage to the material properties of the damage zoe, that is, the material modulus part of the damage area will chage, the specific degradatio will be described i the lower part. With the advet of damage, structural stress ad strai will reassig, the equatio (5) is calculated. The process is repeated util there is o ew damage occurs. 3.3 Material degradatio rule With the icreasig of load, the fibre ad matrix of the lamiated plate are gradually destroyed. The structural rigidity is decreasig, ad the expasio of the damage ca be characterized by the degradatio of the elastic costats of the material. Ta [6] used parameters degradatio to aalysis the cumulative damage process as show i Table. Table. Material degradatio of Ta s rule Failure Mode Ta s rule Matrix tesile crackig E 0. E, G1 0. G1, G3 0. G3 Matrix compressive crackig E 0. 4E, G1 0. 4G1, G3 0. 4G3 Fiber tesile crackig E E11 Fiber compressive crackig E E11 Fiber-matrix shear out G Delamiatio i tesio E G G G3 G Delamiatio i compressio E Hashi s failure criteria [7] Hashi s failure Criteria is popular i the composite idustry due to its elegat form ad accuracy matchig the experimetal results. The three-dimesioal Hashi guidelie expressio is show below: (1) Fiber tesile crackig ( 11 0 ) (5) F f X t S1 S13 1 (6) 4

6 ICIEM 017 IOP Cof. Series: Materials Sciece ad Egieerig (017) 0107 doi: / x/84/1/0107 () Fiber compressive crackig ( 11 0 ) (3) Matrix tesile crackig ( 0 ) F m 33 Y t 33 F f (4) Matrix compressive crackig ( 0 ) 33 3 X S 11 3 c S 1 1 S (7) (8) F m 1 Yc Yc S 1 1 (5) Fiber-matrix shear out ( 11 0 ) 33 (9) S1 S3 S1 S F m YC 33 S 1 1 S , where i 1,,3 represet the ormal stress i the i directio ad i, j 1,,3 the ii (10) is the shear stress i i, j plae. X t, X c, Y t, Yc represet the tesile stregth ad compressive stregth of the fibre S i, j 1,,3 represet directio ad i the plae perpedicular to the fibre directio, respectively. the shear stregth i-plae. F ( j f, m) is the failure factor which ca be used to characterize the j degree of the material damage. The greater the failure factor is, the more serious the damage is. Whe the failure factor is greater tha 1, the material will be failure. 4. Validatio of the umerical simulatios I order to verify the accuracy of the model, specimes with the same size ad material (CCF300/QY8911) as that used i Ref 8 are calculated ad aalysed. The mechaical properties of the composite are show i Table3. The stackig sequeces for the composite ply is [45/0/-45/90/45/0/- 45/0] s ad the thickess of each ply is 0.15mm. Table 3. Material properties of composite material (CCF300/QY8911). Material properties E 11 (GPa) E (GPa) E 33 (GPa) G 1 (GPa) G 3 (GPa) X t (MPa) G 13 (GPa) CCF300/QY X c (MPa) Y t (MPa) Y c (MPa) S 1 (MPa) S 13 (MPa) S 3 (MPa) v 1 v 13 v I Table 3, E 11, E ad E 33 mea the elastic modulus i logitudial directio ad trasverse directio, respectively. G 1, G13 ad G 3 are the i plae ad the out-of plae shear modulus, respectively. X ad Y are the stregth i directio of warp ad fil, respectively. S is shear stregth. Subscript t ad c mea the tesile ad compressive properties. ν is Poisso s ratio. The material properties of metal plate ad bolt are E GPa, 0.31 ad S 675 MPa, where S represets the shear stregth. The cofiguratio of sigle-lap bolt joit was show i Figure 5. 5

7 ICIEM 017 IOP Cof. Series: Materials Sciece ad Egieerig (017) 0107 doi: / x/84/1/0107 Figure 5. Cofiguratio of sigle-lap bolted joit Figure 6 ad Table 4 show the compariso betwee the results of the simulatio ad experimetal results uder static tesile loadig, respectively. Figure 6. Compariso of FEM simulatio result ad experimet result Table 4. Compariso of testig ad umerical results Items Numerical result Test result [8] Relative Errors/% Failure load(kn) Stiffess(N mm -1 ) Accordig to the compariso preseted above, the error betwee simulated predictio value ad experimetal results is less tha 5%, idicatig that this model ca successfully be applied to predict the stregth of advace composites, as well as to compariso of differet joit cofiguratios ad effect of desig parameters. 5. Failure aalysis for lightweight composite lamiates Use our model to simulate the damage process ad failure of light-weight composite (USN15) lamiates. The material properties are listed i Table.5 ad the ply is [45/0/-45/90/45/0/-45/0] s. Table 5. Material properties of composite material (USN15). [9] Material Properties E 11 (GPa) E (GPa) E 33 (GPa) G 1 (GPa) G 3 (GPa) G 13 (GPa) X t (MPa) USN X c (MPa) Y t (MPa) Y c (MPa) S 1 (MPa) S 13 (MPa) S 3 (MPa) v 1 v 13 v

