Composite materials and bird-strike analysis using explicit finite element commercial codes

Composite materials and bird-strike analysis using explicit finite element commercial codes M. Anghileri, L. M. L. Castelletti, L. Lanzi & F. Mentuccia Dipartimento di Ingegneria Aerospaziale, Politecnico di Milano, Italia Abstract The impact of a bird represents one of the most dangerous events for the safety of the modern aircraft. Statistics report that the consequences can be tragic, especially for the small airplanes. In this work, the impact of a bird against a nacelle inlet is analysed by the means of a proven Explicit Finite Element (FE) code, LSTC LS-Dyna. In particular, the dynamic characterisation of the composite materials used in manufacturing the external skin of the inlet was achieved by referring to the drop-tests carried out using thin walled tubes of the same material. The data collected were used to define and tune the damage law of the numerical model. Once the material constitutive model was validated, an actual bird impact test was numerically analysed. In particular, two different bird models were considered: the customary Finite Element and the nodal masses bird model. The numerical results obtained were eventually qualitatively compared with experimental. In conclusion, some remarks on the results obtained are made. Keywords: bird impact, explicit finite element codes, composite material, inlet. 1 Introduction The impact of birds is one of the most dangerous events for the safety of modern aircrafts. Statistics [1,2] report that, despite the efforts provided, the consequences of a bird impact still represent a menace for flight safety, above all, with regard to the small airplanes of the General Aviation. Accordingly, it is essential for modern aircraft structures to maintain a certain level of functionality even when the attempts to prevent the impact fail and the collision becomes unavoidable. In fact, as it is impossible to built structure fully bird-proof, it is important for cabin crew survivability to ensure that the aircraft is able to land

466 Structures Under Shock and Impact VIII safely after the impact with a bird. The design requirements under FAR and CS/JAR establish the constraints to be adhered to as a function of specific part location and contribution to overall airframe body strength. Therefore, several efforts have been made to improve knowledge of the bird impact problem as well as meeting the specifications that themselves act as a fundamental input to bird proof structural development through the use of non-linear explicit codes based on Finite Element Method (FEM). In order to maintain airworthiness, the risk of catastrophic aircraft failure must be reduced to acceptable levels by increasing the overall structural safety including that of bird strike structural interaction. Computer modelling, in such cases, can be extensively used in predicting perceived and unperceived risks. The bird strike modelling, for instance, can be made for predicting structural failure and for obtaining data for use in the design of safer aircraft. After early unsuccessful attempts, numerical representations of bird impacts started improving at the beginning of eighties when new Explicit Non-linear Finite Element codes initially developed for the analysis of contact/impact problems were used [3]. The drawback of these codes, however, is the huge distortion of the bird mesh that eventually leads to inaccuracy in the calculations or, in the worse cases, to a premature analysis termination. In order to overcome these limitations, techniques based on re-zoning and/or re-starting of the analysis were developed. Nonetheless, these techniques are very time-consuming and not always work properly. Feasible alternatives to the Lagrangian approach are the Eulerian and Arbitrary Lagrangian Eulerian (ALE) approaches which are basically insensitive to the mesh distortion. Unfortunately, these approaches are both quite inaccurate and very time-consuming. Accordingly, modelling techniques able to inherit the benefits of a Lagrangian approach and, at the same time, insensitive to the mesh distortion have been developed. One promising technique of these, that in the past was successfully used to analyse a similar problem, consists of discretising the bird by means of a finite set of nodes endowed with a mass and a velocity [4]. In the following, the impact of a bird against the structure of a modern aircraft turbofan intake is investigated with particular emphasis to the consequences of subsequent penetrations. Indeed, for modern turbofans nacelles, it is mandatory to sustain the bird impact and avoid damage to the engine control system, which is placed inside the nacelle structure. The external skin of modern nacelles is made up of composite materials that permit a light structure along with high stiffness-to-weight and strength-toweight ratios. As already known, the crash behaviour of composite materials is characterised by complex failure events (lamina bending, transverse shear and local fibre buckling) and progressive damage modes, such as inter-laminar or intra-laminar cracks [5]. Therefore, appropriate analytical models and specific Constitutive Laws have to be defined and tuned to predict failure/damage mechanisms and their evolution. Accordingly, the present work is divided into two main parts: in the first part, the composite material model is validated in reference to the data collected during dynamic crash-tests performed on purpose. In the second part, an actual

