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School of Aerospace, Mechanical and Manufacturing Engineering (SAMME) AERO 2362/AERO 2366 Critical Design Report Aerospace Design Project - Fixed Wing Aircraft Alternative Fuels and Environmentally Friendly Aircraft System Group C 14 October 2010

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School of Aerospace, Mechanical and Manufacturing Engineering (SAMME)AERO 2362/AERO 2366 – Critical Design ReportAerospace Design Project - Fixed Wing AircraftAlternative Fuels and Environmentally Friendly Aircraft SystemGroup C14 October 2010Alternative Fuels and Environmentally Friendly Aircraft SystemAERO2362/AERO2366BASE ELEMENTSProgram Code: Project Name: 2010 Design Project – AERO2362/AERO2366 Fixed Wing Aircraft Design - Alternative Fuels and Environmental Friendly Aircra

Transcript of Alternative Fuels and Environmentally Friendly Aircraft System

School of Aerospace, Mechanical and

Manufacturing Engineering (SAMME)

AERO 2362/AERO 2366 – Critical Design Report

Aerospace Design Project - Fixed Wing Aircraft

Alternative Fuels and Environmentally

Friendly Aircraft System

Group C

14 October 2010

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B A S E E L E M E N T S

Program Code: 2010 Design Project – AERO2362/AERO2366 Project Name: Fixed Wing Aircraft Design - Alternative Fuels and Environmental

Friendly Aircraft System Project Advisor: Robert Danaher - [email protected] Team Members:

Jed Guinto (Team Leader) [email protected]

Leon Ballis (Deputy-leader) [email protected]

Nigel So (Deputy-leader) [email protected]

Nathaniel Rodda (Deputy-leader) [email protected]

Wai Hung Frederick Chan [email protected]

Terry Cheong Yip [email protected]

Kin Hei Wu [email protected]

Dae Yoon Lee [email protected]

Shanaka Ranmal Jayasekara [email protected]

Ibrahim Al Najjar [email protected]

Babak Shokoohi [email protected]

Vaigunthan Markkandu [email protected]

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E X E C U T I V E S U M M A R Y

“It is not the strongest of the species that survives, nor the most intelligent that survives. It is the one that is the most adaptable to change.” (Charles Darwin)

In 2007 jet fuel accounted for approximately 26.5% of total operating costs, the single largest expense for airlines. The need for a cheaper, more affordable fuel is of major concern for the aerospace industry. As fossil fuel prices and air travel increase, airlines are required to increase the airfares in order to keep up and maintain revenue to cover operating costs. Based on current trends, ticket prices will become out of reach for many consumers. With the global climate steadily increasing, airlines are being pressured into being more responsible with the level of emissions they contribute to the environment. One of the goals that have been set by “National Aeronautics Research and Development Challenges, Goals and Objectives” is to “protect the environment while sustaining growth in air transportation1” which is exactly what this report has accomplished.

The aerospace industry has seen its share of downfalls due to emission issues. Environmental, social and political parties are now placing limits for acceptable levels of emissions for aircrafts and airports. So in order for the entire aerospace industry to continue, they must adapt by designing aircrafts and procedures which can meet and exceed these limits. The projected aircraft is an A320 and B737 replacement aircraft. It is cleaner, smarter, safer, cheaper and most importantly, more environmentally friendly than any other aircraft in its class.

The projected aircraft utilizes alternate fuels known as bio-jet fuel which is much cleaner than the regular Jet-A aviation fuel used on previous aircrafts. As this new fuel is a mixture of biofuel and jet fuel it is compatible with the selected engine as well as the infrastructure of the airport allowing for its handling, storage and maintenance.

The use of 30-35% composite structures by weight assisted in providing a maximum take-off weight saving of approximately 15% compared with that of the Boeing B737-900NG. This resulted in improved weight fractions and lift-to-drag ratios. This weight saving is necessary to further reduce fuel usage and emissions whilst matching the same flight range as aircrafts in the same class.

Its performance is leaps and bounds over current aircraft models through the use of a new engine, the Pratt and Whitney PW1000G, which is more efficient, quieter and has a 15% fuel burn reduction than the old CFM56-7 engine used on the B737’s. An increase in by-pass ratio has been met as requested from the RFP, increasing from 6-to-1 to 12-to-1.

Laminar Flow technologies such as the Co-Flow Jet Nozzle, Winglets and Wing Waggle increased the lift-to-drag ratios of the projected aircraft by 36% (optimum) compared with the B737-900NG. Through the use of these drag reducing and laminar flow technologies the aircraft will be geometrically cleaner thereby reducing the necessary fuel required.

Study of the A350’s, A380’s and the B787 Dreamliner’s advanced systems gave an insight into further improving the efficiency of the aircraft. Updated avionics, actuation methods, navigation equipment and airline management systems will provide the most efficient flight plans as well as reduced aircraft downtime, thereby reducing costs, fuel usage and overall carbon footprint.

The combination of these meet all RFP objectives and requirements.

1 National Aeronautics Research and Development Policy, (December 2006), Page 8

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A C K N O W L E D G E M E N T S

I would personally like to thank each and every one of my team members for their year long

contribution and dedication to this final year design project. Without the assistance and guidance that

you have all provided me throughout the year I believe I would not have been able to produce a report

of this standard.

I would also like to make a special mention to my deputy leaders; Leon, Nathaniel and Nigel. Thank

you for assisting me in managing the team.

On behalf of the entire team, I would also like to thank Mr. Robert Danaher for his continued

guidance without whom this report would not be possible.

Jed Guinto (Team Leader)

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C O N T E N T S

1 Introduction ........................................................................................................................................................... 11

2 AIAA – Request For Proposal .......................................................................................................................... 12

3 Aircrafts Due For Replacement ...................................................................................................................... 14

3.1 B737 Family .................................................................................................................................................. 14

3.2 A320 Family.................................................................................................................................................. 15

4 Preliminary Research ......................................................................................................................................... 16

4.1 Needs Analysis............................................................................................................................................. 16

4.2 Compliance Report .................................................................................................................................... 16

4.3 Feasibility Study ......................................................................................................................................... 16

5 Preliminary Design .............................................................................................................................................. 19

5.1 Work Breakdown Structure and Timeline....................................................................................... 19

6 SECTION 1 – Aircraft Systems, Support and Technologies................................................................. 20

6.1 Laminar Flow, Noise Reduction and Lift-To-Drag Technologies ........................................... 20

6.1.1 Laminar Flow Technology - Co-Flow Jet Flow Control ..................................................... 20

6.1.2 Noise Reduction Technologies .................................................................................................... 22

6.1.3 Lift-To-Drag Technology ............................................................................................................... 26

6.2 Aircraft systems .......................................................................................................................................... 30

6.3 Back-Up Systems ........................................................................................................................................ 43

6.4 Cockpit Avionics ......................................................................................................................................... 46

6.5 Material Selection – Composites, Aluminium and Titanium .................................................... 50

6.5.1 Composites .......................................................................................................................................... 50

6.5.2 Aluminium ........................................................................................................................................... 67

6.5.3 Titanium ............................................................................................................................................... 70

6.6 Ground and Service Support.................................................................................................................. 72

6.6.1 Alternate Fuels - Bio-Jet Fuel ....................................................................................................... 72

6.6.2 Airport Management Systems ..................................................................................................... 76

6.7 Engines – Pratt and Whitney PW1000G ........................................................................................... 83

6.8 Operational and Maintenance Costs ................................................................................................... 84

6.9 Risk Analysis ................................................................................................................................................ 87

7 SECTION 2 – Aircraft Sizing ............................................................................................................................. 89

7.1 Weight Estimate .......................................................................................................................................... 89

7.2 V-n Diagram .................................................................................................................................................. 97

7.3 Wing Structure ............................................................................................................................................ 99

7.3.1 Wing Systems ................................................................................................................................... 102

7.3.2 X-Foil Model Design ....................................................................................................................... 106

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7.3.3 Sweep Angle Comparison – 25o and a 30o Swept Wing .................................................. 108

7.3.4 Wing Sizing calculations .............................................................................................................. 111

7.3.5 Patran Analysis ................................................................................................................................ 145

7.4 Fuselage and Cabin Sizing..................................................................................................................... 160

7.5 Empennage Sizing .................................................................................................................................... 162

7.6 Cockpit Sizing ............................................................................................................................................. 168

7.7 Landing Gear .............................................................................................................................................. 169

8 SECTION 3 – Aircraft Performance ............................................................................................................. 172

8.1 Matching Chart .......................................................................................................................................... 172

8.2 Structural Drag Estimates ..................................................................................................................... 175

8.2.1 CL vs. CL/CD...................................................................................................................................... 178

8.3 Critical Flight Conditions ...................................................................................................................... 181

8.3.1 Takeoff ................................................................................................................................................ 181

8.3.2 Ground Run ....................................................................................................................................... 181

8.3.3 Cruise Conditions............................................................................................................................ 186

8.3.4 Landing ............................................................................................................................................... 187

8.3.5 RFP Performance Requirements .............................................................................................. 189

8.4 Weight and Balance ................................................................................................................................. 190

8.5 Stability – Longitudinal, Lateral and Directional ........................................................................ 192

9 SECTION 4 – Projected Aircraft Specifications ...................................................................................... 199

9.1 Comparisons ............................................................................................................................................... 203

10 Comparison Discussion ................................................................................................................................... 207

11 SECTION 5 – Total Environmental Impact .............................................................................................. 208

12 Conclusion ............................................................................................................................................................. 209

13 References ............................................................................................................................................................. 210

14 Bibliography ......................................................................................................................................................... 211

15 Appendix 1 - Detailed Design Timeline .................................................................................................... 220

16 Appendix 2 - Detailed Design Work Breakdown Structure .............................................................. 221

17 Appendix 3 – Aircraft Specifications – A320 Family ........................................................................... 222

18 Appendix 4 – Aircraft Specifications – B737 Family (737-100 to Next Generation) ............. 223

19 Appendix 5 - Aircraft Specifications – B737 Family (Next Generation Family) ....................... 224

20 Appendix 6 – Aircraft Dimensions – Scale Drawing ............................................................................ 225

21 Appendix 7 – Needs Analysis ........................................................................................................................ 226

22 Appendix 8 – Feasibility Report ................................................................................................................... 229

23 Appendix 9 – Compliance Report – B737-900NG ................................................................................ 246

24 Appendix 10 – Compliance Report - A320-200 ..................................................................................... 247

25 Appendix 11 – Density Calculation ............................................................................................................. 248

26 Appendix 14 - Airfoil Geometry ................................................................................................................... 249

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27 Appendix 15 - Matching Chart Calculations ............................................................................................ 250

T A B L E O F F I G U R E S

Figure 1- Baseline Schematic for CFJ Airfoil Showing Pump Concept ..................................................... 20 Figure 2 - Flow Field for Baseline NACA 2415 CFJ Airfoil at High Angle of Attack ............................ 20 Figure 3 - Measured Lift vs. Angle of Attack for NACA 0025 and CFJ0025-065-196 Airfoil .......... 21 Figure 4 - Measured Drag Polar for NACA 0025 and CFJ0025-131-196 Airfoil .................................. 21 Figure 5 - Historical Progress in Aircraft Noise Reduction (Courtesy of Boeing Co.) ....................... 22 Figure 6 - Effective Perceived Noise (EPN) vs. Distance from Trailing Edge of Rotor Blades to OGV ...................................................................................................................................................................................... 23 Figure 7 - Jet Noise Test Rig with Notched Nozzle, Chevron Nozzle and Base Nozzle...................... 24 Figure 8 - Comparison of Sideline Noise During Takeoff .............................................................................. 24 Figure 9 - Noise Reduction of Notched Nozzle and Chevron Nozzle ........................................................ 24 Figure 10 - Basic Concept of Pylon-mounted Jet Noise Suppressor (Side View) ................................ 25 Figure 11 - Basic Concept of Pylon-mounted Jet Noise Suppressor (Top View) ................................. 25 Figure 12 - Wing Waggle Technology in Action ................................................................................................ 28 Figure 13 - No Bleed System ..................................................................................................................................... 31 Figure 14 - ACtuation System ................................................................................................................................... 33 Figure 15 - Fly-By-Wire Technology Applications ........................................................................................... 34 Figure 16 - Fly-By-Wire System .............................................................................................................................. 35 Figure 17 - Systems architecture of our airplane ............................................................................................. 37 Figure 18 - Eliminated components of the APU System with the no-bleed architecture ................ 38 Figure 19 - In-Wheel Motors ..................................................................................................................................... 41 Figure 20 - In-Wheel Landing Technology .......................................................................................................... 41 Figure 21 - A future look in the design of avionics in the cockpit ............................................................. 47 Figure 22 - Future cockpit geometry with implemented EFB system and LCD touch screen ....... 48 Figure 23 - Cockpit Displays ..................................................................................................................................... 48 Figure 24 - Aircraft Material Density Comparison .......................................................................................... 51 Figure 25 - Typical Sandwich Composite Lay-up ............................................................................................. 51 Figure 26 - Tensile Modulus of Different Materials......................................................................................... 52 Figure 27 - Tensile Strength of Different Materials ......................................................................................... 52 Figure 28 - Fatigue Resistance of different Materials ..................................................................................... 53 Figure 29 - Composite Wing Structure ................................................................................................................. 54 Figure 30 - Composite Fibre Construction .......................................................................................................... 56 Figure 31 - tensile Strength ....................................................................................................................................... 56 Figure 32 - Compressive Strength .......................................................................................................................... 56 Figure 33 - Composite Materials Applications on Airbus A320 ................................................................. 62 Figure 34 - Composite Materials Applications on Boeing 777.................................................................... 63 Figure 35 - Composite Materials Applications on Boeing 787.................................................................... 63 Figure 36 - Growth of composite use in aircraft design as a percentage of weight ........................... 64 Figure 37 - Internal and external components of composite structures ................................................ 65 Figure 38 - Material Selection on projected aircraft (Side View) .............................................................. 66 Figure 39 - Material Selection on proposed aircraft (Top View) ............................................................... 66 Figure 40 - Fuselage Structural Composition .................................................................................................... 68 Figure 41 - Stress-Strain Curve - 2024-T3 Aluminium Alloy ...................................................................... 69 Figure 42 - Increased use of titanium in passenger aircraft ........................................................................ 70

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Figure 43 - Current fuel distribution system ..................................................................................................... 73 Figure 44 - Depiction of asset viewer with available connection types .................................................. 77 Figure 45 - Fuel cell system ....................................................................................................................................... 80 Figure 46 - System Diagram ...................................................................................................................................... 80 Figure 47 - Risk Matrix ................................................................................................................................................ 88 Figure 48 - Mission Profile ......................................................................................................................................... 89 Figure 49 - Table 2.1 from Roskam (Part 1) ....................................................................................................... 91 Figure 50 - Vn Diagram ............................................................................................................................................... 97 Figure 51 – AAA input data ....................................................................................................................................... 98 Figure 52 – Right side of wing (x-foil) .................................................................................................................. 99 Figure 53 - Wingbox structure (Ribs and shear webs) ................................................................................ 100 Figure 54 - WIngbox Structure (Shear webs and skins) ............................................................................. 100 Figure 55 – matching chart ...................................................................................................................................... 101 Figure 56 - Fuel tank layout inside the aircraft. .............................................................................................. 102 Figure 57 - Integral fuel tank inside the wing. ................................................................................................. 103 Figure 58 - Single Slotted Flap ............................................................................................................................... 103 Figure 59 - Flap Location on wing ........................................................................................................................ 104 Figure 60 - Slats ............................................................................................................................................................ 104 Figure 61 - LE Flaps .................................................................................................................................................... 105 Figure 62 - Spoilers and Speed Brake locations.............................................................................................. 105 Figure 63 - CL/CD vs CL graph for B737-900NG wing .................................................................................... 106 Figure 64 - CL/CD vs CL graph for Designed wing ........................................................................................... 106 Figure 65 - 3D wing analysis for B737-900NG wing configuration ........................................................ 107 Figure 66 - 3D wing analysis for Designed wing configuration ............................................................... 107 Figure 67 - Pressure contour and transition position for B737-900NG with 25.02 degrees sweep angle .................................................................................................................................................................................. 108 Figure 68 - Pressure contour and transition position for B737-900NG with 30 degrees sweep angle .................................................................................................................................................................................. 108 Figure 69 - Pressure contour and transition position for Designed model with 25.02 degrees sweep angle .................................................................................................................................................................... 109 Figure 70 - Pressure contour and transition position for Designed model with 30 degrees sweep angle .................................................................................................................................................................................. 109 Figure 71 - Scale dimensioning of optimized design for z-stringer skin in single cell design ..... 113 Figure 72 - Scale Dimensioning of Optimal Design for Z-Stringer Skin in Single Cell Design ...... 119 Figure 73 - Demension of the Z-stringer ............................................................................................................ 124 Figure 74 - Scale dimensioning of optimized design for Integrated stringer skin in single cell design ............................................................................................................................................................................... 126 Figure 75 – Design Chart For Unflanged Integral Stiffeners ...................................................................... 127 Figure 76 - Scaled Dimensioning of Optimum Integral Stringer Skin for Single cell Design ........ 132 Figure 77 - Scaled Dimensioning of De-optimized Integral Stringer Skin For Single Cell Design ............................................................................................................................................................................................. 137 Figure 78 - The single cell configuration with the spar caps ..................................................................... 140 Figure 79 - Wing Dimensions ................................................................................................................................. 145 Figure 80 – Lift Loads ................................................................................................................................................ 145 Figure 81 - Lift Distribution .................................................................................................................................... 146 Figure 82 - Distributed Lift loads .......................................................................................................................... 147 Figure 83 - engine loads (2.5g Loading)............................................................................................................. 147 Figure 84 - Fuel loads (2.5g Loading) ................................................................................................................. 148 Figure 85 - Constraints .............................................................................................................................................. 149 Figure 86 - Maximum Displacement .................................................................................................................... 150 Figure 87 - Stress ......................................................................................................................................................... 150

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Figure 88 - Maximum Deflection ........................................................................................................................... 151 Figure 89 - stress ......................................................................................................................................................... 151 Figure 90 - Cruise COndition Stress (Patran) .................................................................................................. 153 Figure 91 - Point Loads ............................................................................................................................................. 153 Figure 92 - Geometry ................................................................................................................................................. 154 Figure 93 – SFD............................................................................................................................................................. 154 Figure 94 - Torque Diagram .................................................................................................................................... 155 Figure 95 - Bending Moment Diagram ............................................................................................................... 156 Figure 96 - Shear FLow Cut ..................................................................................................................................... 157 Figure 97 - Resultant Shear Flows (Uncorrected) ......................................................................................... 158 Figure 98 - Corrected Shear Flows ....................................................................................................................... 159 Figure 99 - In-Cabin Seating Arrangement and Fuselage Length ............................................................ 160 Figure 100 - Fuselage Cross Section .................................................................................................................... 161 Figure 101 - Boeing 767 vertical tailplane ........................................................................................................ 163 Figure 102 - Boeing 767 horizontal stabilizer ................................................................................................. 163 Figure 103 - Conventional tailplane .................................................................................................................... 164 Figure 104 - Definition of empennage spar, rib and stiffener locations ............................................... 165 Figure 105 - B707 Horizontal stabilizer............................................................................................................. 166 Figure 106 - two methods for vertical tail to fuselage attachment ......................................................... 167 Figure 107 - Visibility from port side .................................................................................................................. 168 Figure 108 - Retraction Method ............................................................................................................................ 169 Figure 109 - Wheel Configurations ...................................................................................................................... 170 Figure 110 - Tire Pressures and LCN .................................................................................................................. 170 Figure 111 - Tire Pressures depending on runway surface ....................................................................... 171 Figure 112 - Equivalent Single Wheel Load ...................................................................................................... 171 Figure 113 – CL/CD Comparison ............................................................................................................................ 179 Figure 114 - Takeoff Mission Profile ................................................................................................................... 181 Figure 115 - landing mission profile ................................................................................................................... 187 Figure 116 - Centre of Pressure and Centre of Gravity Locations ........................................................... 190 Figure 117 - Aerodynamic Geometry .................................................................................................................. 193 Figure 118 - Front View ............................................................................................................................................ 199 Figure 119 - Top View ............................................................................................................................................... 199 Figure 120 - Top View (With Seating Arrangement) .................................................................................... 200 Figure 121 - Side View ............................................................................................................................................... 200 Figure 122 - Concept Artwork................................................................................................................................ 201 Figure 123 – fuselage (Seating Arrangement) ................................................................................................. 201 Figure 124 - fuselage cross section ...................................................................................................................... 201 Figure 125 - CATIA Model (View 1) ..................................................................................................................... 202 Figure 126 - CATIA Model (View 2) ..................................................................................................................... 202 Figure 127 - CATIA Model (View 3) ..................................................................................................................... 202

T A B L E O F T A B L E S

Table 1 - RFP Requirements ..................................................................................................................................... 13 Table 2 - Summary of Noise Reduction technology ........................................................................................ 26 Table 3 - Statistical Overview of Fixed Winglets .............................................................................................. 27 Table 4 - Statistical Overview of Wing Waggle Technology ......................................................................... 29 Table 5 - Summary of Lift-To-Drag Technologies ............................................................................................ 29 Table 6 - Hydraulics per system .............................................................................................................................. 34

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Table 7 – Material Properties of Different Composites ................................................................................. 53 Table 8 - Emission Factors ......................................................................................................................................... 59 Table 9 - Aluminium Alloy Application and Type ............................................................................................ 68 Table 10 - Aluminium Alloy Properties ................................................................................................................ 69 Table 11 - Titanium Characteristics ....................................................................................................................... 70 Table 12 - Fuel Database (Jet A, Bio-jet Fuel, Diesel, Biodiesel, Liquid Methane) .............................. 74 Table 13 - Fuel Database (Liquid Hydrogen, Methanol, Ethanol, Electricity) ...................................... 75 Table 14 - System Capabilities ................................................................................................................................. 77 Table 15 - CFM56 vs. PW1000G .............................................................................................................................. 83 Table 16 - Fuel Types with Associated Properties .......................................................................................... 84 Table 17 - Fuel prices - 25 June 2010 .................................................................................................................... 84 Table 18 - Fuel prices - 17 Sep 2010 ..................................................................................................................... 84 Table 19 - The impact on 2010’s fuel bill of the global airline industry ................................................. 85 Table 20 - Base-Line Aircraft Operating Costs - Average per Block Hour (AUS$) .............................. 85 Table 21 - Estimated Aircraft Operating Costs/Hour as at June 2009(AUS$) ...................................... 86 Table 22 - Fuel burn (kg/hr) by phase of flight ................................................................................................ 86 Table 23 - Fuel Burn for LTO by Average Fleet ................................................................................................. 86 Table 24 - Risk Analysis .............................................................................................................................................. 87 Table 25 - Boeing 737 Specifications .................................................................................................................... 90 Table 26 - Weight Estimation Using the Roskam Method ........................................................................... 95 Table 27 – Weight Comparison – B737-900NG ................................................................................................ 96 Table 28 - Weight comparison - A320-200 ......................................................................................................... 96 Table 29 - wing Specifications .................................................................................................................................. 99 Table 30 - Limitations ................................................................................................................................................ 105 Table 31 - The values obtained by simulation ................................................................................................. 107 Table 32 - The values through weight comparison ....................................................................................... 108 Table 33 - The comparison of glide ratio (CL/CD) with 25.02° and 30° swept wing ...................... 109 Table 34 - The comparison of XCp position in 25.02° and 30° swept wing ........................................ 109 Table 35 – Nomenclature ......................................................................................................................................... 111 Table 36 - Basic structural and loading data for calculation ..................................................................... 113 Table 37 – Summary of Single Celled Optimum For Z-Stringers ............................................................. 119 Table 38 - Summary of Single Cell De-Optimized Design for Z-Stringers ............................................ 124 Table 39 - Comparison of Optimum and De-Optimized Designs ............................................................. 125 Table 40 - Optimum Integral Stringer Skin Summary .................................................................................. 132 Table 41 - De-Optimized Integral Stringer Skin Summary ......................................................................... 137 Table 42 - Comparison of Optimum and De-Optimized Designs ............................................................. 138 Table 43 - comparison of de-optimized z-stringer and de-optimized integral stringer ................ 139 Table 44 - Spar cap dimensions ............................................................................................................................. 143 Table 45 - Spar web dimensions ........................................................................................................................... 143 Table 46 - Lift Distribution per node ................................................................................................................... 146 Table 47 - engine loads ............................................................................................................................................. 147 Table 48 - Fuel Loads ................................................................................................................................................. 148 Table 49 - Aluminium alloy properties .............................................................................................................. 149 Table 50 - Wing thicknesses ................................................................................................................................... 149 Table 51 - Load unit COnversion ........................................................................................................................... 154 Table 52 - Torque Calculations .............................................................................................................................. 155 Table 53 - Bending Moment Calculations .......................................................................................................... 156 Table 54 - Corrected Shear Flows ......................................................................................................................... 158 Table 55 - Horizontal stabilizer geometry ........................................................................................................ 162 Table 56 - Vertical tailplane geometry ............................................................................................................... 162 Table 57 - Empennage Construction ................................................................................................................... 167

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Table 58 - Structural Dimensions ......................................................................................................................... 175 Table 59 - Structural drag estimate ..................................................................................................................... 178 Table 60 - CL vs CL/CD values Without WIng Technology ....................................................................... 179 Table 61 - CL vs CL/CD with wing technology ................................................................................................ 179 Table 62 - Lift-To-Drag Comparison .................................................................................................................... 180 Table 63 - RFP Requirements ................................................................................................................................. 189 Table 64 - specifications ........................................................................................................................................... 192 Table 65 - Specifications ........................................................................................................................................... 195 Table 66 - Details ......................................................................................................................................................... 197 Table 67 – B737-900NG vs. Projected Aircraft ............................................................................................... 203 Table 68 – A320-200 vs. Projected Aircraft...................................................................................................... 204 Table 69 – Projected Aircraft vs. RFP Requirements .................................................................................... 205 Table 70 - RFP Requirements ................................................................................................................................. 206 Table 71 - Noise Reduction Technologies ......................................................................................................... 208 Table 72 - Fuel Consumption and Emission Saving Technology ............................................................. 208

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1 INTRODUCTION

Jet fuel is one of the biggest expenses for the aerospace industry. The need for a more affordable and efficient fuel for this day and age is a must considering the importance of emission control. Environmental, social and political parties are now placing limits on the level of emissions which airports can emit. The proposed RFP by the AIAA has this issue in mind. An investigation is to be carried out which clearly analyzes the prospects of a projected aircraft to be available by 2020 which is more environmentally friendly than the aircrafts it is intended to replace; the Boeing B737 and the Airbus A320.

This aircraft will investigate the properties and use of alternate fuel sources with reduced emissions over standard jet fuel, a possible weight reduction through the use of advanced materials and structures, reduced noise and fuel burn through engine and structural technology, improve lift-to-drag using laminar flow technology, new aircraft systems and navigation equipment for more efficient flight patterns and ground control, improved ground support and improved operational procedures. A risk analysis will be performed for the appropriate items which will include necessary items pre and post 2010.

A preliminary design of the aircrafts structure and weight will then be performed. It will analyze major load bearing components such as the wings and the empennage in detail. The structural integrity will be analyzed using appropriate software for a cruise condition as well as a typical maneuvre for a single aisle commercial transport aircraft. Performance will be analyzed in accordance to the RFP requirements and the critical flight conditions will all be investigated using FAR25. Weight and balance will be performed followed by a longitudinal, lateral and directional stability analysis. A general arrangement will be provided using both CAD and CATIA which will include an internal cabin layout for the seating arrangement, cabin cross-sectional area and the aircrafts external structure. CATIA will provide a 3 dimensional image of the projected aircraft. The overall aircrafts effect on the environment in regards to all types of emissions will be summarized and discussed with pre-existing aircrafts within the same class.

A summary of all RFP and FAR25 requirements, both successfully and unsuccessfully met will be discussed and appropriate recommendations will be made.

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2 AIAA – REQUEST FOR PROPOSAL

Project Description Set By RFP

“To design an aircraft by 2020 which incorporates current and new technologies in order to reduce overall emissions and fuel consumption at all stages of its life-cycle”.

This preliminary report will investigate the plausibility of designing an aircraft which is more environmentally friendly in comparison to existing aircraft models in the same class. The following criteria must be met according to the RFP.

1. Date of aircraft operation: 2020

2. Improvement in energy efficiency for long term national goals in aerospace

3. To reduce the environmental impact of the industry in order to continue aviation

research

4. Reduction in fuel consumption and emissions due to ever increasing fuel costs and a

possible carbon footprint tax in the future.

5. Aircraft systems, structures and performance

a. Performance improvement

b. Impact of alternative fuels on aircraft design as well as airline infrastructure

c. Noise reduction

d. Enhanced laminar flow

e. Integration of high by-pass ratio for lower fuel consumptions

f. Advances in materials and structures for an increase in strength to weight ratios

g. 25% improvement of lift-to-drag through the utilization of multidiscipline

configuration and laminar flow technology

h. Weight fractions improvement

6. Incorporation of new technologies, operational procedures and alternate fuels.

7. The design will need to be considered a 737NG/A320 replacement aircraft

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General Requirements Set By RFP

TABLE 1 - RFP REQUIREMENTS

Item Requirement Note

Safety and Airworthiness Regulations FAR25 -

Crew 2 -

Passengers 175 (1 class)

Seating Pitch 32 inches

Seating Width 17.2 inches

Cabin Width 12.5 feet

Cabin Height 7.25 feet

Cargo Capacity 1240 feet3

Take-off distance 8200 feet

Landing Speed 140 KCAS

Maximum Weight Maximum Zero Fuel Weight plus reserves for maximum

range

Cruise Speed 0.8 MACH

Maximum Operating Speed 0.83 MACH

340 KCAS

Initial Cruise Altitude 35000 feet

Nominal Range 1200 nm

Maximum Range 3500 nm

Payload Capability 37000 lbs

Alternate Fuels Biofuels

Assumptions/Constraints

1. It is assumed that the use of composites in aerospace will continue until and after the year 2020.

2. It is assumed that the systems, avionics and structural technology that are projected to be available in the future will arrive as stated by the manufacturer.

3. Access to certain specifications may be unattainable due to them being trade secrets, therefore, assumptions are made where necessary.

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3 AIRCRAFTS DUE FOR REPLACEMENT

3.1 B737 FAMILY

In 1958 Boeing announced that a twin engine commercial transport aircraft design was being studied. In February 1965 the first order was placed and the project began. Since then the B737 has become the best selling commercial aircraft in aviation history with more than 5900 orders from 225 customers, (for now). It is said that a B737 is taking off somewhere around the world every 5 seconds. Its configuration is similar to the A320 family in that it has two engines mounted on the wing, conventional wings and approximately the same seating capacity depending the class.

Advantages

1. The Next Generation Family has 33% fewer parts than the Classic versions thus reducing production time allowing for the production of 21 aircraft per month.

2. The fuselage, wing and tail sections have all been strengthened and performance has increased significantly from the classic models.

Purchase Price

In 2008 the cost of the B737NG ranged from US$51.5 to US$87 million depending on the class.

For detailed specifications, see “Appendix 4 – Aircraft Specifications – B737 Family (737-100 to Next Generation)and “Appendix 5 - Aircraft Specifications – B737 Family (Next Generation

Family)”

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3.2 A320 FAMILY

Airbus A320 is a twin engine short to medium range commercial aircraft. Launched in March 1984, it first flew on 22 February 1987. It was the world’s first series aircraft with fly-by-wire and side-stick controls allowing the aircraft to land virtually unassisted. It is powered by a either a CFM International CFM56-5 or International Engines (IAE) V2500 Engines. The aircraft is a narrow body design which is capable of various passenger capacities ranging from 107 to 220 passengers depending on the class. It used to require 8 months to produce a single aircraft and in 2008 the A320 was being produced at a rate of 32 aircraft per month making them one of the quickest and most successful aircrafts to be produced to date.

The aircraft is a low wing cantilever monoplane with a conventional empennage design. It has a tricycle wheel configuration and is powered by two wing mounted turbofans. Its structure consists of kevlar, fibreglass, aluminium and titanium.

Advantages

1. The wings of the A320 are longer and thinner than that of the B737 making them more efficient.

2. The aircraft also has a cabin width of 3.7m, making it slightly larger than that of the B737’s 3.25m. This slight difference made the A320 the most comfortable single-aisle aircraft in the world.

Disadvantages

1. Errors in landing manueuvres due to errors in measuring altitude. 2. Warning systems activate only seconds before an accident. 3. The flight path angles and vertical speed indicator utilise the display format and

sometimes becomes confusing pilots.

Purchase Price

In 2008 the cost of the A320 ranged from US$73.2 to US$80.6 million depending on the class.

For detailed specifications, see “Appendix 3 – Aircraft Specifications – A320 Family”

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4 PRELIMINARY RESEARCH

4.1 NEEDS ANALYSIS

In order to understand the requests of the customer, the RFP requirements must be analysed further.

To see the detailed version of the needs analysis see “Appendix 7 – Needs Analysis”

4.2 COMPLIANCE REPORT

Compliance reports are used to determine which existing aircrafts are the most suitable to use as a base reference design. Compliancy was performed for the B737-900NG and the A320-200 as they meet more design criteria as any other model within their families. Our aim is to determine which requirements have been met, which have not and by how much.

The report showed that the Boeing 737-900NG and the A320-200 are almost identical, with only slight differences. The B737-900NG however, is slightly closer to the design requirements set out by the RFP and the needs analysis and therefore was chosen to be the base design. Another feature which helped with the decision is the fact that the Boeing 737 showed a higher range capability.

To see the detailed version of the compliance report see “Appendix 9 – Compliance Report – B737-900NG” and “Appendix 10 – Compliance Report - A320-200”

4.3 FEASIBILITY STUDY

NOTE: Below is a summarized version of the feasibility study. To see in detail see “Appendix 8 – Feasibility Report”

Aircraft Structural Configuration – Engine Attachment

The structural configuration was analyzed for a conventional aircraft with the engines attached to the wings and engines attached to a t-tail section. It was found that the t-tail is only viable for business jet aircrafts as the loads applied by the engines onto the structure are high. The conventional design with engines attached to the wings show much more promise and historical evidence in regards to its success as a commercial transport design.

Wing Configuration – Conventional vs. Blended Wing Body

The use of a low-wing, swept-back, dihedral wing shape was analyzed for use on a commercial aircraft.

Low Wing - It was found that the low wing is the optimized design for commercial aircrafts as it allows room for more passengers, room for landing gear, more structurally sound, easily maintained, easily refueled and more aerodynamically stable due to the possible landing gear configuration. Another advantage is its ability to protect the fuselage from structural damage due to debris upon landing and take-off.

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Swept-back – This design reduces drag during transitions between different Mach numbers through a delay of the formation of shockwaves. This design is a solution for aircrafts that operate just below the speed of sound.

Dihedral – This adds lateral stability which is needed for low wing passenger aircrafts. It also allows for more engine clearance.

This design proved to be more viable than a blended wing body approach. Even though the blended wing body has advantages over the conventional design, it is not economically viable at this time.

Engine Selection – PW1000G vs. Leap-X

The PW1000G and the Leap-X don’t vary too much in terms of specifications and performance. Operational and maintenance costs as well as compatibility with existing aircrafts and airport infrastructure are very similar. The main difference is the release date. As these two aircrafts are new in terms of commercial use, comparatively, very little is known about them.

The PW1000G will be released for commercial use in late 2013 whereas the Leap-X is to be fully certified in 2016. This 3 year gap provides the necessary time to fully assess the PW1000G’s capabilities on a commercial aircraft. It also allows for more time to find a replacement engine in the event that it does not meet the projected performance capabilities. By analysing a second engine, we have also provided a contingency design in the event that the PW1000G becomes unsuitable for our projected design.

Materials – Composite vs. Aluminium

The properties of composites are very desirable when trying to reduce fuel usage. The possible weight saving has been seen in the Airbus A350 and the Boeing 787 Dreamliner designs. However, as composites have many problems associated with it such as recycling, production costs, raw material costs, repair and lightning strike, as well as the relative low knowledge engineers have with it, a high level (i.e. ~50% of design weight) of use on a medium sized commercial aircraft has shown to be infeasible. Lower levels (~ 15% of design weight) of composites however, have shown enough promise in order for it to be utilized on the projected aircraft.

A majority use of aluminium will be used on the aircraft in order to reduce maintenance costs and production as well as provide adequate recycling for the future.

Alternative Fuel Sources – Bio-jet Fuel, Hydrogen, Liquid Methane

Bio-jet fuel – showed the most promise out of the three. It is a renewable energy source made from plants crops, oils or animal fats. It is a combination of Jet fuel and biofuel and has successfully been tested on the B737-800 model and provided excellent results. Tests by the “National Renewable Energy Laboratory” have shown that it substantially reduces carbon dioxide emissions. Problems associated with biofuel production however are the extremely large land mass required to produce it. By combining it with Jet fuel, costs and land mass are reduced. Storage and infrastructure require little change as the handling is similar to current practices. This blended fuel source is also compatible with existing engines.

Hydrogen – this fuel source produces fewer emissions during production and considerable weight reduction compared with current jet fuel. However, there are geographic limitations and only a few countries in world that are capable of producing this fuel. It also requires a substantial infrastructure and ground support modification in order to utilize it.

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Liquid Methane – Methane is a natural gas that is stored within the earth which could potentially offer hundreds of years of fuel. It could be a viable option as there are vast quantities trapped within the earth. However, the use of methane is new and complex at all stages of its utilization. Also, it is not compatible with existing engines or aircraft designs. Ground support and infrastructure costs will be high and therefore infeasible.

Ground Support and Infrastructure – Airport Management Vehicles

One method of reducing overall emissions is by focusing on the external factors of the aircraft, the airport management systems. Reducing emissions at airports also contributes to the overall reduction of emissions of our aircraft design. The use of hydrogen fuel cells will substantially reduce ground emissions and the use of new systems such as the “Zebra Enterprise” and “Airport Visualiser” can increase ground efficiency when refuelling or taxiing as well as reduce aircraft downtime.

Operational and Maintenance Costs

Some methods of cost reduction include producing biofuels and hydrogen at a closer proximity to the airport. This will reduce transportation costs and emissions. Using stronger and smarter materials we hope to increase the intervals between maintenance thereby reducing maintenance costs. We also hope to reduce production times by outsourcing components to other manufacturers around the world. The reason for this is to make an incentive to these countries to purchase our aircrafts.

At the present time we are assuming the price of our aircraft will be around the same as that of the existing Boeing 737-900NG and A320-200 models which is between US$51.5-87M. Even if our aircraft is slightly more expensive, the amount of potential savings which come from a reduction in manufacturing time, fuel efficiency and low maintenance will cover this slight increase in price.

We are assuming that operational and maintenance costs are similar to existing aircrafts within the same class. In terms of maintenance, a twin engine narrow-body aircraft is US$515/hour and a twin engine wide-body is US$787/hour. These values are based on the Bureau of Transportation Statistics, 1999-1998. Delays cost approximately $10000 per hour and operational costs, based on the Boeing 737-800 is US$1665 per hour.

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5 PRELIMINARY DESIGN

Introduction

Preliminary design is made up of 5 main parts;

Section 1 – Aircraft, Systems, Materials, Support and Technologies o This section will contain the internal and external systems of the aircraft such as

laminar flow technology, noise reduction technology, lift-to-drag technology, aircraft systems, back-up systems, cockpit avionics, materials, ground and service support, engine selection, operational and maintenance costs and a risk analysis. It will also contain a detailed fuels database which includes densities, weights and energies.

Section 2 - Aircraft Sizing o This will contain a preliminary weight estimate using Roskam followed by a V-n

diagram win which the sizing of all major aircraft components such as the wing sizing and structural testing for a maneuvre, fuselage and cabin sizing, empennage sizing, landing gear sizing and cockpit sizing will be based upon.

Section 3 - Aircraft Performance o Aircraft performance will contain a matching chart, structural drag estimates, an

analysis on critical flight conditions, weight and balance and stability.

Section 4 - Projected Aircraft Specifications o This section will summarize all specifications, capabilities, weights and other

details. It will also provide comparisons between the projected aircraft against RFP requirements, the A320-200 and the B737-900NG

o Geometrical CAD image’s will also be provided with a top view, side view, front view, internal cabin layout and the cabin cross-sectional area.

Section 5 - Total Environmental Impact o It will sum up the effect of all technologies regarding noise, fuel burn, drag

reduction and lift improvement on emission levels as per requested by the RFP.

5.1 WORK BREAKDOWN STRUCTURE AND TIMELINE

The Work Breakdown Structure for this report has been broken into 10 components. When all components have been completed it will result in the final report which will include all areas mentioned in the introduction. For a diagrammatical view see “Appendix 2 - Detailed Design Work Breakdown Structure”

The timeline for this report holds a focus on the major areas of study first. That being, stability, performance, wing sizing and aircraft systems. For the gantt chart see “Appendix 1 - Detailed Design Timeline.”

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6 SECTION 1 – AIRCRAFT SYSTEMS, SUPPORT AND TECHNOLOGIES

6.1 LAMINAR FLOW, NOISE REDUCTION AND LIFT-TO-DRAG TECHNOLOGIES

6.1.1 Laminar Flow Technology - Co-Flow Jet Flow Control

RFP Requirement: An improvement in laminar flow over the wing through airfoil technology whereby the transition point is delayed.

The Co-Flow Jet (CFJ) airfoil is a design which can improve the flow over the airfoil. To develop this airfoil technology, the flow control technique contains four main aspects.

1. Its effectiveness to enhance lift and stall margin whilst reducing drag. 2. Its penalty on the propulsion system energy efficiency. 3. Its projected weight increase. 4. Ease of implementation.

Concept

The concept of the CFJ airfoil is to inject air from a slot at the leading edge and then suck it using a suction slot near the trailing edge as seen in Figure 1. It works basically by mixing different pressure gradients, i.e. the flow from the injection slot and the boundary layer flow over the airfoil. By doing so it transfers energy from the CFJ to the main flow allowing laminar flow to remain attached to the surface at very high angles of attack. This significantly increases the stall margin and at the same time the high momentum from the jet drastically increases circulation of airflow. This then augments lift, reduces drag and can possible increase thrust.

FIGURE 1- BASELINE SCHEMATIC FOR CFJ AIRFOIL SHOWING PUMP CONCEPT

FIGURE 2 - FLOW FIELD FOR BASELINE NACA 2415 CFJ AIRFOIL AT HIGH ANGLE OF ATTACK

At high angles of attack, the turbulent layer field is displayed at the top part of Figure 2. As you can see, the CFJ airfoil performs much better with almost no separation or turbulence.

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Application

The CFJ airfoil will have a beneficial effect on the aircraft at all critical flight conditions. During takeoff the CFJ provides strong injection to increase the lift-to-drag ratio while a very weak injection at cruise conditions. At approach, high lift and low drag can be produced due to its ability to operate at high angles of attack.

In wind tunnel testing it was compared with an uncontrolled NACA airfoil. Tests revealed that the maximum lift coefficient for the CFJ was 5.02 whilst the NACA airfoil was at 1.52. This 70% increase was possible due to its ability to operate at high angles of attack. The NACA airfoil was seen to stall at 19o while the CFJ stalled at an impressive 44o, a 57% increase. It was also seen that drag was significantly reduced. This can all be seen in Figure 3 and Figure 4 below.

FIGURE 3 - MEASURED LIFT VS. ANGLE OF ATTACK FOR NACA 0025 AND CFJ0025-065-196

AIRFOIL

FIGURE 4 - MEASURED DRAG POLAR FOR NACA 0025 AND CFJ0025-131-196 AIRFOIL

This technology is indeed beneficial for our aircraft and will be implemented. It has shown considerable promise in wind tunnel testing design. Further analysis of this technology will be carried out in the Risk Analysis section of the report.

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6.1.2 Noise Reduction Technologies

RFP Requirement: Noise reduction counts as a type of emission and therefore an investigation is to be carried out to investigate possible ways to reduce noise.

On the surrounding communities and the travelling public, the impact of noise generated from aviation has been a significant issue and a public concern since the beginning of commercial jet powered aviation. The major purpose of noise reduction is the environmental improvement for noise levels outside the airport perimeter in a well-populated urban environment.

As noise is a type of emission, it is absolutely vital to investigate the possible sources of noise and methods to reduce or eliminate them. Noise production will dictate the regions of operation for the aircraft.

Sources of High Noise

1. Exhaust nozzle jet noise at takeoff. 2. Fan noise at takeoff. 3. High lift system from flaps and slats at landing. 4. Wake noise of the airplane at takeoff and landing. 5. The long takeoff/landing distance and shallow climb/descend angles. 6. Landing gear noise at landing.

Noise Technologies

This can be broken down into three main categories: engine noise, structural noise and cabin noise.

6.1.2.1 Engine Noise

Three different engine types are investigated here. The reason for this and not just a study of our projected engine is in the event that our selected engine becomes unavailable. This study then provides a contingency plan in the field of noise reduction.

The noise by jet engines has been reduced with the development of new engine technology. Figure 5 below shows how over time the effective perceived noise level (EPNL) has steadily decreased as new designs are released.

FIGURE 5 - HISTORICAL PROGRESS IN AIRCRAFT NOISE REDUCTION (COURTESY OF BOEING CO.)

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Fan Noise Reduction Technology – Environmental Compatible Engine (ECO)

Typically, the fan of an engine consists of rotating blades (rotor blades), outlet guide vanes (stationary blades) and a frame strut. The new Environmentally Compatible Engine (ECO) is basically the same design except that the outlet guide vanes and the frame strut have been incorporated into one single structure, the Integrated OGV as seen in Figure 6 below. This OGV is composed of 3 innovated designs; the sweeplean design, the swept rotor blade and the short axial spacing between fan rotor blade and OGV.

So by reducing the spacing between the rotor blades and the OGV, perceived noise can be reduced. Figure 6 below shows how the sound is reduced as the spacing is reduced giving a total noise reduction of 3.8dB.

The Integrated OGV is not implemented on our aircraft as the PW1000G engine is already designed. However, in the event that the PW1000G becomes unavailable the ECO is a possible engine replacement system component which can assist in reducing emissions.

FIGURE 6 - EFFECTIVE PERCEIVED NOISE (EPN) VS. DISTANCE FROM TRAILING EDGE OF ROTOR BLADES TO OGV

Jet Noise Reduction Technology – Notched Nozzle

For jet noise reduction, the fundamental method consists of the high temperature, high pressure gas from the turbine and compressed air are mixed inside of the exhaust nozzle, accompanied by a mixture promotion device upstream of the flow. This method however is difficult to implicate as the engine structure is very complicated but also presented some cost and weight issues.

Present research has shown that a notched nozzle has a reduction capability of 2.2EPNdB during takeoff. Testing has already begun on the B787-9 Dreamliner by making the contours of the exit portion of the exhaust nozzle saw-edged. However, the reduction in thrust during flight has yet to be addressed.

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Figure 8 shows how for the same flight time how noise is reduced between a notched nozzle and a base nozzle. Figure 9 shows the noise reduction between the notched nozzle and the chevron nozzle.

FIGURE 7 - JET NOISE TEST RIG WITH NOTCHED NOZZLE, CHEVRON NOZZLE AND BASE NOZZLE

FIGURE 8 - COMPARISON OF SIDELINE NOISE DURING TAKEOFF

FIGURE 9 - NOISE REDUCTION OF NOTCHED NOZZLE AND CHEVRON NOZZLE

Again, this technology is not implemented into our design as the engine has already been designed. It is a possible replacement technology in the event that the PW1000G becomes unavailable.

6.1.2.2 Structural Noise

Pylon-Based Jet Noise Suppressors

The pylon, found in most commercial jet aircrafts is used to mount the engine to the wing as seen in Figure 10 and is a source of high noise generation. For reducing the noise level, the flap implement is installed in sideward and downward directions between the core nozzle and aft of pylon as shown Figure 10 and Figure 11. During testing the fine-perforated flap provided a noise reduction of 2.1dB in the downward direction and 1.0 dB in the lateral direction. This technology will be implemented onto the projected aircraft.

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FIGURE 10 - BASIC CONCEPT OF PYLON-MOUNTED JET NOISE SUPPRESSOR (SIDE VIEW)

FIGURE 11 - BASIC CONCEPT OF PYLON-MOUNTED JET NOISE SUPPRESSOR (TOP VIEW)

Absence of High Lift Devices on Noise

The Co-Flow Jet airfoil as mentioned earlier is capable of reducing noise by avoiding the need of high lift devices. It can do this because it does not require any flaps or slats to produce high lift and therefore takeoff and landing distances can be reduced. These flaps and slats are sources of high noise emission.

6.1.2.3 Cabin Noise

Noise levels within the cabin are now relatively low due to the many advances in engine technology. The insulation of sound and sound proofing material has greatly assisted in noise damping. There is still however room for improvement. Noise is still being produced during takeoff, cruise and landing. Vibrations from the outside are produced from the engine as well as air streaming along the skin. These transmit through the cabin structure via skins, frames, glass wool thermal insulation and trim panels. Simply adding more damping material is not a viable option due to the added weight.

Solution 1 – Actuators and Sensors

A possible solution is to replace the conventional passive mounts which attach the panels to the fuselage structure. Installing three actuator units, three acceleration sensors and an electrical power unit to the panel which is all connected to a digital processor can greatly assist in damping. Vibrations are sensed by the sensors, sent to the processor, and the necessary damping effect is sent back to the actuators.

Solution 2 – Integrated Piezo-Actuators

Another possible solution is the use of integrated piezo-actuators inside the panels of the honeycomb structure during manufacturing. With acceleration sensors on the panels and actuators driven by electronic means, the noise transmissions are minimized with the counter vibrations induced by local flexing of the panels by the actuators.

Each solution was tested and provided good results. Noise reductions of up to 20dB (down to 1% of the initial level) for discreet frequencies and up to 10dB (down to 10% of the initial level) for broadband noises were found.

Application

Solution 1 will be implemented but in the event it does not become compatible with the aircraft

structure solution 2 will be implemented.

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Summary of Noise Technology

TABLE 2 - SUMMARY OF NOISE REDUCTION TECHNOLOGY

Noise Technology Improvement

Environmental Compatible Engine (ECO) 3.8dB reduction

Notched Nozzle 2.2dB reduction

Pylon-Based Jet Noise Suppressors 2.1dB reduction

Cabin Noise Technology 20dB reduction

TOTAL 28.1dB reduction

6.1.3 Lift-To-Drag Technology

RFP Requirement: The RFP states that a 25% increase in lift-to-drag is desired. Because this improvement is a combination of geometry, engines and systems, it is important to consider all these aspects as a whole which contributes to the overall lift-to-drag. So although some of the technologies below will only increase lift-to-drag by a few percent each, in combination with all the other modifications and technologies, the goal of 25% lift-to-drag improvement may be possible.

It is also worth mentioning that because our aircraft is based on the Boeing 737 design, compatibility will be based on the B737-900NG design.

6.1.3.1 Fixed Winglets

Winglets are the 8 feet, carbon graphite wingtips that allow an airplane to extend its range, carry as much as 6,000 lbs more payload from takeoff-limited airports and save on fuel. Winglets are wing tip extensions which reduce lift induced drag and provide some extra lift. There are two ways to implement winglets and they are during the production of the actual aircraft as well as during service, i.e. a retrofit.

Compatibility

Winglets are available on 85% of all new 737s now built, particularly the -800 and -900 series. Some aircrafts do not possess winglets as they cost about US$725,000 and require about 1 week to install which costs an extra US$25,000-80,000. Once fitted, they add 170-235kg to the weight of the aircraft, depending upon whether they were installed during production or a retrofit. The fuel cost of carrying this extra weight will take some flying time from each sector to recover, although this is offset by the need to carry less fuel due to the increased range.

Advantages

1. Environmentally friendly – it can reduce noise production by 6.5% and Nitric Oxide emissions by 5%. This can also provide savings on airport noise quotas.

2. Improved climb gradient – this enables a higher allowable takeoff weight from climb limited airports.

3. A reduced required climb thrust of 3% which can extend engine life and reduce maintenance costs.

4. A reduced required cruise thrust of 6% giving substantial savings of fuel and an increase in range.

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5. Improved cruise performance – winglets allow the aircraft to reach higher altitudes sooner. This means that a direct climb to 41,000 feet is now possible thus avoiding traffic congestion and taking advantage of direct routings and shortcuts which would otherwise be impossible.

6. In terms or regional operational, blended winglets provide added flexibility to fleet operations and route selection.

7. Aesthetically pleasing – it basically looks good which gives the aircraft a modern feel and improves the customer’s perception of the airline.

8. Winglets lower drag and improve aerodynamic efficiency, thus reducing fuel burn; where blended winglets can improve cruise fuel mileage up to 6 percent depending on the range.

9. The addition of Aviation Partners Blended Winglets to the 737 Next Generations has demonstrated drag reductions of 5-7%.

Statistical Overview

Percentage increases in Table 3 below are in comparison between the same aircraft; one with winglets and one without.

TABLE 3 - STATISTICAL OVERVIEW OF FIXED WINGLETS

Item Value

Noise Reduction 6.5%

Emission Reduction 5%

Climb Thrust Reduction 3%

Cruise Thrust Reduction 6%

Fuel Saving 6%

Drag Reduction 5-7%

Purchase Price US$725,000

Implementation Cost US$25-80,000

Weight Increase 175-235 kg

6.1.3.2 Possible Retrofit - Morphing Winglets

Cant angle

These winglets have moveable fins located at the end of a wing where they appear as small upward extensions to a plane’s wings. Essentially, they are meant to reduce the drag experienced by the wing and hence reduce the thrust required. This of course will reduce fuel usage.

Research has also taken place on a winglet which is especially fixed at the cant angle (angle which is approximately 25o from the vertical). This unique location can provide a fuel reduction of between 3-5%.

Flexible Structures

Another modification that is yet to be made on the morphing winglets, which are at the moment in a structurally fixed position, is flexible and moveable winglets which can alter their direction by different angles. They can, for instance, be designed to change their angle when the plane is taking off, climbing, cruising or landing. These improvements are not only geared towards further reducing fuel consumption, but they also aim to minimize noise made by an aircraft during landing.

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Shape Memory Alloys (SMA’s)

Another slight modification that Boeing is considering, and has actually filed a patent application, is a winglet whose movement is facilitated using shape memory alloys (materials which change shape upon an electrical or thermal stimulus). This winglet is to be developed in a completely flattened position in order to give the wing extra lift even when the plane is moving at low speeds. The flattened position of the winglet also minimizes engine thrust and hence facilitating a quieter landing approach.

Experimental Data

Morphing winglet technology is currently undergoing testing in one of Continental Airline’s first winglet equipped aircrafts, the Boeing 757-300. It is postulated that this modification will allow the aircraft’s twin-engine stretched tapered body to operate on longer routes. Continental Airlines has also estimated that the modification could help cut down on the fuel cost of each aircraft by up to $164,000 per year. Presently, continental has included winglets on over 270 of its 757 and 737 fleet.

Application

The projected aircraft will be fitted with the Fixed Winglets at this stage. The morphing winglets will be a future retrofit in the event that they come into commercial use.

6.1.3.3 Wing Waggle

The wing waggle technology reduces mid-flight drag and turbulence by using minute air-powered jets which redirect air so that it flows in a new direction that is perpendicular to the direction of motion. When an aircraft moves, there are hundreds of thousands of jets of air circulating over the aircraft’s surface. Most of these jets of air are normally felt by passengers in the form of large scale airline turbulence. The turbulence experienced by the aircraft must be overcome by the engine if the aircraft is to takeoff and remain in flight. The process of overcoming turbulence creates a lot of drag which in turn burns extra fuel.

FIGURE 12 - WING WAGGLE TECHNOLOGY IN ACTION

Wing Waggle has proven that when an oscillation in airflow is created in a perpendicular direction it actually leads to a net reduction in the level of turbulence near the aircraft. Subsequently, the aircraft experiences minimum drag, leading to a significant reduction in the fuel consumption.

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Moreover, if air flow is moved in a direction that is perpendicular to the direction of travel, and if this movement is maintained back and forth, the possibilities of reducing drag are significantly increased. If the drag can be minimized by as little as 20%, there would be a subsequent reduction in fuel consumption by 20%, thus reducing emissions. The tiny air powered jet operates on Helmholtz resonance principle. According to this principle, “when air is forced into a cavity the pressure increases, which forces air out and sucks it back in again, causing an oscillation”.

Even though research in the development of the micro-jet systems is still in its conceptual stages, it is set to minimize skin fraction drag by up to 4%. Skin drag fraction has been found to be a major cause of mid-flight drag.

Feasibility

It is estimated that wing waggle technology upon completion will cost approximately US$723,000 and their installation will take an extra cost of about US$20-50,000. The installation process is likely to take a week. Once the waggle wings are fitted it is anticipated that they will add an extra 160-220kg to the aircraft’s weight. This will depend on whether the wings were installed at production or whether they were a retrofit. The fuel costs as a result of carrying this extra weight will automatically be offset by reduced fuel consumption over an increased range.

From these estimations Wing Waggle technology is likely to benefit aviation companies whose average sector length is long (more than one hour) as opposed to those whose average sector length is shorter. However, companies with shorter average sector length can use the waggle winged technology if their primary objective is to operate easily on obstacle limited runways and to reduce noise.

Statistical Overview

Percentage increases in Table 4 below are in comparison between the same aircraft; one with winglets and one without.

TABLE 4 - STATISTICAL OVERVIEW OF WING WAGGLE TECHNOLOGY

Item Value

Fuel Saving 20% (optimum)

Drag Reduction 20% (optimum)

Skin Friction Reduction 4%

Purchase Price US$723,000

Implementation Cost US$20-50,000

Weight Increase 160-220 kg

TABLE 5 - SUMMARY OF LIFT-TO-DRAG TECHNOLOGIES

Item Fixed Winglet [%] Wing Waggle [%] TOTAL [%]

Noise Reduction 6.5 - 6.5

Emission Reduction 5 - 5

Climb Thrust Reduction 3 - 3

Cruise Thrust Reduction 6 - 6

Fuel Saving 6 20 26

Drag Reduction (optimum) 7 20 27

Skin Friction Reduction - 4 4

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6.2 AIRCRAFT SYSTEMS

No-Bleed Air Systems

Saving Fuel and Enhancing Operational Efficiencies

In the design of our new environmentally friendly aircraft we will feature a unique systems architecture that will offer numerous advantages during operation. The primary differentiating factor in the systems architecture of new aircraft will be its emphasis on using electrical systems. The electrical systems will replace most of the pneumatic systems found on traditional commercial airplanes like our base model the B737NG. One of the main advantages of using No-Bleed electrical systems architecture is the greater efficiency which will be gained in terms of reduced fuel burn; a fuel savings of about 3% is expected. Other benefit of using this system is the increased operational efficiencies due to the advantages of electrical systems compared to pneumatic systems in terms of weight reduction and reduced lifecycle costs.

Reasons Behind The Move To A More Electric Airplane

Recent advances in technology will allow us to incorporate a new no-bleed systems architecture in our airplane which will eliminate the traditional pneumatic system and bleed manifold. This will allow the airplane to convert the power source of most functions formerly powered by bleed air to electric power (for example, the air-conditioning packs and wing anti-ice systems). The no-bleed systems architecture will offers a number of benefits, including:

1. Improved fuel consumption, due to a more efficient secondary power extraction, transfer, and usage.

2. Reduced maintenance costs, due to elimination of the maintenance-intensive bleed system.

3. Improved reliability due to the use of modern power electronics and fewer components in the engine installation.

4. Expanded range and reduced fuel consumption due to lower overall weight. 5. Reduced maintenance costs and improved reliability because the architecture uses

fewer parts than previous systems. 6. Overall greener and more efficient aircraft with reduced emissions and greenhouse

effects.

It is expected that the new no-bleed systems architecture will allow the airplane’s engines to produce thrust more efficiently because the high-speed air produced by the engines goes straight to thrust. Pneumatic systems that divert high-speed air from the engines rob conventional airplanes of some thrust and increase the engine’s fuel consumption. It is expect that by using electrical power it will be more efficient than engine-generated pneumatic power, and therefore the new architecture should extract as much as 35% less power from the engines. Conventional pneumatic systems generally develop more power than is needed in most conditions, causing excess energy to be dumped overboard.

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FIGURE 13 - NO BLEED SYSTEM

In Figure 13 above, the No-bleed electrical systems architecture which will be used in our aircraft can be seen. The electrical systems architecture is very similar to the new High-Tec electrical systems used in the Boeing 787s.

In our new airplanes systems design, virtually everything that has traditionally been powered by bleed-air from the engines will be transitioned to an electric architecture. The systems which will be electrified include:

1. Engine starting 2. Auxiliary power unit (APU) start 3. Wing ice protection 4. Powering the cabin environmental and pressurization control system 5. Driving the high-capacity hydraulic pumps

Though there will be at least one bleed system remaining; the anti-ice system for the engine inlets.

In this architecture, the power sources for the electrical system are engine-driven and auxiliary power unit (APU) driven generators, while the power sources for the hydraulic system are engine-driven and electric-motor driven hydraulic pumps. The engine-driven hydraulic power sources in the no-bleed architecture are similar to those in the traditional architecture.

In the no-bleed architecture, electrically driven compressors provide the cabin pressurization function, with fresh air brought onboard via dedicated cabin air inlets. This approach is significantly more efficient than the traditional bleed system because it avoids excessive energy extraction from engines with the associated energy waste by pre-coolers and modulating valves. There is no need to regulate down the supplied compressed air. Instead, the compressed air is produced by adjustable speed motor compressors at the required pressure without significant energy waste. This will result in significant improvements in engine fuel consumption.

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While much can be said regarding the efficiency gains achieved by changing the means of extracting power for airplane systems from the engines, the new airplanes no-bleed systems architecture will also bring with it some significant maintenance cost and reliability advantages as well. By eliminating the pneumatic systems from the airplane, it will boast a notable reduction in the mechanical complexity of airplane systems.

The list below highlights just a few of the components eliminated as a result of new electrical systems being used:

1. Pneumatic engine and APU start motors 2. APU load compressor 3. Pre-coolers 4. Various ducts, valves, and air control systems 5. Leak and overheat detection systems

Engines

In more traditional systems architecture, like our base model the B 737, the engines provide the majority of secondary airplane systems power needs in pneumatic form. In the new aircrafts engines system architecture, the engines will provide the majority of airplane systems power needs in electrical form via shaft-driven generators. The traditional airplane pneumatic bleed system architecture results in less than optimum engine efficiency. By eliminating the pneumatic bleed systems, the result will be more efficient engine operation due to reduced overall airplane level power requirements. This means the airplane does not draw as much horsepower off the engine in cruise, so it doesn’t burn as much fuel. The corresponding predicted improvement in fuel consumption, at cruise conditions, is in the range of about 2%.

Also, by using the new no-bleed system for the aircraft, it will allow significant simplification in engine build-up due to the elimination of the pneumatic system and associated pre-coolers, control valves, and required pneumatic ducting.

Hydraulic System

The hydraulic system to be used in our new aircraft systems architecture is similar to the one in the traditional architecture in the Boeing 737. There are three independent systems — left, centre and right that collectively support primary flight control actuators, landing gear actuation, nose gear steering, thrust reversers and leading/trailing edge flaps. The primary power source for the left and right systems are engine-driven pumps mounted on the engine gearbox. In addition, the left and right systems are each powered by an electric-motor driven hydraulic pump for peak demands and for ground operations.

The key difference between the traditional and new hydraulic system is the power source for the centre system. In the traditional architecture, the centre system is powered by two large air-turbine-driven hydraulic pumps, which operate at approximately 50 gallons/minute (gpm) at 3,000 pounds per square inch (psi) to meet peak hydraulic demands for landing gear actuation, high lift actuation and primary flight control during takeoff and landing. During the remainder of the flight, two small (approximately 6 gpm) electric-driven hydraulic pumps power the centre system.

In our new no-bleed system architecture, the centre hydraulic system is powered by two large electric-motor-driven hydraulic pumps. One of the pumps runs throughout the entire flight and the other pump runs only during takeoff and landing. By using 2 higher pressure pumps, the

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new hydraulic system will enable our airplane to use smaller hydraulic components, saving both space and weight.

Primary Flight Control Actuation System

The design, integration and certification support for the Primary Flight Control Actuation System for the newly designed passenger aircraft is based on the High-Tec system used on the Boeing 787. The ‘Moog system’ controls all the primary flight control surfaces on the airplane. The system will control all flight surfaces and will include a mix of electrohydraulic (EH) and electromechanical (EM) servoactuators and all associated control electronics.

The flight control system includes EH servoactuators with remote loop closure electronics for the ailerons, inboard and outboard spoilers, elevator and rudder. The horizontal stabilizer and in-board spoilers employ EM servoactuators with associated motor drive control.

Key features of the system include:

1. Smart actuators with on-board loop closure electronics 2. 5000 psi system operating pressure 3. High power motor controllers with active front ends 4. Advanced materials for weight-performance optimization

FIGURE 14 - ACTUATION SYSTEM

Fly-By-Wire System

Fly-by-wire systems will be used for the aircraft for the main systems as well as for the back-up systems which will be mentioned later in the report.

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The hydraulically powered flight control surfaces are tabulated in Table 6 below followed by a diagrammatical view of the areas consisting of fly-by-wire technology in :

TABLE 6 - HYDRAULICS PER SYSTEM

Electrical Control

System No. of Hydraulics

Elevators 2

Ailerons 2

Slats 10

Speed Brakes 6

Roll Spoilers 8

Tailplane Trim 1

Flaps 4

Lift Dumpers 10

FIGURE 15 - FLY-BY-WIRE TECHNOLOGY APPLICATIONS

The aircraft has three independent hydraulic power systems: blue, green and yellow. Figure 16 below shows how the three systems respectively power the hydraulic flight actuators. There will be a total of seven computers of three types to undertake the flight control computation task.

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System 1 – Elevator/Aileron Computers (ELAC): There are two of these computers to control the aileron and elevator actuators, as seen in Figure 16 below.

System 2 – Spoiler/Elevator Computers (SEC): There are three SEC’s which will control all of he spoilers and in addition provide secondary control to the elevator actuators. The various spoiler sections as seen in Figure 16 will have different functions for the following modes:

1. Ground spoiler mode 2. Speed brake mode 3. Load alleviation mode 4. Roll augmentation

System 3 - Flight Augmentation Computers (FAC’s): There are two FAC’s which will provide conventional yaw damper function, interfacing only with the yaw damper actuators.

In the highly unlikely event of the failure of all computers it will still be possible to fly and land the aircraft. In the case of this happening, the tailplane horizontal stabilizer (THS) and rudder sections would be controlled by mechanical trim tabs, which would then allow pitch and lateral control of the aircraft to be maintained.

FIGURE 16 - FLY-BY-WIRE SYSTEM

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Electrical Power Requirement

Our new aircrafts electrical system will be a hybrid voltage system consisting of various voltage types to cope with the no-bleed electrical system, where greater power demands are needed.

The voltage types include:

235 volts alternating current (VAC), 115 VAC, 28 volts direct current (VDC) and ±270 VDC.

The 115 VAC and 28 VDC voltage types are traditional, like on the Boeing 737, while the 235 VAC and the ±270 VDC voltage types are the consequence of the No-Bleed electrical architecture that results in a greatly expanded electrical system generating twice as much electricity as the base airplane, the Boeing 737.

The system will includes four generators: one per engine and two per APU operating at 235 VAC for reduced generator feeder weight. The system will also include ground power receptacles for airplane servicing on the ground without the use of the APU.

The generators will be directly connected to the engine gearboxes and therefore operate at a variable frequency (360 to 800 hertz) proportional to the engine speed. This type of generator is the simplest and the most efficient generation method because it does not include the complex constant speed drive, which is the key component of an integrated drive generator (IDG). As a result, the generators are expected to be more reliable, require less maintenance, and have lower spare costs than the traditional IDGs.

The electrical system will feature two electrical/electronics (E/E) bays, one forward and one aft, as well as a number of remote power distribution units (RPDU) for supporting airplane electrical equipment and motors. The system will save weight by reducing the size of power feeders. The majority of our airplanes electrical equipment will be either 115 VAC or 28 VDC, and will be supported by the forward E/E bay and RPDUs, (as shown schematically in Figure 17 below). The RPDUs are largely based on solid-state power controllers (SSPC) instead of the traditional thermal circuit breakers and relays. The system will also feature two forward 115 VAC external power receptacles to service the airplane on the ground without the APU and two aft 115 VAC external power receptacles for maintenance activities that require running the large-rated adjustable speed motors.

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FIGURE 17 - SYSTEMS ARCHITECTURE OF OUR AIRPLANE

Engine and APU Start

The projected aircrafts engine-start and APU-start functions will be performed by extensions of the method that has been successfully used for the APU in the Next-Generation 737 airplane family (our base model). In this method, the generators are run as synchronous starting motors with the starting process being controlled by start converters. The start converters provide conditioned electrical power (adjustable voltage and adjustable frequency) to the generators during the start for optimum start performance. Unlike the air turbine engine starters in the traditional architecture that are not used while the respective engines are not running, the start converters will be used after the respective engine is started. The engine and APU-start converters will function as the motor controller for cabin pressurization compressor motors.

Normally, both generators on the APU and both generators on the engine are used for optimum start performance. However, in case of a generator failure, the remaining generator may be used for engine starting but at a slower pace. For APU starting, only one generator is required.

The power source for APU starting may be the airplane battery, a ground power source or an engine-driven generator. The power source for engine starting may be the APU generators, engine-driven generators on the opposite side engine or two forward 115 VAC ground power sources. Also, the aft external power receptacles may be used for a faster start, if it is desired.

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Environmental Control System

In the projected aircrafts electrical architecture, the output of the cabin pressurization compressors will flow through low-pressure air-conditioning packs for improved efficiency. The adjustable speed feature of electrical motors will allow further optimization of airplane energy usage by not requiring excessive energy from the supplied compressed air and later regulating it down through modulating valves resulting in energy loss.

Avoiding the energy waste associated with down regulation results in improvements in engine fuel consumption, and the environmental-control system air inflow can be adjusted in accordance with the number of airplane occupants to achieve the lowest energy waste while meeting the air-flow requirements.

Wing Ice Protection

In the traditional architecture of the base model the Boeing 737, hot bleed air is extracted from the airplanes bleed system and distributed through the areas of the wing leading edge that need ice protection. Since the projected airplane will have traditional bleed-air systems eliminated, it will use an electro-thermal ice protection scheme, in which several heating blankets are bonded to the interior of the protected slat leading edges. The heating blankets can then be energized simultaneously for anti-icing protection or sequentially for de-icing protection to heat the wing leading edge. This method is significantly more efficient than the traditional system because no excess energy is exhausted. As a result, the required ice protection power usage is approximately half that of pneumatic systems. Also, because there are no-bleed air exhaust holes, airplane drag and community noise will be improved relative to the traditional Boeing 737 pneumatic ice protection system.

Auxiliary Power Unit

The APU provides an excellent illustration of the benefits of the more electric architecture. One of the primary functions of a conventional APU is driving a large pneumatic load compressor. By replacing the pneumatic load compressor with starter generators, (as seen in Figure 18 below), this will result in significantly improving the start reliability and power availability. The use of starter generators reduces maintenance requirements and increases reliability due to the simpler design and lower parts count. In terms of in-flight start reliability, the new APU which will be used on the projected airplane is expected to be approximately four times more reliable than conventional APU s with a pneumatic load compressor.

FIGURE 18 - ELIMINATED COMPONENTS OF THE APU SYSTEM WITH THE NO-BLEED ARCHITECTURE

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Therefore, the key benefits expected from the projected aircrafts no-bleed system architecture is improved fuel consumption as a result of more efficient engine cycle and more efficient secondary power extraction, power transfer, and energy usage. Eliminating the maintenance-intensive bleed system is also expected to reduce airplane maintenance needs and improve airplane reliability because there are fewer components on the engine installation. There are no IDGs, pneumatic ducts, pre-coolers, valves, duct burst protection and over-temperature protection and there is no compressed air from the APU, resulting in a simplified and more reliable APU.

The projected aircrafts No-Bleed architecture also features modern power electronics and motors that will provide increased overall reliability, decreased costs and improved performance. This will result in reduced airplane weight, reduced part count and simpler systems installation.

Brakes

Another innovative application of the more electric systems architecture on the projected airplane is that electric brakes will be used, replacing the conventional hydraulically actuated brakes used on the 737NG. Electric brakes will significantly reduce the mechanical complexity of the braking system and eliminate the potential for delays associated with leaking brake hydraulic fluid, leaking valves and other hydraulic failures. The electric brake system which will be used will be modular meaning four independent brake actuators per wheel will be used. This will significantly reduce performance penalties compared with dispatch of a hydraulic brake system with a failure present.

In general, electric systems are much easier to monitor for health and system status than hydraulic or pneumatic systems. The electric brakes will take full advantage of this. Continuous onboard monitoring of the brakes will provide airlines with a number of advantages, such as:

1. Fault detection and isolation 2. Electrical monitoring of brake wear 3. Ability to eliminate scheduled visual brake wear inspections 4. Extended parking times

The braking force applied of the new electric brakes will be able to be monitored even while parked, resulting in extended parking brake times through monitoring and automatically adjusting its parking brakes as the brakes cool.

Electrical Power Generation

Another fundamental architectural change possible on the projected airplane is the use of variable frequency electrical power and the integration of the engine generator and starter functions into a single unit. This change enables elimination of the constant speed drive (also known as the integrated drive generator, IDG), greatly reducing the complexity of the generator. In addition, by using the engine generator as the starter motor (an approach used with great success on the Next-Generation 737 APU), the airplane will be able to eliminate the pneumatic starter from the engine. This provides substantial weight reductions as well as easier and quicker installations as numerous units are combined into one.

By using the new starter generator on the projected airplane, it is expected to have a mean time between faults (MTBF) of 30,000 flight hours. A 300% reliability improvement compared to previous started generators used on the B737NG.

Landing Technology To Save Fuel – NextGen Transport System

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In traditional landing, pilots glide the plane through a series of predetermined elevations in their approach to the runway. This practice makes noise and uses more fuel as planes increase throttle to level off at each of the elevation intervals. With continuous descent, the airplane relies on GPS guidance to essentially coast on idle in a direct line from its cruising elevation to the tarmac.

Certification and Testing

The GPS system that makes continuous descent possible is part of the FAA's Next Generation Air Transportation System, nicknamed NextGen, which also enables planes to fly in straight lines rather than following the twisting paths of the current radar-based system (another fuel saver). In addition, NextGen allows planes to fly closer together, land more planes in quicker succession and be rerouted more quickly. With some 7,000 planes aloft over the USA at any one moment, the new system should also ease congestion and flight delays. The FAA has approved allowing continuous descent tests in Los Angeles, Phoenix, and Salt Lake City.

Costs

The GPS systems are everything but cheap. Every plane needs reprogramming of its onboard computers and other systems, putting the cost to refit some older planes at US$300,000. There are ground GPS units too, though at $200,000 per unit (several are needed for each airport). This may sound pricey but may be considered a bargain compared with million-dollar radar units.

While the FAA and trade groups estimate that NextGen could reduce fuel consumption by more than 10%, it is unlikely to be in place until 2025 due to costs and technology issues. Also, airplanes using continuous descent need not increase throttle. The noise output is 30% lower than that of a traditional landing.

Application

Although this aircraft will not be available until 5 years after the design deadline of 2020, it can be considered a future modification.

Aircraft In-Wheel Motors

An aerospace company has engineered an electric motor and generator system that eliminate the use of jet engines for taxiing and other ground manoeuvres. The technology will potentially save millions in fuel costs and reduce CO2 emission levels.

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FIGURE 19 - IN-WHEEL MOTORS

Much like hybrid vehicle systems that capture energy during braking and store it in batteries, this technology will capture the kinetic energy of the aircraft during a landing event. It does this by converting the kinetic energy into electrical power which may be stored and later used to power the wheel hub motor/generators within the aircraft wheels for taxiing and ground manoeuvres. This could potentially save millions of dollars per year per aircraft in reduced fuel burn.

The technology uses in-wheel electric motor/generators that are capable of producing sufficient power to effectively manoeuvre an aircraft of any weight on the ground. Not only that, they can provide safer and more effective braking of the landing gear wheels thus replacing the old friction disk technology. This technology will be implemented onto the projected aircraft.

FIGURE 20 - IN-WHEEL LANDING TECHNOLOGY

Magnetic Braking and Ground Manoeuvring System

Currently, when taxiing aircrafts it is a very slow process due to the consecutive pushing and pulling required by the management vehicles. This process can cause excessive wear and/or damage to the tug, the management vehicle and most important to the aircraft.

Advanced Magnetic Braking and Ground Manoeuvring System is a new technology for aircraft that replaces the current friction based braking system within wheel motor/generators and will be used on the projected aircraft. This technology is quite simple actually. Without

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going into too much detail, the magnetic braking system does not require any power input and is practically maintenance and fuel usage free. When part of the braking system, (the reaction plate) moves through a permanent magnet eddy currents are produced, which in turn produces forces that oppose the direction of the reaction plate. The braking capability is dependent on the velocity of the reaction plate going through the permanent magnet. This results in a very smooth and controlled deceleration.

As this method requires no power from the aircraft engines to decelerate the aircraft, than ground handling is basically emission free from the aircraft point of view. This reduces maintenance costs as well as increases aircraft life. Also, this design can be considered a failsafe design as it requires no electricity to run allowing for safer and more effective braking and ground handling.

Currently, jet and turbofan aircrafts require tow motors or tugs to push the aircraft backward into the designated taxiway. By eliminating the need to attach and detach tow motors or tugs, aircrafts can enter and exit gates much faster. This reduces between-flight turnaround times, which directly impacts airline and airport capacity utilization rate.

It's projected to provide fuel savings of US$2.4-3.4M a year per aircraft. The long terms savings are projected to be 3 to 4 times this value. The environmental benefits of the technology for airport neighborhoods will be substantial reduced due to the significant reduction in greenhouse gases.

Pre-Spinning Tires

During landing events the landing gear tires are pre-spun to match the relative ground speed thus eliminating the sliding friction tire wear. During normal landing the tires generate particulate pollution which is ingested by the following landing aircraft. This in turn affects the clean performance of the succeeding aircraft as it ingests the deposited rubber and releases it all over the runway. This excess rubber reduces the friction coefficients necessary when braking in rainy weather conditions making it very dangerous if not collected and disposed of. This technology will be implemented onto the projected aircraft.

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6.3 BACK-UP SYSTEMS

BUSS System

In the projected aircraft the back-up system which will be implemented is the Back-up speed scale. The "Backup Speed Scale" or BUSS is a tool which pilots use when speed indications cannot be used. To use the BUSS, the crew must first disconnect the three ADRs (air data reference - anemometric stations). Once these have been disconnected, the crew can no longer use them during the flight.

With the BUSS system, speed is no longer calculated by the pitot probes, but by the aircraft's incidence probes. The speed indication, which is less precise is presented in the form of green, amber and red stripes. In a high turbulence situation at high altitude, the speed indication given is very unstable and difficult to use.

If the pitot tube system malfunctions or the system is choked due to icing or tiny pieces of volcanic ash, the “Back-Up Speed Scale” system steps in, where both altitude and speed are measured by a global positioning system or GPS and an accurate speed figure is indicated on the speedometer. Furthermore, the pilot is warned with a red signal against any value lower or above the required speed. Thus, speed positioning is maintained and the flight continues without dolphining, in other words, without up and down movements.

The system, which was initially developed for the A380 model, has been modified to fit Airbus models A320 and A330/340 as well. The system is installed on older planes as they are serviced while it has become a standard feature for all Airbuses produced in the last six months.

The newly designed aircraft will feature various backup systems for each type of system operating during flight. Firstly, the APU generator may be used to supply primary power on ground and will serve as backup for either engine generator in flight. The APU generator is identical to the engine generators but has no generator drive unit since the APU itself is governed and will maintain a constant generator speed.

Fly-By-Wire Systems For Back-Up

A flight control system consists of the flight control surfaces, cockpit controls, connecting linkage, and necessary operating mechanisms to control the aircraft in flight.

Generally the cockpit controls are arranged like this:

1. Control yoke for roll (ailerons) 2. Control column for pitch (elevators) 3. Rudder pedals for yaw (rudder)

Flight Control Systems Classification

Mechanical - Mechanical flight control systems are the most basic design used. They were used in early aircraft designs and currently in small airplanes where the aerodynamic forces are not excessive. The flight control system uses a collection of mechanical parts such as rods, cables, pulleys and sometimes chains to transmit the forces of the cockpit controls to the control surfaces.

Hydro-Mechanical - Hydro-mechanical was introduced due to the drawbacks of a purely mechanical flight control system as its complex and weight of the flight control systems

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increases considerably with size and performance of the airplane. The use of hydraulic power overcomes these limitations and consequently aircraft size and performance are only limited by economics rather than technology. Hydro-Mechanical works in a way where the pilot commands will cause the mechanical circuit to signal the servo valves in the hydraulic circuit to power the appropriate actuators which will then move the appropriate control surfaces.

Types of Fly-By-Wire

Analogue Fly-By-Wire – Analogue fly-by-wire flight control systems dispenses all the complexity of the mechanical circuit of the hydro-mechanical flight control systems and replaces it with an electrical circuit. The cockpit controls now operate signal transducers which generate the appropriate commands. The commands are processed by an electronic controller. The autopilot is now part of the electronic controller.

Digital Fly-By-Wire – Digital is similar to its analogue counterpart, the main difference being that the signal processing is done by digital computers. This increases flexibility as the digital computers can receive input from any aircraft sensor. Digital fly-by-wire will be used for the aircraft due to the following reasons:

1. Digital fly-by-wire is safer than mechanical because of built-in redundancies 2. Less vulnerable to battle damage in military aircraft than old fashioned hydraulics 3. Digital fly-by-Wire is more maneuverable because computers generate adjustments

more frequently than manually by the pilot which means commercial flying is smoother and the travel experience more pleasurable with lower human error

4. Digital fly-by-wire is fuel efficient as the hardware is compact and lightweight

Advantages of Digital Fly-By-Wire System

1. Digital fly-by-wire systems reduce overall aircraft weight and fuel consumption 2. Weight savings in designing for new hydraulics, landing gear and mechanicals parts. 3. Less maintenance is required with fly-by-wire and there are no hydraulic systems to

lubricate, ‘oil’ to change, tension adjustments of cables etc 4. Weather prediction is greatly improved as is the reliability and accuracy 5. Automated peripherals are believed to have prevented hundreds of accidents and saved

many lives 6. Engine shutdowns have been greatly reduced 7. Advanced fighter designs that are inherently unstable could be thoroughly tested and

evaluated for possible production and combat capability 8. Pilot decisions that could expand problematic parameters (air frame stress, conditions

that might induce a stall) can be blocked by the computer system and autopilot 9. Pilot fatigue can be compensated

Disadvantages of Digital Fly-By-Wire System

1. Many long buses (or wires) to collect and distribute data and commands 2. Increased vulnerability to damage from a single hazardous event if it were to occur in or

near the avionics bay 3. Software changes are more difficult, where the impact of the change could effect a

relatively large part of the software. 4. Failure of a digital system could completely disable a digital device

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Future Upgrade Technology – Intelligent Flight Control System

A newer flight control system, called Intelligent Flight Control System is an extension of modern digital fly-by-wire flight control system. The aim of intelligent flight control system is to intelligently compensate for aircraft damage and failure during mid-flight, such as automatically using engine thrust and other avionics to compensate for severe failures such as loss of hydraulics, loss of rudder, loss of ailerons, loss of an engine, etc. Several demonstrations were made on a flight simulator where a Cessna-trained small-aircraft pilot successfully landed a heavily-damaged full-size jet operated by intelligent flight control system without prior experience with large-body jet aircraft. It is reported that intelligent flight control system is mostly a software upgrade to an existing fully computerized digital fly-by-wire flight control system.

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6.4 COCKPIT AVIONICS

RFP Requirement: RFP requirements state that advancements in navigational equipment is to be investigated.

The design of avionics control systems in commercial aircrafts cockpits are continually being upgraded and redesigned to improve overall use and efficiency. This ensures that controlling the avionics is more user friendly, increases pilots comfort and eliminate time consuming activities which in the long run reduces any serious problems during flight occurring. The latest technology on the market and still developing is the EFB (Electronic Flight Bag).

Electronic Flight Bag (EFB)

This new technology, which both Airbus and Boeing are very keen on investing is ideal for the projected aircraft and will be incorporated into the projected aircraft. EFB’s created to eliminate paper and remove many of the buttons in the cockpit but offer more capabilities such as airport moving maps (AMMs), satellite weather updates, electronic charting, and software compatibility with the Federal Aviation Administration’s (FAA's) Next Generation Air Transportation System (NextGen).

EFBs help to reduce a pilots workload, simplify flight operations and provide greater situational awareness. This in turn improves pilot decision making and enables the crew to easily connect in real-time with the rest of the world.

There are currently 3 classes of the EFB and in the design of the projected aircraft class 2 EFB’s will be used. Class 2 EFB is approved for operation in all phases of flight, it is portable, removable by the pilot without tools and has more software functionality than a Class 1 (such as moving maps, real time satellite weather updates, etc.).

The EFB has the ability to:

1. Provide airport moving maps 2. Engine performance applications 3. Report weather 4. present ADS-B (dependent surveillance broadcast) information

It also has valuable features including:

1. Open framework, enabling easy integration of third party software and data across multiple aircraft types and hardware platforms

2. Share information between crew members to reduce crew workload 3. Chart revisions through wireless or portable memory stick technology 4. Record air traffic control taxi instructions.

Since the electronic flight bags run with a computer behind the monitors and have all the properties and capabilities of a computer they can be upgraded and designed to run different software and applications.

The latest in software development for class 2 EFBs which will feature in the projected aircraft are:

1. The ability to interface with aircraft systems to allow for the deployment of AMM applications with own-ship position shown,

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2. The capability to provide an access point to multiple forms of wireless communications pipelines for text and/or data transfer. Satellite communications for anytime connectivity or Wi-Fi or cellular (3G) for on-the-ground connectivity.

In the future, the deployment of the wireless capability will create an extremely independent way to have on-the-ground connectivity that is separate and apart from the airport infrastructure. This connectivity will give pilots a way to perform their briefings and pre-flight tasks right from the cockpit. And it will also give the crews the additional capability of being able to implement last minute changes all without leaving the flight deck. This helps to save valuable time between all the crews, pilot and ground support operators.

FIGURE 21 - A FUTURE LOOK IN THE DESIGN OF AVIONICS IN THE COCKPIT

The projected aircraft will implement the use of EFB technology and the use of LCD screens and computers in the cockpit. Cockpit design will consist of a liquid crystal display (LCD) and a based multifunction control display unit (MCDU) for improved brightness and readability.

Other technological upgrades include an FMS landing system (FLS) feature, which presents Instrument Landing System (ILS)-like vertical and lateral guidance on the primary flight display using data from GPS, VOR, DME and inertial reference system. This will enable autonomous non-precision approaches. The new design will also support precision global navigation satellite system (GNSS) landing system (GLS) approaches, provided that local area augmentation systems are in place. It will adapt the air data inertial reference unit, dual integrated standby instruments, and integrated surveillance system (ISS).

Also, the information which is displayed on the electronic flight bag's (EFB's) airport moving map helps the pilot maintain situational awareness. The aircraft also will support an "auto pull-up" function, based on the terrain awareness warning system (TAWS). If the pilot does not respond to a TAWS alert, the autopilot will take the lead and manage the recovery of the situation automatically.

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Projected Aircraft’s Avionics Summary:

1. Electronic Flight Bag control system – Class 2 (with load alleviation system) 2. Main displays (6-by-6-inch, video-capable) 3. New design light management system 4. VHF data radios (VDL-2) 5. Integrated surveillance system:

o Weather radar o Terrain awareness warning system o Traffic alert collision avoidance system o Mode S transponder

6. Vertical display on navigator screen 7. Air data inertial reference unit 8. Dual integrated standby instrument system (ISIS), 9. Dual, LCD-based, head-up displays 10. Digital radio altimeter 11. Onboard information terminal (OIT) 12. Airport navigation system

Figure 22 - Future cockpit geometry with implemented EFB system and LCD touch screen

FIGURE 23 - COCKPIT DISPLAYS

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Conclusion

As a consequence of the overall reduction in complexity, airlines operating the projected airplane will experience reduced airplane-level maintenance costs and improved airplane-level dispatch reliability.

By moving to electric systems on the airplane, it is expected to cut about a third of the schedule interrupts compared to a B737 for the systems affected by the no-bleed/more-electric architecture. Other benefits include improved health monitoring, greater fault tolerance, and better potential for future technology improvements.

The key benefit expected from using a no-bleed electrical systems architecture is improved fuel consumption as a result of more efficient engine cycle and more efficient secondary power extraction, power transfer, and energy usage.

Eliminating the maintenance-intensive bleed system is also expected to reduce airplane maintenance needs and improve the airplanes reliability because there will be fewer components on the engine installation. There will be no integrated drive generators, pneumatic ducts, pre-coolers, valves, duct burst protection, and over-temperature protection; and there will be no compressed air from the APU, resulting in a simplified and more reliable APU.

With the use of the Class 2 Electronic Flight Bags system, the design of the avionics in the cockpit will help to reduce pilot workloads, simplify flight operations, and provide greater situational awareness which in turn improves pilot decision making and enables the crew to easily connect in real-time with the rest of the world.

The new to be used no-bleed architecture and Electronic flight bag system also features modern power electronics and motors that will provide increased overall reliability, decreased costs, and improved performance. Finally, the architecture means reduced airplane weight, reduced part count and simpler systems installation.

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6.5 MATERIAL SELECTION – COMPOSITES, ALUMINIUM AND TITANIUM

RFP Requirement: An investigation into the use of light-weight composites is necessary and the structural and material properties are to be analyzed.

In achieving an environmentally friendly, “green” passenger jet, the reduction of emissions and reduced fuel burn is necessary. One of the most effective ways in achieving this is a reduction is in aircraft weight. Therefore, in selecting the materials which are to be used in the new passenger jet, the reduction of weight whilst still achieving full structural integrity is the main factor considered.

There are a number of important properties with materials which need to be considered depending on the application. Properties to be considered include yield and ultimate strength, stiffness, density, fracture toughness, fatigue crack resistance, creep, corrosion resistance, temperature limits, ease of production, cost and repair. This section will lay out the materials which will be most suited to each aircraft component with the reasons for choice discussed.

6.5.1 Composites

The aerospace industry and manufacturers’ unrelenting passion to enhance the performance of commercial and military aircraft is constantly driving the development of improved high performance structural materials. Composite materials are one such class of materials that play a significant role in current and future aerospace components. Composite materials are particularly attractive to aviation and aerospace applications because of their exceptional strength, stiffness-to-density ratios and superior physical properties. For this reason, throughout our new passenger aircraft there will be high percentage of composite materials used.

A composite material typically consists of relatively strong, stiff fibres in a tough resin matrix. Man-made composite materials, used in the aerospace and other industries, are carbon and glass fibre reinforced plastic (CFRP and GFRP respectively) which consist of carbon and glass fibres. Both of which are stiff and strong (for their density), but brittle in a polymer matrix, which is tough but neither particularly stiff nor strong. By combining materials with complementary properties in this way, a composite material with most or all of the benefits (high strength, stiffness, toughness and low density) is obtained with few or none of the weaknesses of the individual component.

Advantages

There are many advantages to using composites. Below is a description of the major advantages of using composite in passenger aircraft;

Light-Weight

Composite materials are extremely light weight when compared to metals such as steel and aluminium. This reduced total structural weight leads to a fuel reduction and a more environmentally friendly aircraft to operate. Composites structures can be 20-80% lighter than steel or aluminium structures of the same design. Refer to the table below for a comparison of the density of materials used on passenger aircraft.

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FIGURE 24 - AIRCRAFT MATERIAL DENSITY COMPARISON

Light-Weight – Sandwich Structure

Sandwich composite materials are even lighter than laminates due to the use of a foam core inserted in between composite laminate plys as seen in the figure below. In passenger airliners the core material used in sandwich composites is usually PVC foam or phenolic honeycomb. The densities of these core materials are; PVC foam core is less than 0.15 g/cm3 and phenolic honeycomb is less than 0.10 g/cm3.

FIGURE 25 - TYPICAL SANDWICH COMPOSITE LAY-UP

Typical Weight Savings

Typically, the weight saving achieved by replacing an aluminium vertical fin box & inboard aileron with a carbon/epoxy composite, results in a weight reduction of 28%. Also, by substituting the aluminium parts with carbon/epoxy composites, there will be fewer assemblies, parts & fasteners needed.

High Stiffness (Modulus)

Composite materials possess a high Young’s modulus which allows them to be suitable for use in passenger aircraft structural components. When compared to aerospace materials, composites achieve a similar or higher Young’s modulus with the bonus of extra light weight.

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FIGURE 26 - TENSILE MODULUS OF DIFFERENT MATERIALS

High Strength (Strength)

The tensile strength of composites used in aircraft is also higher than many aerospace metals as seen in the figure below.

FIGURE 27 - TENSILE STRENGTH OF DIFFERENT MATERIALS

In the stiffness and strength graphs above, the lowest properties for each material are associated with simple manufacturing processes and material forms (e.g. spray lay-up glass fibre) and the higher properties are associated with higher technology manufacture (e.g. autoclave moulding of unidirectional glass fibre prepreg) such as would be found in the aerospace industry.

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For the other materials shown, a range of strength and stiffness (modulus) figures is also given to indicate the spread of properties associated with different alloys.

The above figures clearly show the range of properties that different composite materials can display. These properties can best be summed up as high strengths and stiffness’s combined with low densities. It is these properties that give rise to the characteristic high strength and stiffness to weight ratios that make composite structures ideal for so many applications. This is particularly true of applications which involve propulsion and movement such as passenger aircraft since lighter structures result in the aircraft operating with much more efficiency.

TABLE 7 – MATERIAL PROPERTIES OF DIFFERENT COMPOSITES

Material Density (g/cm3)

Tensile Strength

(GPa)

Specific Tensile

Strength

Tensile Modulus

(GPa)

Specific Tensile

Modulus

Graphite/epoxy (type 1)

1.6 0.93 0.58 213 133

Graphite/epoxy (type 2)

1.5 1.62 1.01 148 92

Aramid/epoxy 1.45 1.38 0.95 58 40

Glass/epoxy 1.9 1.31 0.69 41 22

Mild Steel 7.8 0.99 0.13 207 27

Aluminium Alloy 2.8 0.46 0.17 72 26

Titanium 4.5 0.93 0.21 110 24

Fatigue Resistance

Composite materials have varied fatigue resistance properties ranging from good to excellent. In the case of aerospace grade composites, they achieve outstanding tensile fatigue resistance due to the supreme properties of these materials and the high technology manufacturing techniques used. This is a major reason why composites are used in passenger aircraft.

From the graph below, it can be seen that the fatigue resistance for graphite (carbon)-epoxy composites is exceptionally good, especially when compared to the most used aerospace grade aluminium, 2024-T3.

FIGURE 28 - FATIGUE RESISTANCE OF DIFFERENT MATERIALS

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Tailored Design

The load-bearing capacity of composite structures can be tailored to meet the desired engineering requirements due to the stress and loads involved in particular areas of the structure. By using composite materials, it is possible to align the fibres along principle load directions to maximise stiffness and load-bearing capacity. For example, an aircrafts wing structure strength can be increased through the various ply directions used in specific locations to combat different loads and stresses. The figure below shows the potential of weight saving through the use of composites for a wing.

FIGURE 29 - COMPOSITE WING STRUCTURE

Low Tooling Costs

The costs involved in making moulds and other tools for fabricating composite structures can be much cheaper than fabricating machinery for metals. To manufacture composite a simple mould can be used which can be made from virtually any material. This is much cheaper when compared to the manufacture of metals where expensive stamp presses and extruders are often used.

Complex Shapes

When manufacturing composite materials, most reinforcing fabrics can be easily formed into complex shapes and structures. This not only reduces the number of parts making up a given component, but also reduces the need for numerous fasteners and joints. This is suitable for passenger aircrafts as there are numerous complex shaped structures used throughout the body of the aircraft.

Pre-pregs, which are extensively used in these aircrafts, can be easily formed into the desired shapes prior to consolidation and curing. The benefit of this is that the amount of post-fabrication machining will be significantly reduced. Composites also have good dimensional stability, even if the composite is manufactured into a complex structure and put under various loading scenarios.

Corrosion Resistance

One of the most impressive benefits of using composites in passenger airliners is their outstanding resistance to seawater & chemical corrosion. Aircraft materials are constantly

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exposed to atmospheric moisture as well as internal and external sources such as salt-water spray, aircraft fuel, oils, hydraulic fluids, battery acid, engine exhaust products and even leaking toilets. Therefore using composites will keep the aircraft flying longer and more safely without any problems from corrosion.

Composites are also not prone to stress corrosion cracking which unlike aerospace aluminium alloys is a considerable problem. Therefore, the higher the use of composites within an aircraft structure, the less maintenance and repairs needed due to corrosion of the material. This however, does not mean that 100% composites results in a perfect aircraft design. It simply means that maintenance and repairs will be significantly reduced.

Low Temperature (Cryogenic) Properties

The mechanical properties of composites are not affected when exposed to low temperatures. Passenger airliners benefit from this due to the fact that they fly at high altitude and freezing cold temperatures. On the other hand, the toughness of many steels and even some aluminium’s is severely reduced at freezing temperatures ranging from below -20o to -40oC.

Good Vibration Damping

Composite materials display good vibration and noise damping properties due to the low modulus of core materials.

Disadvantages:

Costs

One of the main disadvantages of composite usage is the costs. The raw materials necessary composite use is much higher than aluminium alloy as seen below.

Note: This is a very approximate cost of composites & some metallic materials.

1. Carbon fibre composite: US$75 to 200 per kg 2. Glass fibre composites: US$1.5 to 10 per kg 3. Boron fibre composites: More than US$800-1000 per kg 4. Aluminium alloys: US$2 to 10 per kg 5. Titanium alloys: US$40 to 200 per kg 6. Steel: US$1 to 10 per kg

Fabrication Processes

The manufacturing and fabrication process of composite materials can be expensive despite the relatively low tooling costs. When producing a high quality composite, the fabrication process is slow. Processes such as wet hand lay-up, resin transfer molding and autoclaving may take several hours to perform because of the time needed to manually lay-up the plies and cure the composite. These days, with new technology, automated composite lay-up, resin infusion and cure can be performed to produce high quality composites.

Anisotropic Material Properties

The mechanical & electrical properties of composites are anisotropic. Therefore properties such as Young’s modulus, strength, fracture toughness and fatigue resistance are dependent on the direction of the applied load relative to the orientation of the fibres. As a result of this, the properties in the through-thickness direction are much lower than the properties in the fibre direction.

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FIGURE 30 - COMPOSITE FIBRE CONSTRUCTION

FIGURE 31 - TENSILE STRENGTH

FIGURE 32 - COMPRESSIVE STRENGTH

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Poor Impact Damage Tolerance

Composite materials are susceptible to impact & bird strike damage. Impact from any foreign object causes delaminations & other problematic damage which reduces the composites mechanical properties. Often this damage is extensive but “invisible” (also known as Barely Visible Impact Damage (BVID)) due to the damage being done to the inner fibres rather than the outer (visible layer) which is a big problem for aerospace conditions.

Repair

Repair of composite structures may be a problem for our projected design. This is one of the many reasons that extensive use is trying to be avoided. Repair for composites is not as simple as removing the section and applying a new one like that of aluminium structural repair. Damage for composites is very difficult to find as damage is usually invisible.

Barely visible impact damage (BVID) is one of the many issues plaguing composite use. The fibres within the resin sometimes fail before any signs of structural failure. This is however advantageous for us as the whole structure must fail (resin and majority of fibres) before an entire system failure. Aluminium does not have this luxury. However, repairing invisible composite fibres to reduce the chances of failure is still a major problem.

Recycling

The recycling process of composites is still being researched. Simply melting the composites like aluminium releases very harmful pollutants and emissions, thus defeating the purpose of an “environmentally friendly aircraft design”.

Environmental Effects of Composites

Below is a detailed study on the environmental effects of manufacturing composites. Also included are techniques which can be applied to reduce the pollution and environmental effects in the manufacturing process. These techniques can be applied in our aircraft design to help reduce pollution in the initial stages before the aircraft even takes flight.

Emissions During Composite Manufacturing

The monomers used in composites manufacturing allow for the chemical cross linking of polymers in the resin. This is what causes the resin to become a strong and durable product.

Unfortunately, any time that uncured resin is exposed to the air, some monomer may evaporate. Once the curing stage is complete, essentially all of the monomers have reacted, and there will be no more evaporation.

Monomers Used And Emitted

Monomers are a class of chemicals that can react with each other to form long chains or that can react with certain polymers to cross-link them together. It is this cross-linking action of monomers that make them essential in composites manufacturing. While there are many different chemicals that are used as monomers, only a few are used in composites and only one is used extensively which is styrene.

Styrene is a clear liquid with a distinctive odour. Styrene is used so extensively because it can be mixed with a wide variety of polymers to form workable resins. It reacts readily and predictably to cross-link the polymers to form solid products and because it imparts important chemical

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and physical properties to the products. Styrene is also a major industrial chemical and so is lower in price than other monomers.

Another monomer that is occasionally used in composites, sometimes in the outer skin of aerospace applications, is methyl methacrylate. It is used for products needing better resistance to sunlight and for applications that require low smoke generation if exposed to fire. However, methyl methacrylate has a higher evaporation rate than styrene and it is much more expensive than styrene.

Since styrene is more widely used and emitted by composites manufacturing operations at a much higher rate than the other monomers, the remainder of this report will address only styrene emissions.

Information on Styrene

Styrene is a clear, colorless liquid that is a component of materials used to make thousands of everyday products for home, school, work, and play.

Styrene is derived from petroleum and natural gas by-products. Styrene helps create thousands of remarkably strong and lightweight composite products, representing a vital part of aerospace applications.

Levels of Styrene Emitted

The amount of styrene emitted by a composites manufacturing facility is determined by the following factors:

1. What manufacturing process and resin application technique is used 2. How much resin is processed 3. What the styrene content of the resin is 4. What pollution prevention or capture-and-control technologies are used

The most accurate way to determine how much styrene is emitted by a facility is to measure it. There are various procedures for measuring emissions, however, direct emission measurements (called "stack tests") are expensive, difficult to reproduce and are seldom used for composites manufacturing operations. Instead of measuring emissions, most composites manufacturers and regulators use emission factors to estimate emissions. In general, emission factors are formulas that produce an emission estimate based on information about a manufacturing operation, like styrene content in the resin, use of pollution prevention techniques or capture-and-control, amount of resin processed, etc. Different emission factors are available for several of the different composites manufacturing processes.

For example:

250,000 35 87,500

87,500 5.5 ( ) 4,813

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TABLE 8 - EMISSION FACTORS

Production Type Emission factor [%] Average [%] Mass [kg] Emission [kg]

Open-Moulding 8-18% 13 87500 11375

Closed-Moulding 1-3% 2 87500 1750

Casting 1-3% 2 87500 1750

Pultrusion 4-7% 5.5 87500 4812.5

Continuous Lamination 4-7% 5.5 87500 4812.5

Health of Workers Exposed To Styrene The health of workers in plants making or using styrene has been monitored for many years. Studies looking for long-term health effects related to styrene exposure have examined health records of over 50,000 workers exposed to styrene, going back nearly 50 years. These studies have not shown any statistically significant increases in long-term health problems of any kind attributable to styrene in these workers.

In most industrialized countries, there are strict regulations protecting worker health. In 1989, a safe exposure standard for styrene of 50 parts per million (ppm) over an eight-hour day was established. In years past, before effective monitoring systems were available, worker exposure to styrene (as well as other materials) often was greater than current exposure levels.

Minimization of Emissions

Two options can be considered when attempting to reduce the emissions from any manufacturing process:

Pollution Prevention

Engineering the manufacturing process so that it emits less is called "pollution prevention." As the name implies, with pollution prevention the manufacturing process is modified so that it generates fewer emissions. Pollution prevention technologies can have a number of benefits: reducing emissions at the process is often less expensive than capturing and controlling them later; pollution prevention can reduce occupational exposures in addition to reducing environmental releases.

However, pollution prevention technologies can have one major limitation. Since pollution prevention requires changing the manufacturing process, the feasibility of it will depend on whether the modified process still makes good products. Many pollution prevention technologies are widely used throughout the composites industry, but all have their limitations. They all change the properties of the products being made. Whether these changes in properties can be tolerated will depend on the nature of the products and their ultimate use.

Capture and Control

"Capture-and-Control" technologies, on the other hand, do not interfere with the manufacturing process. In theory, capture-and-control means that you simply capture whatever happens to be emitted from a manufacturing operation and then you direct these emissions to a device that converts them to some innocuous, non-polluting substance. However, these technologies can be very expensive to install and operate, they can consume large amounts of natural gas (a non-renewable natural resource) and they can have significant secondary environmental impacts, such as the generation of new pollutants.

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Low Styrene Resins

One way for some moulders to minimize the loss of styrene from their process is to use resins formulated with less styrene. Resin suppliers offer a variety of these resins.

While the use of low-styrene resin sounds as if it should be relatively simple, this technology is not for everyone. First, for many moulding processes, such as pultrusion and compression moulding, reducing the monomer content of a resin may not result in significant reductions in emissions. Products moulded using low-styrene resins will have different chemical and physical properties. Changing the amount of styrene in a resin, or changing the type of monomer used, will change the properties of the product, sometimes unacceptably.

Fillers

A second way to reduce styrene emissions is to use less resin. This might mean engineering a product so that it uses only as much resin as needed to provide the required properties, but more often this approach means adding an inert filler to the formula.

Fillers reduce the amount of resin needed (thereby reducing monomer usage and consequently emissions), while possibly increasing product stiffness or fire resistance (depending on the type of filler used).

Low Emitting Resin Application Technologies

While the emission reduction technologies discussed above work by decreasing the amount of styrene used to make a product, low emitting resin technology works by reducing the contact of resin with air.

For example, the various non-atomized application technologies like pressure-fed rollers and flow-applicators (also called flow coaters) work to reduce emissions by minimizing the contact of resin to air. In traditional spray application of resin, very small "atomized" droplets of resin travel from the spray gun to the mould or to the floor, walls or filters of the spray booth creating very high surface areas of wet resin and consequently high emissions.

Another low-emitting resin application technology is resin impregnation where glass mats are mechanically fed through a resin bath and then applied to the mould. Again, there is no spray (or stream) of resin through the air and more of the resin is delivered to the mould.

These technologies have their limitations. None can practically be used for gel coat, since this material needs to be applied in a very uniform layer that can be accomplished only with atomized spray.

Suppressants

A second way of reducing contact of resin with air is through the use of suppressants. These materials, which are typically waxes, form a layer on the surface of the resin while it is curing. This layer reduces the evaporation of monomer. Suppressants work only when the resin is left undisturbed, so that the layer of wax can be formed on the surface.

Controlled spray

For the atomized spray application of resin or gel coat, the use of a controlled spray program is one of the most effective ways of reducing the surface area of the wet resin, thereby reducing contact of resin with air, and consequently reducing emissions. A controlled spray program consists of three components:

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1. Proper calibration of spray guns, so that the lowest tip pressure that still delivers an adequate fan pattern is used (higher pressures result in much more "misting" of resin and evaporation of monomer)

2. Training operators to handle spray guns and apply resin using techniques designed to minimize overspray

3. Use of containment flanges at mould perimeters, which prevent the deposition of high surface area resin layers on the floor and walls.

Airflow Management

Air has a limited ability to hold monomer vapor and once this limit is reached no more monomer will evaporate, unless the monomer-laden air is replaced with fresh air. So, in theory, allowing a monomer-laden layer of air to remain at the resin surface can reduce monomer evaporation. In practice, airflow must be reduced to very low levels to achieve any real impact on monomer evaporation. Tests have shown that reducing airflow across wet open moulds does not significantly reduce emissions in open molding, probably because airflow cannot be practically reduced enough over an open mould.

Closed Moulding

Closed molding is a family of composites molding processes where liquid resin is not exposed to the air. This group includes resin transfer molding, resin infusion molding, compression molding and injection molding.

These technologies can have very low emissions. However, process economies often severely limit the application of these technologies. Compression molding is feasible only when large production runs are needed (say 50,000 parts per year) and where the capital is available to pay for the large presses and complex molds.

Resin infusion and resin transfer molding are sometimes useful and economically feasible for products made with open molding. These processes find use when customers demand and are willing to pay for higher quality products and where high volume production is not needed.

Oxidation

Oxidation is the most common capture-and-control technology used by composites manufacturing facilities. While there are different types of oxidizers available, they all use heat to convert monomer vapor to carbon dioxide (CO2) and water. While CO2 is not considered toxic, it does increase the retention of heat by the atmosphere and contributes to global warming.

In addition, most of the oxidizers consume large amounts of natural gas. Burning natural gas creates nitrogen oxides, one of the primary constituents of smog.

Application of Composites on Past Designs

In the RFP, the design team is required to design a narrow-body passenger aircraft which is due to replace the Airbus A320 or Boeing 737 jet planes.

Therefore research on passenger aircraft which are currently in use was necessary to determine where available use of composites are and a description of possible weight savings achieved.

The use of composite materials in commercial transport aircraft is increasingly as seen in the figure below due to their attractive outcome in reducing airframe weight. This in turn enables better fuel economy and therefore lowers operating costs. The first significant use of composite

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material in a commercial aircraft was by Airbus in 1983 in the rudder of the A300 and A310, and then in 1985 in the vertical tail fin. In the latter case, the 2,000 parts (excluding fasteners) of the metal fin was reduced to fewer than 100 for the composite fin, lowering its weight and production cost. Later, a honeycomb core with CFRP faceplates was used for the elevator of the A310.

Following these successes, composite materials were used for the entire tail structure of the A320, which also featured composite fuselage belly skins, fin/fuselage fairings, fixed leading- and trailing-edge bottom access panels and deflectors, trailing-edge flaps and flap-track fairings, spoilers, ailerons, wheel doors, main gear leg fairing doors, and nacelles. In addition, the floor panels were made of GFRP. In total, composites constitute 28% of the weight of the A320 airframe, as seen in Figure 33 below.

FIGURE 33 - COMPOSITE MATERIALS APPLICATIONS ON AIRBUS A320

The A340-500 and 600 feature additional composite structures, including the rear pressure bulkhead, the keel beam and some of the fixed leading edge of the wing. The last is particularly significant as it constitutes the first large-scale use of a thermoplastic matrix composite component on a commercial transport aircraft. Composites enabled a 20% saving in weight along with a lower production time and improved damage tolerance.

The A380 is about 20-22% composites by weight and also makes extensive use of GLARE (glass-fibre-reinforced aluminium alloy), which features in the front fairing, upper fuselage shells, crown and side panels and the upper sections of the forward and aft upper fuselage. GLARE laminates are made up of four or more 0.38 mm (0.015 in) thick sheets of aluminium alloy and glass fibre resin bond film. GLARE offers weight savings of 15-30% over aluminium alloy along with very good fatigue resistance. The top and bottom skin panels of the A380 and the front, centre and rear spars contain CFRP, which is also used for the rear pressure bulkhead, the upper deck floor beams, and for the ailerons, spoilers and outer flaps. The belly fairing consists of about 100 composite honeycomb panels.

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The Boeing 777, whose maiden flight was about 15 years ago, is around 20% composites by weight, with composite materials being used for the wing’s fixed leading edge, the trailing-edge panels, the flaps and flaperons, the spoilers and the outboard aileron. They are also used for the floor beams, the wing-to-body fairing, and the landing-gear doors. Using composite materials for the empennage saves approximately 1,500 lbs in weight.

FIGURE 34 - COMPOSITE MATERIALS APPLICATIONS ON BOEING 777

The latest Boeing development in passenger liners is the Boeing 787 Dreamliner. This highly efficient wide-body passenger aircraft will leverage extensive use of composite materials roughly 50% by weight, which gives the aircraft very high fuel efficiency and performance with reduced weight. The development of the Boeing 787 achieves a fuel reduction of 20% with an increase life cycle due to the low maintenance need of the composite materials.

FIGURE 35 - COMPOSITE MATERIALS APPLICATIONS ON BOEING 787

In terms of a single aisle aircraft, the benefits of using a 50% plus composite design is just not feasible unlike a larger two aisle aircraft like the Boeing 787. This is due to the fact that overall

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benefits of large amounts of composites used in a single aisle aircraft is not enough to outweigh the increased manufacturing cost of composites and problems which may occur during operation.

FIGURE 36 - GROWTH OF COMPOSITE USE IN AIRCRAFT DESIGN AS A PERCENTAGE OF WEIGHT

Therefore the decision to use a select amount of composite materials, around 30-35% by weight as seen in Figure 36 is justifiable since they will provide the aircraft with enough weight-saving to increase the aircrafts efficiency while still ensuring a safe reliable passenger aircraft in all flight conditions over its intended life.

This value has indeed increased from 15% as mentioned in the feasibility report to the now revised 30-35%. This new value is justifiable due to the increase in knowledge of this area.

Application of Composites on Projected Aircraft

Fibreglass

Centre & outboard flap track fairings, fin/fuselage fairings, the wing-to-body fairings floor panels and floor beams,l

Kevlar

Radome, part of the engine fan cowls, inboard track faring (behind engine), nose gear doors.

Carbon-Epoxy (laminate and sandwich composite)

Wings, including fixed trailing-edges, trailing-edge flaps, spoilers, ailerons, thrust reverser cowls, winglets, front, centre and rear spars. Also the undercarriage main and centre landing gear doors and fairings, fuselage belly skins, Pylon fairings and Nacelles cowlings, Central torsion box, rear pressure bulkhead and the keel beam. The empennage section including vertical and horizontal stabilizer, rudder, elevators, tail cone, dorsal of vertical stabilizer will be carbon-epoxy with the use of composite honeycomb core with carbon-epoxy faceplates for the

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elevator and other control surfaces. The belly fairing will also consist of carbon faceplate composite honeycomb panels. Lastly, the aft section of the fuselage skin will be made from carbon-epoxy composites.

Engines

The use of composites in engines in high-load components such as compressor and main fan blades has been proven successful in the General Electric GENx turbofan (Boeing 747-8 and 7). Utilizing composites in the engine fan blades, containment casing and cowling will also be sued.

FIGURE 37 - INTERNAL AND EXTERNAL COMPONENTS OF COMPOSITE STRUCTURES

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FIGURE 38 - MATERIAL SELECTION ON PROJECTED AIRCRAFT (SIDE VIEW)

FIGURE 39 - MATERIAL SELECTION ON PROPOSED AIRCRAFT (TOP VIEW)

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6.5.2 Aluminium

Aluminium alloys are widely used in the construction of aircraft structures, such as wings and fuselages due to their superior metal fatigue resistance and low weight (when compared to most other metallic materials). The strength-to-weight ratio of aluminium alloys is proficient enough to be capable of supporting the aircraft under all loading conditions. With its high strength and excellent fatigue resistance, it is used to advantage on structures and components on the aircraft with the knowledge that they won’t fail during flight. Aluminium alloy is excellent for aircraft structures which are under tension such as the bottom of wings.

In cases where aluminium alloy components are damaged during flight or when on the ground, the repair and replacement of the parts is relatively cheap and easy. Most basic aircraft maintenance workshops would have the necessary equipment and tools needed to achieve a successful repair to an aluminium alloy component.

Aluminium alloys have been widely used in the aerospace industry since the early 1900’s and due to this there is an adequate amount of information on their structural properties and loading cases in all flight conditions. With this information it is safe to say that an aluminium alloyed aircraft can fly safely and sufficiently for passenger aircraft.

Aluminium alloyed structures and components are readily available and in constant production around the world. This is due to aluminium being the most widely used material for the production of aircraft, especially passenger airliners. The price of manufacturing the aluminium then producing specific parts is relatively cheap, especially when compared to the price of composite materials. Since aluminium is widely available and recyclable, it is much easier to produce a good component in a shorter amount of time.

Aluminium alloys have many advantages when using them for aerospace applications. In terms of cost, they are one of the cheapest aerospace materials which are widely available in production of aircraft. Due to their good strength-to-weight ratio and fatigue resistance, aluminium alloys will be continued to be used throughout aircraft structures. When disposing of unused or retired aluminium alloyed aerospace structures, it is easily done by recycling the components by melting and re-using the aluminium and alloys for other cheap products like aluminium cans or even reused on new aircrafts.

During flight, when the aircraft is subjected to poor weather conditions like thunderstorms, with the use of aluminium alloys the aircraft will be protected from server lighting strike. Unlike composite materials, the aluminium alloys will dampen the lightning strike leaving minimal damage.

Application of Aluminium

The projected aircraft is a single aisle, narrow body aircraft. For this reason aluminium will be used in the majority of the fuselage and most of the structural components due to its sufficient strength-to-weight ratio and its cost to manufacture. These parts of the aircraft will be constructed of aluminium alloys, mainly alloyed with copper and magnesium. Composite will still be used in components other then the main part of the fuselage.

The fuselage of the aircraft is a semi-monocoque structure. It will be made from different types of aluminium alloys which will be used for different areas of the aircraft depending upon the characteristics required. The alloys are mainly aluminium, zinc, magnesium & copper but also contain traces of silicon, iron, manganese, chromium, titanium, zirconium and most likely

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several other elements that remain trade secrets. The different alloys are mixed with different ingredients to give different properties as shown in Table 9 below:

TABLE 9 - ALUMINIUM ALLOY APPLICATION AND TYPE

Section Material Material Description

Majority of Fuselage Skin (areas primarily loaded in tension)

Aluminium Alloy 2024 (Aluminium and Copper)

Good fatigue performance, fracture toughness and slow

propagation rate

Portion of frames, stringers and wing ribs

Aluminium alloy 7075 (Aluminium & zinc)

High mechanical properties and improved stress corrosion

cracking resistance

Minimal use in wing spars & beams

Aluminium alloy 7178 (Aluminium, zinc,

magnesium & copper)

High compressive strength to weight ratio

Landing gear beam Aluminium alloy 7175

(Aluminium, zinc, magnesium & copper)

A very tough, very high tensile strength alloy

FIGURE 40 - FUSELAGE STRUCTURAL COMPOSITION

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There will be a range of aluminium alloys used on the projected aircraft, with the majority of the fuselage skin comprising of aluminium alloy 2024. Table 10 below shows aluminium alloy 2024-T3’s properties. The other aluminium alloys present in the aircraft are minimal but help improve the strength of the aircraft where it is necessary.

TABLE 10 - ALUMINIUM ALLOY PROPERTIES

2024-T3 Aluminium Alloy Properties

Ultimate Yield Strength (psi) 58000-62000

Yield Strength (psi) 39000-40000

0.2% Proof Stress (MPa) 322.1

Modulus of Elasticity (lbf/in2) 10.6 x 106

Density (lbf/in3) 0.1

Shear Modulus (ksi) 4060

Shear Strength (lbf/in2) 41000

FIGURE 41 - STRESS-STRAIN CURVE - 2024-T3 ALUMINIUM ALLOY

Aluminium Alloy 2024 is widely used for construction of aircraft structures, such as wings and fuselages. The superior metal fatigue resistance of this alloy makes it excellent for aircraft structures which are under tension, such as the case of our new aircraft in the fuselage. 2024 aluminium alloy has copper as the alloying element and is one of the best known alloys for high strength. With its high strength and excellent fatigue

resistance, it is used to advantage on structures and parts where good strength-to-weight ratio is desired. 2024-T3 Aluminium Alloy is readily machined to a high finish. It is readily formed in the annealed condition and may be subsequently heat treated.

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6.5.3 Titanium

The aerospace industry is the single largest market for titanium products primarily due to the exceptional strength to weight ratio, elevated temperature performance and corrosion resistance. Titanium applications are most significant in jet engine and airframe components that are subject to temperatures up to 590°C and for other critical structural parts. Usage of titanium is widespread in most commercial and military aircraft. Significant characteristics of titanium can be seen in Table 11 below.

TABLE 11 - TITANIUM CHARACTERISTICS

Characteristic Description

Lightness Its low density means that it weighs only around 56% as much as steel.

Strength-to-Weight Ratio

Titanium is the highest of any of today's structured metals.

Flexibility With its low modulus of elasticity (14.9 x 106 psi), about half that of steel,

titanium is not only extraordinarily flexible, but it also springs back very strongly after it has been stressed

Corrosion and Erosion

Highly resistant to corrosion and erosion

Thermal Conductivity

Very high thermal conductivity

Low Coefficient of Expansion

Titanium's low coefficient of expansion makes it much easier to use in combination with ceramics, composites and glass than most other metals.

In the current aircraft industry, new titanium products, alloys and manufacturing methods are being highly employed. The use of precision castings and new alloys such as and Ti-3AL8V-6Cr-4Zr-4Mo are making it possible for titanium to displace alternate, less efficient structural materials in a wide spectrum of aerospace applications. Since titanium was first introduced to the aviation industry in the 1960’s, it has steadily increased use in passenger aircraft as seen in Figure 42 below. The projected aircraft will feature approximately 10% in weight titanium.

FIGURE 42 - INCREASED USE OF TITANIUM IN PASSENGER AIRCRAFT

After extensive research, engineers found that while only titanium and steel had the ability to withstand high aviation operating temperatures, aged Ti-13V-11 Cr-3AL titanium weighed half as much as stainless steel per cubic inch and its ultimate strength was about equal to stainless. Using "conventional" fabrication techniques, fewer parts are needed with Ti-13V-11 Cr-3AL

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than with steel. Therefore the projected aircraft will make use of titanium in specific areas where high strength and low weight is necessary.

Application of Titanium

Engines

The largest single use of titanium in our aircraft will be in the gas turbine engines. In the new PW-1000 gas turbine engines, titanium-based alloy parts will make up around 25- 30% of the dry weight, primarily in the compressor. Applications include blades, discs or hubs, inlet guide vanes and cases. Titanium is perfect for use in these areas since these engine parts operate up to 590°C and titanium can still operate safely without failure.

Airframes

Titanium alloys effectively compete with aluminium, nickel, ferrous alloys and composite materials in the design of the aircraft and its structures. Applications of titanium in the airframes structural members will be minimal, mainly taking the form in small critical fasteners, springs and hydraulic tubing.

Selection of titanium in both airframes and engines is based upon titanium's basic attributes; weight reduction due to high strength to weight ratios coupled with exemplary reliability in service attributable to outstanding corrosion resistance compared to alternate structural metals.

Conclusion

The material selection portion in the design of a new aircraft is a vital part due to the fact that the materials play a big role in determining the weight and strength of the aircraft. It must be stresses that the projected aircraft is at a very early stage in terms of design, thus material selection, location of implementation and quantity are all subject to change.

After intensive analysis is completed, the results will determine the final outcome on the material which each section will consist of. The materials which are selected in this report are proven to perform successfully in other previous aircraft and therefore safe to incorporate them into the design.

According to the RFP, it was desired to make use of a high percentage of composites in the new passenger aircraft. This design has been accomplished by extensive research into the benefits and feasibility of using composite materials in specific locations. The combination of composites, aluminium, titanium and other materials is the best option for a more environmentally friendly, practical, safe and realistic passenger aircraft.

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6.6 GROUND AND SERVICE SUPPORT

6.6.1 Alternate Fuels - Bio-Jet Fuel

RFP Requirement: RFP requirements state that the utilization of alternate fuels for emission reduction is one of the major targets of this design. As can be seen from the feasibility study and the fuel database available in Table 12 and Table 13, there are many fuels to choose from. The use of biofuel in aerospace is still under development. Biofuel is a renewable resource. Production depends on crop production, land mass availability, oil production and the extraction process. Research shows that aircrafts can run safely using a 50-50 blend of conventional Jet-A fuel and Biofuel known as Bio-Jet Fuel.

Availability

As bio-jet fuel is still under development, it is not available in the market at present. However, it is predicted to be available by 2020. Boeing has been testing biofuel tailored for aviation purpose, airports may purchase the fuel from Boeing directly or other common aviation fuel providers such as Shell if they could adopt the technology to produce it. Moreover, B20 (20% biodiesel blend) for ground vehicles is available from most fuel providers, so there is an expectation that bio-jet fuel will be available from current aviation fuel suppliers in the future.

Compatibility - Airport and Aircraft Infrastructure

Storage Facilities

Biofuel blend can be stored in conventional jet fuel storage tanks without significant modification. B20 blends, which is already available for ground vehicles, can be stored in and dispensed from the same refueling equipment used for conventional diesel fuel. Research shows that B20 is compatible with most refueling tanks and dispensing systems with minor modifications. Based on the current properties of biodiesel, higher concentration of blends such as the future bio-jet fuel may be a solvent. Dispenser filters may collect residues dissolved by bio-jet fuel and should be replaced before filling with bio-jet fuel. There is also a potential danger that particles dissolved may go through the filter, which could damage the engine fuel system. Regular removal of dissolved residues and particles is important for using existing fuel tanks. In terms of costs however, using existing systems is always cheaper than a complete infrastructure upgrade. Biodiesel can dissolve more water than petroleum diesel fuel and has the characteristic to attract and absorb moisture directly from air. Bio-jet fuel may have the same properties. This is particularly important for tanks being used above ground and in a humid environment. Water absorbed or dissolved will sink and create an additional volume on the bottom of the tank known as water bottoms. Water layer formed will support biological growth and have an effect on the quality of the fuel. Microbial growth in bio-jet fuel storage tanks may occur and affect fuel quality. Therefore special fuel monitoring systems may be required. Removal of water bottoms could help control the biological growth. However, if bio-jet fuels are to be stored for long periods, regular samples must be inspected to ensure the stability of the fuel. During the switching process from conventional jet fuel to bio-jet fuel, existing tanks need to be cleaned and de-watered before the first fill of biofuel. Further cleaning after filling may be required to remove any tank residue that may have been dissolved from the tank. As the bio-jet fuel is still under development, the final properties of the fuel are unknown. They

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would depend on the crops and ingredients used to produce the fuel. As a result, other specific tank maintenance procedures and upgrades to existing storage tanks may be needed.

New Storage Tanks

New storage tanks can be installed if required. For above ground tanks, they would be available when bio-jet fuel is available in the market. For underground storage, existing underground storage tanks could be easily modified to be used for bio-jet fuel.

Fuel Transport and Distribution

The current fuel distribution system shown in Figure 43 would remain the same as bio-jet fuel and would be a “drop in” replacement to conventional aviation fuel. Minor modifications may be required for trucks, tanks and pipelines.

FIGURE 43 - CURRENT FUEL DISTRIBUTION SYSTEM

Fuel Injection

“Diesel Fuel Injection Equipment (FIE) Manufacturers Common Position Statement” of 2009 states: that the use of blends up to 5% biodiesel is tolerated. However, the FIE Manufacturers are not responsible for any failures caused by using fuels which the products were not designed for. Due to the fact that bio-jet fuel is an experimental product, it could have side effects on the fuel injection system. In terms of fuel injection equipment, manufacturers would most likely not take responsibility for any malfunctioning or damages caused by using bio-jet fuel. Therefore, new maintenance procedures and monitoring system on fuel injection systems and equipments would be required to prevent any damage and loss caused by using bio-jet fuel.

Fuel Handling and Safety

As described in the previous section, although the final properties of bio-jet fuel are still unknown, it could be a solvent. Therefore bio-jet equipments such as hoses, seals and gaskets must be resistant to the fuel. Bio-jet fuel may be toxic or non-toxic, in either case, inhalation, ingestion and contact with the eyes and skin should be avoided.

Biodiesel has a higher freezing point than conventional aviation fuels. Bio-jet fuel has a similar problem but research has been conducted to overcome this difficulty. The final properties of the bio-jet fuel are expected to be very similar to conventional aviation fuel. Therefore, the temperature handling of bio-jet fuel would be very similar to conventional aviation fuel.

6.6.1.1 Fuel Database

A detailed comparison of the different candidate fuels has been provided on the following page.

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TABLE 12 - FUEL DATABASE (JET A, BIO-JET FUEL, DIESEL, BIODIESEL, LIQUID METHANE)

Fuel Jet A Bio-jet Fuel (by

Boeing) Diesel Biodiesel (Typical) Liquid Methane

Specific Energy (MJ/kg)

43.2 48 (approx.) 42.8 38.9 50

Density at 15˚C (kg/100 MJ)

0.808 NA (reportedly

lighter than Jet A) NA 0.87 0.424

Energy Density (MJ/L)

34.9 NA 35.8 33.9 21.2

Physical State

Liquid Liquid Liquid Liquid Liquid

Main Fuel Source

Kerosene, Gasoline

Crops and animal fats

Crude Oil Crops and animal fats Underground reserves

Types of vehicles available today

Most aircrafts No vehicles are

available for commercial use

Many types of vehicles classes

Any Vehicle that runs on diesel today—no modifications are needed for up to 5% blends. Many engines also compatible with up to 20%

blends.

-

Environmental Impacts

Produce harmful

emissions

Produce less harmful emissions

than Jet-A

Produce harmful

emissions

Produce less harmful emissions than Diesel fuel

Produce less harmful emissions than Jet-A, but liquid methane itself is a type of greenhouse gas

Energy Security Impact

Non-renewable Renewable Non-renewable Renewable Renewable

Fuel Availability

Available from most fuel suppliers

Not Available. Predicted to be ready by 2020

Available from most fuel suppliers

Available from most fuel suppliers Not commonly used but

available

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TABLE 13 - FUEL DATABASE (LIQUID HYDROGEN, METHANOL, ETHANOL, ELECTRICITY)

Fuel Liquid Hydrogen Methanol Ethanol Electricity

Specific Energy (MJ/kg)

120 19.9 27.2 -

Density at 15˚C (kg/100 MJ)

0.071 0.796 0.794 -

Energy Density (MJ/L)

8.4 15.9 21.6 -

Physical State Liquid Liquid Liquid -

Main Fuel Source Natural Gas, Methanol,

and other energy sources Natural gas, coal, woody

biomass Corn, Grains, or agricultural

waste Coal, nuclear, and renewable

resources

Types of vehicles available today

No vehicles are available for commercial use

Mostly Heavy-duty buses are available

Light-duty vehicles, medium and heavy-duty trucks and buses runs on E85(ethanol)

gasoline

Electric Vehicles, Bicycles, Light-duty vehicles, medium and heavy-

duty trucks and buses.

Environmental Impacts

Liquid Hydrogen produced from water or

nuclear does not produce C02 emissions

M-85 ground vehicles can demonstrate a 40%

reduction in ozone-forming emissions

E-85 ground vehicles can demonstrate a

25% reduction in ozone-forming emissions

Electricity produces no emissions during operations. However

emissions can be contributed to power generation

Energy Security Impact

Can be produced from natural gas or renewable

sources Renewable Renewable

Electricity is generated mainly through coal fired power plants.

However it can also be generated through renewable sources.

Fuel Availability Not commonly used but

available Blends are available from

most fuel suppliers Blends are available from

most fuel suppliers Available from power stations all

over the world

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6.6.2 Airport Management Systems

Ground Support Systems

The development of an environmentally friendly aircraft is not solely based on the aircraft itself, it is just as important that the support systems in place are improved as well as they can have a dramatic effect on the aircraft’s overall environmental footprint. Development into improved ground support systems is becoming an important issue as the improvements on the actual aircraft become harder to achieve. This section will lay out an improved ground support system to be introduced at airports to improve the overall carbon footprint of the aircraft.

Equipment Fleet Management System

The way that that ground support vehicles are managed can have a significant impact on the efficiency of the ground support system. If the ground support system is poorly managed, with vehicles sitting idle for long periods of time or constantly driving across the airport when not necessary, it can lead to high emissions as well as an unnecessary high operating costs. It is therefore important that the management of the system is streamline to reduce the waste of energy.

6.6.2.1 Zebra Enterprise Solutions Ground Management System

There are a number of system for improving the efficiency of ground support systems, however after careful analysis of the available systems the product produced by Zebra Enterprise Solutions proved to be the most viable system, giving the best overall efficiency without a huge increase in implementation cost.

The system works by installing a telemetry system into each asset in the airport. This allows for a real-time, visual representation of the status and workload of each asset at the airport. This will allow the system and airport managers to monitor each asset in real time with the system giving detailed information such as fuel levels, maintenance schedules, current use and position in airport accurately from a central location. This in turn will allow for a better distribution of the work load of each asset, reducing the idle times of assets and aircraft and thus reducing the overall running costs of the airport.

Figure 44 shows a simple way that the systems can be monitored. This entails a plan view of the airport showing the current location of each of the assets and aircraft. While there are many view types of the system to give detained information to each asset individually this view allows for the overall monitoring of the entire airport accurately from a single control room.

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FIGURE 44 - DEPICTION OF ASSET VIEWER WITH AVAILABLE CONNECTION TYPES

Advantages

The ground support system produced by Zebra Enterprise Solutions gives many advantages over conventional airport management systems. Table 14 are the systems improvements when the system is installed at an airport. This information is available from the ‘Equipment Fleet Management for Aviation DATASHEET’ produced by Zebra Enterprise Solutions.2

TABLE 14 - SYSTEM CAPABILITIES

Operation Function

Status & Location Monitoring

Provides the essential information to locate and visualize an asset in real-time and to detect its status, such as running time and idle time. All data is stored for additional analysis.

Maintenance Forecasting

Calculates the due date for the next preventative maintenance job, inspection or other planned workshop activity.

Condition Monitoring

Condition Monitoring obtains error signals from the telemetry device and converts them into notifications for the maintenance and operation teams.

Access Control

Includes a set of access control mechanisms to ensure that only authorized staff can operate the vehicle.

Geofencing

Geofencing allows the customer to setup special zones where a GSE or an operator is not allowed to enter. Whenever such a zone is entered, the telemetry device on the equipment will take action. Such actions might for instance depend on the authorization of the equipment, the authorization of the operator, the work state and the current time. Geofencing works both in a 'connected' as well as in a ‘non-connected’ mode.

2 http://zes.zebra.com/pdf/products-datasheets/ds_equipment_fleet_manager.pdf

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Operation Function

Safety and Security Audit

The safety checklist enforces the safety of vehicle operations through location replay. Impact sensing allows for an analysis of the safety practices during operations. The safety checklist data is stored for future reference and analysis.

Battery and Fluid Monitoring

Signals and alerts of battery status and/or the fuel level, de-icing liquids or any other parameter that can be remotely monitored by the telemetry device.

Reports and Key Performance Indicators

At the heart of the system, it can produce the essential information to evaluate the utilization patterns of the fleet, the degree to which the operations are efficient or wasteful. Using the stored historic data, it produces the information needed to operate the fleet efficiently and at a higher safety level.

Billing Support

Provides evidence of the actual use of equipment and the basis for appropriate billing.

Third Party System Integration

Includes tools to share information with billing, scheduling and ERP systems used by airports.

Disadvantages

Like any new system there will also be some disadvantages. The foremost of these will be the cost of installing and setting up the system. This could be quite significant as each asset requires modification to allow the system to work. However due to the vast improvements that the system can offer including the reduced operating cost, the system will break-even within a number years. Another issue could be the down time of assets as they are being upgraded to the new system, as this may require them to be out-of-action for some time as they are being modified.

There is also the issue of training with any new system as all personal will need to be retrained in the use of the new system. This may take some time and thus reduce the efficiency of the airport during this period. However, once the new training is completed the airport will run more efficiently.

6.6.2.2 Hydrogen Powered Ground Support Vehicles At Airports

Introduction

The need for a alterative green fuel source for airport ground support vehicles has become an important issue if the aviation industry is going to reach the reduces CO2 emissions targets set down by governing bodies. This section will lay out the two most viable answers to solve this problem. The advantages and disadvantages of each will be discussed and the reason for the selected fuel will be outlined and justified.

Fuel Selection

There are a number of alternative fuels available to power ground support vehicles. These include: hydrogen, electric, solar, LPG and natural gas. The two main contenders for a viable fuel source is hydrogen and electric. The choice to select the hydrogen power system over the

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electric system was made after carefully weighing up all the advantages and disadvantages of each system and then matching them to the needs of the ground support vehicles.

Electric

Electric cars or battery electric as they can also be known as, have been around for some time. However, only in the last twenty years have they started to become a viable form of propulsion. This is due to the latest development of batteries. They use these batteries to store electricity which is then used to power an electric motor which in turn drives the wheels. These batteries can be recharged from any mains power source. As with all mains power the electricity is mostly produced by burning fossil fuel such as coal. However, in the last ten to fifteen years there has been a large development in green power such as wind and solar. Another option to produce this power is from nuclear. While there are many arguments against this form it would make a cleaner stopgap until another less destructive form of power could be produced.

The electric car is very efficient, normally between 80-90%, this come from not having to convert the electricity to another form for storage and use. There are two major flaws when it comes to electric power; they have limited range as batteries are not super-efficient at present and need to be charged after only short period of time. This may not be such a problem in an airport when any support vehicle running low on power does not have far to travel to receive a top-up. This however leads to the second and significant problem of the charge time. It takes a number of hours to charge the batteries. While this may not be as big a problem for the auto industry where drivers can leave there car on charge overnight, it is however a problem for an airport where any down time of service vehicles will cost the airport significant amounts of money.

Hydrogen

Hydrogen as a power source has been around since the beginning of the universe. It is the fuel source of the sun and is the most common element in the universe. However, it is not found in its raw state on earth and therefore has to be extracted from other elements such as water or natural gas. Once the hydrogen has been extracted it can then be used to power vehicles. There are two ways that it can achieve this. This first is through the Internal Combustion Engine (ICE), this is the same system as the current petrol engine with modifications to allow for the hydrogen fuel. The second system is the fuel cell and electric motor system. This uses a fuel cell that combines the hydrogen and oxygen from the air to produce electricity, which then powers an electric motor to drive the wheels. In this respect it is very similar to the electric system. Figure 45 below shows the fuel cell system diagram.

There are a number of ways to produce hydrogen. The current way to produce hydrogen is through steam reforming. This is when the hydrogen particles are striped from hydrocarbons. The most common of these is natural gas with approximately 80% efficiency. This method however uses fossil fuels which are contrary to the aim of reducing reliance on said fossil fuels. It also produces greenhouse gasses. The second and most desirable way to produce hydrogen is by electrolysis. This is the process of splitting water atoms by passing electricity through the water. While this is not as efficient as steam reforming it is a lot better for the environment as no by-products are produced. This efficiency can be increased by both heating and pressurizing the water. This also has added benefits of not needing to pressurize the hydrogen after it is produced as it is already pressurized. This can add as much as 4% to the efficiency due to not needing a pressure pump. Using these methods a total system efficiency of 50- 60% can be achieved when coupled with a hydrogen fuel cell. However, it is expected to increase in the next 10-15 years as new techniques currently under development get introduced and certified for mass production. Some of the other ways to produce hydrogen are through thermolysis (using

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micro-organisms to break down waste to release hydrogen) and biomass (extracting the hydrogen from living or dead material, normally plants). However, both these are in their infant stage of development and are not currently viable for mass production.

FIGURE 45 - FUEL CELL SYSTEM

Selected System – Hydrogen Fuel Cell

The selected system for use in the ground support vehicles will be utilizing the hydrogen fuel cell system. However, it will also be using batteries instead of just a fuel cell. This system will use the batteries to power the electric motor. The batteries in turn will be recharged by the hydrogen fuel cell. This system is very similar to hybrid cars (Toyota Prius) currently available, the main difference being a hydrogen fuel cell instead of a petrol engine. This system will be a more efficient system then a straight fuel cell system. The use of batteries will reduce the wear on the system, thus reducing the maintenance intervals. It achieves this by smoothing out the load on the system. The electric motor will draw varying current from the batteries and the hydrogen fuel cell will recharge the batteries at a steady rate, this will reduce the strain on the fuel cell as the fuel cell is the most expensive component of the system. This reduced strain on this component will increase the life of the cell and thus decrease maintenance cost. Figure 46 below shows the schematic diagram of the ground support systems fuel cell system.

FIGURE 46 - SYSTEM DIAGRAM

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The airport infrastructure is also an important part of the system. The airports will produce the hydrogen on site through high pressure/temperature electrolysis. This will greatly reduce the logistics cost as the hydrogen will not have to be transported in cryogenic pressure vessels around the country from the production site. This will dramatically reduce the operating cost of the system. An important aspect of the system is that the airport get its electricity from green sources such as wind or solar. Nuclear is also an option until more green energy is developed. It may not be possible to get all energy from green sources, it is important then that if the airport does get some of its energy from coal power it must strive to reduce it dependency as new green sources are developed and introduced.

Advantages

The selected system has a number of advantages over the current and other systems currently available. The first advantage of the selected system is the fact that the system is completely green when the power used to create the hydrogen comes from non-greenhouse producing sources such and wind, solar or nuclear. This is an important aspect when it comes to environmentally friendly technology. If the power comes from a coal power station it defeats the purpose of an environmentally friendly and sustainable system.

The major advantage the selected system has over the electric system is its refill times. With it only taking between five to ten minutes to fill a hydrogen tank compared to three to seven hours to recharge a battery system, it has a substantial ability to reduce down time of ground support vehicles due to the time it would take to refill the system. This will lead to more productivity and thus reduce the operating costs as the system can do more work in a shorter period of time. The batteries can also act as a backup if the hydrogen system fails or the fuel tank runs out. It will allow the vehicle to finish its current job and return to fix the problem without any delays in productivity.

The combination of both a fuel cell and batteries will reduce the wear on the system. This is an important aspect when it comes to maintainability of the system. This reduced wear will increase the maintenance times and thus decrease the operating cost of the system.

Hydrogen can be produced at relatively the same price as petrol at the moment. With the price of fuel increasing and the new technology allowing for cheaper production of hydrogen the operating cost of the system will actually decrease from its current level. This will look very attractive as the airline will also be utilizing a renewable source of energy and thus not contributing to global warming, while at the same time reducing their operation costs.This system will also be able to operate in enclosed spaces and will not exhibit a health risk to drivers or surrounding people. This is not currently achievable by the current system as poisonous gases are produced in the petrol combustion engine. This is a very important aspect when it comes to OH&S which should be at the for-front of the airlines concerns with the current system.

Disadvantages

The system also has some issues as well as the positives mentioned above. The foremost of these is the procurement and installation cost of the system. The system does require the current ground support system to be replaced with the new system. This cost could be reduced if the new system was phased in allowing both systems to run in conjunction which will reduce the amount of delay and thus reduce the cost of implementing the new system. This system is a long term investment and when this is coupled with the renewable side of the system, the airports will no longer need to purchase carbon offsets or credits and the system is carbon neutral. The long-term benefits outweigh the short term set-up costs.

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The parts for the fuel cell will be the largest item in the maintenance cost. Using the system setup described to reduce the load on the fuel cell the costs can be kept to a minimum and thus not have a significant impact on the bottom line of the entire system over its life.The hydrogen system is not as efficient as a pure electric system. However, when the refueling times and cost of idle vehicles are considered the hydrogen system is by far the better choice. It will allow the vehicles to operate for longer with little downtime due to refueling. As this is an important aspect to busy airports, hydrogen becomes the propulsion system of choice.

Future improvements

With any system it is important to look at the upgradability of the system. Can the system be upgraded easily if a new improved technology came along or is it fixed where improvement would require the introduction of a new system? This hydrogen system has room for improvements. New technology being developed now such as a hydrogen system that produced hydrogen in the vehicle from water, would eliminate the need for a fuelling station at the airport with only a small modification to the existing vehicles.

Conclusion

The need for an alternative green fuel source for airport ground support vehicles is an important issue, with the need to reduce CO2 emissions as the main objective. The hydrogen fuel system coupled with a battery system has been found to be the best option. It offers an efficient system while at the same time not compromising productivity. The system is carbon neutral especially if the power to create the hydrogen comes from clean sources, this supports any goals to reduce carbon emissions for the future.

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6.7 ENGINES – PRATT AND WHITNEY PW1000G

RFP Requirement: In any aircraft design the engine plays a major role in emissions, noise and fuel consumption. RFP requirements state that engine efficiency must be investigated and improved.

The Pratt & Whitney PW1000 is the first design from the Pure-Power Engines design line and is to be certified in 2013. It is a geared turbofan engine which provides a thrust range of 14,000–23,000 lbf. The geared system separates the engine fan from the low pressure compressor and turbine, allowing each of the modules to operate at their optimum speeds. This enables the fan to rotate slower while the low pressure compressor and turbine operate at a high speed. This increases engine efficiency and delivering significantly lower fuel consumption, emissions and noise. This type of engine is selected by Bombardier to power the 110/130 Seat C-Series and by Mitsubishi to power the Mitsubishi Regional Jet.

Compared to the current engine used on the B737-900NG, the CFM International CFM56, the Pratt & Whitney PW1000 offers a fuel burn reduction of 15%. The PW1000 also has a maintenance benefit. It is due to the geared turbo fan which has four less life-limited disks and 1,500 less airfoils. Each of these disk cost about US$100,000. Pratt & Whitney vice-president claimed the cost of these disks represent a savings of $20 or more per engine flight hour. These life-limited parts plus lower airfoils will enable the PW1000 to have greater than 20% engine maintenance cost advantage over the CFM56.

In terms of emissions it is estimated to have 15dB reduction in noise pollution and a significant CO2 reduction. This improvement helps a further 2-3% reduction in airline operating costs by avoiding the airport noise fee and the CO2 emission penalties.

Flight tests have already been conducted on a 737SP on July 2008. The test results were outstanding and gave good values on noise, emissions and fuel burn improvements. In terms of noise emission, carbon emission thrust capacity and other improvements. The PW1000G is still currently undergoing research to further increase its performance capabilities by 2013.

TABLE 15 - CFM56 VS. PW1000G

Engine Type CFM56 Pratt & Whitney

PW1000

Thrust 20,000 – 27,000

lfb 14,000–23,000 lbf

Fuel burn reduction - 15%

Noise reduction - 15db

Carbon Emission reduction per aircraft per year (tonnes) comparison

3000

By Pass ratio 6-to-1 12-to-1

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6.8 OPERATIONAL AND MAINTENANCE COSTS

RFP Requirements: RFP requirements state the need for an investigation of operating and maintenance costs based on current in-service aircrafts.

The aircrafts costs will be similar to the aircrafts mentioned below.

Type of Aircraft Fuel (Civil Jet Fuels)

Jet Fuel is used in the aviation industry to power turbo prop and jet engine civil aircrafts. Currently there are two main grades of turbine fuel in use in civil commercial aviation: Jet A-1 and Jet A both of which are kerosene type fuels. There is another grade of jet fuel known as Jet B which is a wide cut kerosene (a blend of gasoline and kerosene) but it is rarely used except in very cold climates. Detailed descriptions are provided in Table 16 below.

TABLE 16 - FUEL TYPES WITH ASSOCIATED PROPERTIES

Fuel Type Description

JET A-1 Jet A-1 is a kerosene grade of fuel suitable for most turbine engine aircraft. It is

produced to a stringent internationally agreed standard. Has a flash point above 38°C (100°F) and a freeze point maximum of -47°C. It is available worldwide.

JET A Jet A is a similar kerosene type of fuel, produced to an American Society for Testing and Materials (ASTM) specification and normally only available in the USA. It has the same

flash point as Jet A-1 but a higher freeze point maximum (-40°C).

JET B

Jet B is a distillate covering the naphtha and kerosene fractions. It can be used as an alternative to Jet A-1 but because it is more difficult to handle (higher flammability), there is only significant demand in very cold climates where its better cold weather

performance is important.

Fuel Cost3

Jet Fuel Price Monitor For A Given Day

The global weekly average price paid for jet fuel and year-to-date average fuel price per barrel is provided by the IATA Jet Fuel Price Monitor

TABLE 17 - FUEL PRICES - 25 JUNE 2010

Percentage change

25-June-2010 Index* $/b cts/gal $/mt 1 week ago 1 month ago 1 year ago Jet Fuel Price 245.2 89.7 213.5 706.9 -1.9% 7.7% 15.3%

Compared to year 2000, where Index*=100 (87 cts/gal)

TABLE 18 - FUEL PRICES - 17 SEP 2010

Percentage change

17-Sep-2010 Index* $/b cts/gal $/mt 1 week ago 1 month ago 1 year ago

Jet Fuel Price 243.3 89.0 211.9 701.5 -0.5% 1.9% 12.8%

Compared to year 2000, where Index*=100 (87 cts/gal)

3 IATA Jet Fuel Price Monitor (Viewed on 23, September, 2010)

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Jet Fuel Price Monitor For The Year

TABLE 19 - THE IMPACT ON 2010’S FUEL BILL OF THE GLOBAL AIRLINE INDUSTRY

25-June-2010 New fuel price average for 2010 Impact on 2010 fuel bill

$88.5/b +$18 billion

17-Sept-2010 New fuel price average for 2010 Impact on 2010 fuel bill

$88.2/b +$17 billion

Table 19 above calculates the ‘new fuel price average’ for the current year by extrapolating this week's price out to the end of the year, at this week's level. This price is multiplied by the gallons of jet fuel consumed by the global commercial aviation industry last year. This total is then subtracted from last year’s fuel bill in order to estimate the impact of the change in average fuel prices on the current year’s fuel bill.

Aircrafts Operating Cost4

The International Civil Aviation Organization (ICAO) provides base-line average aircraft operating costs per block hour. These values are provided, per aircraft type and are shown in Australian dollars below:

TABLE 20 - BASE-LINE AIRCRAFT OPERATING COSTS - AVERAGE PER BLOCK HOUR (AUS$)

Aircraft Type Typical Fuel

Consumption (Litres)

Other Costs Cost of Fuel Total Operating

Costs*

A300-600 7071.1 $3,733.60 $2,961.52 $6,695.11

A319 3107.8 $2,719.07 $1,301.72 $4,020.80

A320 3353.9 $2,834.70 $1,404.30 $4,239.00

A321 3505.3 $3,237.53 $1,467.70 $4,705.23

A330-200 6669.9 $4,108.45 $2,793.67 $6,902.12

A330-300 7082.5 $4,127.10 $2,965.25 $7,092.35

A340-300 8229.5 $4,149.48 $3,446.40 $7,595.88

A340-600 9781.5 $4,949.54 $4,095.40 $9,044.93

B-727-200 4027.7 $3,494.89 $1,389.38 $4,884.26

B-737-200 3013.2 $2,666.86 $1,038.77 $3,705.62

B-737-200C 4300.2 $3,694.44 $1,482.62 $5,177.06

B-737-300/700 2611.9 $2,816.05 $900.76 $3,716.81

B-737-400 3051.0 $3,144.28 $1,051.82 $4,196.10

B-737-800 2135.0 $2,215.54 $736.65 $2,952.19

B-737-500 3043.5 $2,821.65 $1,049.96 $3,871.60

B747-100 15175.7 $9,231.43 $5,233.01 $14,464.43

B-747-200 15228.7 $10,458.55 $5,251.66 $15,710.21

B-747-400 14168.8 $7,722.69 $4,886.13 $12,608.82

B-757-200 3406.9 $3,832.44 $1,174.91 $5,007.35

B-767-200 4606.8 $4,410.57 $1,588.92 $5,999.49

B-767-300 4909.7 $4,643.69 $1,693.36 $6,337.05

B-777-200 7302.1 $5,219.95 $2,517.66 $7,737.61 *These operating costs do not consider landing costs at Australian airports. These charges can vary from location to location and by aircraft type. The costs were originally stated in 2000 US dollars however, they have been inflated to 2009 costs and converted to Australian dollars using the PPP conversion factor. US gallons were also converted to litres at a rate of 1 US gallon being equal to 3.8 litres.

4 ICAO Aircraft Operating Costs – Base line (Viewed on 23, September, 2010)

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Visual Flight Rules/Instrument Flight Rules - Estimated Aircraft Operating Costs

The average estimated aircraft operating costs for VFR/IFR aircraft are provided in Table 21 below by aircraft category. These costs are based on fuel and incremental operating costs.

TABLE 21 - ESTIMATED AIRCRAFT OPERATING COSTS/HOUR AS AT JUNE 2009(AUS$)

Aircraft category Estimated Aircraft Operating Costs/Hour AUS$

(excluding fixed annual overheads)

VFR General Aviation Operators $172-$344

Moderate sized IFR traffic $1,148

Larger IFR jet $3,443

Rate of Fuel Burn5

The term “Rate of Fuel Burn” describes the average rate that an aircraft burns fuel. The data for fuel burn (kg/hr) by phase of flight and aircraft type is shown in Table 22 below.

TABLE 22 - FUEL BURN (KG/HR) BY PHASE OF FLIGHT

Phase APU only

Stationary ground

Active taxi out

En–route Arrival management

Load (% max

payload weight)

n/a From 50%

to 80%

From 50% to

80% 50% 65% 80% 50% 65% 80%

Low Base High Low Base High

B737-300 115 690 900 2,355 2,436 2,523 2,656 2,731 2,814

B737-400 115 690 900 2,337 2,410 2,498 2,504 2,588 2,676

B737-500 115 690 900 2,169 2,224 2,288 2,483 2,530 2,584

B737-800 115 690 900 2,485 2,572 2,668 2,038 2,187 2,229

B757-200 150 820 1,000 3,195 3,311 3,417 2,685 2,789 2,867

B767-300ER

150 1,120 1,400 4,514 4,726 4,941 3,735 3,908 4,093

B747-400 280 2,700 3,400 9,484 9,809 10,125 7,198 7,421 7,647

A319 120 630 720 2,240 2,304 2,374 1,791 1,854 1,919

A320 120 630 720 2,279 2,355 2,429 2,002 2,074 2,151

A321 120 730 840 2,695 2,788 2,885 2,524 2,625 2,728

Average Rate of Fuel Burn Per Landing Take-Off6

Table below displays fuel burn for Landing-Takeoff (LTO) by average fleet.

TABLE 23 - FUEL BURN FOR LTO BY AVERAGE FLEET

Average fleet kg/LTO

Domestic B737-400 825

International

B767 1,617

B737-400 (short distance) 825

B747-400 (long distance) 3,400

5 EUROCONTROL Cost Benefit Analyses (Viewed on 26, September, 2010)

6 EUROCONTROL Cost Benefit Analyses (Viewed on 26, September, 2010)

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6.9 RISK ANALYSIS

TABLE 24 - RISK ANALYSIS

No. Item/Product Risk Risk Level Risk Control

Action Plan Matrix Code High Med Low

1 Co-Flow Jet Flow Control

Technology becomes too expensive to implement or does not become compatible with the

projected aircraft

x

Implementation will be

cancelled C2

2 Pylon-Based Jet Noise Suppressors

Compatibility with the projected aircraft

x

Implementation will be

cancelled D1

3

Actuators and Sensors Technology for Cabin Noise Reduction

Compatibility with the projected aircraft

x

Implementation will be

cancelled C1

4 Fixed Winglets Too costly to implement

x Implementation

will be cancelled

E1

5 Wing Waggle Compatibility as well as cost to

implement x

Implementation will be

cancelled B2

6 No-Bleed Air Systems

Technology becomes unavailable or is not compatible

with the projected aircraft

x

Implementation will be

cancelled and will revert back

to original system

C2

7 Fly-By-Wire Systems

Compatibility

x Implementation

will be cancelled

E1

8 NextGen Transport System

A relatively new system. The change of any transport system could damage the operational

systems instead of improve them.

x

Implementation will be

cancelled D3

9 Aircraft In-Wheel Motors

Compatibility and Cost to implement

x

Implementation will be

cancelled D1

10

Magnetic Braking and Ground Manoeuvring System

Compatibility and Cost to implement

x

Implementation will be

cancelled D2

11 Pre-Spinning Tires

Compatibility and Cost to implement

x

Implementation will be

cancelled D2

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12 BUSS System Compatibility and Cost to

implement x

Implementation will be

cancelled C3

13 Electronic Flight Bag

Compatibility and Cost to implement

x Implementation

will be cancelled

D2

14 Materials – Composites

Cost and the use of composites on an aircraft design which has

not yet been tested for such high levels.

x

Conversion to mostly

aluminium D1

15 Alternate Fuel – Bio-Jet Fuel

Cost and compatibility x

Fuel usage will return to Jet-A

Fuel B3

16

Zebra Enterprise Ground Management Systems

A relatively new system. The change of any transport system could damage the operational

systems instead of improve them.

x

Implementation will be

cancelled B2

17

Hydrogen Powered Ground Support Vehicles

Compatibility and Cost to implement

x

Implementation will be

cancelled B2

18

Engines – Pratt and Whitney PW1000G

Compatibility with aircraft and alternate fuels and Cost to

implement

x

Engine will be replaced with

the LEAP-X engine due for arrival in 2016.

D2

FIGURE 47 - RISK MATRIX

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7 SECTION 2 – AIRCRAFT SIZING

7.1 WEIGHT ESTIMATE

The weight estimate is critical in design as it is the basis for many factors. Achieving a substantially lighter aircraft is a major benefit as it reduces overall fuel consumption and operating costs.

The weight estimation was conducting using 3 main sources:

1. Airplane Design, Part 1: Preliminary Sizing of Airplanes by Dr. Jan Roskam. 2. Aircraft Design: A Conceptual Approach, AIAA Education series, Daniel P. Raymer 3. Boeing 737-600, -700, -800, -900 design specifications. (See appendix)

Approach

This particular calculation is based on the following mission profile:

1. Engine start up and warm up 2. Taxi 3. Take off 4. Climb to cruise altitude and

accelerate to cruise speed 5. Cruise 6. Loiter 7. Descent 8. Fly to alternate and descend 9. Landing, taxi and shutdown

The aircraft is assumed to be similar to the Boeing 737-900 Next Generation aircraft and therefore some characteristics have been used as a reference. These are:

B737-900NG Specifications

1. Maximum Take-off Weight (MTOW) = 85130 kg = 187679.52 lbs 2. Cruise Velocity (VCruise) = 228.61 m/s = 444.37 kts 3. Wing Area (S) = 125 m2 4. Aspect Ratio (AR) = 9.45 5. Cruise Altitude = 41000 ft 6. Oswald Factor (e) = 0.85 (Assumption) 7. Density at Cruise Altitude (ρ41000 ft) = 0.30 kg/m3 (Calculated, see appendix) 8. Profile Drag (CDO) = 0.03 (See Drag Estimates)

Lift-to-Drag Estimate Using B737-900

FIGURE 48 - MISSION PROFILE

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TABLE 25 - BOEING 737 SPECIFICATIONS

Boeing 737-900NG Characteristics For Initial Weight Estimate

Mass 85130 kg

Weight 835125.3 N

Cruise Velocity 228.61 m/s

Density At Cruise 41000 ft) 0.30 kg/m3

Wing Area 125 m2

Coefficient of lift 0.854 -

Oswald Factor 0.85 -

Aspect Ratio 9.45 -

Profile Drag 0.03 -

Coefficient of drag 0.059 -

Lift-To-Drag 14.5 -

Using the Roskam Method of preliminary sizing7 requires 7 key steps. It must be stated however, that the preliminary weight estimate used is the B737-900NG maximum take-off weight. This weight will be reiterated once a final value of empty weight has been derived. This empty weight will incorporate the use of composites and therefore an estimate of 10% empty weight saving is assumed. The reduction of 15% fuel usage is accounted for based on the PW1000G manufacturing specifications.

The important outputs produced using this method are the following:

1. Preliminary maximum takeoff weight estimate 2. Empty weight 3. Fuel required for specified mission profile 4. Landing Weight 5. The method of sizing is provided below. 6. Aircraft Sizing for 200 passengers and 3500 Nautical Mile Range

Step 1 – Payload Weight Estimate

1. Number of passengers = 200 2. Individual passenger weight = 180 lbs (assumption) 3. Baggage Weight = 30 lbs (assumption)

Step 2 – Maximum Take-off Weight Estimate

As described above, the B737-900NG maximum takeoff weight will be used for the first iteration.

7 Roskam, part 1, chapter 2, 2.6.2 Example 2: Jet Transport, page 54

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Step 3 – Determination of Fuel weight, WF for Mission Profile.

FIGURE 49 - TABLE 2.1 FROM ROSKAM (PART 1) 8

Phase 1 - Engine Start and Warm-up

The fuel fraction for this stage using figure 39 for a twin engine is typically 0.992. Therefore:

Phase 2 – Taxi

Using Figure 49, the twin engine taxiing fuel fraction is typically 0.996. Therefore:

Phase 3 – Take off

Using Figure 49, the twin engine taxiing fuel fraction is typically 0.996. Therefore:

Phase 4 – Climb to cruise altitude and accelerate to cruise speed

Using Figure 49, the twin engine taxiing fuel fraction is typically 0.990.

Phase 5 - Cruise

8Roskam, part 1, chapter 2, Table 2.1 Suggested Fuel-fractions for several mission phases, page 12

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An average climb velocity is 275 knots and a climb rate of 2500 ft/min (assumptions). Therefore the range covered during the climb is as follows:

Using Breguet’s range equation9 and the following constants we can find the cruise fuel fraction.

EQUATION 1

1. VCruise = 444.27 kts 2. cj = 0.5 lb/hr/lb10 3. L/D = 14.5 (from “Lift-To-Drag Calculations” of B737-900NG above) 4. Rcr = 3500 – 75.17 = 3424.83 nm (Range required – range covered during climb)

Rearranging the formula gives:

Phase 6 – Loiter

To find the fuel fraction for this phase Breguet’s endurance equation11 is used with the following constants:

EQUATION 2

1. cj = 0.5 lb/hr/lb 2. L/D = 10 (assumption12) 3. Endurance loiter time is assumed to be 1 hour

Rearranging the formula gives:

9 Roskam, part 1, chapter 2, Equation (2.10), page 15

10Roskam, Table 2.2 Suggested values for L/D, cj, ηp and for cp for several mission phases, part 1,

chapter 2, page 14, (Transport aircraft) 11

Roskam, part 1, chapter 2, Equation (2.12), page 15 12

Roskam, part 1, chapter 2, Table 2.2 Suggested values for L/D, cj, ηp and for cp for several mission

phases, page 14, (Twin engine average)

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Phase 7 – Descent

Using Figure 49, the twin engine taxiing fuel fraction is typically 0.992. Therefore:

Phase 8 – Fly to alternate and descend

According to FAR requirements the cruise speed can be no more than 250 kts. Using Breguet's range equation once again and the following constants yields:

1. L/D = 10 (assumption – for the cruise to alternate a value for L/D of only 10 can be achieved13)

2. cj = 0.9 lb/hr/lb (assumption) 3. VCruise = 250 kts (B737-900NG Specifications) 4. Range = 100 nm (assumption)

Phase 9 – Landing, Taxi and Shutdown

Using Figure 49, the twin engine taxiing fuel fraction is typically 0.992. Therefore:

Mission Fuel Fraction is as follows

Fuel used14 is as follows

EQUATION 3

This calculation takes fuel reserves into account. Therefore the used fuel is equal to the fuel required. Hence:

We can now apply the 15% fuel efficiency provided by the PW1000G engines. Therefore

13

Phase 8, chapter 2, page 57 14

Roskam, part 1, chapter 2, Equation (2.14), page 16

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Step 4 – Tentative Operating Empty Weight

Using the tentative operating empty weight formula15 is calculated:

EQUATION 4

Step 5 – Tentative Empty Weight

Crew weight is based on B737-900NG configuration and RFP requirements. The following values were used to find total crew weight.

1. Number of pilots = 2 2. Number of flight attendants = 4 3. Human weight = 180 lbs 4. Baggage weight = 30 lbs

We can now calculate the tentative empty weight16 using the formula below.

,

EQUATION 5

where Wtfo is the unusable fuel and oil (assumed to be 0.5% of guess takeoff weight)

Step 6 – Empty Weight

To find empty weight we use the “A” and “B” values for table 2.1517 for a transport aircraft and the following equation:

A = 0.0833

B = 1.0838

A composite weight saving of 10% can now be taken into account. Hence, this yields

15

Roskam, part 1, chapter 2, Equation (2.4), page 7 16

Roskam, part 1, chapter 2, Equation (2.5), page 7 17

Table 2.15 Regression line contents A and B of Equation (2.16), part 1, chapter 2, page 47,

(Transport aircraft)

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Note that the empty weight and the tentative operating weight have a difference of 5756.22 lbs. Therefore, further iteration must be conducted in order for these two values to be within a certain tolerance (for this analysis we assume tolerance of 1%). By adjusting the estimated take-off weight; WTO-Guess of 85130 kg (187679.52 lbs) to 73750 kg (162589.25 lbs), the difference in weights reduces to approximately 1%. The new values are displayed in Table 26 below.

Now that the new values have been iterated the landing weight can found. The landing weight formula18 is WL = 0.85WTO, therefore:

TABLE 26 - WEIGHT ESTIMATION USING THE ROSKAM METHOD

Roskam Weight Analysis for MTOW, Empty Weight and Landing Weight

Item Symbol Value Unit

Payload weight WPL 42000 lbs

Weight Estimate WTO-Guess 162589.25 lbs

Engine Start and Warm up W1/WTO 0.992 -

Taxi W2/W1 0.996 -

To W3/W2 0.996 -

Climb To Cruise Alt and Acc. To cruise espeed W4/W3 0.99 -

Cruise W5/W4 0.77 -

Loiter W6/W5 0.95 -

Descent W7/W6 0.992 -

Fly to alternate and descend W8/W7 0.96 -

Landing, Taxi and Shutdown W9/W8 0.992 -

Mission fuel fraction Mff 0.70 -

Fuel Used WF-Used 48921.21 lbs

Fuel used with efficiency WF 41583.03 lbs

Tentative operating empty weight WOE-TENT 79006.22 lbs

Crew Weight WCREW 1260 lbs

Tentative empty weight WE-TENT 76933.27 lbs

Allowable Empty Weight WE-ALLOW 86824.41 lbs

Empty Weight with composites contribution WE 78141.97 lbs

Tolerance 1 %

Landing Weight WLAND 138200.86 lbs

18

Roskam, part 1, chapter 3,3.3.4 – Example of FAR 25 Landing Distance Sizing, page 113

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TABLE 27 – WEIGHT COMPARISON – B737-900NG

Aircraft Weight Comparison (B737-900NG)

Weight Type B737-900NG Projected Aircraft Saving (%)

MTOW (lbs) 190742 165244 13.37%

Empty Weight (lbs) 98492.7 79417.1 19.37%

Maximum Landing Weight (lbs) 148688 140456 5.54%

TABLE 28 - WEIGHT COMPARISON - A320-200

Aircraft Weight Comparison (A320-200)

Weight Type A320-200 Projected Aircraft Saving (%)

MTOW (lbs) 174766.80 165244.25 5.45%

Empty Weight (lbs) 95449.56 79417.11 16.80%

Maximum Landing Weight (lbs) 144518.70 140456.06 2.81%

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7.2 V-N DIAGRAM

FIGURE 50 - VN DIAGRAM

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FIGURE 51 – AAA INPUT DATA

Modelling The V-n Diagram

The V-n diagram in Figure 50 was developed using the input data in Figure 51. This input data includes the following specifications:

1. Aircraft Weight Estimate = 723487.5 N 2. Aircraft Wing Area (Preliminary Sizing) = 108.26 m2 3. Velocity (equivalent) = 235 m/s 4. Altitude = 12497 m = 41000 ft

Some values were assumed according to the recommendations of the AAA software for a twin engine commercial aircraft. Once these values were placed into the AAA Software, the necessary outputs were automatically calculated and the V-n diagram was developed.

Flight Envelope

Load Factor, n shows that maximum maneuvering occurs at 2.5g and therefore the wing calculations for structural integrity must accommodate for these loadings. The right side of the maneuvre envelope is the “never-exceed” airspeed. It is a performance parameter and flying beyond this can cause flutter, stability issues, control issues or a host of other problems with potentially drastic consequences. At low velocities the maximum load factor is constrained by the aircrafts maximum lift co-efficient. At higher velocities the load factor may be restricted as specified by the FAR25 regulations.

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7.3 WING STRUCTURE

From the preliminary report of wing sizing and selection, much of the data was revised to meet the RFP requirements and to further reduce weight in the final aircraft sizing. No major changes were amended, with only few variations to the parameters.

TABLE 29 - WING SPECIFICATIONS

Item Dimension

Wing Area (m2) 108.26

Wing Span (Without Winglets) (m)

31.98

Wing Span (With Winglets) (m) 33.98

Aspect Ratio 9.45

Taper Ratio 0.159

Dihedral (Degrees) 6

Sweep Angle (Degrees) 25.02

MAC (m) 3.38

Tip Chord (m) 1.25

Root Chord (m) 7.877

Although the majority of parameters are still based on the B737-900NG wing, the final wing design took into consideration the new materials and systems. The overall net effect resulted in higher lift to drag ratio while meeting a weight reduction. It is noticeable that the wing span will be shorter than its predecessor. A suitable fuel tank selection was also analyzed and integrated into our final design, including various wing systems such as leading and trailing edge devices.

FIGURE 52 – RIGHT SIDE OF WING (X-FOIL)

Wing Structure

The wings will primarily be made up of epoxy composites and a mixture of fibre glass with dual-path, fail-safe, two-spar structure. Bending Loads are carried by the front and rear spars and shear stresses are carried by the upper and lower skin panels. These four surfaces form the wing box which also serves as an integral fuel tank. The spars are reinforced by vertical stiffeners and the skin panels are reinforced by span wise integral stringers.

The original wing was designed to be short to keep weight to a minimum, but be large enough to carry fuel for the designed range. It also had to have a short chord to accommodate the engine which was to be flush mounted rather than on a pylon because of ground clearance. It also had

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the exact sweep than its predecessor of 25o because speed was not considered important compared to the majority of RFP objectives. As a result the average cruise speed will be Mach 0.78 which is used as a reference for the final design.

Having the engines wing-mounted also helped alleviate the bending moment of the wings from the lift, which allowed the spar weights to be reduced. However since the thickness of the wing is rather thin, the spars were strengthened and incorporated into our final design to overcome the design loads.

The basic wing geometry will remain the same as on the 737-900NG with 1m wingtip extensions and a slightly modified aerofoil for the LE slats. The wingtip extension came about when a flutter boom was being designed to eliminate flutter between the wing and the new power plant.

FIGURE 53 - WINGBOX STRUCTURE (RIBS AND SHEAR WEBS)

FIGURE 54 - WINGBOX STRUCTURE (SHEAR WEBS AND SKINS)

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Matching Chart

FIGURE 55 – MATCHING CHART

From the preliminary report, the matching chart provided a set of parameters and limitations for the final wing design. In order to achieve the optimum performance, some main factors were compromised; these were cruise speed and weight of the aircraft. From these governing restrictions, the final set of crucial parameters is calculated below:

Wing Area Calculation

1. Aircraft Weight, (W) = 708772.5 N (Conceptual Design weight Estimate Section) 2. Density at sea-level, ( ) = 1.225 kg/m3 (Assumed) 3. Stall Velocity, (Vst) = 77.06 m/s (B737-900NG) 4. Lift-Coefficient, (CL-MAX) = 1.8 (Roskam16)

EQUATION 6

Wing Span

EQUATION 7

9.45 108.26 31.99

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Mean Aerodynamic Chord

EQUATION 8

Wing Loading

EQUATION 9

7.3.1 Wing Systems

Fuel Tanks

From the RFP, integral tanks will be utilized for fuel storage in the aircraft. In comparison to other systems available, integral tanks provide the greatest weight saving and allows a combination of aircraft stability advantages.

FIGURE 56 - FUEL TANK LAYOUT INSIDE THE AIRCRAFT.

The tanks will be situated mainly in the wings internal structure and belly of the fuselage. However due to the load limit on the wing and thinner design, fuel will also be situated in the tail of the plane. In addition to the fuel tanks situating at the tail section, this would provide some stability advantages for the aircraft.

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Fuel Load Distribution

50% - Wing internal structure

30% - Belly of fuselage

20% - Tail internal structure

FIGURE 57 - INTEGRAL FUEL TANK INSIDE THE WING.

Since these tanks are part of the aircraft structure, they cannot be removed for service or inspection. Inspection panels will be provided to allow internal inspection, repair, and overall servicing of the tank.

Trailing Edge Devices - Trailing Edge Flaps

Based on the B787 configuration, the aircraft will use the simple single slotted flaps. The flaps will be designed at approximately 30% of the chord with 3 default angle settings. They will be located near the inboard section of the wing under the ground spoilers and flight spoilers.

The single slotted flap is basically a section of the wing at the trailing edge that extends rearward as it rotates downward. Airflow from the lower surface of the wing passes through it on to the upper surface of the flap to produce higher lift. Other purposes include reducing the stall speed and promote good low-speed handling qualities.

FIGURE 58 - SINGLE SLOTTED FLAP

Default Angles:

0 degree angle – Cruise condition

20 degree angle – Take Off condition

40 degree angle – Landing condition

Compared to the triple slotted flaps used in the 737-900NG, the single slotted flaps were chosen entirely based on the RFP requirements. As flaps during landing contributes over 40% of noise of a modern airliner, the single slotted flaps is proven to reduce noise level by up to 23% while producing the same induced drag to the initial configuration. This configuration also leads to mechanical simplicity and less risk of mechanical breakdown resulting in low maintenance cost.

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FIGURE 59 - FLAP LOCATION ON WING

Leading Edge Devices - Slats

Similar to the 737-900NG, the final wing design comprise of 8 slats outboard of the engines. The purpose of the slats will give 4% improvement in maximum lift to drag, a 0.02 Mach increase at maximum lift to drag, and a 3% reduction in block fuel at flight range over 1500 nautical miles. The slats will be located from the engine pylon to wingtip and give an average chord increase of 4% over the whole wing and will have similar approach speeds to the 737-900NG.

FIGURE 60 - SLATS

Leading Edge Flaps

There will be four leading edge Krueger flaps, two inboard of each engine. Their only function is either fully extended or retract. When the flap is retracted, the folding nose section rotates and is stored under the wing as. These flaps are used primarily for landings to increasing the wing camber and maximum coefficient of lift.

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FIGURE 61 - LE FLAPS

Spoilers and Speed Brakes

Like a majority of the onboard wing systems, the spoilers and speed brake will be retrofitted from the 737-900NG. The flight and ground spoilers are powered by hydraulic systems and can be continued to operate with speed brakes deployed.

FIGURE 62 - SPOILERS AND SPEED BRAKE LOCATIONS.

Limitations

In the final design there will be limitations for wing operations. These limitations are predominately standards set by Boeing for its 737-900NG variant. Although this final design is relatively different, the same guidelines will be followed as most of these systems are tested for optimum performance and safety.

TABLE 30 - LIMITATIONS

Condition Magnitude

Speed limit If only 1 leading edge device remains extended. M0.65

Speed limit if more than 1 leading edge device remains extended. 230kts

Do not deploy speed brakes in flight at radio altitudes less than: 1000ft

Holding in icing conditions with flaps extended is prohibited. -

Do not use speed brake when flaps are beyond 15. -

Max flap extension altitude 20,000ft

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7.3.2 X-Foil Model Design

This sections objective is to simply model the projected aircrafts wing geometry as well as the B737-900NG’s wing geometry and to compare results. This comparison will verify whether the projected aircrafts model is working correctly. This model will then be used to analyze other aspects of wing analysis such as performance, wing loading, wing sizing, load distribution etc.

Wing configuration analysis was performed by using a program called XFLR5 with capabilities in aerodynamic analysis for 2D and 3D models. The program is based on the XFOIL algorithm.

The XFLR5 program was used to compare the projected aircrafts wing structure and a benchmark model (Boeing 737-900NG) in order to show the improvement of the wing. XFLR5 is an aerodynamics program which rapidly solves airfoil research and relatively complex wing design problems in 2D and 3D. In addition, it produces good graphic result and helps modify and enhance 2D and 3D wing designs with great effect.

For XFLR5 simulation, some assumptions were applied to get more precise results.

In 3-dimension wing simulations, the flow state of the air was in the same condition as stall condition. That means that velocity, density, kinematic viscosity and sound of speed are used in the same values; 73.1 (m/s), 1.225 (kg/m3), 1.5 x 10-5 (m2/s) and 340 (m/s) respectively. Mach number and the range of Reynolds number were employed as 2.15 x 105 and 6 x 106 to 4.6 x 107.

Results

FIGURE 63 - CL/CD VS CL GRAPH FOR B737-900NG WING

FIGURE 64 - CL/CD VS CL GRAPH FOR DESIGNED WING

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FIGURE 65 - 3D WING ANALYSIS FOR B737-900NG WING CONFIGURATION

FIGURE 66 - 3D WING ANALYSIS FOR DESIGNED WING CONFIGURATION

TABLE 31 - THE VALUES OBTAINED BY SIMULATION

α = 4° (deg) B737-900NG Projected Aircraft

CL 0.4079 0.404

CD 0.0117 0.0116

CL/CD 34.8624 34.6843

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TABLE 32 - THE VALUES THROUGH WEIGHT COMPARISON

B737-900NG Projected Aircraft

Weight (kg) 85100 73750

This model will now be used to investigate the effect of a 30o swept wing design on laminar flow as well lift-to-drag.

7.3.3 Sweep Angle Comparison – 25o and a 30o Swept Wing

FIGURE 67 - PRESSURE CONTOUR AND TRANSITION POSITION FOR B737-900NG WITH 25.02 DEGREES SWEEP ANGLE

FIGURE 68 - PRESSURE CONTOUR AND TRANSITION POSITION FOR B737-900NG WITH 30 DEGREES SWEEP ANGLE

Based on Figure 67 and Figure 68 above, the co-efficient of pressure is much lower for the 30o swept wing as you approach the wing tips. Meaning that the transition point has been pushed back towards the trailing edge of the wing.

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FIGURE 69 - PRESSURE CONTOUR AND TRANSITION POSITION FOR DESIGNED MODEL WITH 25.02 DEGREES SWEEP ANGLE

FIGURE 70 - PRESSURE CONTOUR AND TRANSITION POSITION FOR DESIGNED MODEL WITH 30 DEGREES SWEEP ANGLE

In Figure 69 and Figure 70 the same characteristics occurred as with the B737-900NG swept wing. Pressure is reduced as you approach the wing tips.

TABLE 33 - THE COMPARISON OF GLIDE RATIO (CL/CD) WITH 25.02° AND 30° SWEPT WING

(At CL≈0.4)

B737-900NG Designed model

θ 25.02° θ 30° θ 25.02° θ 30°

CL/CD 34.8624 35.2491 34.6843 34.8548

TABLE 34 - THE COMPARISON OF XCP POSITION IN 25.02° AND 30° SWEPT WING

(At CL≈0.4) B737-900NG Designed model

θ = 25.02° θ = 30° θ = 25.02° θ = 30°

XCp 5.562 6.428 5.643 6.256

Note: XCp is the centre of pressure’s streamwise (chordwise)position

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From Table 34 above the centre of pressure is clearly moving as the sweep is increased. This

means that laminar flow is becoming much smoother and occurring for much longer than the

25o condition. The 30

o sweep wing is indeed beneficial for a design. However, it has not yet

been implemented into the projected aircraft as research into this field was the initial

objective. Also, as sweep is increased the structure undergoes some changes whereby

bending moments and torque increase. When torque and bending moments increase it results

in the spars having to be much thicker in order to absorb and carry the added load. So overall

the laminar flow may be better, achieving it through the same structural configuration is the

problem.

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7.3.4 Wing Sizing calculations

TABLE 35 – NOMENCLATURE

Description Symbol Units

Distributed Compression Load ω lbf/in

Distributed bending Load ω1 lbf/in

Density ρ lbm/in3

Longest Side of Panel when calculating K a In

Inboard Spar Cap Area Ainboard in2

middle Spar Cap Area Amiddle in2

Outboard Spar Cap Area Aoutboard in2

Rib Area AR in2

Optimum Rib Area ARO in2

Stringer Area AS in2

Optimum Stringer Area ASO in2

Shortest Side of Panel when calculating K b in

Stringer Pitch b in

Optimum Stringer Pitch bo in

Wing Box Chord Length c in

Wing Chord Length cAERO in

Wing Box Height d in

Z-Stringer Width d in

Optimum Z-Stringer Width do in

Elastic Modulus E lbf/in2

Rib Modulus ER lbf/in2

Tangent Modulus ET lbf/in2

Compressive Stress fc lbf/in2

Compressive Surface Stress fs lbf/in2

Working Stress fw lbf/in2

Stringer Efficiency Factor F

Stringer Height h in

Optimum Stringer Height ho in

Moment of Inertia I in4

Rib Moment of Inertia IR in4

Optimum Rib Moment of Inertia IRo in4

Buckling Co-Efficient Ks Rib Pitch L in

Minimum Rib Pitch LMIN in

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Optimum Rib Pitch Lo in

Bending Moment M lbf in

Number of Ribs NR

Number of Stringers NS

Load on Flanges P lbf

Shear Flow q lbf/in

Wing Box Span Lw in

Skin Thickness t in

Equivalent Skin Thickness t*e in

Equivalent Thickness of Skin and Rib te in

Spar Cap Thickness tf in

Optimum Skin Thickness to in

Optimum Equivalent Skin Thickness teo* in

Rib Thickness tR in

Optimum Rib Thickness tRO in

Stringer Thickness tS in

Optimum Stringer Thickness tSO in

Web Thickness tW in

Weight per Panel W lbm

Weight per Panel per unit area WPUA lbm

Optimized Rib Weight WRO lbm

Rib Weight WR lbm

Stringer Skin Panel Weight WSP lbm

Stringer Weight WS lbm

Optimized Stringer Weight WSO lbm

Stringer Skin Panel Weight WSTRINGER lbm

Total Wing Weight WT lbm

Optimized Total Wing Weight WTO lbm

Web Weight WW lbm

Assumptions

The assumptions used in the stringer skin design are as follows:

The panels are simply supported All parts of the wing are constructed with the material Aluminum 2024 Maximum Allowable Stress is 0.2% proof Stress Spar caps (flanges) and spar webs would be manufactured by extrusion Ribs are to be manufactured by stamping or folding. Top and bottom skin are made from the same panels.

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The wing structure is assumed to be a box-structure Mean aerodynamic chord would be considered as a chord length Structural chord would be 60% of Mean Aerodynamic Chord The minimum boom area is 0.25 in2 Z-stringer efficiency is 0.95 Integral efficiency is 0.81 Flanges are of L-section The bending load distribution is 70% of MTOW on the inboard and 30%MTOW on the

middle. is assumed equal to Fitting factor is 1.15

TABLE 36 - BASIC STRUCTURAL AND LOADING DATA FOR CALCULATION

Data Value Unit

CStru 79.84236 in

Average height 11.811 in

Single wing span 629.72 in

ET 10600000 lbf/in2

ER 10600000 lbf/in2

8000 lbf/in

Safety factor 1.5 -

Fitting factor 1.15 -

0.2% stress 46716.7 lbf/in2

F (Z stringer) 0.95 -

F(Integral Stringer) 0.81 -

Density 0.1 lbf/in3

7.3.4.1 Z-Stringer Skin

Geometry and Dimensioning

FIGURE 71 - SCALE DIMENSIONING OF OPTIMIZED DESIGN FOR Z-STRINGER SKIN IN SINGLE CELL DESIGN

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Single Cell Dimensioning

A comparison will be made between the optimized and actual de-optimized design. The de-optimized design will take into consideration such constraints as maximum allowable material stresses and material gauge sizes. The following data is used for the calculation. As will be the case in all the stringer and spar sections, the distributed loading will be split equally between the structures. The loading is optimised in excel so as to approach to 0.2% proof stress strength with the consideration of safety factor=1.5 and fitting factor=1.15. Hence is the loading that will be carried by the skins of the aircraft. The other basic data used for the calculations are below.

Optimum Design

The optimum design will assume that maximum allowable material stress and gauge sizing of material sheets are not a limiting factor in the design criteria. Instead the dimensions created will be based entirely on the loading given and the material properties. The dimensions for the Ribs and stringer skin will be calculated in this section

Optimum Rib Pitch

An Optimized Rib Pitch is fundamental in obtaining the minimum weight for a given loading requirement. From this the optimum stringer skin design will be created. The rib pitch is expressed as:

EQUATION 10 - RIB PITCH

Where for an optimum Z-Stringer Skin:

Since the assumption of simply supported ribs is used:

EQUATION 11 - C VALUE

Solving for these, assuming that the material for the Ribs and the skins are constructed from the same material, Aluminium 2024 and that = =

From this the optimum rib pitch can be found. The distributed loading is increased be a safety factor of 1.5:

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The rib pitch will be divided equally along the wing box span. As there will not be an exact fit using the following spacing, some de-optimization of the results will have to be made. As this section assumes an optimized solution, the even spacing of the ribs will be ignored.

Optimum Rib Dimensions

Once the Rib pitch is attained, the optimum rib thickness can be calculated using the following method (Farrar, 1949):

EQUATION 12 - RIB OPTIMUM THICKNESS

In this case the distributed loading is also assumed in include the safety factor. Therefore:

As this is an optimum design, no common gauge thickness will be assumed. The equivalent thickness

can be found simply using the optimal relationship that exists between the

optimum rib thickness and the equivalent thickness:

EQUATION 13 - OPTIMUM EQUIVALENT THICKNESS

The equivalent thickness of the optimum rib and stringer skin is given by:

EQUATION 14 - EQUIVALENT RIB THICKNESS

From this the optimum rib area can be calculated:

EQUATION 15 - AREA OF RIB

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Moment of inertia:

EQUATION 16 - MOMENT OF INERTIA OF RIB

Once this has been calculated the working stress acting on the stringer skin can be calculated:

EQUATION 17 - WORKING STRESS ON STRINGER SKIN

And the compression surface stress can also be found:

EQUATION 18 - COMPRESSION SURFACE STRESS ON STRINGER SKIN

This is well above the compressive stress of the material, which is . This design cannot be used and a de-optimized design must be used for the rib spacing. However, for the sake of weight comparison the optimum stringer dimensions will be calculated.

To begin the calculation of the weight of the ribs, the number of ribs must first be calculated. This will have to be rounded up to the nearest whole number.

EQUATION 19 - NUMBER OF RIBS

Once rounded up, the weight of the optimum rib is given as:

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EQUATION 20 - WEIGHT OF OPTIMUM RIB

Optimum Stringer Dimensions and Weight

The optimum stringer pitch can be calculated as (Farrar, 1949):

EQUATION 21 - STRINGER PITCH

The optimum skin thickness can then be calculated(Farrar, 1949):

EQUATION 22 - STRINGER SKIN THICKNESS

From this the optimum stringer thickness can be calculated using the relationship

:

EQUATION 23 - STRINGER THICKNESS

The optimum stringer height is given as (Farrar, 1949):

EQUATION 24 - STRINGER HEIGHT

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From this the optimum stringer flange width can be calculated where the following relationship is known (Farrar, 1949):

EQUATION 25 - STRINGER FLANGE WIDTH

With this relationship the formula for the optimum stringer flange width is:

EQUATION 26 - FORMULA FOR OPTIMUM FLANGE

Given all the dimensioning and spacing the total weight of the stringer skin can now be calculated (Farrar, 1949).

EQUATION 27 - TOTAL WEIGHT OF STRINGER SKIN

Therefore the final weight is a combination of the stringer skin weight and the total rib weight:

EQUATION 28 - FINAL WEIGHT OF STRINGER AND RIB CONFIGURATION

Summary of Optimum Z Stringer Design

The final dimensions and geometry of the Optimized Z-Stringer Skin design is summarized below:

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TABLE 37 – SUMMARY OF SINGLE CELLED OPTIMUM FOR Z-STRINGERS

Description Symbol Values Units

Rib Pitch Lo 33.61 in

Rib Thickness tRO 0.0551 in

Rib Area ARO 1.85 in2

Rib MOI IRO 64.53 in4

Skin Equivalent thickness to* 0.22 in

Equivalent thickness teo 0.275 in

Working Stress fW 62670 lb/in2

Compression Surface Stress fS 62670 lb/in2

Skin Thickness to 0.0941 in

Stringer Thickness tso 0.0988 in

Stringer Pitch bo 2.52 in

Stringer Height ho 2.11 in

Stringer Flange Width do 0.745 in

Rib Weight WRO 102.45 lb

Stringer Skin Weight WST 2375.19 lb

Total Weight WTO 2477.64 lb

FIGURE 72 - SCALE DIMENSIONING OF OPTIMAL DESIGN FOR Z-STRINGER SKIN IN SINGLE CELL DESIGN

De-Optimized Design

The de-optimized design will take into consideration such constraints as maximum allowable material stresses and material gauge sizes. Z-stringer efficiency is 0.95. Ribs and stringer skin are

considered to be made of the same material, Aluminium 2024 Alloy, therefore same Elastic Modulus

is used for both sections ( ) and "Stiffness based" ribs used for the calculations.

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De-Optimized Rib

The minimum rib pitch for the working stress using the 0.2% proof stress criteria is given as

EQUATION 29 - MINIMUM RIB PITCH FOR THE MAX ALLOWABLE WORKING STRESS

to obtain an even solution for the rib pitch over the wing semi-span, we will select to be 66.28. This allows for an exact number of ribs to be evenly spaced along the wing.

The working stress is calculated by using:

EQUATION 30 - WORKING STRESS FOR DE-OPTIMUM DESIGN

The resultant design solution will account for critical stress limitations of the material. As it is under the 0.2% proof stress of the material it is an acceptable design case.

From this the optimum equivalent thickness can be calculated:

EQUATION 31 - OPTIMUM EQUIVALENT THICKNESS FROM WORKING STRESS

Then by using:

EQUATION 32 - EQUIVALENT THICKNESS

Solving for the equivalent thickness where

:

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The standard gauge equivalent thickness is 0.58 in (Gauge 000000)

The Rib thickness is:

EQUATION 33 - RIB THICKNESS

The stand gauge rib thickness is 0.2294 in (Gauge 3)

By calculating rib thickness, the de-optimized geometry can be calculated. The rib cross sectional area is given by:

EQUATION 34 - RIB AREA

in2

And the Moment of Inertia by:

EQUATION 35 - RIB MOMENT OF INERTIA

in4

Now the skin thickness can be calculated using (Mileshkin & Bayandor, 2009):

EQUATION 36 - RIB THICKNESS

in

Weight of the Ribs is:

EQUATION 37 - WEIGHT OF RIBS

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De-Optimized Stringer

For a better design configuration we de-optimize the stringer section as follows:

First the stringer height is calculated:

EQUATION 38 - STRINGER HEIGHT

Then the stringers pitch:

EQUATION 39 - STRINGER PITCH

Now the de-optimized area using the following formula:

EQUATION 40 - STRINGER AREA

in2

Referencing to Farrar for analysis the de-optimised skin thickness is:

EQUATION 41 - RIB THICKNESS WITH FARRAR FACTOR

Standard gauge stringer skin thickness is 0.1443 in (Gauge 7)

So therefore the flange width can be calculated:

EQUATION42 - STRINGER FLANGE

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in

Now the weight per panel is calculated:

EQUATION43 - WEIGHT PER UNIT AREA

lbm

Then we get the weight of one of the stringer skin panel that is located between two ribs

EQUATION 44 - WEIGHT OF STRINGER SKIN PANEL

lbm

So now for the de- optimum design we got a total weight of 83.557 pounds. And then we can calculate the total weight of the stringer skins, which is given by:

EQUATION 45 - TOTAL WEIGHT OF STRINGER SKINS

=4516.32

The total weight is that of the design is the combination of the total weight of the ribs and stringer skin from Equation 28:

EQUATION 46 - TOTAL OF WEIGHT OF RIBS AND STRINGER

Summary of De-Optimum Z Stringer Design

The final dimensions and geometry of the De-Optimized Z-Stringer Skin design is summarized below:

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TABLE 38 - SUMMARY OF SINGLE CELL DE-OPTIMIZED DESIGN FOR Z-STRINGERS

Description Symbol Value Units

Minimum Rib Pitch 66.28

Working Stress 44629.79

Compression Surface Stress 46,716

Equivalent Thickness 0.58

Rib Thickness 0.214 De-Optimized Geometry 15.2 2

Moment of Inertia 530.26 4

Skin Thickness

Stringer Thickness ts 0.1388 in

Stringer Height

Stringer Pitch 3.97 2

De-Optimized Area 2

Flange Width Weight Per Panel

Weight of Ribs 432.66

Weight of Stringer Skin 163.63

Total Weight of Stringer Skins 4616.318

Total Weight 4948.975

FIGURE 73 - DEMENSION OF THE Z-STRINGER

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7.3.4.1.1 COMPARISON OF OPTIMUM AND DE-OPTIMIZED Z STRINGER DESIGNS

Now with the completed designs it is possible to compare the dimensions and weights for each. Table 39 shows the comparison.

TABLE 39 - COMPARISON OF OPTIMUM AND DE-OPTIMIZED DESIGNS

Description Optimum De-Optimized

Units Symbol Value Symbol Value

Rib Pitch 33.61 66.28 in

Equivalent Surface Thickness 0.22

0.309 in

Rib Thickness 0.0551 0.214 in

Equivalent thickness for ribs and skins 0.275 0.58 in

Rib Cross Sectional Area 1.85 15.2 in2

Rib Moment of Inertia 64.53 530.26 in4

Number of Ribs 20 20 n/a

Skin Thickness 0.0941 in

Stringer Thickness 0.0988 0.1388 in

Stringer Pitch 2.52 3.97 in

Stringer Height 2.11 in

Working Stress 62670 44629.79

Compression Surface Stress 62670 46,716

Rib Weight 102.45 432.66 lbm

Stringer Weight 2375.19 4616.318 lbm

Total Weight 2477.64 4948.975 lbm

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7.3.4.2 Integrated Stringer Skin Design

Geometry and Dimensions

FIGURE 74 - SCALE DIMENSIONING OF OPTIMIZED DESIGN FOR INTEGRATED STRINGER SKIN IN SINGLE CELL DESIGN

This section will investigate the use of an integral stringer skin. The design for the integral stringer skin will be based on the theory found in the journal entry of Catchpole (Catchpole)

Optimum Integrated Stringer Skin Design

Ribs

Initially the Rib Pitch will need to be calculated as this is the critical design for the wing. This sizing will in turn determine the overall panel size. This optimum design will determine the minimum possible weight for both the surface and the ribs for the given loading. The feasibility of the design will not be considered at this stage as a de-optimized design may be required.

The rib pitch is defined as:

EQUATION 47 - OPTIMUM RIB PITCH

Where:

Then for the one-cell configuration it is assumed that all ribs will be simply supported.

Assuming this, it can be stated that

(Mileshkin & Bayandor, 2009). Additionally it is

assumed that the ribs and stringer skin are constructed from the same material. Due to this it can be stated that .

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The using the theory from Catchpole it is possible to find the value for Farrar’s efficiency ( ) for the integral stiffened skin. As the goal for this design is to find the optimum conditions it is aimed to find the minimum possible weight, which occurs when “F” reaches its maximum(Farrar, 1949). This value is then found from the graph relating the integral stiffeners to the Farrar’s efficiency; this is taken from theory in Catchpole and can be seen below (Catchpole)Using this graph in Figure 75 below it can be stated that:

FIGURE 75 – DESIGN CHART FOR UNFLANGED INTEGRAL STIFFENERS

Then it can then be stated that the is the distributed load on the wing. This load will also be multiplied by a safety factor of 1.5 and fitting factor of 1.15.

Hence using these parameters it is possible to calculate the rib pitch where:

With the rib pitch known it is possible to now calculate the corresponding dimensions of the rib. The first dimension to be sized is the optimum rib thickness ( ). This can be found from the thickness equations where:

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ω

EQUATION 48 - OPTIMUM RIB THICKNESS

From this it is possible to calculate the equivalent surface thickness, that will coincide with

the rib thickness. This can be found from:

EQUATION 49 - EQUIVALENT SURFACE THICKNESS

Using this, an equivalent thickness for the ribs and skin can be found, where:

EQUATION 50 - EQUIVALENT THICKNESS

Additionally it is possible to calculate the optimum geometry of the wing. This will include calculating both the cross sectional area of the rib, and its moment of inertia, this can be found where:

EQUATION 51 - OPTIMUM RIB AREA

EQUATION 52 - OPTIMUM RIB MOMENT OF INERTIA

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The ribs will be subjected to two main forces that they will need to withstand. These are the working stresses and the compression surface stresses . The working stress will once again be based on the Farrar’s efficiency factor and can be found where:

EQUATION 53 - WORKING STRESS

And the compression surface stress can be found where:

EQUATION 54 - COMPRESSION SURFACE STRESS

This indicates that the maximum force that the rib would be required to take would be though as can be seen from Table 36 the 0.2% proof stress is only

. This means that the stresses applied are too high and the structure will fail. This would mean a de-optimization of the ribs would be required, though the analysis of the optimum rib condition will continue so comparisons in the final results can be seen. Additionally all of the thickness will not be altered and changed to standard gauge thickness for the optimum solution as this would effectively be de-optimizing the structure.

In order to calculate the weight of the ribs it is important to determine the number of ribs that will be required This can be found where:

EQUATION 55 - NUMBER OF RIB

This means that the required number of ribs is 40. By knowing the number of ribs it is possible to calculate the corresponding optimum rib weight, .

EQUATION 56 - OPTIMUM RIB WEIGHT

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Where the density of the material is

Stringers

Now that the ribs have been sized it is now important to size the stringers. The initial dimension will be to size the optimum skin thickness for the integral stringer skin configuration. This can be found where:

EQUATION 57 - OPTIMUM SKIN THICKNESS

in

Using this it is then possible to find a relationship for the stringer thickness where:

EQUATION 58 - OPTIMUM STRINGER THICKNESS

Additionally it is possible to calculate the stringer pitch where:

EQUATION 59 - STRINGER PITCH

Utilizing this value it is then possible to calculate the stringer height where:

EQUATION 60 - STRINGER HEIGHT

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It is also possible to calculate the weight of the stringer skin. This is initially done by finding the weight per unit area which if found from:

EQUATION 61 - WEIGHT PER UNIT AREA

Using this it is then possible to calculate the total stringer weight which if found by multiplying the weight by then area of the panel, where:

EQUATION 62 - TOTAL STRINGER WEIGHT

The total weight is found from sum of weight of panel and stringer, where:

EQUATION 63 - TOTAL WEIGHT OF PANEL AND STRINGER

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Summary of Optimum Integral Stringer Design

With the information found on the optimum rib and stringer skin it is possible to develop a summary that can be seen in Table 40 below. Additionally in Figure 76 below it is possible to see a scaled sketch of the stringer skin geometry.

TABLE 40 - OPTIMUM INTEGRAL STRINGER SKIN SUMMARY

Optimum Design Parameters Symbol Value Units

Rib Pitch Lo 33.61 in

Rib Thickness tRO 0.0551 in

Optimum Surface Thickness to* 0.22 in

Equivalent thickness for ribs and skins teo 0.275 in

Rib Cross Sectional Area ARO 1.85 in2

Rib Moment of Inertia IRO 64.53 in4

Number of Ribs NRO 20 -

Skin Thickness to* 0.1048 in

Stringer Thickness tso 0.2358 in

Stringer Pitch bo 3.53 in

Stringer Height ho 2.29 in

Working Stress fwo 53434.61 lb/in2

Compression Surface Stress fso 62670.22 lb/in2

Rib Weight WRO 102.45 lbm

Stringer Weight WSO 2596.99 lbm

Total Integral Stringer Weight WO 2699.43 lbm

FIGURE 76 - SCALED DIMENSIONING OF OPTIMUM INTEGRAL STRINGER SKIN FOR SINGLE CELL DESIGN

De-Optimized Integrated Stringer Skin Design

As was stated during the Optimized design the forces on the structure would cause the wing to fail. Due to this some changes in the design parameters will need to be made. This would de-optimize the design but make it possible to carry the loads required. Additionally when the previous design was analysed all thickness above the minimum thickness requirements were assumed acceptable though in reality making thickness to these specific sizes would be very expensive. Due to this as the design is going to be de-optimized standard gauge thickness will be used. This will add to the weight to the structure though in turn will mean the design is much cheaper to construct.

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Ribs

As with the optimum design the initial parameter that needs to be calculated is the rib pitch. Though in order to size this rib pitch to withstand the loads, a minimum rib pitch, will need to be calculated. This minimum rib pitch will be the based on ensuring the working loads to not exceed the 0.2% proof stress. The relationship is found where:

EQUATION 64 - MINIMUM RIB PITCH

in

Although this is the minimum possible rib pitch that will correspond to the capable loads of the structure, being equal to that of the 0.2% proof stress, this length will not divide into the overall length of the structure. The mission of this project is to design 20 ribs on the wing. Hence the minimum rib pitch is 66.28 in. Using this chosen rib pitch it is possible to calculate the working stress, where:

EQUATION 65 - WORKING STRESS

lb/in2

As this value is now less than the 0.2% proof stress of the material the structure would be capable of withstanding the forced on it. This would mean that this design would be feasible.

It is then assumed that the thickness of the skin will be adequate to handle the compressive loads on the ribs, so it can be stated that:

Using this relationship ship it is then possible to calculate the equivalent thickness, of the rib where:

EQUATION 66 - EQUIVALENT THICKNESS OF RIB

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Utilizing this value it is then possible to calculate the calibrated rib thickness, where:

EQUATION 67 - RIB THICKNESS

As this thickness it above the minimum allowable thickness of 0.02 in it would be acceptable. Additionally as this is now the de-optimized design this will be changes to a standard gauge thickness. The closest gauge thickness that would be allowable would correspond to 0.0907 in, (Gauge 9)

From this it is then possible to calculate the equivalent thickness (t*) for the ribs and skin where:

EQUATION 68 - EQUIVALENT THICKNESS FOR RIBS AND SKIN

this is now the de-optimized design this will be changes to a standard gauge thickness. The closest gauge thickness that would be allowable would correspond to 0.5165 in, (Gauge 0)

Using these recalculated thicknesses it is possible to calculate the now de-optimized geometry of the wing. This included calculating both the cross sectional area of the rib, and the moment of inertia, where:

EQUATION 69 - CROSS SECTIONAL AREA OF RIB

EQUATION 70 - RIB MOMENT OF INERTIA

In order to calculate the weight of the ribs it would first be necessary to calculate the number of ribs, that would be required in order to withstand the loads on the wing.

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EQUATION 71 - NUMBER OF RIBS

Stringers

Now that the ribs have been sized for the de-optimized model it is important to size the stringers. The initial dimension will be to size the skin thickness, for the integral stringer skin configuration. This can be found where:

EQUATION 72 - SKIN THICKNESS

This skin thickness will then be changed to match a standard gauge thickness. The closet gauge thickness that would be appropriate for this thickness would be 0.1443 in (Gauge 9)(Carpenters Inc.). By then using this value it is possible to find a relationship for the stringer thickness, where:

EQUATION 73 - STRINGER THICKNESS

This thickness then would be converted to a standard gauge thickness of 0.3648 which corresponds to (gauge 10)

Additionally it is possible to calculate the stringer pitch, where:

EQUATION 74 - STRINGER PITCH

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Utilizing this value it is then possible to calculate the stringer height, where:

EQUATION 75 - STRINGER HEIGHT

It is also possible to calculate the weight of the stringer skin. This is initially done by finding the weight per unit area, which if found from:

EQUATION 76 - WEIGHT PER UNIT AREA

Using this it is then possible to calculate the total stringer weight which if found by multiplying the weight by then area of the panel, where:

EQUATION 77 - TOTAL STRINGER WEIGH

lbm

The total weight of integral stringer is

EQUATION 78 - TOTAL WEIGHT OF INTEGRAL STRINGER

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Summary of De-optimum Integral Stringer Design With the information found on the rib and stringer skin it is possible to develop a summary that can be seen in Table 41 below. Additionally in Figure 77 below it is possible to see a scaled sketch of the stringer skin geometry.

TABLE 41 - DE-OPTIMIZED INTEGRAL STRINGER SKIN SUMMARY

De-Optimized Design Parameters Symbol Value Units

Rib Pitch Lmin 43.98 in

Assumed Rib Pitch L 66.28 in

Equivalent Surface Thickness t* 0.363 in

Rib Thickness tR 0.0907 in

Equivalent thickness for ribs and skins te 0.453 in

Rib Cross Sectional Area AR 6.012 in2

Rib Moment of Inertia IR 209.65 in4

Number of Ribs NR 20 -

Skin Thickness t 0.1472 in

Stringer Thickness ts 0.33 in

Stringer Pitch b 3.527 in

Stringer Height h 2.29 in

Working Stress fw 38052.77 lb/in2

Rib Weight WR 171.06 lbm

Stringer Weight WS 3646.75 lbm

Total Integral Stringer Weight WT 3817.81 lbm

FIGURE 77 - SCALED DIMENSIONING OF DE-OPTIMIZED INTEGRAL STRINGER SKIN FOR SINGLE CELL DESIGN

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7.3.4.2.1 COMPARISON OF OPTIMUM AND DE-OPTIMIZED DESIGNS

Now with the completed designs it is possible to compare the dimensions and weights for each of them. In Table 42 below a full comparison can be seen. Overall it is shown that the forces on the de-optimized design are reduced, though as a result the thickness are increased and the weights are increased.

TABLE 42 - COMPARISON OF OPTIMUM AND DE-OPTIMIZED DESIGNS

Description Optimum De-Optimized

Units Symbol Value Symbol Value

Rib Pitch 33.61 66.28 In

Equivalent Surface Thickness 0.22 0.316 In

Rib Thickness 0.0551 0.0907 In

Equivalent thickness for ribs and skins 0.275 0.453 In

Rib Cross Sectional Area 1.85 6.012 in2

Rib Moment of Inertia 64.53 209.65 in4

Number of Ribs 20 20 n/a

Skin Thickness 0.1048 0.1472 In

Stringer Thickness 0.2358 0.33 In

Stringer Pitch 3.53 3.527 In

Stringer Height 2.29 2.29 In

Working Stress 53434.61 38052.77

Compression Surface Stress 62670.22 46716

Rib Weight 102.45 171.06 Lbm

Stringer Weight 2596.99 3646.75 Lbm

Total Integral Stringer Weight 2699.43 3817.81 Lbm

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7.3.4.2.2 COMPARISON OF DE-OPTIMIZED DESIGNS ON Z STRINGER AND INTEGRATED STRINGER

Now with the completed designs it is possible to compare the dimensions and weights for each of them. In Table 43 below a full comparison can be seen.

TABLE 43 - COMPARISON OF DE-OPTIMIZED Z-STRINGER AND DE-OPTIMIZED INTEGRAL STRINGER

Description De-Optimized Integrated De-Optimized Z

stringer

Units

Symbol Value Symbol Value

Rib Pitch 66.28 66.28 in

Equivalent Surface

Thickness

0.316 0.309 in

Rib Thickness 0.0907 0.214 in

Equivalent thickness for ribs

and skins

0.453 0.58 in

Rib Cross Sectional Area 6.012 15.2 in2

Rib Moment of Inertia 209.65 530.26 in4

Number of Ribs 20 20 N/A

Skin Thickness 0.1472 in

Stringer Thickness 0.33 0.1388 in

Stringer Pitch 3.527 3.97 in

Stringer Height 2.29 in

Working Stress 38052.77 44629.79 lb

Compression Surface Stress 46716 46,716 lb

Rib Weight 171.06 432.66 lbm

Stringer Weight 3646.75 4616.318 lbm

Total Weight 3817.81 4948.975 lbm

7.3.4.3 Spars and shear webs

Spar caps

The maximum distributed load is determined by Maximum Takeoff Weight = 162645 lbf = 723481 N. The distribution of weight on the wing structure is assumed as 70% on the inboard STN 0- 232 and 30% on the middle STN 232-463.

Load Distribution

L-section shape is employed in order to design the geometry of spar cap as shown in Figure 78 below.

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FIGURE 78 - THE SINGLE CELL CONFIGURATION WITH THE SPAR CAPS

Inboard (Station 0 – 232)

The load, ω of distributed flange in this section is calculated by the equation as:

EQUATION 79 - DISTRIBUTED LOADING ON SPAR

The bending moment, with respect to the distributed load is equal to:

EQUATION 80 - BENDING MOMENT

The force, of concentrated flange is equal to:

EQUATION 81 - CONCENTRATED LOAD

The cross-section area of spar cap designed to get over the concentrated force is equal to:

EQUATION 82 - AREA OF SPAR CAPS

The thickness of spar cap is equal to:

EQUATION 83 - THICKNESS OF SPAR CAPS

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The weight of spar caps in this section is equal to:

EQUATION 84 - WEIGHT OF SPAR CAPS

Middle (Station 232 – 463)

The load, of distributed the flange in this section is calculated as follows:

The bending moment, with respect to the distributed load is calculated as follows:

The force, P of concentrated flange is:

The cross-section area of spar cap designed to get over the concentrated force is equal to:

The minimum area of spar cap is 0.25 in2

The thickness of spar cap is:

At this middle section, the spar caps are tapered. It means that the cross-section area is determined as the average area between inboard and middle spar cap. Therefore, the weight of spar caps in this middle section is:

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Outboard (Station 463 – 630)

The load, of distributed flange in this outboard section is zero.

The cross-section area of outboard spar cap is selected as minimum gauge value:

The thickness of outboard spar cap is equal to:

As the spar caps are tapered, the weight of spar caps in this outboard section is also calculated with the average value of the middle and outboard area.

