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ISTANBUL TECHNICAL UNIVERSITY

FACULTY OF AERONAUTICS AND ASTRONAUTICS

UCK 328E-STRUCTURAL DESIGN

Instructor: Prof. Dr. Zahit Mecitoglu

Term Project: Structural Design And Analyses Of A Wing Of ATA

Unmanned Aircraft

110090052 Kaan Berki KARABAY

110090053 Cemre UNAL

110110083 Selahattin GOKCEN

Spring, 2013


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INDEX

Page No

INDEX ..................................................................................................................................... 2TABLE LIST .......................................................................................................................... 2

FIGURE LIST ........................................................................................................................ 3

1. INTRODUCTION .............................................................................................................. 5

2. GEOMETRY and MATERIAL ........................................................................................ 6

2.1 Geometry ....................................................................................................................... 6

2.2 Material ....................................................................................................................... 10

3. LOADING CONDITIONS .............................................................................................. 11

4. FINITE ELEMENT METHOD ...................................................................................... 16

5. RESULTS OF ANALYSES ............................................................................................. 20

6. EVALUATION ................................................................................................................. 34

BIBLIOGRAPHY ................................................................................................................ 36

ATTACHMENTS...................................................................................................................37

TABLE LIST

Page No

Table 2.1 : Material Numbers..........................................................................................9

Table 2.2 : Properties Of Components...........................................................................10

Table 2.3: Mechanical Properties of Balsa ...............................................................10

Table 2.4: Mechanical Properties of Carbon Pipe ....................................................11

Table 3.1: The Parameters Of The ATA Aircrafts.........................................................12

Table 5.1: Convergency Of Meshes...............................................................................28

Table 5.2 :Vibration Frequencies Of 5 Modes Of The Wing(Hz)................................28

Table 5.3: The Result Of Buckling Analysis For First Modes.......................................31

Table 5.4: The Reaction Forces In Y Direction..............................................................33


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FIGURE LIST Page No

Figure 1.1 : A Sample Image Of ATA Unmanned Aircraft .................................6

Figure 2.1: Top View Of The Wing And Dimensions Of First Geometry...................6

Figure 2.2 : Side View Of The Wing And Dimensions Of First Geometry]................7

Figure 2.3 : The General Image Of Wing................................................................7

Figure 2.4: Top View Of The Wing And Dimensions Of Final Geometry..................8

Figure 2.6: Cross Section Of Carbon Pipe (Thickness : 1 mm)................................8

Figure 2.7 : Cross Section Of Spar (Thickness : 4 mm)...........................................9

Figure 2.8 :Display of Components According To Real Constant Numbers...............9

Figure 3.1: Lift Distribution Of The Single Wing..................................................13

Figure 3.2 : The Calculation Of Parasite Drag Coefficient......................................14

Figure 3.4 : The Pressure Distribution On Wing....................................................15

Figure 3.5: Pressure Distribution On The Rib At The Root.....................................15

Figure 4.1: SHELL 63 Elastic Element ...........................................................16

Figure 4.2: The Meshed Carbon Pipe...................................................................17

Figure 4.3: The Meshed Spar...............................................................................17

Figure 4.4: The Meshed Structure Containing All Components...............................18

Figure 4.6 : Boundry Conditions..........................................................................19

Figure 5.1: Displacement Values Of First Design (Max:53 mm).............................20

Figure 5.2: Von Mises Stress Values Of First Design(Max:40 MPa)........................20

Figure 5.3: Tsai-Wu Failure Criteriation Values Of First Design.............................21

Figure 5.4: Displacement Values ( Max:67 mm)....................................................22


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FIGURE LIST Page No

Figure 5.5: Tsai-Wu Failure Criteria (Max: 0,87)...................................................22

Figure 5.6: Von Mises Stress Values Of Carbon Pipe For Nodes..............................23

Figure 5.7 : Von Mises Stress Values Of Carbon Pipe For Elements.........................24

Figure 5.8 : Von Mises Stress Values For Spars For Nodes ( Max:12 MPa)...............24

Figure 5.9 : Von Mises Stress Values Of Ribs For Elements.....................................25

Figure 5.10: Von Mises Stress Values Of The Spar For Nodes(Max:6.3MP................26

Figure 5.11: Von Mises Stress Values Of The Spar For Elements(Max:6.3 MPa.........26

Figure 5.12: General Image Of Distribution Of Von Mises Stress Values For

Nodes (Max:46 MPa).........................................................................................27

Figure 5.13: General Image Of Distribution Of Von Mises Stress Values For

Elements (Max:47 MPa).........................................................................................28

Figure 5.14: Image Of 1th Mode Of Wing (Max Displacement: 207.3 mm).................29

Figure 5.15: Image Of 2nd Mode Of Wing (Max Displacement: 197 mm)...................29

Figure 5.16: Image Of 3rd Mode Of Wing (Max Displacement: 199 mm)....................30

Figure 5.17: Image Of 4th Mode Of Wing (Max Displacement: 199 mm)....................31

Figure 5.18: Image Of 5th Mode Of Wing (Max Displacement: 277 mm).....................31

Figure 5.19: Front View Of Buckling Analysis For 1st Mode.........................................32

Figure 5.20: Isometric View Of Buckling Analysis For 1st Mode...................................32

Figure 5.20: Side View Of Buckling Analysis For 1st Mode............................................3


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1.INTRODUCTION

An unmanned aerial vehicle (UAV), commonly known as a drone, is an aircraft

without a human pilot on board. Its flight is controlled either autonomously by computers in

the vehicle, or under the remote control of a pilot on the ground or in another vehicle. [1]

ATA unmanned aircraft was designed and built for the competition named Desing

/Build/Fly that was organized by the supports of American Cessna Aircraft Company and

Rahytheon Rocket Systems in 2012, April. ATA Team was all consisted of students of

Istanbul Technical University, Faculty of Aeronautics and Astronautics. In this competition,

ATA Team became 4th that was the best success ever among very popular univercities like

MIT, Illnois, Virginia Technical Univercity.

In this study, the wing of the aircraft was designed and analyzed in CATIA V5 and

ANSYS 14.5, respectively. All the calculations were done in Microsoft Office EXCEL.

First geometry of wing was designed. Because of importance of lightness, balsa and

carbon are choosen as materials for skin and ribs, respectively. SHELL 63 Elastic Shell

element was used for finite element model. Lift and drag forces was calculated by appropriate

equations. These processes were discussed in next chapters in details. Von Mises stress values

were compared to allowable stresses values. The convergency of meshes was examined and

found proper. After providing stress and displacement conditions, Tsai-Wu failure criteriation

and modal analysis under just gravitational load were done. At the end, total weight of the

wing was calculated. Results of analyses were examined in evaluation part. The sources that

were researched were indicated in the bibliography part. A sample image of ATA unmanned

aircraft was shown in Figure 1.1 .


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Figure 1.1 : A Sample Image Of ATA Unmanned Aircraft [2]

2. MATERIAL AND GEOMETRY

2.1. Geometry

2.1.1 First Geometry

For airfoil geometry, MH114 was chosen by designer . With the given data, first geometry

was created. The first geometry of the wing was shown in Figure 2.1, 2.2, 2.3 .

Figure 2.1: Top View Of The Wing And Dimensions Of First Geometry

65 mm 715 mm

59.2 mm

177.3 mm

65 mm

RootTip

61.1 mm


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Figure 2.2 : Side View Of The Wing And Dimensions Of First Geometry

Figure 2.3 : The General Image Of Wing

2.1.2. Final Geometry

After analyzing the first wing, there were found unnecessary parts and some parts were

substracted, then the length of carbon pipe was shorten. The final geometry of the wing was

shown in Figure. There used totally 13 ribs.

10 mm

177,3 mm59,2 mm

Carbon Pipe

Rib

Spar


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Figure 2.4: Top View Of The Wing And Dimensions Of Final Geometry

Figure 2.5: Side View Of The Wing And Dimensions Of Final Geometry

Chord length and wing span were 236.5 mm and 780 mm, respectively.According to

these dimensions, wing area equals to 0,185 m. The carbon pipe had offset. Because the

offset was mounted to fuselage and connection point for the wing.

Figure 2.6: Cross Section Of Carbon Pipe (Thickness : 1 mm)

The cross sections of the carbon pipe and spar were shown in Figure 2.6 and 2.7,

respectively.

This part was substracted. The new length

of carbon pipe was 260 mm.

36.4 mm 23 mm 29.1 mm 48 mm 100 mm

RootTip

9 mm 10 mm15 mm

7 mm

10 mm

8 mm

61.1 mm


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Figure 2.7 : Cross Section Of Spar (Thickness : 4 mm)

The cross section areas were considered same for all the structure because of easiness in

designing process.The thickness of the skin was also considered 1 mm for all the structure.

The properties of the components were shown in Table 2.1 , 2.2 and Figure 2.8.

Table 2.1 : Material Numbers

Material Number 1 Balsa

Material Number 2 Carbon Tube

Figure 2.8 :Display of Components According To Real Constant Numbers

4 mm24.97 mm

1

2

3


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Table 2.2 : Properties Of Components

Structure Real Constant Number Thickness

Rib 1 3 mm

Skin 2 1 mm

Spar 3 4 mm

2.2.Material

Model aircraft had to be very light and resistent because of the missions of the competition.

Because of that carbon tube was choosen for the pipe where the wing connected to the

fuselage. The other parts of the wing was made from Balsa that had very high elasticity

modulus in x direction. Safety factor was considered 1,3 because of the importance of

lightness. The mechanical properties of the carbon tube and balsa were shown in Table 2.3

and Table 2.4, respectively. Also the allowable stresses in different directions were calculated

in these tables.

Table 2.3: Mechanical Properties of Balsa [3]

Elasticity Modulus in X Direction 4600000 MPa

Elasticity Modulus in Y Direction 110 MPa

Elasticity Modulus in Z Direction 110 MPa

Poisson Ratio in XY Plane 0,3Poisson Ratio in YZ Plane 0,00717

Poisson Ratio in XZ Plane 0,00717

Shear Modulus in XY Plane 600 Mpa

Shear Modulus in YZ Plane 600 Mpa

Shear Modulus in XZ Plane 360 Mpa

Density 150 kg/m Allowable Stress

Tensile Yield Strength in X Direction 32,5 MPa 32,5/1,3=25 MPa

Tensile Yield Strength in Y Direction 27,5 MPa 25,7/1,3=19,8MPa


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Tensile Yield Strength in Z Direction 27,5 MPa 25,7/1,3=19,8MPa

Compression Yield Strength in X Direction 19,5 MPa 19,5/1,3=15 MPa

Compression Yield Strength in Y Direction 16,5 MPa 16,5/1,3=12,7MPa

Compression Yield Strength in Z Direction 16,5 MPa 16,5/1,3=12,7MPa

Table 2.4: Mechanical Properties of Carbon Pipe [4]

Elasticity Modulus in X Direction 142000 Mpa

Elasticity Modulus in Y Direction 10300 MPa

Elasticity Modulus in Z Direction 10300 Mpa

Poisson Ratio in XY Plane 0,27

Poisson Ratio in YZ Plane 0,0195

Poisson Ratio in XZ Plane 0,0195

Shear Modulus in XY Plane 7200 MPa

Shear Modulus in YZ Plane 7200 MPa

Shear Modulus in XZ Plane 4824 MPa

Density 1500 kg/m Allowable Stresses

Tensile Yield Stress in X Direction 1035 MPa 1035/1,3=796,2 MPa

Tensile Yield Stress in Y Direction 41 MPa 41/1,3= 31,6 MPa

Tensile Yield Stress in Z Direction 41 MPa 41/1,3=31,6 MPa

Compression Yield Stress in X Direction 689 MPa 689/1,3= 530 MPa

Compression Yield Stress in Y Direction 117 MPa 117/1,3=90 MPa

Compression Yield Stress in Z Direction 117 MPa 117/1,3= MPa

3.LOADING CONDITIONS

Because of the manueveirs of the aircraft the load factor was considered 1.5 . Load factor

is defined as the as the ratio of the lift of an aircraft to its weight. The total lift for for the one

wing can be calculated from Formula 1. The total weight of the structure was 18 N.

(1)


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The lift force was for two wings. Since one wing was analyzed, the lift force was divided

by 2. So the total lift force was 13,5 N.

First, some parameters were needed. The calculated parameters were shown in Table 3.1:

Table 3.1: The Parameters Of The ATA Aircrafts

The lift distribution of lift and drag distributions were calculated as follows:

Lift DistributionThe wing span and the magnitute of circulation an the root of a three dimensional

wing were 2s and 0. The formulas used in calculations were below:

With the data of coordinates given, circulation was found. Then inserting the lift force

formula lift force was found. The details of the calculations were given in attachments. The

graph of the lift distribution was shown in Figure 3.1.


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Figure 3.1: Lift Distribution Of The Single Wing

Drag DistributionFor a wing with an eliptical lift distribution, induced drag is calculated as follows:

[6]


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For a wing with an eliptical lift distribution, parasite drag was calculated and shown

in Figure 3.2 .

Figure 3.2 : The Calculation Of Parasite Drag Coefficient

The total drag force was calculated from the formula given below:

For all the calculations, the freestream density was taken 1,226 kg/m3 and total drag was

calculated as 2,5653 N. The wing was divided into 12 zones and these zones were divided

into 4 areas. The areas were shown in Figure 3.3.

Figure 3.3 : The Different Divided Areas


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The distributions of drag and lift forces were applied as pressures and shown in Figure

3.4 and 3.5.

Figure 3.4 : The Pressure Distribution On Wing

Figure 3.5: Pressure Distribution On The Rib At The Root


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4.FINITE ELEMENT MODEL

After determining the geometry and appropriate element was choosen as SHELL 63 Elastic

Element. SHELL63 is well suited to model linear, warped, moderately-thick shell structures.

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 deformation shapes are linear in

both in-plane directions. For the out-of-plane motion, it uses a mixed interpolation of tonsorial

components.The geometry, node locations, and the coordinate system for this element are

shown in Figure 4.1.

Figure 4.1: SHELL 63 Elastic Element [7]

The element is defined by four nodes, four thicknesses, and the orthotropic material

properties. A triangular-shaped element may be formed by defining the same node number for

nodes K and L as described in Triangle, Prism and Tetrahedral Elements. The element has

plasticity, creep, stress stiffening, large deflection, and large strain capabilities.The meshed

Figures are shown in Figure 4.2, 4.3 and 4.4.


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Figure 4.2: The Meshed Carbon Pipe

Figure 4.3: The Meshed Spar


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Figure 4.4: The Meshed Structure Containing All Components

Figure 4.5: The Meshed Ribs


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In the wing structure there used totally 19307 nodes and 19961 elements. Because of

limitation on the number of elements, there could not be used another mesh. For that reason

the most appropriate mesh was choosen and element length was 5 mm. So the convergency of

different meshes could not be examined also.

For fixing the sutructure spar and carbon pipe were connected to wing. For that reason,

boundry conditions were applied on spar and carbon pipe. The degree of freedom or the

surfaces were limited in every direction. It was shown in Figure 4.6.

Figure 4.6 : Boundry Conditions


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5.RESULTS OF ANALYSIS

Results Of Analyses Of First Design

Figure 5.1: Displacement Values Of First Design (Max:53 mm)

Figure 5.2: Von Mises Stress Values Of First Design(Max:40 MPa)


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Displacement and Von Mises Stress values of first design were shown in Figure 5.1

and 5.2, respectively. The results were found appropriate. But because of the competition

rules and the missions that ATA unmanned aircraft would do, the lightness was very

important property. For that reason, the length of the carbon pipe whose density was very high

compared to balsa material was shorten and the unnecessary parts of the ribs were substracted.

After these prosesses, desired geometry was created. Also Tsai-Wu failure criteriation of first

design was shown in Figure 5.3. The structure was stable that failure criteriation value did not

exceed 1 under the loading conditions.

Figure 5.3: Tsai-Wu Failure Criteriation Values Of First Design


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Results Of Analyses Of Final DesignIn this study, static and modal analysis of the components of the wing structure were

examined. For each component the analysis results were shown below.

Figure 5.4: Displacement Values ( Max:67 mm)

Figure 5.5: Tsai-Wu Failure Criteria (Max: 0,87)


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The displacement values are appropriate for the structure. Maximum displacement

occured at the tip and the value was found 67 mm. After that Tsai-Wu criteriation was

examined and found maximum 0,87. Tsai-Wu criteriation values was under 1 and there was

no problem in the wing structure. The displacement and Tsai-Wu results were shown in

Figure 5.4 and 5.5 respectively.

Figure 5.6: Von Mises Stress Values Of Carbon Pipe For Nodes

Figure 5.6 and 5.7 showed the Von Mises Stress values for nodes and elements,

respectively. The maximum stress values were found 46 MPa and 47 MPa. Maximum stresses

were under the allowable stress values and occured at near the middle of the tube as expected.


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Figure 5.7 : Von Mises Stress Values Of Carbon Pipe For Elements

Figure 5.8 : Von Mises Stress Values For Spars For Nodes ( Max:12 MPa)


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Figure 5.9 and 5.10 showed the Von Mises Stress values of spars for nodes and elements,

respectively. The maximum stress values were found 12 MPa and 17.9 MPa. Maximum

stresses were under the allowable stress values and occured at the hole of the root rib as

expected. The great difference between nodal and element solution is the cause of mesh. As

mentioned before, because of limitation of the number of elements there could not be used

more elements. Here could not be found exact value. But the locations where the maximum

stress occured were same. Also the convergency of the meshes were not appropriate for the

results for ribs. In fact, the geometry of the ribs was the main reason for the irregularity of the

meshes.

Figure 5.9 : Von Mises Stress Values Of Ribs For Elements


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Figure 5.10: Von Mises Stress Values Of The Spar For Nodes(Max:6.3MPa)

Figure 5.11: Von Mises Stress Values Of The Spar For Elements(Max:6.3 MPa)


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Von Mises stress values of the spar for nodes and elements were shown in Figure 5.10 and

5.11 respectively. The maximum stress values were found 6.3 MPa and same. It did not

exceed the allowable values. As the geometry was rectangle prism and smooth, the meshes

was excellent. Because of that the convergency of the meshes was good.

Figure 5.12: General Image Of Distribution Of Von Mises Stress Values For Nodes (Max:46 MPa)

As understood from the figures none of the Von Mises stress values exceeded the

allowable values. The values of the carbon pipe were the same with general distribution. So

the maximum stress occured on the carbon pipe. Because of that reason the skin thickness was

considered as thin as possible and 1 mm. In addition, the rib on the tip of wing was not same

with the others. It remained same with first geometry for easiness in the covering prosess.

There was one hole that carbon pipe connected as understood from Figure 5.12 and 5.13.


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Figure 5.13: General Image Of Distribution Of Von Mises Stress Values For Elements (Max:47 MPa)

Table 5.1: Convergency Of Meshes

Component Element Solution(MPa) Nodal Solution (MPa) Convergency(%)

Carbon Pipe 47.2 45.9 2.75

Rib 17.9 11.9 33

Spar 6.37 6.37 0

The convergency of meshes were shown in Table 5.1. After Von Mises stress analyses,

modal analyses of 5 modes of the wing were examined and the vibration frequencies were

found as shown in Table 5.2. The displacements for 5 modes were shown in Figure 5.14, 5.15,

5.16, 5.17, 5.18.

Table 5.2 :Vibration Frequencies Of 5 Modes Of The Wing(Hz)


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Figure 5.14: Image Of 1th

Mode Of Wing (Max Displacement: 207.3 mm)

Figure 5.15: Image Of 2

nd

Mode Of Wing (Max Displacement: 197 mm)


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Figure 5.16: Image Of 3rd





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After examining the modal analysis, also buckling analysis of the wing was done for one

mode. The reason of the negative value in the frequency was reasearched and consulted to

instructors, but could not be found. The value was shown in Table 5.3 and the images of the

buckling analyses were shown in Figure 5.19, 5.20 and 5.21.

Table 5.3: The Result Of Buckling Analysis For First Mode


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Figure 5.19: Front View Of Buckling Analysis For 1st

Mode

Figure 5.20: Isometric View Of Buckling Analysis For 1st

Mode


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Figure 5.20: Side View Of Buckling Analysis For 1st

Mode

Maximum displacement for buckling analysis for first mode was found 1mm. It was

considered appropriate.

Weight Of The StructureTable 5.4: The Reaction Forces In Y Direction

With Gravity Without Gravity

First Design -12.208 N -13.530 N

Final Design -12.501 N -13.530 N

As understood from Table 5.4, in the first design, the weight of the structure was

found 1.32N, 137.6 g. After idealizing the wing total weight was found 1.029N, 105 gr.


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6.EVALUATION

In this study, design and analyses of the wing of ATA unmanned aircraft that

participated in 'Design/Build/Fly' competition that was held by the supports of American

Cessna Aircraft Company and Rahytheon Rocket Systems in 2012, April.

Design, analyses and calculations were done in CATIA V5 , ANSYS 14.5 and

Microsoft Office EXCEL program, respectively.

Firstly, the wing was designed then analyses were done.Lightness was very important

parameter for ATA aircraft.Because of that load and safety factor were taken 1.5 and

1.3,respectively.Under these conditions, Displacement and Von Mises stress values were

found 53 mm and 40 MPa. Results were considered appropriate. But that was not ideal one.

For finding the desired design the wing was optimized. Because of importance of lightness,

balsa and carbon are choosen as materials for skin and ribs, respectively. SHELL 63 Elastic

Shell element was used for finite element model.Structure was modeled with 5 mm-length

elements. There used 19307 nodes and 19961 elements. Because of limitation on the number

of elements, there could not be used another mesh. Total lift and drag forces were 13,5 N.