Structures Under Shock and Impact VIII 467 bird impact against an inlet is investigated specifically using two different already validated bird models: namely the Finite Element (FE) and the nodal masses model. 2 Material model and experimental validation The numerical model, developed in the first stage of the work, was validated by means of experimental vertical crashes tests performed using cylindrical composite tubes. Two different sets of staking sequences were considered. All the cylindrical shells tested were made up of the material and have the same nominal geometry. In particular, a carbon fibre reinforced plastic (CFRP) woven with resin volume fraction of 42% was used. The main static properties of this composite material are reported in Table 1. The specimens had a nominal height of 300 mm and a nominal internal diameter of 68 mm. The upper edge of the specimens was chamfered in order to provide a suitable trigger. Two different kind of stacking sequences were considered: the first one consisting of four [0 /45 ] S oriented layers and the second one consisting of five [0 /45 /0 /45 /0 ] oriented layers. Table 1: Material properties. E 11 [N/mm 2 ] E 22 [N/mm 2 ] ν G [N/mm 2 ] ρ [kg/m 3 ] 69100 62600 0.05 5900 1600 2.1 Experimental tests The composite shell specimen was tested at the LA.S.T. Labs - Dipartimento di Ingegneria Aerospaziale of the Politecnico di Milano. The test facility consists of a trolley constrained to move vertically and lifted by an electric winch. The trolley is released via an electro-mechanical system. The weight of the impacting mass can be changed and the impact velocity can be chosen changing the drop-height. During the tests, the vertical displacement is measured with an incremental encoder while the vertical acceleration of the trolley is measured with a 200 g accelerometer. The reaction force of the specimen is computed as being the measured trolley acceleration multiplied by the impacting mass. A picture of the test facility is shown in Figure 1 together with a composite specimen before and after the test. Table 2: Summary of some meaningful test data. Lay-up Mean force [N] Absorbed energy [J] Test_11 (a) [0 /45 ] S 15636 2238 Test_12 (a) [0 /45 ] S 15254 2235 Test_21 (b) [0 /45 /0 /45 /0 ] 20492 2385 Test_22 (b) [0 /45 /0 /45 /0 ] 22858 2630 (a) Mean force and absorbed energy computed considering a shortening of 150 mm. (b) Mean force and absorbed energy computed considering a shortening of 120 mm.

468 Structures Under Shock and Impact VIII In the test considered, an impacting mass of 110 kg with an impact velocity of about 8 m/s was used. The signals acquired were filtered and post-processed and, hence, both the force peaks and the mean crush force were evaluated. The experimentally observed crash behaviour is consistent and experimental tests were both reliable and repeatable. Test results are summarised in Table 2 with regard to the two specimen typologies. Accordingly, the data collected seem to be suited to develop the numerical model. Figure 1: Test facility and a cylindrical shell before and after the test. 25 20 Axial load [kn] 15 10 5 0 Test_11 Test_12 Numerical 0 40 80 120 160 Shortening [mm] Figure 2: Load-shortening curves of the [0 /45 ]S specimens. 2.2 Numerical model and experimental correlation The cylindrical shells were modelled with four-node shell elements. A single integration point is defined for each ply throughout the thickness. After a sensitivity analysis on the element size carried out to avoid mesh effects on the numerical results, the characteristic length of the shell elements was fixed to 3 mm. Consequently, the final Finite Element model consists of 3916 shell elements. Boundary conditions, impact velocity and impacting-mass were

