Numerical Analysis of Carbon Fibre Reinforced Aircraft Win

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648International Journal of Earth Sciences and Engineering

ISSN 0974-5904, Volume 04, No 06 SPL, October 2011, pp 648-651

#020410343 Copyright 2011 CAFET-INNOVA TECHNICAL SOCIETY. All rights reserved

Numerical Analysis of Carbon Fibre Reinforced Aircraft WingSanya Maria Gomez

M.Tech Student, Structural Engineering, M A College of Engineering, Kothamangalam 686 666, [email protected]

Alice MathaiProfessor, Civil Engineering Department, M A College of Engineering, Kothamangalam 686 666

ABSTRACT: Hypersonic Aircraft is used to transport satellites or humans to space. Wing of a hypersonic aircraft is one

of the crucial components which determine the performance of the aircraft. The main group of materials used in aircraftconstruction has been wood, steel, Aluminium alloys and more recently, titanium alloys and fibre reinforced composites.

Several factors influence the selection of material of which strength allied to lightness is the most important. Composite

materials are well known for their excellent combination of high structural stiffness and low weight. Composite materialconsists of glass (GFRP) or carbon (CFRP) set in a matrix of plastic or epoxy resin, which is mechanically or chemically

protective. CFRP is seen to have a modulus twice & strength three times that of Aluminium alloy, the conventional

material used in aircraft construction. In the present work the aircraft wing components like ribs, spars and panels areanalysed considering both isotropic and composite materials. Since each laminate in the composite material can have

distinct fibre orientations which may vary from the adjoining laminates, the optimum ply orientation is also obtained as a

result of the parametric study conducted using ANSYS finite element package by varying the orientation sequence in the

composite. From the studies conducted regarding the weight reduction, it is estimated that replacement of Al. alloy by

CFRP results in 34.82% saving in the total structural weight of the aircraft wing.

KEY WORDS: Hypersonic Wing, Static Analysis, CFRP, Ply Orientation, Finite Element Model.

INTRODUCTION

As an aircraft moves through the air, the air moleculesnear the aircraft are disturbed and move around the

aircraft. How the air re-acts to the aircraft depends uponthe ratio of the speed of the aircraft to the speed of sound

through the air [20]. Because of the importance of this

speed ratio, aerodynamicists have designated it with a

special parameter called the Mach number. Mach

Number was named after the Austrian physicist Ernst

Mach. Mach 1 is the speed of sound, which isapproximately 760 miles per hour at sea level. An airplane

flying less than Mach 1 is travelling at subsonic speeds,

faster than Mach 1 would be supersonic speeds and Mach2 would be twice the speed of sound [22]. For aircraft

speeds which are much greater than the speed of sound,

the aircraft is called hypersonic.Weight reduction is highly desirable for aircraft vehicles

since light vehicles have improved range, fuel savings,

and increased payload. Aircraft structures are typified byarrangements of thin, load bearing skins, frames and

stiffeners, fabricated from lightweight, high strength

materials of which aluminium alloys are the most widely

used examples. The increased complexities in the flight

regime and loading conditions led to the search of acompetitive material for wing structures. Carbon fiber

reinforced polymer-[CFRP] has over the past two decades

become an increasingly notable material used in structuralengineering applications. Much of the fuselage of the new

Boeing 787 Dreamliner and Airbus A350 XWB will be

composed of CFRP, making the aircraft lighter than acomparable aluminum fuselage, with the added benefit ofless maintenance, superior fatigue resistance and high fuel

efficiency. Due to the high ratio of strength to weight,

CFRP is widely used in micro air vehicles (MAVs).FRPs are commonly used in the aerospace, automotive,

marine, and construction industries [18]. Manufacturers of

not only commercial airplanes but also military planes and

helicopters have developed various usage of compositematerial. In every case, objectives of using composite

material have been to reduce weight of planes and to havehighly performing flying machines. Composite material

also has contributed to those secondary objectives assaving of assembling manpower.

Typical speeds for hypersonic aircraft are greater than

3000 mph and Mach number M > 5. Thus, it will beexperiencing the flight environment of an upper stage in

launch vehicle during its ascend phase and the flight

conditions of an aircraft during its descend phase. Sincelift and drag depend on the square of the velocity,

hypersonic aircraft do not require a large wing area.

CONFIGURATION OF HYPERSONIC AIRCRAFT

WING

The wing of hypersonic aircraft considered in the presentstudy is double delta in plan form. The cross section of

wing is reflex aerofoil. The sweep back angle of wing is

45 degree. The wing has two control surfaces at the tailingedge. The wing has a root chord of 2085 mm and tip

chord of 290 mm. [10]. The various structural components

of a hypersonic aircraft wing structure considered are:

1. Panels Top and bottom skin2. Spars - Front and Rear spar3. Ribs Root rib, Rib 1, rib 2 and Tip ribThe air loads act directly on the wing cover, whichtransmits the loads to the ribs. The ribs transmit the loads

to the spar webs and distribute the load between them in

proportion to the web stiffness. The use of several sparspermit a reduction in rib stresses and also provides a bettersupport for the span wise bending.

