ISTANBUL TECHNICAL UNIVERSITY FACULTY OF …

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ISTANBUL TECHNICAL UNIVERSITY FACULTY OF AERONAUTICS AND

ASTRONAUTICS

PERFORMANCE EVALUATION OF MORPHING WING STRUCTURES AND

THEIR APPLICATIONS ON WINGLETS

GRADUATION PROJECT

Hasan Furkan ARBAZ

Department of Aeronautical Engineering

Anabilim Dalı : Herhangi Mühendislik, Bilim

Programı : Herhangi Program

Thesis Advisor: Yrd. Doç. Dr. Demet BALKAN

February, 2021

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ISTANBUL TECHNICAL UNIVERSITY FACULTY OF AERONAUTICS AND

ASTRONAUTICS

PERFORMANCE EVALUATION OF MORPHING WING STRUCTURES AND

THEIR APPLICATIONS ON WINGLETS

GRADUATION PROJECT

Hasan Furkan ARBAZ

110170505

1

Department of Aeronautical Engineering

Anabilim Dalı : Herhangi Mühendislik, Bilim

Programı : Herhangi Program

Thesis Advisor: Yrd. Doç. Dr. Demet BALKAN

February, 2021

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Hasan Furkan ARBAZ ,student of ITU Faculty of Aeronautics and Astronautics

student ID 110170505, successfully defended the graduation entitled “Performance

Evaluation of Morphing Wing Structures and Their Applications on Winglets”,

which he prepared after fulfilling the requirements specified in the associated

legislations, before the jury whose signatures are below.

Thesis Advisor: Yrd. Doç. Dr. Demet BALKAN ..............................

İstanbul Technical University

Jury Members: Prof. Dr. Halit Süleyman TÜRKMEN .............................


Prof. Dr. Vedat Ziya DOĞAN ..............................


Date of Submission: 01.02.2021

Date of Defense: 08.02.2021

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To my beloved father and to my family,

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FOREWORD

This study, in which effects of variable cant angle morphing winglets were evaluated,

discussed, designed and analyzed, with my academic knowledge and experience, and

my thesis advisor, who supported me throughout my studies, I would like to thank

Assoc. Prof. Dr. Demet BALKAN.

I would like to thank my dear father whom I recently lost, for dedicating everything he

had to see me having a good future. He was always with me and he will always be. I

also thank my mother and brother who have supported me during my whole education

life.

February, 2021 Hasan Furkan ARBAZ

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TABLE OF CONTENTS

Page

FOREWORD ....................................................................................................... vii

ABBREVIATIONS ................................................................................................ x

LIST OF TABLES ................................................................................................ xi

LIST OF FIGURES ............................................................................................. xii

SUMMARY ......................................................................................................... xiv

ÖZET .....................................................................................................................xv

1. INTRODUCTION .......................................................................................... 1

1.1 What is Morphing Wing? ........................................................................... 3

2. EFFECTS OF WINGLETS ON OVERAL WING PERFORMANCE AT

VARIABLE CANT ANGLES ............................................................................... 6

3. WINGLET DESIGN ......................................................................................13

3.1 Preliminary Design ...................................................................................13

3.2 CAD Drawing of The Design ....................................................................13

4. PLAUSIBLE SKIN MATERIAL CANDIDATES FOR MORPHING

STRUCTURES .....................................................................................................17

4.1 Materials That Are Evaluated Through .....................................................17

4.2 Comparison of The Materials ....................................................................19

5. ANSYS ANALYSIS OF SELECTED MATERIAL FOR DIFFERENT

CANT ANGLES....................................................................................................22

5.1 CFD Analysis (with Fluent) ......................................................................22

5.2 FSI (Fluid-Structure Interaction) Analysis.................................................44

6. RESULTS AND CONCLUSIONS ................................................................55

7. REFERENCES ..............................................................................................56

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ABBREVIATIONS

UAV : Unmanned Aerial Vehicle

NACA : National Advisory Commitee for Aeronautics

AOA : Angle of Attack

CFD : Computational Fluid Dynamics

CAD : Computer Aided Design

CAM : Computer Aided Manufacturing

SMM : Shape Memory Material

SMP : Shape Memory Polymer

SMPC : Shape Memory Polymer Composite

PAM : Pneumatic Artificial Muscle

CFD : Computational Fluid Dynamics

FSI : Fluid Structure Interaction

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LIST OF TABLES

Page

Table 1: Comparison of materials after uniaxially tested ........................................ 20 Table 2: Comparison of materials for strain and hold experiment ........................... 21

Table-3: Mechanical Properties of Tecoflex 80A ................................................... 53 Table-4: Analysis data for each cant angle configuration........................................ 54

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LIST OF FIGURES

Page

Figure 1: An example of winglet structure ............................................................... 1

Figure 2 : Winglet-like bird feathers for enhanced flight efficiency ......................... 2 Figure 3: An Aircraft’s Drag Profile ........................................................................ 3

Figure 4: A morphing wing with different camber configurations............................ 5 Figure 5: Winglet Cant Angle ................................................................................. 6

Figure 6: 90o,60o and 45o cant angle configurations ................................................. 6 Figure 7: Lift Coefficient (CL) For Different Winglet Cant Angles ......................... 7

Figure 8: Drag Coefficient (CD) For Different Winglet Cant Angles ........................... 7

Figure 9: CL/CD Values For Different Winglet Cant Angles ................................... 8 Figure 10: CL/CD Graph For Different Winglet Cant Angles .................................. 8

Figure 11: Predicted particle path lines in case wing without winglet at angle of

attack 12o ................................................................................................................ 9

Figure 12: Predicted particle pathlines, WCSA-45-00 at (a) AoA 0o, (b) AoA of 12o 9 Figure 13: CL values of winglet configurations with various AoAs .........................10

Figure 14: CD values of winglet configurations with various AoAs ........................11 Figure 15: L/D values of winglet configurations with various AoAs .......................12

Figure 16: Coordinates exported from Excel to CATIA for Lrn-1015 il profile (150

mm chord) ..............................................................................................................14

Figure 17: Winglet design having 0o cant angle (no winglet configuration) ............14 Figure 18: Winglet design having 15o cant angle ....................................................15

Figure 19: Winglet design having 30o cant angle ....................................................15 Figure 20: Winglet design having 45o cant angle ....................................................16

Figure 21: Dimensions for servo pocket .................................................................16 Figure 22: Graphic showing the elastic modulus change with respect to temperature

...............................................................................................................................19 Figure 23: Project Schematic in ANSYS Workbench for the analysis .....................22

Figure 24: Flow volume for the CFD analysis of 0o cant angle ...............................23 Figure 25: Cross sectional are of flow volume for 0o cant angle .............................23

Figure 26: Named selection for winglet airfoil for 0o cant angle .............................24 Figure 27: Surface mesh parameters for 0o cant angle.............................................24

Figure 28: Surface mesh for 0o cant angle ..............................................................25 Figure 29: Visualization of flow volume in finite elements for 0o cant angle ..........25

Figure 30: Boundary layer mesh for 0o cant angle ..................................................25 Figure 31: Mesh volume fill and min. cell length of the mesh for 0o cant angle ......26

Figure 32: Boundary scope and other parameters of mesh for 0o cant angle ............26 Figure 33: Flow analysis case for 0o cant angle ......................................................27

Figure 34: Viscous model selections for 0o cant angle ............................................28 Figure 35: Solution methods for 0o cant angle ........................................................29 Figure 36: Residual monitors for 0o cant angle .......................................................29

Figure 37: Run calculation for 0o cant angle ...........................................................30 Figure 38: Pressure distribution over airfoil for 0o cant angle .................................30

Figure 39: Pressure distribution over winglet for 0o cant angle ...............................31 Figure 40: Streamlines for 0o cant angle .................................................................31

Figure 41: Velocity vectors around airfoil for 0o cant angle ....................................32 Figure 42: Flow volume for the CFD analysis of 15o cant angle .............................32

Figure 43: Named selection for winglet airfoil for 15o cant angle ...........................33

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Figure 44: Visualization of flow volume in finite elements for 15o cant angle ........ 33

Figure 45: Boundary layer mesh for 15o cant angle ................................................ 34 Figure 46: Flow analysis case for 15o cant angle .................................................... 34

Figure 47: Pressure distribution over airfoil for 15o cant angle ............................... 35 Figure 49: Streamlines for 15o cant angle ............................................................... 36

Figure 50: Velocity vectors around airfoil for 15o cant angle.................................. 36 Figure 51: Flow volume for the CFD analysis of 30o cant angle ............................. 37

Figure 52: Named selection for winglet airfoil for 30o cant angle ........................... 37 Figure 53: Visualization of flow volume in finite elements for 30o cant angle ........ 38

Figure 54: Boundary layer mesh for 30o cant angle ................................................ 38 Figure 55: Pressure distribution over airfoil for 30o cant angle ............................... 39

Figure 56: Pressure distribution over winglet for 30o cant angle ............................. 39 Figure 57: Streamlines for 30o cant angle ............................................................... 40

Figure 58: Velocity vectors around airfoil for 30o cant angle.................................. 40 Figure 59: Flow volume for the CFD analysis of 45o cant angle ............................. 41

Figure 60: Named selection for winglet airfoil for 45o cant angle ........................... 41 Figure 61: Visualization of flow volume in finite elements for 45o cant angle ........ 42

Figure 62: Pressure distribution over airfoil for 45o cant angle ............................... 42 Figure 63: Pressure distribution over winglet for 45o cant angle ............................. 43

Figure 65: Velocity vectors around airfoil for 45o cant angle.................................. 44 Figure 66: Pressure distribution on upper surface of airfoil for 0o cant angle .......... 45

Figure 67: Pressure distribution on lower surface of airfoil for 0o cant angle .......... 45 Figure 69: Total deformation distribution on lower part of airfoil for 0o cant angle 46

Figure 70: Equivalent stress distribution on upper part of airfoil for 0o cant angle .. 47 Figure 71: Equivalent stress distribution on lower part of airfoil for 0o cant angle .. 47

Figure 72: Pressure distribution on upper surface of airfoil for 15o cant angle ........ 47 Figure 73: Pressure distribution on lower surface of airfoil for 15o cant angle ........ 48

Figure 74: Total deformation distribution on upper part of airfoil for 15o cant angle

............................................................................................................................... 48 Figure 75: Total deformation distribution on lower part of airfoil for 15o cant angle

............................................................................................................................... 48 Figure 76: Equivalent stress distribution on upper part of airfoil for 15o cant angle 49 Figure 77: Equivalent stress distribution on lower part of airfoil for 0o cant angle .. 49

Figure 78: Pressure distribution on upper surface of airfoil for 30o cant angle ........ 49 Figure 85: Pressure distribution on lower surface of airfoil for 45o cant angle ........ 52

Figure 88: Equivalent stress distribution on upper part of airfoil for 45o cant angle 53

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PERFORMANCE EVALUATION OF MORPHING WING STRUCTURES

AND THEIR APPLICATIONS ON WINGLETS

SUMMARY

For the past few years, the focus on multi role aircrafts have significantly

increased. To provide multi role feature to aircrafts, one of the cleverest ideas, is to be

able change the shape of airfoil and the other extensions of wing during the course of

a flight. To be able to manage that, interest in morphing structures increases every day.

