Modeling and analysis of composite hinge for aircraft spoiler using finite element analysis.

Master of Science in Mechanical Engineering

Faculty of Mechanical Engineering

MODELING AND ANALYSIS OF COMPOSITE HINGE FOR

AIRCRAFT SPOILER USING FINITE ELEMENT ANALYSIS

Wai Chee Mun

MODELING AND ANALYSIS OF COMPOSITE HINGE FOR AIRCRAFT SPOILER USING FINITE ELEMENT ANALYSIS

WAI CHEE MUN

A thesis submitted

in fulfillment of the requirements for the degree of Master of Science in Mechanical Engineering

Faculty of Mechanical Engineering

UNIVERSITI TEKNIKAL MALAYSIA MELAKA

DECLARATION

I declare that this thesis entitle “Modeling and Analysis of Composite Hinge for Aircraft Spoiler using Finite Element Analysis” is the result of my own research except as cited in the references. The thesis has not been accepted for any degree and is not concurrently submitted in candidature of any other degree.

Signature : ... Name : WAI CHEE MUN Date : ...

APPROVAL

I hereby declare that I have read this dissertation/report and in my opinion this dissertation/report is sufficient in terms of scope and quality as a partial fulfillment of Master of Science in Mechanical Engineering.

Signature : ………... Supervisor Name : AHMAD RIVAI

DEDICATION

ABSTRACT

This thesis is concerned with the modeling and analysis of a composite hinge bracket for Airbus A320 aircraft using Finite Element Method. The increasing demand of composite materials in aircraft structural design is due to the high strength and stiffness to weight ratio of the composite materials. The reduction in structural weight by the composite materials reduces the fuel consumption of the aircraft. Since the hinge bracket of the A320 aircraft is made of metallic materials, this research aims to redesign the hinge bracket with laminated composite plates. However, the design of structures with composites is not as easy as the design of metallic structures. Composites are heterogeneous and anisotropic in nature thus, complicating the structural analysis. The best practice of laminated composite design in the industry is simplified to accommodate the limitation of this research. The simplified design process is a simple three stages design process which focused on using Finite Element Anlysis for unidirectional laminated composite structural design. It begins with the determination of stress distribution in a metallic structure. The stress distribution allows the designer to identify areas of different stress distribution of the structure. Once the areas are identified, the direction of principle stresses is determine to decide on the fiber orientation and stacking sequence of the composite laminates for each area. Finally, the analysis of the laminated composite structure is carried out using Classical Laminate Plate theory and the structural integrity is accessed by the calculation of margin of safety. The simplified design process has been applied to redesign the metallic hinge bracket of A320 aircraft using IM7/8552 composite laminates. MSC Patran/Nastran and Abaqus CAE were used to carry out the finite element analysis of the hinge bracket. The composite hinge bracket has a margin of safety of 0.041. The weight savings by this composite hinge bracket is estimated to be 42.29 percent of the original metallic hinge bracket of A320. The success of this composite hinge design has proven that the simplified design process proposed in this research is indeed feasible for the purpose of preliminary design of laminated composite structures. The advantage of this simplified design process is that it approaches the design of composite structures in a systematic manner thus, allowing the identification of factors influencing the strength the laminated composite structural design. Since it is based on Finite Element Method, the results are at best approximations but sufficient to boost the level of confidence in the preliminary design stage.

ABSTRAK

Tesis ini adalah berkenaan dengan pemodelan dan analisis pendakap engsel komposit untuk pesawat Airbus A320 dengan menggunakan Kaedah Unsur Terhingga. Permintaan yang semakin meningkat daripada bahan komposit dalam reka bentuk struktur pesawat adalah disebabkan nisbah kekuatan dan kekakuan kepada berat bahan komposit yang tinggi. Pengurangan berat struktur oleh bahan-bahan komposit mampu mengurangkan penggunaan bahan api pesawat. Penyelidikan ini bertujuan untuk mereka bentuk semula pendakap engsel pesawat A320 dengan menggunakan plat komposit berlapis untuk menggantikan pendakap engsel lama pesawat A320 yang diperbuat daripada bahan logam. Walau bagaimanapun, reka bentuk struktur komposit tidak semudah reka bentuk struktur logam. Sifat bahan komposit yang anisotropik merumitkan analisis struktur. Amalan terbaik reka bentuk komposit berlapis dalam industri dipermudahkan untuk menampung had kajian ini. Proses reka bentuk yang dipermudahkan terdiri daripade tiga peringkat asas yang memberi tumpuan kepada menggunakan Analysis Unsur Terhingga untuk mereka bentuk struktur daripada plat komposit berlapis satu arah.. Ia bermula dengan penentuan agihan tegasan dalam struktur logam. Agihan tegasan membolehkan pereka mengenal pasti kawasan agihan tegasan yang berbeza dalam struktur. Sebaik sahaja kawasan dikenal pasti, arah tegasan prinsip digunakan untuk menentukan orientasi gentian dan susunan urutans lapisan komposit bagi setiap kawasan. Akhir sekali, analisis struktur komposit berlapis dijalankan dengan menggunakan teori Klasik lamina Plat dan integriti struktur diukur melalui pengiraan margin keselamatan. Proses reka bentuk yang dipermudahkan telah digunakan untuk mereka bentuk semula pendakap engsel logam pesawat A320 dengan menggunakan IM7/8552 komposit berlapis. MSC Patran/Nastran dan Abaqus CAE telah digunakan untuk menjalankan analisis unsur terhingga pendakap engsel tersebut. Pendakap engsel komposit mempunyai margin keselamatan 0.041. Penjimatan berat badan oleh pendakap engsel komposit ini dianggarkan 42.29 peratus daripada berat badan asal pendakap engsel logan pesawat A320. Kejayaan reka bentuk engsel komposit ini telah membuktikan bahawa proses reka bentuk yang mudah yang dicadangkan dalam kajian ini memang boleh digunakan untuk tujuan reka bentuk awal struktur komposit berlapis. Kelebihan model ini ialah ia membolehkan pengenalpastian faktor yang mempengaruhi kekuatan reka bentuk struktur komposit berlapis dalam proses reka bentuk struktur. Memandangkan ia adalah satu kaedah berdasarkan Kaedah Unsur Terhingga, keputusan analysis merupakan satu anggaran sahaja tetapi cukup untuk meningkatkan tahap keyakinan dalam reka bentuk dan pembuatan komposit.

ACKNOWLEDGEMENTS

I would like to thank Centre for Research and Innovation Management (CRIM) and Faculty of Mechanical Engineering (FKM) Universiti Teknikal Malaysia Melaka for funding this research through PJP/2012/FKM(18C)S1112 Research Fund and having interest in this area of engineering. I hope that the contents of this thesis can benefit them in designing and developing laminated composite structures in the future.

I would like to extend my appreciation to my supervisors, Associate Professor Ahmad Rivai and Mr Omar Bapokutty, for their suggestions and guidance. Thank you, Mr. Omar, for willing to join my research at a late notice.

I would also like to thank my parents, family members, friends who were involved directly or indirectly in the completion of my thesis. Thank you for your morale support throughout my graduate studies.

Last but not least, I would like to thank God for answering my prayers in my darkest time. The completion of this thesis would not been possible without God’s will. I humbly thank you for guiding and instilling faith in me.

TABLE OF CONTENTS

PAGE DECLARATION

DEDICATION

ABSTRACT i

ABSTRAK ii

ACKNOWLEDGEMENTS iii

TABLE OF CONTENTS iv

LIST OF TABLES vi

LIST OF FIGURES viii

LIST OF APPENDIXES xi

LIST OF ABBREVIATIONS xii

LIST OF SYMBOLS xiv

LIST OF PUBLICATIONS xvi

CHAPTER

1 INTRODUCTION 1

1.1 Background 3

1.2 Problem Statement 4

1.3 Objectives 4

1.4 Scope 5

1.5 Overview of Thesis

2 LITERATURE REVIEW 6

2.1 Structural Design 6

2.1.1 The Design Process 7

2.1.2 Stress Analysis 9

2.1.3 Factor and Margin of Safety 11

2.1.4 Accuracy 13

2.1.5 Materials Selection 14

2.2 Composite 15

2.2.1 Composite in Aircraft Design 16

2.2.2 Laminated Composite 18

2.2.3 Mechanics of Laminated Composite 20

2.2.4 Failure Criteria of Laminated Composite 26

2.2.5 Composite Design Method 28

2.3 Finite Element Method and Analysis 29

2.3.1 Finite Element Procedures in Engineering Analysis 31 2.3.1.1 Verification and Validation of Simulation Models 33 2.3.2 Software Packages for Finite Element Analysis 36

2.3.2.1 MSC Patran & Nastran 37

2.3.2.2 Abaqus CAE 38

2.4 Chapter 2 Overview 39

3 METHODOLOGY 40

3.1 Working Method for Composite Design 40

3.2 Preliminary Analysis 42

3.3 Discretization 44

3.4 Composite Laminate Modeling 45

4 MODELING AND VERIFICATION 46

4.1 3D Hinge Bracket Modeling 46

4.2 2D Hinge Bracket Modeling 49

4.3 Verification of Modeling 52

5 RESULTS AND ANALYSIS 56

5.1 Preliminary Analysis of Metallic Hinge Bracket 56

5.1.1 Stress Analysis 57

5.2 Discretization of Composite Hinge Bracket 63

5.3 Original Composite Hinge Bracket Analysis 66

5.4 Modified Composite Hinge Bracket 75

5.5 Brief Economy Analysis 81

6 CONCLUSION AND RECOMMENDATIONS 82

6.1 Conclusion 82

6.2 Recommendations 84

REFERENCES 86

APPENDIXES 99

LIST OF TABLES

TABLE TITLE PAGE

4.1 Mechanical Properties of AA7075-T651 49

4.2 Mechanical Properties of IM7/8552 Lamina 51

4.3 Von Mises Stresses of W460x74 Beam 53

4.4 Deflection of W460x74 Beam at Different Beam Length 53

4.5 Elastic Properties of Each Ply 54

4.6 Ply by Ply Stress Results of Quasi-isotropic Laminate 55

5.1

Summary of Load Cases Considered in FEA for Spoiler Hinge Bracket

56

5.2

Comparison of Von Mises Stresses at the Top Surface of the Right Side of the Metallic Hinge Bracket

61

5.3

Composite Laminate Layup for Each Zone of A320 Composite Hinge Bracket

64

5.4

Summary of Worst MS for Each Zone of the Composite Hinge Bracket

68

5.5

Layer by Layer Longitudinal Stress of Zone A for Case of Spoiler Close withNegative Wing Bending

70

5.6

Layer by Layer Longitudinal Stress of Zone A for Case of Spoiler Open withPositive Wing Bending

70

5.7 Displacement Analysis of Zone A at Node 1 74

5.8

Composite Laminate Layup of the Modified Composite Hinge Bracket

77

5.9

Summary of Worst MS by Zones for the Modified Composite Hinge Bracket

79

5.10

Comparison of Worst MS at Each Zone between Original and ModifiedComposite Hinge Bracket

80

5.11

Comparison of Weight Savings between Different Hinge Bracket Designs

81

LIST OF FIGURES

FIGURE TITLE PAGE

1.1 Rolling Motion Cause by Deploying Right Spoiler 1

1.2 FACC Carbon Composite Center Hinge Fitting 2

1.3 Composite Structural Weight Development in Airbus 3

2.1 Design Model 7

2.2 Fiber Arrangement Patterns in a Lamina 18

2.3 Composite Laminate Drawing 19

2.4 Sign Convention of Positive and Negative Fiber Orientation 20

2.5 Finite Element Method Model 31

2.6 The Process of Finite Element Analysis 32

2.7 Verification and Validation Activities and Outcomes 34 3.1 Schematic Picture of Composite Working Method 40

3.2 Aero Shaped Segment 41

3.3 Unidirectional Laminated Composite Design Model 42 3.4 CAD Drawing of Airbus A320 Outer Hinge Bracket 43 3.5 Discretized Zones in Airbus A320 Hinge Bracket 44 4.1 CAD Model of A320 Hinge Bracket for 3D FE Modeling 46 4.2 Finite Element Modeling of 3D Hinge Bracket 47 4.3 CAD Model of A320 Hinge Bracket for 2D FE Modeling 49

4.4 Finite Element Modeling of 2D Hinge Bracket 50

4.5 W460x74 Beam 52

4.6 Composite Laminate Setup for Tension Coupon Test 54

5.1

Stress Distribution of Metallic Hinge Bracket for Case of Spoiler Close

56

5.2

Stress Distribution of Metallic Hinge Bracket for Case of Spoiler Open

56

5.3 Stress Concentration at Bolt Holes of Metallic Hinge Bracket 58

5.4

Stress Distribution for Metallic Hinge Bracket for Case of Spoiler Close

59

5.5

Comparison of Stress Concentration at Lug of the Metallic Hinge Bracket

60

5.6

Nodes Location at the Top Surface of the Right Side of the Hinge Bracket

61

5.7

Comparison of Stresses Obtained by Nastran and Abaqus at the Top Surface of the Sides of the Metallic Hinge Bracket

62

5.8

Comparison of Stresses between Six Different Load Cases at the Top Surface of the Sides of the Metallic Hinge Bracket

62

5.9

Discretization of A320 Hinge Bracket by Areas of Different Thickness

64

5.10

Discretization of 2D FE Model of A320 Hinge Bracket by Areas of Different Stress Distribution and Laminate Thickness

64

5.11 Difference in Global Axis and Local Material Axis in Zone A 66

5.12

Comparison of Von Mises Stresses (MPa) between Metallic Hinge Bracket and Composite Hinge Bracket for the Case of Spoiler Close

67

With negative Wing Bending

5.13

Comparison of Von Mises Stresses (MPa) between Metallic Hinge Bracket and Composite Hinge Bracket for the Case of Spoiler Open With Positive Wing Bending

67

5.14 Comparison between 3D and 2D FE Model of A320 Hinge Bracket 69

5.15 Laminate Stacking Sequence of Zone A 69

5.16 Left Side View of Zone A of the Composite Hinge Bracket 70

5.17

Average Longitudinal Stress in MPa of the Composite Laminate in Zone A

72

5.18

Layer by Layer Analysis of Longitudinal Stress of Zone A of the Composite Hinge Bracket for the Case of Spoiler Close with Negative Wing Bending

72

5.19

Layer by Layer Analysis of Longitudinal Stress of Zone A of the CompositeHinge Bracket for the Case of Spoiler Open with Positive Wing Bending

73

5.20 Deformation of Composite Laminate in Zone A 74

5.21 Composite Hinge Bracket 75

5.22 Discretization of Modified Composite Hinge Bracket 76

5.23

Comparison of Von Mises Stresses between Original and Modified Composite Hinge Bracket for the Case of Spoiler Close with Negative Wing Bending

78

5.24

Comparison of Von Mises stresses between Original and Modified CompositeHinge Bracket for the Case of Spoiler Open with Positive Wing Bending

78

LIST OF APPENDIXES

APPENDIX TITLE PAGE

A Airbus A320 Original Hinge Bracket Drawing 99

B Airbus A320 Hinge Loads 100

C Selected Properties of Matrix and Fibrous Materials 102 D Hexcel Datasheet for HexPly 8552 Epoxy Matrix 103 E Calculation of IM7/8552 Lamina Properties 104 F Weight Calculation of Airbus A320 Hinge Bracket 106

G Mesh Refinement 110

H Calculation of Principle Stresses 113

LIST OF ABBREVIATIONS

1D - One dimensional 2D - Two dimensional 3D - Three dimensional

ASM - American Society for Metals

ASME - American Society of Mechanical Engineers CAD - Computer Aided Design

CAM - Cylindrical Assemblage Model CLT - Classical Laminate Theory cm3 - Centimeter cube

DoFs - Degree of Freedoms

DTA - Damage Tolerance Analysis etc. - et cetera

FAA - Federal Aviation Administration

FACC - Fischer Advanced Composite Components FE - Finite Element

FEA - Finite Element Analysis FEM - Finite Element Method FPF - First ply failure

FRP - Fiber Reinforced Polymer FS - Factor of Safety

g - Gram

g/cm3 - Gram per centimeter cube GUI - Graphic-user interface IBC - International Building Code

in - Inch

lb/in - Pounds per square inch mm - millimeters

mm2 - millimeters square mm3 - millimeters cube MPa - MegaPascal MS - Margin of Safety

Msi - Megapounds per square inch

N - Newton

Nm - Newton meter

NASA - The National Aeronautics and Space Administration PLM - Patran Laminate Modeler

PPF - Progressive ply failure psi - Pounds per square inch RBE - Rigid Body Elements RF - Reserved Factor

RTM - Resin transfer moulding UD - Unidirectional

VAC - Volvo Aero Corporation V&V - Verification and validation

VV&A - Verification, validation and accreditation

LIST OF SYMBOLS

% - Percent

° - Degree

cm3 - Centimeter cube

g/cm3 - Gram per centimeter cube

Ɵ - Angle

Vf - Fiber volume fraction Vm - Matrix volume fraction Wf - Fiber weight fraction Wm - Matrix weight fraction

Ρ - Density

ρc - Lamina density ρf - Fiber density ρm - Matrix density

E1 - Lamina longitudinal modulus Ef - Fiber elastic modulus

Em - Matrix elastic modulus v12 - In-plane Poisson’s ratio vf - Fiber Poisson’s ratio vm - Matrix Poisson’s ratio E2 - Lamina transverse modulus

ζ - Empirical parameter

G12 - Lamina in-plane shear modulus Gf - Fiber shear modulus

Gm - Matrix shear modulus.

G23 - Lamina out of plane shear modulus

σ1 - Maximum local stress along fiber direction

σ2 - Maximum local stress transverse to the fiber direction τ12 - Maximum shear stress in the principal material direction Fft - Fiber tensile strength

F1c - Lamina longitudinal compressive strength F2t - Lamina transverse tensile strength

Fmt - Matrix tensile strength

F2c - Lamina transverse compressive strength Fmc - Matrix compressive strength

F6 - Lamina in-plane shear strength Fms - Matrix shear strength

ασ - Standard deviation of fiber misalignment Vv - Void volume fraction

П - A constant with the value of 3.142

LIST OF PUBLICATIONS

Journal:

Putra, A., Rivai, A., Wai, C.M., and Muhammad, N., 2014. Preliminary Investigation on the Perforated Panel Mobility using Finite Element Method. Applied Mechanics and Materials, 471, pp.341 – 346.

Wai, C.M., Rivai, A., and Omar, B., 2013. Modeling Optimization Involving Different Types of Elements in Finite Element Analysis. IOP Conference Series: Materials Science and Engineering, 50 (1), pp. 1 – 8.

Wai, C.M., Rivai, A., and Omar, B., 2013. Effects of Elements on Linear Elastic Stress Analysis: A Finite Element Approach. International Journal of Research in Engineering and Technology (IJRET), 10 (2), pp.561 – 567.

Wai, C.M., Ahmad, R., and Omar, B., 2013. Design and Analysis of an Aircraft Composite Hinge Bracket using Finite Element Approach. Applied Mechanics and Materials (Accepted manuscript).

Putra, A., Cheah, Y.M., Muhammad, N., Rivai, A., and Wai, C.M., 2014. The Effect of Perforation on the Dynamics of a Flexible Panel. Advances in Acoustic and Vibration (Accepted manuscript).

CHAPTER 1

1 INTRODUCTION

1.1 Background

Commercial aircrafts, such as Airbus A380 and Boeing E787, have spoilers attached to the top surface of their wings. Spoilers are plate like structures used to assist the landing and descending of an aircraft from higher to lower altitude as well as generate rolling motion of the aircraft (Dawson, 2006; NASA, 2010). Figure 1.1 shows the rolling motion generated by deploying the right spoiler of an aircraft.

Figure 1.1 : Rolling Motion Cause by Deploying Right Spoiler (NASA, 2010)

As the name suggest, aircraft spoilers are used to “spoil” the airflow over the wing. When both the spoilers are deployed, the drag increases while the lift decreases (Dawson, 2006). The decrease in the lift enables the aircraft to descend from a higher to lower altitude and slow down before landing. Spoilers are also useful in assisting the breaking of the aircraft on the runway (NASA, 2010).

The upward and downward movements of the aircraft spoilers are made possible by a set of hinges. A hinge is a mechanism which allows the rotational motion in one axis (Leet et al., 2008). Hinges usually consists of two brackets. Some hinges have bearings fitting to reduce frictional force between the pin and two brackets. The brackets of the hinges are supportive structures which are usually made of metals. In 2006, Fischer Advanced Composite Components (FACC) AG successfully designed a composite center hinge fitting for Airbus, shown in Figure 1.2, by using resin transfer molding (RTM). The composite hinge is said to be able to withstand 20 ton air load and save 25 percent in weight per set of hinge (Dawson, 2006).

Figure 1.2 : FACC Carbon Composite Center Hinge Fitting (Dawson, 2006): (a) Metal Bushings and Bearings Installed in the Fitting Lugs; (b) Hinge Fitting Molded via RTM

In aircraft construction, composites are more commonly used for secondary structure, wings, fuselages, and even smaller parts such as spoilers (Dawson, 2006). However, the use of composite materials in commercial aircraft has been increasing exponentially since the last decade due to their capability in reducing the overall structural weight of the aircraft effectively (Basavaraju, 2005; Mallick, 2007). Companies such as Airbus and Boeing have been replacing their metallic aircraft parts with composite materials. Figure 1.3 shows the composite structural weight development in Airbus commercial aircrafts. Though the development of composites has mostly been derived

from military and The National Aeronautics and Space Administration (NASA) prototypes (Cole, 2002), it is not too farfetch to say that there will be an all composite aircraft one day.

Figure 1.3 : Composite Structural Weight Development in Airbus (Bold, 2007)

1.2 Problem Statement

Existing hinges for aircraft spoilers, especially the old aircraft models, are mostly made of metals. For example, the Airbus A320 aircraft spoiler hinges are still made of metal. The replacement of the existing metallic hinges with composite materials may save up quite a considerable amount of the hinges structural weight. The decrease in the aircraft structural weight will reduce the fuel consumption of the aircraft. However, composite is still considered new compared to the well established monolithic metals (Maligno, 2007; Nguyen, 2010). Therefore, more information and a different approach to design composite structure are required (Mallick, 2007). The industry has developed their own best practices to design composite structures but the educational model flow to design composite structures, especially laminated composite structures, is still lacking in the engineering textbooks.

LIST OF SYMBOLS

% - Percent

° - Degree

cm3 - Centimeter cube

g/cm3 - Gram per centimeter cube

Ɵ - Angle

Vf - Fiber volume fraction Vm - Matrix volume fraction Wf - Fiber weight fraction Wm - Matrix weight fraction

Ρ - Density

ρc - Lamina density

ρf - Fiber density

ρm - Matrix density

E1 - Lamina longitudinal modulus Ef - Fiber elastic modulus

Em - Matrix elastic modulus v12 - In-plane Poisson’s ratio vf - Fiber Poisson’s ratio vm - Matrix Poisson’s ratio E2 - Lamina transverse modulus

ζ - Empirical parameter

G12 - Lamina in-plane shear modulus Gf - Fiber shear modulus

Gm - Matrix shear modulus.

G23 - Lamina out of plane shear modulus

σ1 - Maximum local stress along fiber direction

σ2 - Maximum local stress transverse to the fiber direction

τ12 - Maximum shear stress in the principal material direction Fft - Fiber tensile strength

F1c - Lamina longitudinal compressive strength F2t - Lamina transverse tensile strength

Fmt - Matrix tensile strength

F2c - Lamina transverse compressive strength Fmc - Matrix compressive strength

F6 - Lamina in-plane shear strength Fms - Matrix shear strength

ασ - Standard deviation of fiber misalignment Vv - Void volume fraction

П - A constant with the value of 3.142

LIST OF PUBLICATIONS

Journal:

Putra, A., Rivai, A., Wai, C.M., and Muhammad, N., 2014. Preliminary Investigation on the Perforated Panel Mobility using Finite Element Method. Applied Mechanics and Materials, 471, pp.341 – 346.

Wai, C.M., Rivai, A., and Omar, B., 2013. Modeling Optimization Involving Different Types of Elements in Finite Element Analysis. IOP Conference Series: Materials Science and Engineering, 50 (1), pp. 1 – 8.

Wai, C.M., Rivai, A., and Omar, B., 2013. Effects of Elements on Linear Elastic Stress Analysis: A Finite Element Approach. International Journal of Research in Engineering and Technology (IJRET), 10 (2), pp.561 – 567.

Wai, C.M., Ahmad, R., and Omar, B., 2013. Design and Analysis of an Aircraft Composite Hinge Bracket using Finite Element Approach. Applied Mechanics and Materials (Accepted manuscript).

Putra, A., Cheah, Y.M., Muhammad, N., Rivai, A., and Wai, C.M., 2014. The Effect of Perforation on the Dynamics of a Flexible Panel. Advances in Acoustic and Vibration (Accepted manuscript).

CHAPTER 1

1 INTRODUCTION

1.1 Background

Commercial aircrafts, such as Airbus A380 and Boeing E787, have spoilers attached to the top surface of their wings. Spoilers are plate like structures used to assist the landing and descending of an aircraft from higher to lower altitude as well as generate rolling motion of the aircraft (Dawson, 2006; NASA, 2010). Figure 1.1 shows the rolling motion generated by deploying the right spoiler of an aircraft.

Figure 1.1 : Rolling Motion Cause by Deploying Right Spoiler (NASA, 2010)

As the name suggest, aircraft spoilers are used to “spoil” the airflow over the wing. When both the spoilers are deployed, the drag increases while the lift decreases (Dawson, 2006). The decrease in the lift enables the aircraft to descend from a higher to lower altitude and slow down before landing. Spoilers are also useful in assisting the breaking of the aircraft on the runway (NASA, 2010).

The upward and downward movements of the aircraft spoilers are made possible by a set of hinges. A hinge is a mechanism which allows the rotational motion in one axis (Leet et al., 2008). Hinges usually consists of two brackets. Some hinges have bearings fitting to reduce frictional force between the pin and two brackets. The brackets of the hinges are supportive structures which are usually made of metals. In 2006, Fischer Advanced Composite Components (FACC) AG successfully designed a composite center hinge fitting for Airbus, shown in Figure 1.2, by using resin transfer molding (RTM). The composite hinge is said to be able to withstand 20 ton air load and save 25 percent in weight per set of hinge (Dawson, 2006).

Figure 1.2 : FACC Carbon Composite Center Hinge Fitting (Dawson, 2006): (a) Metal Bushings and Bearings Installed in the Fitting Lugs; (b) Hinge Fitting Molded via RTM

In aircraft construction, composites are more commonly used for secondary structure, wings, fuselages, and even smaller parts such as spoilers (Dawson, 2006). However, the use of composite materials in commercial aircraft has been increasing exponentially since the last decade due to their capability in reducing the overall structural weight of the aircraft effectively (Basavaraju, 2005; Mallick, 2007). Companies such as Airbus and Boeing have been replacing their metallic aircraft parts with composite materials. Figure 1.3 shows the composite structural weight development in Airbus commercial aircrafts. Though the development of composites has mostly been derived

from military and The National Aeronautics and Space Administration (NASA) prototypes (Cole, 2002), it is not too farfetch to say that there will be an all composite aircraft one day.

Figure 1.3 : Composite Structural Weight Development in Airbus (Bold, 2007)

1.2 Problem Statement

Existing hinges for aircraft spoilers, especially the old aircraft models, are mostly made of metals. For example, the Airbus A320 aircraft spoiler hinges are still made of metal. The replacement of the existing metallic hinges with composite materials may save up quite a considerable amount of the hinges structural weight. The decrease in the aircraft structural weight will reduce the fuel consumption of the aircraft. However, composite is still considered new compared to the well established monolithic metals (Maligno, 2007; Nguyen, 2010). Therefore, more information and a different approach to design composite structure are required (Mallick, 2007). The industry has developed their own best practices to design composite structures but the educational model flow to design composite structures, especially laminated composite structures, is still lacking in the engineering textbooks.