8 ICIEM 017 IOP Cof. Series: Materials Sciece ad Egieerig (017) 0107 doi: / x/84/1/0107 Figure 7 illustrates the iitial stress states of the composite lamiate. It ca be see from the stress compoet cotour diagram that the lattice aroud the bolt hole of the composite material chages more obviously, especially i the directio of -45,which idicates that the fiber orietatio of the -45 layer would firstly reach the maximum bearig stregth. I high compressio zoe, due to high values of S33, i.e. of priciple stress alog lamiate thickess, there exists a strog potetial of delamiatio due to compressio, thus causig a typical bearig failure. It has also bee observed that shearig stress i yz-plae (S3) ad xz-plae (Sl3) are very low as compared to shearig stress i xy-plae (S1) ad the trasverse compressive stress of the fiber is lower tha the shearig stress. Figure 7. Stress compoet cloud state variables results 7

9 ICIEM 017 IOP Cof. Series: Materials Sciece ad Egieerig (017) 0107 doi: / x/84/1/0107 Figure 8. Stress distributio of bolt-hole ad progressive damage of lamiate From Figure 8, we visualize the stress ad deformatio progressively. It is observed that most of the applied load is cotributed towards compressio alog fiber directios. Remarkable tesile stress trasverse to the directio of lamiate is ear the hole which may cause the effect of matrix crackig. However, because the regio uder compressive load is reasoably more as compared to regio i tesile stress, there are more chaces of bearig failure of the joit. The iter lamia stresses of both 90 plies with their eighborig plies may cause delamiatio, thus to effect the ultimate failure load. There is a warp betwee the composite lamiates ad the metal plates durig the stretchig process, ad there is a tedecy to be separated, ad the bolt is iclied i the stretchig directio. 6. Coclusios I this study, a three-dimesioal Progressive Damage Model was developed to compute the tesile stregth ad the maximum load of the composite metal joit damage accumulatio failure aalysis model, cosiderig the preload, frictio ad complex cotact area of coectio ad cotact deformatio. Accordig to the study, coclusios are as follows. (1) The tesile stregth ad the maximum load of the composite (CCF300/QY8911)-metal joit are MPa ad 9.414kN, respectively. The results of the experimetatio are 14.5MPa ad 9.kN. The error betwee simulatio ad experimet result is less tha 5%, idicatig that this model ca be applied to predict the stregth of advace composites. ()The failure process ad failure load of the lightweight composite (USN15)-metal joit were obtaied by the established model. The iitial failure crack occurs i the directio of -45 ad 45 of the fiber layer ad the matrix separatio of the stress geerated i the directio of -45, so that the directio of -45 fiber bear the mai load. (3)Sice the error is very stable betwee simulated stregth predictio value ad experimetal value of stregth so the model i the paper ca successfully be applied for compariso of differet joit cofiguratios ad effect of desig parameters for sigle bolted joit but it s ecessary to validate the three-dimesioal fiite model ad calculatio method i this paper i multiple-bolted joits i further study. 7. Refereces [1] Starikov R, Scho J. Quasi-static Behaviour of Composite Joits with Protrudig Head Bolts. Compos Struct 001; 51: [] Saqib Mehmood, Zhao Libi ad Huag Hai. Failure Aalysis of Sigle ad Double Bolted Joits for Composite Lamiates.1th Iteratioal Bhurba Coferece o Applied Scieces & Techology 011;15:

10 ICIEM 017 IOP Cof. Series: Materials Sciece ad Egieerig (017) 0107 doi: / x/84/1/0107 [3] Thoppul S D, Fiega J, Gibso. Mechaics of Mechaically Fasteed Joits i Polymer-matrix Composite Structures -a review. Compos Sci Techo 009; 69: [4] Dao ML, Kamal E, Gedro G. Aalysis of Bolted Joits i Composite Lamiates: strais ad bearig stiffess predictios. Compos Struct 007; 79(4): [5] She GL, Hu GK. Mechaics of Composite Materials[M]. Tsig Hua Uiversity Press 006. [6] Ta SC, Perez J. Progressive Failure of Lamiated Composites with a Hole uder Compressive Loadig. J Reif Plas Compos 1993; 1: [7] Hashi Z. Failure Criteria for Uidirectioal Fibre Composites. Applied Mechaics 1980; 47(): [8] Liu X K, Zhag WW, Li Y Z. Ivestigatio o the Deformatio Failure of Sigle-bolted Metal to Composite Joits. Sciece Techology ad Egieerig 014; (14): [9] Seog MS, Kim TW, Nguye KH, Kweo JH ad Choi JH. A Parametric Study o the Failure of Boded Sigle Lap Joits of Carbo Composite ad Alumiium. Compos Struct 008; 86: