Crashworthiness Simulation Research of Fuselage Section with Composite Skin

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ScienceDirect Procedia Engineering 80 (2014) 59 – 65

3rd International Symposium on Aircraft Airworthiness, ISAA 2013

Crashworthiness Simulation Research of Fuselage Section with Composite Skin MOU Haolei a*, ZOU Tianchuna, FENG Zhenyua, REN Jiana Tianjin Key Laboratory of Civil Aircraft Airworthiness and Maintenance, Civil Aviation University of China, Tianjin 300300, P.R.China

Abstract For the demand of crashworthiness design and certification of composite fuselage section, the influences of composite skin on composite fuselage section crashworthiness have been studied. A finite element model of fuselage section with sub-floor waved-plates is developed, and the dynamic responses of composite fuselage section subjected to vertical impact velocity of 6.67m/s are analyzed in LS-DYNA, and the effects on composite fuselage section crashworthiness are further investigated by changing the ply numbers and ply angles. The failure modes and acceleration history curves of composite fuselage section under different conditions are given. The results show that the composite fuselage section crashworthiness can be effectively improved by selecting appropriate composite ply numbers and ply angles. © 2014 Published by Elsevier Ltd. This is an open access article under the CC BY-NC-ND license

© 2013 Published by Elsevier Ltd. Selection and peer-review under responsibility of ENAC (http://creativecommons.org/licenses/by-nc-nd/3.0/). Selection and peer-review under responsibility of Airworthiness Technologies Research Center, Beihang University/NLAA. Keywords: Crashworthiness; Composite skin; Fuselage section; Dynamic response

1. Introduction In recent twenty years, crashworthiness design and certification have been and will continue to be the main concern in aviation safety. The fuselage section should be designed to have the crashworthy capabilities to improve the occupant survivability during a crash [1, 2, 3]. Composite materials are being increasingly used in commercial aviation because of their high strength/weight and stiffness/weight ratios and reduction in manufacturing time compared with traditional metallic materials. Along with these advantages, composite materials also have excellent energy-absorbing performance during impact and

* Corresponding author. Tel.: +86-022-24092312; fax: +86-022-24092311. E-mail address: [email protected].

1877-7058 © 2014 Published by Elsevier Ltd. This is an open access article under the CC BY-NC-ND license

(http://creativecommons.org/licenses/by-nc-nd/3.0/). Selection and peer-review under responsibility of Airworthiness Technologies Research Center, Beihang University/NLAA. doi:10.1016/j.proeng.2014.09.060

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crash events [4, 5, 6]. However, the increasing use of composite materials in aircraft structures also brings further challenges for the aircraft crashworthiness design and certification [7]. The crashworthiness design and certification of composite aircrafts have been studied by NASA and AIRBUS earlier. The full-scale crash tests of composite aircrafts, such as Beech Starship, CirrusSR-20 and Lancair etc, were carried on by NASA Langley Research Center in 1990s. The tests results showed that these composite aircrafts had perfect crashworthiness [8, 9]. The EU also had conducted a lot of work on crashworthiness design and certification of composite aircrafts. AIRBUS participated in the CRASURV (The Design for Crash Survivability) program donated by the EU, and put forward a complete set of crashworthy test methods for composite aircraft, and evaluated the crashworthiness of A380 based on these test methods [10]. Although some researches have been conducted on the impact responding characteristics of composite fuselage section in the U.S. and EU so far, the crashworthiness researches of composite fuselage section are relatively fewer and simple, and have not investigated the effects of composite ply parameters on the aircraft crashworthiness. The finite element method was used to research the crashworthy performance of composite fuselage section with different composite ply parameters. The finite element model of composite fuselage section was built in HyperMesh, and the nonlinear finite element code LS-DYNA was used to dynamically simulate the drop test of composite fuselage sections. The failure modes and acceleration responses of composite fuselage section were obtained and analyzed based on our researches. The results gave some advice regarding crashworthiness design and certification of composite fuselage sections. 2. Finite element model of composite fuselage section A finite element model of fuselage section with sub-floor waved-plates was developed. The finite element model of fuselage section with a double elliptical section consists of the cabin and cargo. The radiuses of cabin and cargo are 1700 mm and 1600 mm, respectively. The length of fuselage section is 2200 mm. The fuselage section has five frames and three rows of seats. The overall finite element model consists of 609367 nodes and 616625 shell elements. Key components of finite element model are shown in Fig.1, which including composite skin, frames, stringers, seat tracks, cabin floor and oblique struts, cargo floor and waved-plates. Because of the large size and complexity of real fuselage section structure, the finite element model of fuselage section was reasonably simplified during the modeling process based on the simplified principles described by Adams and Lankarani [11] and Ikuo et al. [12].