Structures Under Shock and Impact VIII 469 carefully considered and reproduced according to the experimental tests. An initial trigger was also realised selectively decreasing of about one-fifth the thickness of some elements in the first two shell rings. 25 20 Axial load [kn] 15 10 5 0 Test_21 Test_22 Numerical 0 40 80 120 160 Shortening [mm] Figure 3: Load-shortening curves of the [0 /45 /0 /45 /0 ] specimens. The material model used, MAT 58 [5,8], is developed specifically for laminated composite material. Basically, it is an elastic damage model developed around the idea that damages introduce micro-cracks and cavities into materials and that these defects primarily cause merely stiffness degradation with rather small permanent deformation unless material undergoes rather high loading and is not close to deterioration. A non-smooth failure surface is assumed and, in order to allow an almost uncoupled failure of an arbitrary composite, all failure criteria are taken to be independent of each other. The numerical results obtained on the two specimen typologies are reported in Table 3. The comparisons between the load-shortening curves numerically and experimentally obtained considering the specimen of four [0 /45 ] S oriented layers and the specimen of five [0 /45 /0 /45 /0 ] oriented layers are shown in Figure 2 and 3 respectively. The numerical results show a good correlation both in terms of load and absorbed energy values as well as in terms of the curve shape. The parameters, which characterise the material model, after being validated with reference to the experimental data, are in turn used in the following section to investigate the bird-strike against a typical composite inlet. Table 3: Summary of meaningful numerical results. Lay-up Mean force [N] Absorbed energy [J] 4 layers (a) [0 /45 ] S 13986 2085 5 layers (b) [0 /45 /0 /45 /0 ] 20464 2378 (a) Mean force and absorbed energy computed considering a shortening of 150 mm. (b) Mean force and absorbed energy computed considering a shortening of 120 mm.

470 Structures Under Shock and Impact VIII 3 Bird impact onto a nacelle inlet The composite model validated in the first stage of the work was successively used to characterise the inlet skin of a modern aircraft nacelle. Without reducing the generality of the results obtained in the first stage of the work, it was considered a bird impact test, which eventually results in one of the most severe certification tests for an engine nacelle. In particular, in the test, the bird impacted the nacelle near the exhaust with a velocity of about 330 kts. Also, after the impact, the airframe structure collapsed and the bird also penetrated into the airframe itself. A bird-strike event is characterised by loads with high intensity and short duration. The materials undergo high strain rates, large elastic and inelastic strains. In addition, a deep interaction exists between the deformation of the structure and the contemporaneous deformation of the impacting bird. The mutual influence of impact loads and response of the structure is so deep that only with the development of explicit codes based on Finite Element Method made possible to numerically analyse the event with a degree of accuracy. 3.1 Inlet Finite Element model The inlet used in the test is from an executive-like aircraft. The Finite Element mesh was built on the original inlet geometry: the final inlet model consisted of more than sixty thousand four-node shell elements (Figure 4). The inner barrel aft part was constrained as in actual test: so that displacements and rotations are prevented. Figure 4: FE model of the nacelle. The inlet structure consists of fifty-three different parts made up with customary Aerospace Industry materials. Of these, thirty parts are manufactured in composite material using sandwich technology. The remaining parts are