The morphing wing is a birdlike wing that has the ability to adapt to accommodate

multiple flight regimes or to obtain better flight performance. With the help of

morphing structures, need for high lift devices will be vanished. Thus, the aircraft will

get rid of extra structural weight and drag.

In this study, a morphing winglet structure has been proposed to optimize fuel

consumption and flight range by avoiding any extra wing tip vortices during the flight.

With the help of the design that is proposed, the winglet of the aircraft will be able to

change its cant angle for different flight regimes. The benefits of this morphing

structure have been discussed.

After the discussion of the benefits of this morphing structure, a variable cant

angle winglet design is made via CATIA software. In this design the winglet is

morphed in 3 stages. Total profile has the thickness of 180mm.

After the design is made, an adequate skin material selection has been

investigated to be used with this morphing winglet structure. 4 different skin material

candidates which are; polyurethane, copolyester, shape memory polymer and woven

materials are inspected and different studies are discussed to select the most suitable

skin material for the proposed morphing winglet design. After the material selection,

an ANSYS analysis conducted to see how winglet acts at the cant angles of 0o, 15o ,

30o , 45o

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ŞEKİL DEĞİŞTİREBİLEN KANAT YAPILARININ PERFORMANS

DEĞERLENDİRMESİ VE BU YAPILARIN KANATÇIK TASARIMI

ÜZERİNDE UYGULAMALARI

ÖZET

Son yıllarda çok rollü hava araçlarına olan odak artmaktadır. Bir hava aracına

çok rollülük kazandırmanın en akıllıca yollarından birisi, hava aracının kanat profilinin

ve kanada bağlı diğer uzantılarının şeklini değiştirme özelliğini hava aracına

kazandırmaktır. Bunu başarabilmek adına şekil değiştirebilen yapılara olan ilgi gün

geçtikçe artmaktadır. Şekil değiştirebilen kanatlar; kuş kanadına benzer yapıda olup

birden fazla uçuş rejimine uyum sağlayabilen veya uçuş esnasında performansı

optimize edebilen kanatlardır. Şekil değiştirebilen bu yapılar sayesinde, kaldırma

kuvvetini artırmaya yardımcı uçuş aletlerine olan ihtiyaç ortadan kalkacaktır.

Böylelikle hava aracı ilave ağırlık ve sürüklemeden kurtarılacaktır.

Bu çalışmada, uçuş sırasında ilave kanat ucu girdaplarının etkisini azaltacak,

yakıt tüketimini ve uçuş menzilini optimize etmeye yardımcı olacak, şekil

değiştirebilen bir kanatçık tasarımı önerilmiştir. Önerilen bu kanatçık tasarımıyla, hava

aracının kanatçık bükülme açısı, değişik uçuş rejimlerine göre ayarlanabilecektir. Şekil

değiştirebilen bu kanatçık tasarımının faydaları, çalışma boyunca tartışılmıştır.

Şekil değiştirebilen bu yapının faydaları tartışıldıktan sonra, değişebilir bükülme

açısına sahip bir kanatçık tasarımı, CATIA V5 yazılımıyla tasarlanmıştır. Bu

tasarımda kanatçık, 3 aşamada biçimlendirilmiştir. Toplam profil kalınlığı 180

mm’dir.

Tasarım işleminin ardından, şekil değiştirebilen bu kanatçık yapısı için uygun

bir yüzey malzemesi araştırması yapılmıştır. Poliüretan, kopoliester, şekil hafızalı

polimer ve dokunmuş malzemeler olmak üzere 4 farklı yüzey malzemesi adayı

incelenmiştir. Bu alanda yapılan çeşitli çalışmaların ışığında, önerilen kanatçık

tasarımı için bir malzeme seçimi çalışması yapılmıştır. Malzeme seçiminin ardından,

kanatçığın 0o, 15o , 30o ve 45o burulma açılarındaki davranışları ANSYS programıyla

analiz edilmiştir.

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1. INTRODUCTION

Winglets are one of the most commonly used parts of an aircraft’s wing, which

have great effect on diminishing fuel consumption. Wing tip vortices highly effect an

aircraft’s performance during flight due to induced drag created. Winglets are designed

to eliminate the effects of induced drag as much as possible. These wingtip devices

artificially increase the span of the wing. Increasing span means reducing the lift-

induced drag.

Some studies showed that, using winglets as wingtip device in an aircraft can

decrease the cost of fuel consumption up to 4-6 %. They also help to reduce takeoff

distance and increase climb rate. Winglets help to increase effective aspect ratio

without actually needing to increase span.

Figure 1: An example of winglet structure [1]

Today, most of the winglet designs are for only cruise flight condition. In cruise

flight, winglets have considerable effect on increasing L/D ratio and diminish the

negative effects of drag. But throughout the whole flight mission, aircrafts don’t only

fly in cruise flight condition. There are several phases of flight conditions during a

mission such as takeoff, climbing, descending, landing etc. During non-cruise flight

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conditions, conventional winglet types don’t seem to increase the overall performance

of the aircraft. This means conventional winglet designs do not provide maximum

efficiency during takeoff, climbing and landing in terms of fuel consumption

reduction. Non-cruise flight condition has a large fraction during whole flight which

means winglet designs must be reconsidered and optimized for non-cruise conditions,

as well.[2]

When we observe the nature, we can see that, also birds have different feathers for

their different flight conditions. They adapt their wings with their feathers for each

flight condition. This is an inspiring aspect for human being to design the air vehicles

in such way. [2]

Figure 2 : Winglet-like bird feathers for enhanced flight efficiency [2]

As explained before, winglets have great effects on aerodynamic performance.

But while we were mentioning the positive effects of winglets, they might have some

negative effects, as well. The increment of root moment of the wing is one of the few

drawbacks of using winglets. Also, they cause some skin friction and pressure drag

which cause to increase parasite drag. As we recall total drag equals to the sum of

parasite drag and induced drag. Winglets will only be beneficial only when the

reduction of induced drag is greater than the increment in the parasite drag which is

because of winglets.