Left outside

Right outside Left inside

Right inside

Fig. 1. The finite element model of fuselage section; (a) Fuselage section and rigid surface; (b) Skin of fuselage section; (c) Main frame of fuselage section;

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Four kinds of materials are used in finite element model of composite fuselage section, which including Al-2024-T3, Al-7075-T6, Al-7150-T77511 and composite material T300/QY8911. Al-2024-T3 is used for the cargo floor. Al-7075-T6 is used for frames, stringers and wave-plates. Al-7150-T77511 is used for seat tracks, floor beams and oblique struts. These three materials adopt MATL24 material model with the bilinear elastic-plastic properties. The composite material T300/QY8911 is only used for fuselage skin and adopts MATL59 material model. The mechanical properties of materials used in the overall fuselage section are shown in Table 1 and Table 2 [13, 14]. Table 1. The Aluminium mechanical properties Material

Al-2024-T3 3

Al-7075-T6

Al-7150-T77511

Density (kg/m )

2760

2794

2823

Modulus of Elasticity (GPa)

66.33

71.02

71.02

Poisson's Ratio

0.33

0.35

0.35

Yield Modulus(MPa)

243

362

538

Enhanced Modulus(MPa)

826.7

1001.8

679

Maximum strain

0.1463

0.0449

0.07

Table 2. Material properties of T300/QY8911 unidirectional laminates Parameter

Value

Description

ȡ, kg/m3

1570

Density

E1, GPa

135

Young’s modulus – longitudinal direction

E2, GPa

8.8

Young’s modulus – Transverse direction

ȝ12, GPa

0.2

Poisson’s ratio

G12, GPa

4.47

Shear modulus of elasticity

Xt, MPa

1548

Longitudinal tensile strength

Xc, MPa

1226

Longitudinal compressive strength

Yt, MPa

55.5

Transverse tensile strength

Yc, MPa

218

Transverse compressive strength

S12, MPa

89.9

Shear strength

The Belytschko-Tsay shell element has been adopted because the shell element usage has the advantages of being able to more accurately and effectively simulate buckling, as well as calculate internal energy absorption during a crash. The impact mass has an important influence on crashworthiness of composite fuselage section, the masses of seats and dummies were accounted as concentrated masses located at the junctions between seats and floor. The concentrated mass of each seat and dummy is defined to 88 kg according to Federal Aviation Regulation 25.562(b), and the fuselage section has twelve concentrated masses. There are three rows of seats in the finite element model, and due to the limit of the constraints, the calculated results of the first row and the third row of seats may be inaccurate. So the acceleration responses of the junction between middle row of seats and floor are output. The four locations of acceleration results on seat tracks are shown in Fig.1(a). The vertical crash direction of the model is parallel to the normal direction of rigid wall, and the vertical impact velocity of 6.67 m/s is selected to evaluate the crashworthiness of composite fuselage

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section without considering aerodynamic force. Based on the penalty algorithm, Automatic_Single_ Surface_Contact was used to define the contact relationships among composite skin, fuselage frames and rigid ground. 3. Effects of composite ply number on crashworthiness properties of fuselage section 3.1. Effects of composite ply number on failure mode of fuselage section The crashworthiness of fuselage section with different composite ply numbers was researched because it is essential to consider the effects of composite ply number on composite structures failure modes. The ply numbers of composite skin were 18, 22 and 26, respectively. The composite ply angles were [90/45/0/-45/90/45/0/-45/90]s, [90/45/0/-45/90/45/0/-45/90/45/0]s and [90/45/0/-45/90/45/0/-45/90/45/0/45/90]s, respectively, and each layer thickness is 0.12mm. Fig.2 showed the stress clouds of fuselage sections with different composite ply numbers at 100ms. Fig.2 showed that the deformation and failure modes of three composite fuselage sections were similar. The energy was mainly absorbed by the destruction of composite skin, oblique struts, fuselage frame, and the deformation of waved-plates and the bottom of fuselage section. Compared with the other two cases, the concentrated stress of fuselage bottom frame was smaller, and the maximum stress of fuselage frame was larger obviously when the composite ply number was 26. It indicated that the failure mode of the bottom of fuselage section was stable. Therefore, with the increasing of composite ply number, the failure of composite fuselage section was more stable, but the stress distribution value increased correspondingly.

Fig. 2. The stress clouds of fuselage sections with different composite ply numbers, t=100ms; (a) The composite ply number is 18; (b) The composite ply number is 22; (c) The composite ply number is 26;

3.2. Effects of composite ply number on acceleration characteristics In order to compare the acceleration characteristics of fuselage section with different composite skin ply numbers, the acceleration responses of the junctions between left outside/right inside seats and floor were obtained because the finite element model of composite fuselage section was symmetrical. Fig.3 showed the acceleration history curves of three fuselage sections with different composite ply numbers. The peak accelerations of the junctions between outside seats and floor were larger than those of the junctions between inside seats and floor, respectively. Fig.3(a) showed that the initial peak acceleration was at maximum, and the maximum peak acceleration was relatively smaller when the composite ply number was 22. The peak acceleration came earlier a bit with the increasing of composite ply number. In

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Acceleration/g

Acceleration/g

general, the crashworthiness of composite fuselage section can be effectively improved by selecting a suitable composite skin ply number.