Dividing the wing 14 zones, again dividing zones into 4 areas, the pressures were applied.

The wing was attached from spar and carbon pipe to fuselage. Because of that the movement

of areas that were connected was limited in translation and rotation in x, y, z directions. In the

results part, displacement values were found suitable for 76cm wing. Then none of Von Mises

values exceeded the allowable stress values for different components. The convergency of

meshes were good except mesh of ribs. Because of the geometry of ribs meshes were irregular

and convergency was too high. Tsai-Wu failure criteriation values were under 1 as expected.

Then 5 modes of modal analyses were done. The vibrations were evaluated as normal. Then

buckling analysis was done for one mode and maximum displacement for that mode was 1


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mm. The sources used and calculations were given in bibliography and attachments,

respectively. For under all these conditions, the structure can be tought as resistent and could

do missions successfully.

This study was just an approximation to real case. In theory, under some assumptions,

the structure had no problem. But to get exact results, the aircraft had to be flied and tested by

experts.


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BIBLIOPGRAPHY

[1]Url-1 Taken Date : 01.05.2013

[2]Url-2 Taken Date : 01.05.2013

[3]Url-3 Taken Date : 06.05.2013

[4]Url-4

[5]Url-5

[6]Url-6

[7] ANSYS 14.5 Help Topics

ATTACHMENTS


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ATTACHMENTS

Area(mm2) Pressure (Mpa) Applied Pressure(Mpa)

1. Area (upper) 8071 0,00008784927 0,000114204

2.Area (upper) 8018 0,00008727239 0,000113454

3. Area (Lower) 7760 - 0,0000263554.Area (Lower) 7825 - 0,000026182

1. Area (upper) 8071 0,00008614798 0,000111992

2.Area (upper) 8018 0,00008558227 0,000111257

3. Area (Lower) 7760 - 0,000025844

4.Area (Lower) 7825 - 0,000025675

1. Area (upper) 8071 0,00008538880 0,000111005

2.Area (upper) 8018 0,00008482807 0,000110276

3. Area (Lower) 7760 - 0,000025617

4.Area (Lower) 7825 - 0,000025448

1. Area (upper) 8071 0,00008419116 0,000109449

2.Area (upper) 8018 0,00008363830 0,0001087303. Area (Lower) 7760 - 0,000025257

4.Area (Lower) 7825 - 0,000025091

1. Area (upper) 8071 0,00008247669 0,000107220

2.Area (upper) 8018 0,00008193509 0,000106516

3. Area (Lower) 7760 - 0,000024743

4.Area (Lower) 7825 - 0,000024581

1. Area (upper) 8071 0,00008012876 0,000104167

2.Area (upper) 8018 0,00007960258 0,000103483

3. Area (Lower) 7760 - 0,000024039

4.Area (Lower) 7825 - 0,000023881

1. Area (upper) 8071 0,00007697721 0,0001000702.Area (upper) 8018 0,00007647172 0,000099413

3. Area (Lower) 7760 - 0,000023093

4.Area (Lower) 7825 - 0,000022942

1. Area (upper) 8071 0,00007276948 0,000094600

2.Area (upper) 8018 0,00007229162 0,000093979

3. Area (Lower) 7760 - 0,000021831

4.Area (Lower) 7825 - 0,000021687

1. Area (upper) 8071 0,00006711053 0,000087244

2.Area (upper) 8018 0,00006666983 0,000086671

3. Area (Lower) 7760 - 0,000020133

4.Area (Lower) 7825 - 0,0000200011. Area (upper) 8071 0,00005931621 0,000077111

2.Area (upper) 8018 0,00005892669 0,000076605

3. Area (Lower) 7760 - 0,000017795

4.Area (Lower) 7825 - 0,000017678

1. Area (upper) 8071 0,00004793766 0,000062319

2.Area (upper) 8018 0,00004762287 0,000061910

3. Area (Lower) 7760 - 0,000014381

4.Area (Lower) 7825 - 0,000014287

1. Area (upper) 8071 0,00002714268 0,000035285

2.Area (upper) 8018 0,00002696445 0,000035054

3. Area (Lower) 7760 - 0,0000081434.Area (Lower) 7825 - 0,000008089

9.Zone

10.Zone

11.Zone

12.Zone

8.Zone

3.Zone

4.Zone

5.Zone

6.Zone

7.Zone

1.Zone

2.Zone


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