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Figure 3: An Aircraft’s Drag Profile (source:Leeham Co.)

In this study, we designed a winglet structure which is able to change its cant

angle during flight. But our configuration is not like configurations that contain hinge

structures. In our study, we designed a winglet extension on a morphing wing structure.

This means unity of our coating material of our wing doesn’t change when the winglet

shape is changed. To do this, we studied both on proper cant angle changing structure

design and flexible material selection. Our design will be available for unmanned

aerial vehicles.

1.1 What is Morphing Wing?

The morphing wing is a birdlike wing that has the ability to adapt to

accommodate multiple flight regimes or to obtain better flight performance. (Inman J.

et.al., 2005, p.528) There are several applications of morphing wings to enhance flight

efficiency such as; wing twist, wing span change, wing sweep change, camber change

etc. With the help of morphing structures, need for high lift devices will be vanished.

Thus, the aircraft will get rid of extra structural weight and drag.[3]

As we mentioned, we are going to design an adaptive morphing wing structure

for an unmanned aerial vehicle. The reason behind this selection is all about multi-

tasking. UAVs are mostly used for military applications. A UAV can either be used as

a reconnaissance vehicle or an attacking aircraft. During an operation, the aircraft

sometimes needs to be fast and agile and sometimes it needs to be stable and endurable

while staying in the air. Most of the wing structures of UAVs today do not meet such

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requirements. Because they don’t change span, winglet angle, sweep angle etc. This

causes huge performance, endurance and fuel consumption problems. Adaptive wings

are exactly manufactured for these tasks. They can increase span, alter the winglet cant

angle, change the camber of the airfoil. This brings huge advantages for every UAVs

in terms of multi-tasking capabilities. In this study, we are going to specifically focus

on adaptive winglet structures. We are going to propose an adaptive winglet design for

an unmanned aerial vehicle.

Aircraft morphing is a concept that means a physical and smooth shape change that

makes a significant and desirable performance increment to the flight characteristics

of an aircraft. Nowadays in aircraft industry, there are thousands of aircrafts that fly

for different tasks. With the morphing wing technology, we will be able to adapt

aircrafts to perform more than one task. With that, aircrafts will be multi-tasking and

will be able to complete more than one task. By altering the camber of the airfoil,

changing the shape of the wing or the angle of the winglets, the aircrafts in the future

will not be mission specific, instead they will be able to adapt to any kind of mission

profiles. [4]

Morphing wings also bring structural advantages to the overall aircraft performance.

Morphing wings have unitized construction that makes their manufacturing easier.

This also diminishes complex assembly operations. As we have discussed so far, the

mentality of morphing wings is that they don’t have joints which cause extra weight

penalty. With morphing wing, this weight penalty goes away therefore additional

weight savings are provided. The hinges and joints also incur wear and lubrication

problems. Via this jointless wing structure, these kind of maintenance problems will

also be diminished. [5]

Now we will try to emphasize the benefit of morphing wing with an example. B-2

stealth bombing aircrafts have a long wing span that will allow to do long range

missions. But these bomber aircrafts are not used as fighter plane. Because they are

unable to make quick maneuvers and would be lumbering. On the other hand, F-18

Hornet have a relatively smaller wing span and they are used as fighter jets thanks to

their high maneuver capability. Most of the time, bomber aircrafts like B-2 are escorted

with a fighter plane such as F-18 in a mission. F-18’s high maneuver ability helps B-

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2 to complete its bombing mission without problem. As you can see, each aircraft only

has one mission profile and that makes extra aircraft for every mission and this

increases the cost of the mission. The main reason for this adaptive morphing aircraft

concept is to provide multi-tasking ability by changing the wing shape, chamber and

winglet angle during the mission. Today many engineers are researching, designing

and testing different wing profiles to obtain morphing wing. Most of the aircrafts today

use hinged structures to change the shape of the wing. [6]

Morphing wings can be designed to change into different wing configurations. Via

morphing wing, the wing can easily change its sweep angle, camber of the airfoil, wing

twist and winglet cant angle. Changing the camber of the wing will make the effect of

a flap. With that, induced drag effects will be minimized in cruise flight whereas

increasing the lift force at slow speeds will be provided. [6]

Figure 4: A morphing wing with different camber configurations [6]

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2. EFFECTS OF WINGLETS ON OVERAL WING PERFORMANCE AT

VARIABLE CANT ANGLES

In Essam Khalil’s et al. (2016) study, 3 different scenarios with 45o, 60o and 90o cant

angle winglets are conducted. NACA 2412 is used as airfoil. Taper ratio is 0.5 and

leading-edge sweep angle is 11o. The analyses are made under 0.2 Mach speed which

is quite convenient for our study because we are going to design morphing winglet

structure for unmanned aerial vehicles which operate about this speed.[7]

Figure 5: Winglet Cant Angle [8]

Figure 6: 90o,60o and 45o cant angle configurations [7]

CL analyses of above configurations are as follows:

Wings with winglets have higher CL values which are about 5-12%

At 0o angle of attack, 90o cant angle winglet has the highest CL

As AoA increases, 45o cant angle winglet has the highest CL value.

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WCSA-00-00 is the configuration for 90o cant angle.



Figure 7: Lift Coefficient (CL) For Different Winglet Cant Angles [7]

CD analyses of above configurations are as follows:

At 0o AoA there is not significant winglet effect since there is not much lift-

induced drag

At higher AoAs, 45o cant angle winglet gives about 1.5-2.5 % less CD in

comparison to without winglet configuration.

Figure 8: Drag Coefficient (CD) For Different Winglet Cant Angles [7]

CL/CD analyses of above configurations are as follows:

At all AoA’s from 0o to 12o, 45o cant angle gives the greatest CL/CD ratio which

is approximately 9-11 % more than w/o winglet configuration. 60o gives

8.5-10 %, 90o gives 3.5-6.4 % better performance.

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Figure 9: CL/CD Values For Different Winglet Cant Angles [7]

Figure 10: CL/CD Graph For Different Winglet Cant Angles [7]

During takeoff, wing without winglet creates great number of vortices from wingtips.

That is because of high AoA during takeoff (approximately 12o). When we use a

winglet at the cant angle of 45o during takeoff, the wing tip vortices will be greatly

reduced.

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Figure 11: Predicted particle path lines in case wing without winglet at angle of attack 12o [7]

Figure 12: Predicted particle pathlines, WCSA-45-00 at (a) AoA 0o, (b) AoA of 12o

[7]

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In Beechook and Wang’s study (2013), there performed a wind tunnel test and CFD

analysis for different variation of winglet cant angles. The tests were applied on a wing

without winglet, with horizontal winglet and 60-degree cant angle winglet. The results

proved that CD decreased by 25-30% and CL increased by 10-20 % using bird-like

multiple winglet configuration at 8-degree AoA. Their main goal with their study is to

show that aircraft performance will improve with using variable cant angle winglets

for different flight phases. They applied wind tunnel tests and run CFD analysis on

base wing (without winglet wing), 0o,30o,45o and 60o cant angle winglets. The tests

were run under the flow velocity of less than 0.2 Mach which is a convenient value for

our UAV winglet design. [9]

CL analysis of this study are as follow:

Among all configurations, 45o cant angle winglet gave the highest lift

coefficient.

As the AoA increases, the efficiency of winglets with cant angles increases.

Figure 13: CL values of winglet configurations with various AoAs [9]

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In brief, we can rank the winglet configurations for CL coefficients as follow (from

best to worst):

1. 45o cant angle (highest CL)

2. 30o cant angle

3. 60o cant angle

4. 0o cant angle

5. w/o winglet (lowest CL)

CD analysis of this study are as follow:

Among all configurations, 45o cant angle winglet gave the lowest drag

coefficient.

As the AoA increases, the efficiency of winglets with cant angles increases.

Figure 14: CD values of winglet configurations with various AoAs [9]

In brief, we can rank the winglet configurations for CD coefficients as follow (from

best to worst):

1. 45o cant angle (lowest CD)

2. 30o cant angle

3. 60o cant angle

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4. 0o cant angle

5. w/o winglet (highest CL)

L/D analysis of this study are as follow:

Winglet cant angle of 45o gave the highest L/D ratio

Figure 15: L/D values of winglet configurations with various AoAs [9]

In brief, we can rank the winglet configurations for L/D ratios as follow (from best to

worst):

1. 45o cant angle (highest L/D)

2. 30o cant angle

3. 60o cant angle

4. 0o cant angle

5. w/o winglet (lowest L/D)

As seen, base wing yielded the worst results in terms of CL, CD, and L/D values.

We can infer from the analysis that, winglet configuration will bring positive

aerodynamic efficiency and reduction on drag. As AoA increases, 45o winglet cant

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angle gave the best results. At lower AoAs, the significance of high winglet cant angle

will diminish.

3. WINGLET DESIGN

3.1 Preliminary Design

After making the literature search, we saw that, variable cant angle morphing

winglet will be efficient in terms of performance and fuel consumption reduction. We

decided to change our cant angle in 3 stages. There will be used servos to deflect the

winglet parts to have desired cant angles. 1st servo will be placed at the boundary where

the winglet will start. First servo will be connected to all 3 stages of winglet from

starting to edge via flexible wires. For instance, when a deflection angle of 15o is given

with first servo, all 3 parts of winglet will deflect simultaneously and the winglet will

get the cant angle of 15o degrees. Second servo will be placed 60 mm away from the

winglet root. This servo will deflect the remaining two parts. Last servo will be placed

into last 60 mm of the winglet and it will only deflect the part at the edge.

3.2 CAD Drawing of The Design

Since we are going to deal with a small model, we decided to have the chord

length of 15 cm for winglet extension. The winglet will be tapered in 3 stages as

mentioned in previous section. We selected the airfoil of “Lrn-1015 il” which has very

convenient values of CL, thickness and camberness for a UAV aircraft.[10]

To start CAD design, we firstly took the data of the airfoil via “airfoiltools.com”. The

coordinates are downloaded in an excel file. Then using macro, the coordinates of the

airfoil is exported from excel to CATIA software where we made our CAD design of

the winglet profiles.

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Figure 16: Coordinates exported from Excel to CATIA for Lrn-1015 il profile (150 mm chord)

After the data for the airfoil is taken, the drawing was made by CATIA.

Since we are going to analyze four different cases; 0o, 15o, 30o, 45o, we draw 4

different designs for analysis.

Figure 17: Winglet design having 0o cant angle (no winglet configuration)

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Figure 18: Winglet design having 15o cant angle


16


The pockets in design are left to put servos inside. Other pockets are made to

lighten the structure. The dimensions for servo pocket are arranged in accordance with

the dimensions of the servo. Servo pocket is placed at quarter chord where there is

aerodynamic center to avoid unwanted moments about y-direction.

Figure 21: Dimensions for servo pocket

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4. PLAUSIBLE SKIN MATERIAL CANDIDATES FOR MORPHING

STRUCTURES

4.1 Materials That Are Evaluated Through

In the last chapter we discussed morphing wing and different shape changes in

wings, profiles and winglets. It can be clearly seen that a rigid material like thin metal

material is not suitable for skin applications. The material that is to be used as a skin

for morphing structures in an airplane should be more flexible and elastic. While being

elastic and flexible, it should also have toughness and resistance against abrasion as

well. In a nutshell, candidate skin material should have following characteristics:

elastic, flexible, high recovery, resistant to different weather conditions, resistant to

abrasion and having enough hardness number to be able to handle aerodynamic loads

while flying. Also, the material should recover its original shape after being deformed

to a different wing shape. [6]

M.T. Kikuta (2003) made a study about mechanical properties of candidate

materials for morphing wings. In his study, he inspected 4 candidate materials that can

be used in morphing wings. Those are; polyurethane, copolyester, shape memory

materials (SMM) and woven materials. [6]

In the last few years, engineering advancements in polymer materials took a

different dimension. With the enhancements, those materials become more durable,

more flexible, more elastic and they have better recovery percentage. With these

enhancements in polymer area, reasonable kind of materials can be used as a skin for

morphing wings.

One of the materials that we will discuss is polyurethane. Polyurethane was

invented to be alternative of rubber. After some years of study, engineers obtained

different kinds of polyurethane materials for different kinds of applications.

Polyurethane is a synthetic material so you can obtain different hardness by changing

the formula. One of the main abilities that polyurethanes have is, having the ability to

resist against abrasion and to be able to withstand pressure loads. This material is

resistant against oils, solvents and greases. Since the aircraft will fly in different

weather conditions, the skin material should be able to withstand every kind of weather

incident. Polyurethane is resistant to sunlight, oxygen and ozone. The results of the

18

tests implemented on Tecoflex which is a type of medical thermoplastic polyurethane

is presented in this study. [6]

The other material that we will discuss is copolyester. In Kikuta’s study (2003),

there are two copolyster materials tested: Arnitel and Riteflex. These materials are

thermoplastics. According to the manufacturer of Riteflex; it combines the features of

thermoset elastomers and easy processing capability of plastics. Riteflex seems to meet

many specifications that we need for morphing wing skins. The other material for

copolyester is Arnitel. According to the manufacturer itself; this material has the

advantages of thermoplastics, easy to process with excellent mechanical properties

while having the flexibility of rubbers. Arnitel has the similar specifications of

polyurethanes in terms of; strength, abrasion resistance, chemical resistance and heat

resistance. [6]

Another material inspected for candidate skin material is shape memory

materials. According to Lui et. al. (2002), shape memory materials can be deformed

into a temporary shape under specific temperatures and stress. [11] And these materials

can restore their original stress-free shape with a thermal or electrical actuator. Shape

memory polymers have similar specifications with rubbers but their elasticity

is better above a critical temperature which is controlled by glass transition

temperature (TG). For shape memory polymers, the material changes its characteristic

in relation with the glass transition temperature. Below TG, the material acts like a rigid

solid. Beyond TG the material becomes in rubbery state. When the material is like

rubber, it can easily deform. But most of the shape memory polymers have flow state

where the material loses its ability to recover its original shape after heated extremely.

19

Figure 22: Graphic showing the elastic modulus change with respect to temperature [6]

The last material we will discuss is woven materials made out of fibers and

yarns. These types of materials have the ability to be elastic and recoverable for

morphing wing applications so they were inspected through. Two types of woven

materials were inspected: Spandura and Tru-Stretch.

In a brief to this chapter, requirements for materials that can be used as flexible

skin for morphing wing, are discussed. The materials that will be used for this purpose

must be; flexible, elastic, highly recoverable, having high hardness number, resistant

to weather conditions, chemicals and abrasion and be able to withstand aerodynamic

loads while flying.

4.2 Comparison of The Materials

In this part we are going to evaluate the results from the experiments implemented

on the materials. Kikuta experimented the materials uniaxially, biaxially and in terms

of pressure deflection and prepared a chart to put all results together. Table-1 shows

the results of uniaxial experiment. The results shown in Table-1 are maximum

experimental strain and worst recovery strain. Worst recovery strain is observed

because the availability of the material for morphing wing is closely related with it.

20

Table 1: Comparison of materials after uniaxially tested [6]

Table 2 shows the results of strain and hold. In this table; strain, maximum force,

dissipated force and recovery strains are presented.

21

Table 2: Comparison of materials for strain and hold experiment [6]

Now, when we check Table-1 we can see that Tecoflex 80A gives the highest

and Shape Memory Polymer (SMP) gives the 2nd highest amount of strain without

break. Also, we can clearly see that; maximum force needed to strain Tecoflex and

SMPs are the lowest, 15.91 lbs and 11.09 lbs, respectively. When we inspect the third

column for Table-1 that shows the recovery strain, we can see for both Tecoflex 80A

and SMP that they have two of the lowest recovery strain results. This shows that these

two materials recovered best. When we check Table-2, we can also see similar results

for Tecoflex 80A and SMPs which yield best results in terms of strain, maximum force

and recovery strain results.

In our study, we decided to use Tecoflex 80A which has flexibility, elasticity

and highly recoverability. Those are are the properties we seek for since we are

working on a morphing structure. After choosing the material, we will analyze if

Tecoflex 80A is an appropriate material under aerodynamic loads.

22

5. ANSYS ANALYSIS OF SELECTED MATERIAL FOR DIFFERENT

CANT ANGLES

After designing our model in CATIA V5, we imported our design as “.stp”

extension to ANSYS software. Our analysis completed in 2 steps. Firstly, we did the

CFD analysis to see the pressure distribution on our models. We used ‘Fluid Flow

(Fluent)’ tool on ANSYS Workbench. After distributing pressure on our models, we

analyzed the loads on our models due to pressure distribution via ‘Static Structural’

tool.

Figure 23: Project Schematic in ANSYS Workbench for the analysis

5.1 CFD Analysis (with Fluent)

Firstly, a flow volume is created in where the models are analyzed. The

geometry of the flow volume for 0o cant angle is as shown in Figure 24

23

Figure 24: Flow volume for the CFD analysis of 0o cant angle

Cross sectional area for the flow volume and the model is shown figure 25. Upper,

lower, side surfaces of the wall and airfoil surfaces are selected as wall. Flow comes

inside from inlet part, flows over the winglet airfoil and leaves from the outlet part.

Figure 25: Cross sectional are of flow volume for 0o cant angle

24

Figure 26: Named selection for winglet airfoil for 0o cant angle

After choosing the flow volume, inlet, outlet, walls and airfoil walls; the

volume is meshed to start the analysis. Mesh parameters created for surface mesh is as

shown in figure 27. Mesh minimum size is selected as 0,0002. This will be explained

later.

Figure 27: Surface mesh parameters for 0o cant angle

25

Figure 28: Surface mesh for 0o cant angle

Finite elements visualization for the flow volume is shown in figure 29. In

figure 30, boundary layer mesh is visualized. A 5-layer boundary mesh is made in

accordance with the minimum cell length of 0,0002 m.

Figure 29: Visualization of flow volume in finite elements for 0o cant angle

Figure 30: Boundary layer mesh for 0o cant angle

26

In figure 31 and 32, volume fill type of the mesh, minimum cell length, first

height of the boundary layer mesh, number of layers, boundary scope and other

parameters of mesh are shown.

Figure 31: Mesh volume fill and min. cell length of the mesh for 0o cant angle

Figure 32: Boundary scope and other parameters of mesh for 0o cant angle

27

Case for the fluid analysis is shown in figure 33. In the figure, flow volume

and velocity vectors can be observed.

Figure 33: Flow analysis case for 0o cant angle

In figure 34, viscous model selection is shown. For this analysis, k-epsilon

realizable, standard wall function model selected. This viscous model generally gives

good results for outer flow analysis, like this case. Before, it is told that minimum cell

length is selected as 0,0002 m. This calculation was made according to these formulas:

𝑅𝑒𝑥 =𝜌 𝑈∞ 𝐿

𝜇 𝐶𝑓 =

0,026

𝑅𝑒𝑥1/7 𝜏𝑤𝑎𝑙𝑙 =

𝐶𝑓 𝜌 𝑈∞2

2

𝑈𝑓𝑟𝑖𝑐 = √𝜏𝑤𝑎𝑙𝑙

𝜌 ∆𝑠=

𝑦+ 𝜇

𝑈𝑓𝑟𝑖𝑐 𝜌

28

y+ value [13] for k-epsilon realizable, standard wall function model is 30. Our

inlet velocity for the case is 50 m/s. Chord length is 150 mm. ρ = 1,225 kg/m3 and μ =

0.000018375 kg/m.s. When these values are put in formulas, we obtained ∆𝑠 = 0,0002

m.

Figure 34: Viscous model selections for 0o cant angle

Solution method is shown figure 34. To get more accurate results the analysis

was made using second order.

29

Figure 35: Solution methods for 0o cant angle

In figure 36, residual monitor is shown. All residual criteria are selected as 10-

5. This means that if the results in the iteration change 10-5 from the iteration before,

the iterations will stop. The other criteria for the analysis to stop is number of iterations.

If iteration number exceeds 500, the analysis will stop.

Figure 36: Residual monitors for 0o cant angle

30

Figure 37: Run calculation for 0o cant angle

Pressure distribution contours over airfoil are shown in figure 38. As can be

seen in the figure, upper surface of the airfoil has lower pressure distribution whereas

the lower surface has higher pressure distribution.

Figure 38: Pressure distribution over airfoil for 0o cant angle

31

Figure 39: Pressure distribution over winglet for 0o cant angle

In figure 40 and 41, velocity vectors in the flow volume can be observed.

Figure 40: Streamlines for 0o cant angle

32

Figure 41: Velocity vectors around airfoil for 0o cant angle

Figures were from the analysis of 00 cant angle winglet. The same analysis

were also made for; 150, 30o and 45o cant angle models.

For 15o cant angle winglet:


33



34


Figure 46: Flow analysis case for 15o cant angle

35



36



37

For 30o cant angle winglet:



38



39



40



41

For 45o cant angle:



42



43



44


5.2 FSI (Fluid-Structure Interaction) Analysis

After completing CFD analys is in Fluent, now the loads due to the pressure

distributions will be analyzed. In Fluent, all models (0, 15, 30 and 45 degree cant angle

winglets) were analyzed individually and some outputs were shown in the previous

chapter.

In order to see the static effects of the pressure distribution on our models,

‘Static Structural’ tool is used in ANSYS. All of the models meshed and analyzed

individually. Load input was imported from the Fluent analysis.

The results for 00 cant angle model are in Figure 66, 67, 68, 69, 70 and 71. In

figure 66 and 67, it can be seen that pressure distribution on the upper surface of the

airfoil is lower than the lower surface. Maximum pressure occurs in the middle of

leading edge and the trailing edge.

45

Figure 66: Pressure distribution on upper surface of airfoil for 0o cant angle

Figure 67: Pressure distribution on lower surface of airfoil for 0o cant angle

In figure 68 and 69, total deformation are shown for upper and lower surface

of the airfoil, respectively. As can be seen in the figures, maximum deflections occur

in the region where there is no spar to support. Maximum deflection occurs at the

bottom edge. That is because high pressure region is on the below surface of the airfoil.

It causes a force from lower surface to upper surface and it causes the maximum

deformation to be there.

46


Figure 69: Total deformation distribution on lower part of airfoil for 0o cant angle

In figure 70 and 71, equivalent stresses are shown. Maximum stress occurs

close to the trailing edge where the hollow part ends. There was no edge fillet in that

part so stresses increased there. That problem can be solved by putting an edge fillet

in the design.