Time/ms

Time/ms

Fig. 3. Acceleration responses of the junctions between seats and floor with different composite ply numbers; (a) Left outside seat; (b) Right inside seat;

4. Effects of composite ply angle on crashworthiness properties of fuselage section 4.1. Effects of composite ply angle on failure mode of fuselage section According to FAA AC 20-107B, the effects of composite ply angle also needed to be researched except for the effects of composite ply number on composite structure crashworthiness. The composite skin ply angles are [0/90]18, [±45]18, [90/45/0/-45/90/45/0/-45/90]s, respectively. The composite skin consists of 18 layers and each layer thickness is 0.12mm. Fig.4(a) and (b) and Fig.2(a) showed the stress clouds of fuselage section with different composite skin ply angles [0/90]18, [± 45]18, and [90/45/0/45/90/45/0/-45/90]s at 100ms. The failure modes of three composite fuselage sections were similar. The energy was mainly absorbed by the buckling and destruction of frames, composite skin, oblique struts and waved-plates. Compared with the other two cases, the concentrated stress of fuselage bottom frame was smaller, and the maximum stress of fuselage frame was larger obviously when the composite ply angle was [±45]18, which indicated that the failure of the bottom of fuselage section was stable. When the composite ply angle was selected to [90/45/0/-45/90/45/0/-45/90]s, the concentrated stress resulting from the deformation of fuselage bottom frame was smaller than the composite ply angle was [0/90]18, and the maximum stress value was also smaller relatively. The stress value can be decreased effectively when the composite ply angles were selected to [±45]18 or [90/45/0/-45/90/45/0/-45/90]s. Therefore, the crashworthiness of fuselage section can be improved by selecting suitable composite skin ply angle.

Fig. 4. The stress clouds of fuselage sections with different composite ply angles, t=100ms; (a) The composite ply angle is [0/90]18; (b) The composite ply angle is [f45]18;

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4.2. Effects of composite ply angle on acceleration characteristics To compare the acceleration characteristics of fuselage section with different composite ply angles, the acceleration responses of the junctions between left outside/right inside seats and floor were obtained in our work. Fig.5 showed the acceleration history curves of the fuselage section with different composite ply angles, corresponding to[0/90]18, [±45]18, and [90/45/0/-45/90/45/0/-45/90]s, respectively. The peak accelerations of the junctions between outside seats and floor were larger than those of the junctions between inside seats and floor, respectively. Figure 5(a) showed that the peak acceleration was larger when the composite ply angle was [0/90]18 or [±45]18. When the composite ply angle was [90/45/0/-45/90/45/0/-45/90]s, the peak acceleration was minimal and the energy-absorbing process was the most stable. Therefore, it was feasible to improve the crashworthiness of fuselage section by selecting suitable composite ply angle based on our researches.

Time/ms

[0/90]18 [±45]18 [90/45/0/-45/90/45/0/-45/90]s

Acceleration/g

Acceleration/g

[0/90]18 [±45]18 [90/45/0/-45/90/45/0/-45/90]s

Time/ms

Fig. 5. Acceleration responses of the junctions between seats and floor with different composite ply angles; (a) Left outside seat; (b) Right inside seat;

5. Conclusion The influence of composite ply numbers and ply angles on crashworthiness of composite fuselage section were studied in the paper. The main conclusions were listed as follows: (1) The composite ply number can greatly influence on the crashworthiness of fuselage section. With the increasing of composite ply number, the failure of composite fuselage section was more stable, but the stress distribution value increased, and the peak acceleration came earlier. In general, the crashworthiness of fuselage section can be improved by selecting suitable composite skin ply number. (2) The composite ply angle also can greatly influence on the crashworthiness of fuselage section. The peak acceleration was smaller and the energy-absorbing process was more stable when the composite ply angles were selected to ±45° or 0°, ±45°, 90°, and the stress value of fuselage section decreased effectively. Therefore, it is feasible to improve the crashworthiness of fuselage section by selecting suitable composite ply angle. 6. Copyright All authors must sign the Transfer of Copyright agreement before the article can be published. This transfer agreement enables Elsevier to protect the copyrighted material for the authors, but does not relinquish the authors' proprietary rights. The copyright transfer covers the exclusive rights to reproduce and distribute the article, including reprints, photographic reproductions, microfilm or any other

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reproductions of similar nature and translations. Authors are responsible for obtaining from the copyright holder permission to reproduce any figures for which copyright exists.

Acknowledgements This work was supported by the science and technology item from the Civil Aviation Administration of China (NO. MHRD201010) and supported by the Fundamental Research Funds for the Central Universities (NO. ZXH2012B004).

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