Ribs are used to hold the panel to contour shape. The rib

also has another major purpose, to transfer or distributethe loads. The ribs provide stability to spars and panels.

The primary function of the wing skin is to form an

impermeable surface for supporting the aerodynamic

pressure distribution from which the lifting capability ofthe wing is derived.


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649Numerical Analysis of Carbon Fibre Reinforced Aircraft Wing

International Journal of Earth Sciences and EngineeringISSN 0974-5904, Volume 04, No 06 SPL, October 2011, pp 648-651

Structural LayoutThe front spar of aircraft wing is placed about 15% of the

wing chord and inclined to 45 to fuselage longitudinal

axis to facilitate mounting of leading edge and providespace for accommodating landing gear. The rear spar of

aircraft wing is placed about 68% of wing chord to

facilitate mounting of control surfaces. The rear spar is

oriented normal to the fuselage axis for better loadtransfer. Four ribs are connected to the spars to form the

structural frame. The ribs are placed at the root (root rib),

wing strake to basic wing junction (rib-1), inner and outerelevon mounting bracket junction (rib-2) and wing tip (tip

rib). The root rib is placed at the wing root, close to the

fuselage side. The skin of the wing is made in two parts

and placed over the spars and ribs at top and bottom.The primary objective is to develop a light vehicle that

can safely complete the required flights in its specified life

and to accomplish this at minimum cost. Compositematerials consist of strong fibres such as glass or carbon

set in a matrix of plastic or epoxy resin, which is

mechanically and chemically protective. The

incorporation of carbon fibre reinforced polymers

(CFRP), a composite laminate material for the wingpanels effectively reduces the weight of the panels thereby

reducing the total weight of the wing. Weight reduction

and lower production costs are important goals for aircraftstructural engineers and researchers. In recent years, the

use of advanced composite structures has increased to

realize these goals.

Wing Model

The model consists of a hypersonic aircraft wing. Thevarious structural components are top and bottom panels,

four ribs (root rib, tip rib and two intermediate ribs) and

front and rear spars. To study the effects of composite

laminate (CFRP) when compared to the conventional

material (Al. alloy) on wing structures, the panels weremodeled as orthotropic CFRP laminates and the remaining

components were modeled as isotropic material.

The material properties of isotropic and orthotropicmaterials are as given in table 1.

Table 1 Material Properties of Al. Alloy and CFRP

Material Property Al. Alloy(2014-T6)

CFRP (M55j/914prepreg)

Mod. of Elasticity 70 GPa 270 Gpa

Mass Density 2700 kg/m3

1760 kg/m3

Poissons Ratio 0.3 0.365

Tensile Strength 0.380 GPa 1.8 GPa

Shear Strength 0.260 GPa 0.092 GPa

ANSYS is a general purpose finite element modelingpackage for numerically solving a wide variety of

problems. The wing model was first created in CATIAand the analysis was carried out using ANSYS software.

Composite materials can be modeled in ANSYS using

specialized elements called layered elements - Shell 99element. The element has six degrees of freedom at each

node: translations in the nodal x, y and z directions and

rotations about the nodal x, y and z axes. The Shell 63-

Elastic Shell element was used to model the intermediate

components of the wing. The section details of the wing

components are given in table 2.

Table 2 Section Details of Wing Components

Structural Component Section Detail

PanelTop and bottom skin

2mm

Root rib Flanges 5mm

Root rib-Web 3mm

Rib1,Rib2 & tip rib

Flanges & web3mm

Front spar-Flanges & web

Rear spar-Flanges & web

5mm

7.5mm

The primary load on the wing is the aerodynamic

pressure. The finite element mesh of the wing model is

shown in Fig 1

Fig. 1 Finite element mesh of the wing model

The wing model with the elements and boundary

conditions given are shown in Fig. 2. The boundarycondition was given as fixed at the root rib where the

wing attaches to the fuselage.

Fig. 2 Wing model with the elements and boundary

conditions


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650 Sanya Maria Gome, Alice Mathai


The combined wing was then analyzed using the ANSYS

software to obtain the Von Mises stresses and

displacements at the nodal points. Max. values of stresses

and displacements of individual components are given in

table 3

Table 3 Maximum values of stresses and displacements

Wing Component Von-Mises

Stress (N/mm2)

Displacement

(mm)

Root Rib 195.51 3.37

Rib 1 13.17 0.15

Rib 2 14.43 0.09

Tip Rib 99.02 0.57

Front Spar 52.18 1.22

Rear Spar 72.31 5.57

The wing has the configuration of a double delta wing.

The hypersonic aircraft will be boosted to hypersonic

speeds using Solid motor booster and then it separates and

descends using aerodynamic control. Thus, it will be

experiencing the flight environment of an upper stage in

launch vehicle during its ascend phase and the flight

conditions of an aircraft during its descend phase. Thus

the hypersonic aircraft has to survive extreme hot

environment during its flight regime in descend phase.

Therefore, its external surface consisting of the top and

bottom panels are made with carbon fiber reinforced

polymer (CFRP) which has very low coefficient of

thermal expansion. Each laminate can have distinct fiber

orientations which may vary from the adjoining laminates.

The present study focuses on the effect of the ply

orientation on the strength of the panels. The CFRPmaterial was modeled using the finite element software

ANSYS. The CFRP material was modeled using shell 99

element. In this study the top and bottom panels of the

aircraft wing were modeled as CFRP material consisting

of 20 plies each. The thickness of each ply was taken as

0.1mm. In order to study the effect of ply orientation the

following layup sequences were selected [02/902/+/-

/902/02] ns. The angle was varied from 0 degree to 90

degree at intervals of 15 degrees each. The ply sequences

selected for this study are given below in table 4.

Table 4 Ply sequences selected for this studyPly Sequences

[02/902/+0/-0/902/02] ns

[02/902/+15/-15/902/02] ns

[02/902/+30/-30/902/02] ns

[02/902/+45/-45/902/02] ns

[02/902/+60/-60/902/02] ns

[02/902/+75/-75/902/02] ns

[02/902/+90/-90/902/02] ns

The resultant Von Mises Stresses and displacements for

the panels corresponding to the ply orientations were

studied. The Von Mises stress for ply layout sequence -

[02/902/+45/-45/902/02] ns is shown in fig.3

Fig. 3 Von Mises stress for [02/902/+45/-45/902/02] nsThe resultant Von Mises Stresses and displacements for

the bottom panel corresponding to the ply orientations areas given below in table 5.

Table 5 Maximum values of stresses and displacements

for different ply orientations

Ply LayoutSequence

Von-MisesStress (N/mm

2)

Displacement(mm)

[02/902/+0/-0/902/02] ns

209.55 9.57

[02/902/+15/-15/902/02] ns

188.02 6.75

[02/902/+30/-

30/902/02] ns216.39 6.86

[02/902/+45/-45/902/02] ns

136.07 3.89

[02/902/+60/-60/902/02] ns

186.38 5.54

[02/902/+75/-75/902/02] ns

196.05 5.72

[02/902/+90/-90/902/02] ns

251.19 9.92

Von-Mises stresses and displacements obtained for the

various ply layout sequences were studied.

RESULTS AND DISCUSSION

The static analysis of the aircraft wing was carried out inANSYS Software. From the static analysis the stresses

and displacements were found out. The maximum value

of the Von-Mises stress was found to be for the root ribhaving a value of 195.51 N/mm2. The maximum

displacement was found to be for the rear spar. The largest

magnitude of displacement was obtained at the free end ofthe combined wing. The total weight of the aircraft wing

made of the conventional material was seen to be 972kg

while that replaced with CFRP weighed only 633.6kg.

Hence the replacement of Aluminium alloy by CFRP

reduces the total weight of the aircraft wing by 34.82%thereby reducing the total weight of the aircraft

tremendously. From the parametric study conducted by


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651Numerical Analysis of Carbon Fibre Reinforced Aircraft Wing


varying the ply orientations the Von Mises stress value forthe ply sequence [02/902/+45/-45/902/02] ns has the least

value of 136.07N/mm2. The displacement in the bottom

plate corresponding to the ply sequence [02/902/+45/-45/902/02] ns has the least value, which proves that this

ply sequence is seen to have better performance under the

pressure loads given.

CONCLUSIONSWing of a hypersonic aircraft was considered for

structural analysis using finite element method. From the

static analysis the resultant stresses and displacementswere obtained. A study was also conducted to verify the

suitability of carbon fiber reinforced polymer (CFRP) to

be used as the structural material replacing the

conventional material used in aircraft structures aluminium alloy. In the parametric study conducted the

panels of the aircraft wing were modeled as CFRP

material consisting of 20 plies. The thickness of each plywas taken as 0.1mm and the following layup sequences

were selected [02/902/+/-/902/02] ns = [0/0/90/90/+/-

/90/90/0/0] ns. The angle was varied from 0 to 90 at

intervals of 15 each. The following conclusions can be

drawn from the studies conducted.1. The Von Mises stress distribution in the case of wing

is less towards the wings leading and trailing edges and

decreases towards the wing tip.2. The variation in fiber orientation at the same skin

thickness will produce the variation in the Von Mises

stress (increase or decrease).

4. Maximum values of Von-Mises stress was observed atthe support position of the combined wing.

5. The largest magnitude of displacement was obtained atthe free end of the combined wing.

6. The replacement of Aluminium alloy by CFRP reduces

the total weight of the aircraft wing by 34.82%.

7. The parametric study shows that, in the bottom plate the

stress pattern for [02/902/+45/-45/902/02] ns seems to havea more uniform distribution when compared to the other

ply sequences studied.

8. The Von Mises stress value for the ply sequence[02/902/+45/-45/902/02] ns is seen to have the least value

of 136.07N/mm2. The displacement corresponding to the

ply sequence [02/902/+45/-45/902/02] ns is seen to have the

least value, which proves that this ply sequence is seen tohave better performance. Thus it is desirable to adopt the

ply sequence [02/902/+45/-45/902/02] ns for composite

aircraft wings in comparison with the other ply sequencesconsidered in the present study.

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Numerical Analysis of Carbon Fibre Reinforced Aircraft Win

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