47

Figure 70: Equivalent stress distribution on upper part of airfoil for 0o cant angle

Figure 71: Equivalent stress distribution on lower part of airfoil for 0o cant angle

The results for 150 cant angle model are in Figure 72, 73, 74, 75, 76 and 77


48




49



The results for 300 cant angle model are in Figure 78, 79, 80, 81, 82 and 83.


50




51



The results for 450 cant angle model are in Figure 84, 85, 86, 87, 88 and 89.


52




53



After the analysis are completed, we are going to control the results for each

cant angle and compare it with our material Tecoflex 80A and see if it is adequate for

cruise flight condition and different cant angles. Our material properties are as in

Table-4.

Mechanical Properties of Tecoflex 80A

Modulus of elasticity (Mpa) 1,07

Ultimate tensile strength (Mpa) 39,97

Ultimate tensile elongation % 660

Shear modulus (Mpa) 0,3057

Specific gravity 1,04

Table-3: Mechanical Properties of Tecoflex 80A

54

Maximum pressure, equivalent stress and maximum deformation values in 50

m/s are presented in Table-5.

@ 0o @ 15o @ 30o @ 45o

Max. Pressure

(Pa)

37,4493 37,3116 37,4043 37,8739

Max.

Deformation

(m)

1,7 E-3 1,783 E-8 2,11 E-8 2,38 E-8

Max. Stress

(kPa)

10,878 9,9561 9,5073 9,3285

Table-4: Analysis data for each cant angle configuration

As can be seen from the Table-5, 0o cant angle configuration has the maximum

deformation. The other configurations showed similar results.

When the results compared with the mechanical properties of the material, it

can be seen that in cruise flight condition, none of the configurations will fail. This

shows us that Tecoflex 80A is an appropriate material for all of the winglet

configurations.

55

6. RESULTS AND CONCLUSIONS

Morphing winglet application to aircraft wings will be very beneficial in terms of

range improvement and wing tip vortices reduction. But using a fixed winglet at the

tip of a wing is not an optimum solution for the best performance. Throughout our

study we see that, having the ability to change the cant angle of a winglet during the

flight will yield better results in terms of lift increment and drag decrement caused due

to wing tip vortices. With the design we proposed in this study, the cant angle of

winglet will have the ability to change and this will be a significant move for future

studies and applications.

An appropriate skin material application is also discussed in our study. Since our

morphing winglet will change its shape during flight, the skin cannot be rigid. The

coating material should be flexible enough to provide shape changes and should have

enough strength not to be torn apart. Some studies and experiments are inspected and

it is concluded that A type of polyurethane material Tecoflex 80 A can be a good option

for skin material selection.

After material selection, both flow analysis and structural analysis conducted

via ANSYS software. For CFD analysis Fluent; for FSI analysis Static Structural tools

are used. Pressure data obtained from the CFD analysis are imported in FSI analysis

to observe static loads on the proposed winglet designs. After the analysis, none of 4

configurations failed when Tecoflex 80A material is used. It is concluded that Tecoflex

80A can be an appropriate material for the winglet design proposed in this study.

56

7. REFERENCES

[1] Url-1 < https://forums.x-plane.org/index.php?/forums/topic/215357-winglet-tip-

texture-issue/

[2] Guerrero J., Sanguineti M., Wittkowski K. (2018). CFD Study of the Impact of

Variable Cant Angle Winglets on Total Drag Reduction, DICCA, Università degli

Studi di Genova, Via Montallegro , 3 December 2018

[3] Bae S., Seigler M., Inman J. (2005). Aerodynamic and Static Aeroelastic

Characteristics of a Variable-Span Morphing Wing, Journal of Aircraft Vol. 42, No. 2,

March–April 2005

[4] Arrison L. et. al. (2003). 2002-2003 AE/ME Morphing Wing Design, Virginia

Tech Aerospace Engineering Senior Design Project Spring Semester Final Report

1 May, 2003

[5] Kota S. et. al. (2003). Design and Application of Compliant Mechanisms For

Morphing Aircraft Structures, Smart Structures and Materials 2003: Industrial and

Commercial Applications of Smart Structures Technologies, Eric H. Anderson, Editor,

Proceedings of SPIE Vol. 5054 (2003)

[6] Kikuta T. (2003). Mechanical Properties of Candidate Materials for Morphing

Wings, Thesis Submitted to the Faculty of the Virginia Polytechnic Institute and State

University in partial fulfilment of the requirements for the degree of Master of Science

[7] Khalil E. et. al. (2016). Air Craft Winglet Design and Performance: Cant Angle

Effect, Journal of Robotics and Mechanical Engineering Research Vol: 1, Issue: 3

[8] Johansen J., Sorensen N. (2007). Numerical Analysis of Winglets on Wind

Turbine Blades using CFD, January 2007

[9] Beechook A., Wang J. (2013). Aerodynamic Analysis of Variable Cant Angle

Winglets for Improved Aircraft Performance, Proceedings of the 19th International

Conference on Automation & Computing, Brunel University, London, UK, 13-14

September 2013

[10] Url-2 > http://airfoiltools.com/

[11] Lui, C. et al, (2002). Tailored Shape Memory Polymers: Not all SMPs are Created

Equal. Proceedings of The First World Congress on Biomimetics, December 9-11,

Albuquerque, New Mexico, pp. 1-7.

57

[12] Sun J. et. al. (2014). Mechanical Properties of Shape Memory Polymer

Composites Enhanced by Elastic Fibers and Their Application in Variable Stiffness

Morphing Skins, Journal of Intelligent Material Systems and Structures Aug 26, 2014

[13] Url-3 > https://www.pointwise.com/yplus/

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