Structures Under Shock and Impact VIII 471 constructed in Titanium and Aluminium-alloy. The numerical model validated in the first stage of the work was used for composite material parts. Whilst for the mechanical properties of the remaining parts were used the customary values also considering the strain rate dependency via Cowper-Symond coefficients. Particular attention was reserved to both the contact among the parts and the riveted junction model. The real time duration of the simulation was 0.004 s. The duration time was mainly chosen to follow the bird penetrating into the inlet structure and impacting the forward bulkhead placed aft of the nose-lip. 3.2 Bird numerical model The bird used in the test had the customary features prescribed by the accepted international rules. In particular, the bird weight was 4 lb. Two already validated numerical models were considered for the bird: the customary Finite Element and the nodal masses models. The Finite Element model consisted of 1800 eight-nodes solid elements and the Gruniensen s equation of state [5] was defined using the values validated referring to experimental data. The nodal masses model consisted in 1445 node endowed with a mass of 0.627 grams. No equation of state was necessary. The model consists only in independent discrete masses: no interactions among the masses are defined. This quite unconventional modelling technique vaguely inspired to Discrete Element Method (DEM) was firstly introduced to overcome the typical limit of Lagrangian Finite Element model and particularly the large bird mesh distortions [4]. The results obtained are remarkable especially for normal impacts. 3.3 Results obtained The results obtained were firstly compared with the photographic post-test damage documentation. Both models gave a quite accurate description of the event. The nacelle deformations obtained with Lagrangian and nodal mass bird models are shown in Figure 5 and 6 respectively, whilst the experimental one is shown in Figure 7. In the first phase of the impact, the FE model has an advantage on nodal masses model because it is more accurate. However, as the simulation goes on, the FE mesh distortions increase and both the accuracy and the time-step decrease dramatically. Under these conditions, the nodal masses model becomes the only possible way to model the impact. As an added benefit, the same nodal masses model, with the same good results, can be readily implemented in different explicit codes, such as ESI-Group PamCRASH (Figure 8). As a further remark, it is worth noticing that as a main result of the simulation performed it has been demonstrated that the validated composite material model, regardless to the specific problem considered, can be proficiency used to analyse high velocity impacts such as bird impacts.

472 Structures Under Shock and Impact VIII deformation at t = 0.004s Figure 5: Lagrangian bird models. deformation at t = 0.004s Figure 6: Nodal masses bird models. Figure 7: Actual nacelle deformation.

Structures Under Shock and Impact VIII 473 A Figure 8: Nodal masses model: with Ls-Dyna (A) and with PamCRASH (B). Also, it could be interesting to note that, even if a sharp comparison between the two models in terms of required CPU-time is impossible, the simulation performed using the nodal masses model is somewhat faster. Indeed, the nodal masses model is not only completely indifferent to the bird mesh distortion, but also requires a smaller time-per-cycle. 4 Conclusions The impact of a bird represents one of the most hazardous events for the safety of the modern aircraft. In this work, the impact of a bird strike against a nacelle inlet is analysed through a very wide diffuse and proven explicit FE code, namely LSTC LS-Dyna. The analysis of the dynamic characteristics of the composite materials used in manufacturing the external skin of the inlet was achieved after referring to several drop-tests carried on using thin walled tubes of the same material. The validated material was, in turn, used to analyse an actual bird impact test. Two different bird models were considered: the customary Finite Element and the nodal masses bird model. The consequent numerical results, qualitatively compared with experimental evidence, showed the proficiency of the approach proposed. References [1] E. C. Cleary, R. A. Dolbeer, S. E. Wright: Wildlife strikes to civil aircraft in the United States 1990 2002. Federal Aviation Administration, National Wildlife Strike Database. Serial report number 9. Washington, DC. June 2003. B

474 Structures Under Shock and Impact VIII [2] W. J. Richardson, T. West: Serious birdstrike accidents to military aircraft: updated list and summary International Bird Strike Committee 25th Meeting. Amsterdam, the Netherlands, 17-20 April 2000. [3] R. A. Brockman T. W. Held Explicit finite element method for transparency impact analysis Dayton University, OH Final technical report, Sep.1988-Sep.1990. [4] M. Anghileri, C. Bisagni "New model of bird strike against aircraft turbofan inlet", Proceedings of 3rd International KRASH Conference, Arizona State University, Tempe (USA), 2001, p. 8.3.1-8.3.11. [5] J. O. Hallquist LS-DYNA Theoretical Manual Livermore Software Technology Corporation (1998). [6] Schweizerhof K, Weimar K, Munz T, Rottner T. Crashworthiness analysis with enhanced composite material models in LS-DYNA - Merits and Limits. Proceedings of the 5th International LS-DYNA Users Conference, Southfield, 21-22 September 1998.