Now total weight of the spar caps can be found as:

Shear Webs

The shear flow on the wing box structure is assumed from thrust of engine. The thrust of engine is 23000 lbf. The shear flow is calculated with safety factor and fitting factor

EQUATION 85 - SHEAR FLOW ON SHEAR WEBS

The buckling shear coefficient, Ks for the thickness design of the shear web is determined by b/a ratio;

EQUATION 86 - B/A RATIO FOR BUCKLING KS

For

, the value, Ks is:

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The thickness of shear web can be calculated by the equation:

EQUATION 87 - THICKNESS OF SHEAR WEB

Note: for elastic design

From the standard thickness gauge 8,

Total weight of the shear webs can be calculated as:

EQUATION 88 - TOTAL SHEAR WEB WEIGHT

Summary of Shear Web and Spar Cap

For the one-cell configuration, the spar cap and web dimensions are summarized via using the tables.

TABLE 44 - SPAR CAP DIMENSIONS

Spar Cap Area (Ac) Spar Cap Thickness (tƒ) Spar Cap Weight (Wc)

Inboard Section 0.3617(in2) 0.1736 (in) 33.57 (lbm)

Middle Section 0.25 (in2) 0.1443 (in) 28.26 (lbm)

Outboard Section 0.25 (in2) 0.1443 (in) 16.7 (lbm)

Total weight (WT) 78.53 (lbm)

TABLE 45 - SPAR WEB DIMENSIONS

Spar Web Thickness (ts) Spar Web Weight (Wweb)

Dimension 0.1285 (in) 191.15 (lbm)

The total weight of single cell spar configuration is:

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7.3.4.4 Summary of Wing Box Structure

Assessment of the above tables indicates that the integrated stringer skin design produces the lightest weight wing. Although the Z-Section design is the preferred method for manufacture, the Integral stringer is preferred method for the projected aircraft. Therefore this result is good for weight saving.

The total weight of the whole wing box structure is

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7.3.5 Patran Analysis

Introduction

This analysis is performed in order to analyze whether the wing structure can withstand the loads during the cruise condition and a 2.5g loading condition during flight. To analyze these two conditions the lift-distribution loads across one wing obtained from the X-foil analysis will be distributed along the elements of the Patran model. Discussion of results will be provided and recommendations in improving the design will be supplied. It will also analyse maximum deflections and stress concentrations

Modelling the Structure

Wing Dimensions

FIGURE 79 - WING DIMENSIONS

Lift loads

The lift was distributed as seen in Figure 80 below.

FIGURE 80 – LIFT LOADS

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FIGURE 81 - LIFT DISTRIBUTION

The image on the bottom right of Figure 81 above represents the local lift as a percentage of total lift across the wingspan of one wing. This distribution will be averaged out in the Patran model analysis across the top skin elements. There is an assumption that the number of elements is constant from the wing root to the wing tip. From the Patran model there are 120 nodes across the top skin. This is then divided by three in order to determine the number of nodes in each section as seen in Figure 81.

To apply the loads using the lift distribution the minimum and maximum values for every 200 inches were found and the average value was calculated. This value was then divided by the number of elements in that particular station thus providing the distributed load in the Patran analysis. The loads are different for each section and can be seen in and Figure 82 below.

TABLE 46 - LIFT DISTRIBUTION PER NODE

Lift for one Wing [N]

361743.75

Inboard Section Mid Section

Outboard Section

Nodes X-Axis 7 6 6

Nodes Y-Axis 6 6 6

Total Nodes 42 36 36

Lift-Distribution %

47-38 38-26 26-14

Average 153741.0938 115758 72348.75

Force Per Node [N]

3660.502232 3215.5 2009.6875

Note: the nodes in the x-axis are the number of nodes across the top surface of the skin going from the wing root to the wing tip, the nodes in the y-axis are the number of nodes also on the top surface going from the leading edge to the trailing edge.

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FIGURE 82 - DISTRIBUTED LIFT LOADS

Engine Loads

The engine loads were applied at the same location as seen in the geometry of the aircraft and can be seen in Figure 83 below. The engine was attached at the 8th rib from the wing root and in order to determine the engine load distribution, the engine weight was divided by the number of node elements along that rib which was 6.

TABLE 47 - ENGINE LOADS

Cruise Condition 2.5g Loading Condition

Engine Weight [N] 23249.7 58124.25

Number of Nodes 6 6

Force Per Node [N] 3874.95 9687.38

FIGURE 83 - ENGINE LOADS (2.5G LOADING)

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Fuel Loads

In order to first determine the fuel weight some assumptions were made according to the amount of fuel allocated to the wings structure. The total fuel weight is 189300.93 N and the following portions were used:

1. Empennage – 20% = 37860.186 N 2. Fuselage – 30% = 56790.279 N 3. Wings – 50% = 94650.465 N

a. Each Wing = 47325.23 N

Therefore, it is assumed that the fuel will be distributed along two thirds of the structure from the wing root as seen in Figure 84 below.

FIGURE 84 - FUEL LOADS (2.5G LOADING)

The load per node is provided below.

TABLE 48 - FUEL LOADS

Cruise Condition 2.5g Loading Condition

Fuel Weight [N] 47325.23 118313.08

Nodes X-Axis 13 13

Nodes Y-Axis 6 6

Total Nodes 78 78

Force Per Node [N] 606.73 1516.83

Materials

The whole structure was modelled using Aluminium Alloy 2024. The specifications are provided below.

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TABLE 49 - ALUMINIUM ALLOY PROPERTIES

Aluminium Alloy 2024 Ult. Yield strength 60000 Psi

Yield strength 40000 Psi

0.2% proof stress 322.1 MPa

Young’s Modulus 10600000 lbf/in2

Density 0.1 lbf/in3

Shear Modulus 4060 ksi

Shear Strength 41000 lbf/in2

TABLE 50 - WING THICKNESSES

AIRCRAFT WING Number of Ribs 20

[in] [cm] [m] [mm]

Rib Thickness 0.0907 0.230378 0.002304 2.30378

Skin Thickness 0.3648 0.926592 0.009266 9.26592

Wing Span 629.7232 1599.497 15.99497 15994.97

Shear Web Thickness

0.1101 0.279654 0.002797 2.79654

[in2] [cm2] [m2] [mm2]

Spar Cap Area 0.36169 2.333479 0.000233 233.3479

Constraints

Constraints were applied in all 6DOF for the wing root as seen in image below.

FIGURE 85 - CONSTRAINTS

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Results

Cruise Condition – Maximum Displacement

FIGURE 86 - MAXIMUM DISPLACEMENT

The maximum deflection obtained for the cruise condition was 24.9 inches in the vertical direction. This is quite reasonable and the deflection rate across the wingspan is what was expected.

Cruise Condition – Stress Concentration

FIGURE 87 - STRESS

From this image it shows the maximum stresses due to the specified loading condition. It was expected that the shear webs and spar caps would carry the most amount of load as they are responsible for carrying the bending moments of the wing structure. A maximum stress of 38.4 ksi had developed at the wing root which is below the ultimate load of 60 ksi of Aluminium Alloy 2024 meaning the structure is not failing for the cruise condition.

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2.5g Loading Condition – Maximum Displacement

FIGURE 88 - MAXIMUM DEFLECTION

The maximum deflection occurred at the wing tip with a magnitude of 20.9 inches in the vertical direction as expected.

2.5g Loading Condition – Stress Concentration

FIGURE 89 - STRESS

In Figure 89 above the maximum stress of 26.1 ksi developed at the location of engine attachment which is far below the ultimate load of 60 ksi of Aluminium Alloy meaning the structure is not failing. This is expected in the 2.5g loading condition as the weight loads would multiple by a factor of 2.5g’s. The recommendation for this case is, if required, is to optimize the design of the rib at this location, either thickening the rib or changing the engine attachment location. However, as this analysis is simply to determine the areas of maximum stress, optimization is not necessary for this analysis.

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Discussion

This analysis was carried out in order to determine the areas which are in need of improvement for two conditions: cruise and the 2.5g loading condition. Results showed that the areas in need of improvement are the engine mounting locations and spar dimensions. If necessary after future analysis, the shear web thicknesses as well as the number of ribs will be optimized to reduce the weight.

In terms of comparison to the expected real conditions of the structure, a reduced number of ribs is possible as the current model is based on the B737-900NG series which weighs approximately 10% heavier and does not utilize the same level of composites or technologies. Therefore, less ribs may be possible as well as thinner ribs, shear webs and skins allowing for a lighter structure.

Conclusion

1. Computational results show that the wing structure will not fail under both loading conditions for Aluminium Alloy 2024.

2. The structure can be optimized for a lighter wing by reducing the number of ribs of reducing the thickness in either the skins, ribs and/or shear webs.

3. An optimization of the rib where the engine is attached can be thickened more so than the other ribs which will in turn result in the other ribs being much thinner.

4. The use of a higher mesh in the analysis may result in higher accuracy but is not necessary.

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7.3.5.1 Validation

Validation is performed for the first rib section in Figure 90 below.

FIGURE 90 - CRUISE CONDITION STRESS (PATRAN)

Assumptions:

1. A final value which is within the same magnitude as the Patran analysis for stress will be considered accurate.

2. The value determined from the following method gives the average stress value over the rib and not specifically for the corner as seen in Figure 90 above.

Conversion to Point Loads

FIGURE 91 - POINT LOADS

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TABLE 51 - LOAD UNIT CONVERSION

Item Force [N] Shear Force [lbf]

L1 153741 34560.98

WF -47325.23 -10638.71

WE -23249.7 -5226.53

L2 115758 26022.40

L3 72348 16263.83

Note: Dimensions and loads were all derived from CAD dimensions and Figure 43 as seen in the “Patran Analysis” section.

Cross-Sectional Geometry

FIGURE 92 - GEOMETRY

Moment of Inertia

For 4 booms the calculation is as follows.

Shear Force Diagram

FIGURE 93 – SFD

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Torque

TABLE 52 - TORQUE CALCULATIONS

Rib Distance from

origin [in] Lift Force

[N] Fuel Weight

[N] Engine Weight

[N] Torque [lbf.in]

(Innermost Rib) 1

93.68 21963.0 -3640.4 0 385860.7

2 101.70 21963.0 -3640.4 0 804755.2

3 109.72 21963.0 -3640.4 0 1256683.4

4 117.74 21963.0 -3640.4 0 1741645.4

5 125.76 21963.0 -3640.4 0 2259641.1

6 133.78 21963.0 -3640.4 0 2810670.7

7 141.80 21963.0 -3640.4 0 3394733.9

8 149.82 21963.0 -3640.4 -23249.7 3228791.9

9 157.84 19293.0 -3640.4 0 3784184.1

10 165.86 19293.0 -3640.4 0 4367796.3

11 173.88 19293.0 -3640.4 0 4979628.6

12 181.90 19293.0 -3640.4 0 5619680.8

13 189.92 19293.0 -3640.4 0 6287953.1

14 197.94 19293.0 0 0 7146432.0

15 205.96 12058.1 0 0 7704720.9

16 213.98 12058.1 0 0 8284749.3

17 222.00 12058.1 0 0 8886517.3

18 230.02 12058.1 0 0 9510024.8

19 238.04 12058.1 0 0 10155271.8

(Outermost Rib)20

246.06 12058.1 0 0 10822258.4

FIGURE 94 - TORQUE DIAGRAM

0

2000000

4000000

6000000

8000000

10000000

12000000

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20

To

rqu

e [

lbf.

in]

Rib Station

Torque at Rib Locations

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Bending Moment

TABLE 53 - BENDING MOMENT CALCULATIONS

Rib

Distance from origin [in]

Lift Force [N]

Fuel Weight [N]

Engine Weight [N]

Bending Moment [lbf.in]

1 29.85 21963.0 -3640.4 0 1967827.0

2 29.85 21963.0 -3640.4 0 1844868.9

3 29.85 21963.0 -3640.4 0 1721910.8

4 29.85 21963.0 -3640.4 0 1598952.8

5 29.85 21963.0 -3640.4 0 1475994.7

6 29.85 21963.0 -3640.4 0 1353036.6

7 29.85 21963.0 -3640.4 0 1230078.5

8 29.85 21963.0 -3640.4 -23249.7 1107120.4

9 29.85 19293.0 -3640.4 0 1140184.8

10 29.85 19293.0 -3640.4 0 1035144.4

11 29.85 19293.0 -3640.4 0 930104.1

12 29.85 19293.0 -3640.4 0 825063.7

13 29.85 19293.0 -3640.4 0 720023.4

14 29.85 19293.0 0 0 614983.0

15 29.85 12058.1 0 0 485512.9

16 29.85 12058.1 0 0 404594.1

17 29.85 12058.1 0 0 323675.3

18 29.85 12058.1 0 0 242756.4

19 29.85 12058.1 0 0 161837.6

20 29.85 12058.1 0 0 80918.8

FIGURE 95 - BENDING MOMENT DIAGRAM

0

500000

1000000

1500000

2000000

2500000

1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20

Be

nd

ing

Mo

me

nt

[lb

f.in

]

Rib Station

Bending Moment at Rib Locations

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Shear Flow Calculations

FIGURE 96 - SHEAR FLOW CUT

Using the shear flow equations below, the shear flows for each section can be calculated.

Where:

S = Shear Force (lbf) I = Mass Moment of Inertia (in4) An = Effective Area (in2) yn = Distance from Horizontal Centreline (Marked blue in Figure 96) (in) qo = Previous Shear Flow Value (lbf.in)

EQUATION 89 - SHEAR FLOW EQUATION

The shear force for the first rib can be determined from Table 51 above(underlined).

Panel 2-3

Panel 3-4

Panel 4-1

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FIGURE 97 - RESULTANT SHEAR FLOWS (UNCORRECTED)

Shear Flow Correction

To correct the shear flows the torque from Table 52 for the first shear web was used (underlined).

Using the torque equation below the correction shear flow value can be determined.

Where:

AE = Effective Area (in2)

qb = Basic Shear Flow (lbf.in)

q = Corrected Shear Flow Value (lbf.in)

Solve for q:

Add the corrected shear flow value to the shear flows calculated in the previous section (see Figure 97).

TABLE 54 - CORRECTED SHEAR FLOWS

Section Basic Shear Flows [lbf.in] Corrected Shear Flow Value [lbf.in]

1-2 0 54.11

2-3 -397.99 -343.88

3-4 0 54.11

4-1 397.99 452.1

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FIGURE 98 - CORRECTED SHEAR FLOWS

Stress Calculation

Using the equation below the stress in the panel can be determined.

Panel 1-2

Panel 2-3

Comparison

Patran analysis revealed a maximum stress value of 38400 psi and validation showed a maximum of 14931.27 psi. These values are within the same magnitude and due to assumptions such as the absence of taper as well as the averaging effect of the calculations for shear flow, this validation is deemed successful.

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7.4 FUSELAGE AND CABIN SIZING

The fuselage and cabin sizing performed for this analysis is strictly preliminary and subject to change. Basic shape and length will remain relatively the same whereas the in-cabin dimensions and possibly maximum seating capacities may vary.

Fuselage Length

In the process of fuselage sizing, RFP requirements basically provided the minimum parameters for design. It stated that the design must be able to handle a minimum of 175 passengers, a cargo capacity of 1240 ft3 and a seating pitch and width of 32 and 17.2 inches respectively.

So by using these values we can safely determine the necessary floor space for a specific number of passengers. We aimed for a capacity of 200 passengers as our aircraft must be deemed as a replacement aircraft of current models. The B737-900 can currently carry a maximum of 215 passengers whereas the A320-200 can carry a maximum of 180 passengers. Therefore we aimed for approximate midpoint as a starting point.

Most geometry was found during calculations performed of weight, performance and weight and balance. Another method to find dimensions was to use a scaled image of the B737-900NG (see appendix 8). This provided rough values of location and size for sections such as the nose, tails, doors, lavatories, windows, emergency exits, cockpit and galleys.

Using this method of sizing showed that the aircraft is capable of a passenger capacity of 204. This value may be altered at a later stage of design to allow for more passengers. The aisle width was found to be 20 inches by studying other aircrafts of similar size. This allows passengers enough room to transverse the aircraft without compromising the cabin width. The galleys and toilets were later added and thus the minimum overall aircraft length was established. The seating arrangement and fuselage shape are shown below.

It must be noted that during the weight estimate calculations, a passenger capacity of 200 was used as a baseline value before geometry of the aircraft designed. The reason the capacity was neglected were for two main reasons. The first is that difference of 4 passengers at this stage would only provide slightly more accurate values. The other reason is that the aircraft geometry and weight will be reanalysed in the detailed design phase and the change can be made then. Therefore, in terms of time constraints for recalculating weight estimates, it would be more important to have a rough estimate at this stage rather than an incomplete weight and geometric analysis.

FIGURE 99 - IN-CABIN SEATING ARRANGEMENT AND FUSELAGE LENGTH

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In-Cabin dimensions

The in-cabin dimensions were found by analysing the geometry of the B737-900NG. Using CAD we were able to design a proportional design which allows enough room between seats, overhead compartments and head room.

The figure below was also able to assess the available room for cargo as well as cables running longitudinally through the aircraft.

FIGURE 100 - FUSELAGE CROSS SECTION

Cargo Space

From Figure 100 the estimated cargo space can be calculated. As the wing box occupies part of the cargo it is taken into account by subtracted the wing root length from the cabin floor space. The calculation is as follows.

Data

1. Cargo space cross sectional area, (SCargo) = 22.3 ft2 2. Cabin Length, (LCabin) = 93.42 ft 3. Wing Root Length, (CRoot) = 7.877 m = 25.83 ft

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7.5 EMPENNAGE SIZING

The empennage for the projected aircraft was based on the B737-900NG aircraft and is a conventional wing configuration. The specifications can be seen in Table 55 and Table 56 below. This wing structure will be tested in a later section for stability.

TABLE 55 - HORIZONTAL STABILIZER GEOMETRY

Horizontal Stabilizer

Symbol Value Unit Value Unit Condition Notes

Aspect Ratio ARH 6.16

B737-900NG

Thickness tH 0.2 m 0.66 ft

Assumption

Mean Aero Chord

MACH 2.31 m 7.57 ft ( / )

Root Chord CROOT 3.84 m 12.58 ft 2*MAC/(1+λ)

Tip Chord CTIP 1.29 m 4.25 ft 2*MAC*λ/(1+λ)

Taper Ratio λH 0.203

B737-900NG

Span bH 14.21 ( / )

Distance From Fuselage Centreline to MAC

YHS 2.77 m 9.08 ft (λ/6)*(1+2λ)/(1+λ)

Tailplane Area STH 32.78 m2 352.84 ft2

B737-900NG

Elevator Area SE 6.55 m2 70.50 ft2

B737-900NG

Dihedral ΛH 7 degrees

B737-900NG

1/4 Sweep Angle

30 degrees

B737-900NG

TABLE 56 - VERTICAL TAILPLANE GEOMETRY

Vertical Stabilizer (FIN)

Symbol Value Unit Value Unit Condition Notes

Fin Height

7.16 m 23.49 ft

B737-900NG

Fin Area SF 26.44 m2 284.60 ft2

B737-900NG

Rudder Area SRudder 5.22 m2 56.19 ft2

B737-900NG

Aspect Ratio AR 1.91

B737-900NG

Taper Ratio λ 0.271

B737-900NG

Thickness tF 0.2 m 0.66 ft

assumption

Mean Aero Chord MACF 3.72 m 12.21 ft sqrt(SF/AR)

Root Chord CRoot 5.85 m 19.21 ft 2*MAC/(1+λ)

Tip Chord CTip 1.59 m 5.21 ft 2*MAC*λ/(1+λ)

1/4 Chord Sweep Angle

35 degrees

B737-900NG

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Empennage Theory

For the empennage the aspect ratio of either the vertical or horizontal stabilizer tends to be smaller than a wing aspect ratio. However, this low aspect ratio results in less bending moment due to the smaller span. In terms of concentrated loads, only the hinges supporting the movable surfaces must be taken into account. These control surfaces are not constructed with an emphasis of strength only, stiffness of the structure is also important in order to prevent flutter. Large deflections at the empennage main box will cause the movable surfaces to experience severe loads and it may even cause banding at the hinge brackets. Here are some typical constructions:

1. Single spar construction with auxiliary rear spar and all bending material concentrated in the spar cap only. This design is mostly in light aircraft.

2. Two spar construction with all bending materials concentrated in the spar caps. 3. Multi-spar construction with spars resisting all the bending loads such as the DC-10

fin structure.

FIGURE 101 - BOEING 767 VERTICAL TAILPLANE19

FIGURE 102 - BOEING 767 HORIZONTAL STABILIZER20

19

Roskam, part 3, chapter 5, page 283

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The design of the fuselage frames and bulkheads is also important in terms of attachment to the empennage. It must be able to carry the empennage loads and therefore proper attachment of empennage spars is crucial for a safe design.

Conventional Tail

The conventional horizontal stabilizer consists of a left and right outboard sections attached to a centre section (torque box) which is located aft of the fuselage. In large transport planes, the stabilizer is designed to pivot on two-self aligning bushing type hinge joints attached to a heavy bulkhead in the fuselage and the angle of attack is adjusted by means of an electrically driven or manually operated ball nut and jackscrew which is attached to the forward side of the centre section. All vertical load distributions on the stabilizer are reacted at these three above mentioned attachment points.

FIGURE 103 - CONVENTIONAL TAILPLANE21

Drag

Drag of the empennage is approximately 10-20% of the total drag of the aircraft, depending on its size and position relative to the rest of the aircraft.

Summary of Tail Loads22

1. General a. Sonic Fatigue b. Fatigue c. Fail-safe d. Control Surface hinge load due to the stabilizer forced bending e. Hoisting f. Flutter g. Control surface reversal h. Control surface effectivity i. Rib crushing j. Control surface support k. Actuator support l. Concentrated load redistribution

20

Roskam, part 3, chapter 5, page 283 21

Airframe Structural Design, page 358 22

Airframe Structural Design, page 363

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2. Leading Edge a. Hail strike

3. Stabilizer a. Instantaneous elevator b. Fin gust (on t-tail) c. Positive maneuver d. Negative maneuver e. Unsymmetrical spanwise load distribution

4. Fin a. Instantaneous rudder – yaw initiation b. Dynamic overyaw c. Check maneuver d. Engine out e. Fin gust

5. Aft Fuselage a. Redistribute vertical and horizontal stabilizer concentrated laods b. Tail skid load

Longitudinal Control Mechanization

In almost all cases the entire horizontal stabilizer is used for trim and sometimes primary longitudinal control. It usually pivots about a fixed point using one or more actuators.

Configuration and Structural Locations

These values are usually dependent on the type of aircraft being designed as well as the types and magnitudes of loads in which it is to be subjected to. Figure 104 shows some basic configurations.

FIGURE 104 - DEFINITION OF EMPENNAGE SPAR, RIB AND STIFFENER LOCATIONS23

23

Roskam, part 3, chapter 5, page 276

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Empennage spar locations

Most aircraft use a torque box to carry the main loads and it is located at a location which utilizes the airfoil thickness which in turn helps to save weight. It is enclosed similarly to the main wing box by a front and rear spar, upper and lower skins as well as ribs. Multiple spars are often used in the case of fighter aircraft.

Typical spar locations:

1. Front spar: 15-25% of chord 2. Rear spar: 70-75% of chord

Empennage Rib locations

Ribs are used in order to stabilize the torque box of the empennage and also to serve as anchors for the control surface attachment brackets. Whenever point loads are expected a rib will be required. A typical rib configuration can be seen in Figure 105 below.

Typical rib locations

1. Light airplanes: 15 - 30 inches 2. Transports: 24 inches 3. Fighters and trainers: rib spacings vary widely

FIGURE 105 - B707 HORIZONTAL STABILIZER

Empennage Stiffener Spacings

These vary widely depending on the type of aircraft. The relative empennage stiffness determines the number and type of stiffeners required.

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Vertical tail attachment

FIGURE 106 - TWO METHODS FOR VERTICAL TAIL TO FUSELAGE ATTACHMENT24

Empennage Construction for Various Transport Aircrafts

TABLE 57 - EMPENNAGE CONSTRUCTION

Airplane Name

Vertical/Horizontal Tail

Type of Construction

Material Skin/Stringer

Panel Shape

No. Of Spars; Type of Ribs

B747 Vertical Skin-Stringer 7075-T6 2 Spars; Ribs-web

Horizontal Skin-Stringer 7075-T6 2 Spars; Ribs-web

B707 and B737

Vertical Skin-Stringer 2024-T3 2 Spars; Ribs-web

Horizontal Skin-Stringer 2024-T3 2 Spars; Ribs-web

24

Roskam, part 3, chapter 5, page 285

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7.6 COCKPIT SIZING

Determination of Visibility From The Cockpit25

Good visibility is important for the following reasons:

1. During Take-off and landing operations a pilot must have good view of the immediate

surroundings

2. During en-route operations the pilot must be able to observe conflicting traffic

3. In freighters, success in combat may depend on good visibility. Formation Flying is

impossible without it.

As our cockpit is based on the geometry of the B737-900NG and the internal layout has not changed then visibility for the projected aircraft will be exactly the same as the B737-900NG. Therefore, sizing and determining the view angle at this stage of design is not necessary.

FIGURE 107 - VISIBILITY FROM PORT SIDE

25

2.3 Determination of Visibility from the cockpit, Roskam, Part III, Chapter 2, Page 23

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7.7 LANDING GEAR

The landing gear of an aircraft plays a major role in operation. It assists in ground movement, landing shock absorption. Increasing the efficiency of these functions will result in less fuel consumption. Below is a brief description of the landing gear system used as well as some basic sizing.

Landing Gear Functions

1. Absorb landing shocks and taxing shocks 2. Provide ability for ground manoeuvring, taxi, takeoff roll, landing roll and steering 3. Braking capability 4. Allow for airplane towing 5. Protect ground surface 6. Landing Gear type

FIGURE 108 - RETRACTION METHOD

As our aircraft type is designed for more efficient flight at high altitudes and at long range, a fixed landing gear would of course not be an option due to aerodynamic reasons. Therefore, a retractable landing gear system is be used. This is also quite beneficial to us our wing is a low wing configuration, allowing for a storage space for the landing gears as seen in figure 44.

The use of a tricycle gear configuration as seen in figure 45 has also been selected due to the advantages it provides our aircraft. It allows for good visibility over the nose during ground operations as well as steering capabilities. It also allows the cabin floor and the cargo floor to be levelled.

During groundlooping, it provides stable behaviour for the passenger as well as the cargo. The weight of this landing gear is medium compared with the bicycle or tail wheel configurations.

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FIGURE 109 - WHEEL CONFIGURATIONS

Loading

The compatibility of landing gear and runway surface is determined by the allowable wheel loads. The load on the landing gear strut and tire must not exceed its maximum allowable load as it can permanently damage the structure as well as the runway.

Sizing

To determine the maximum load a wheel can absorb the following method was adopted from Roskam. The design weight of our aircraft is approximately 160,000 lbs. From the table below, our aircraft has a resemblance closest to the Boeing 727-200 which has a takeoff weight of 190,000 lbs. This gives a tire pressure and load classification number (LCN) of 160 psi and 80 respectively.

FIGURE 110 - TIRE PRESSURES AND LCN26

The LCN number is the reference number used by major runways around the world. It is a method of designing the landing gear without exceeding the minimum global requirements. The maximum allowable pressures are controlled by the various surfaces as seen in the figure below.

26

Roskam, Part 4, chapter 2, page 17

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FIGURE 111 - TIRE PRESSURES DEPENDING ON RUNWAY SURFACE27

Using the tire pressure and the LCN number mentioned above we can determine the equivalent single wheel load on the landing gear. From the figure below we find that the equivalent load is approximately 64,000 lbs.

FIGURE 112 - EQUIVALENT SINGLE WHEEL LOAD28

27

Roskam, Part 4, chapter 2, page 17 28

Roskam, Part 4, chapter 2, page 18

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8 SECTION 3 – AIRCRAFT PERFORMANCE

8.1 MATCHING CHART

A good method of keeping within the parameters of the RFP design as well as meeting necessary performance requirements is through the use of a matching chart analysis. Below was the method used to derive the matching chart displayed in the wing sizing section of this report (figure 41). The matching chart will be performed for 5 different conditions; takeoff distance, landing distance, cruise, stall and climb.

Roskam Part 1: Preliminary sizing of airplanes was used for the following analysis. All lift coefficient values were obtained from table 3.129 assuming a transport aircraft configuration.

Takeoff Distance

The first step was to incorporate the FAR 25 requirement of 5000 ft maximum takeoff field length (TOFL). Using this method of analysis we are meeting the landing requirement of 8200 ft laid down by the RFP. Therefore the takeoff parameter (TOP25)30 is analysed first.

EQUATION 90 – TAKE OFF PARAMETER

The next step was to find the thrust to weight ratio value as a function of lift coefficient, wing loading and density ratio ( ). As the aircraft is assumed to takeoff at sea-level the density ratio is 1. Therefore using equation 3.731

EQUATION 91 -TAKE OFF PARAMETER

Rearranging this formula gives:

Using a lift coefficient range from 1.6 to 2.2 allowed the development of the takeoff slopes. Wing loading values range from 0 to 300 lbs/ft2.

Landing

The FAR 25 landing distance requirement is 5000 ft. First step is to calculate the approach velocity32.

29

Roskam, Table 3.Typical Values For Maximum Lift Co-efficient, part 1, chapter 3, page 91,

(Transport aircraft - averaged value) 30

Roskam, equation 3.8, Part 1, Chapter 3, Page 98. 31

Roskam, equation 3.7, Part 1, Chapter 3, Page 98.

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EQUATION 92 - APPROACH VELOCITY

The next step is calculating the stall speed for landing33.

EQUATION 93 - STALL SPEED FOR LANDING

Using the stall velocity equation34 below

EQUATION 94 - STALL VELOCITY EQUATION

Rearranging this formula to find the wing loading as a function of lift coefficient gives:

Assuming that the landing weight is 85% of the takeoff weight35 we get:

Using a lift coefficient range from 1.8 to 2.8 allowed the development of the takeoff slopes. Wing loading values range from 0 to 240 lbs/ft2.

Cruise

Using the method provided in “cruise speed sizing of jet airplanes36”we derived the equations to develop the cruise speed curve.

Using equation 3.6037 we developed a relation between the thrust to weight ratio and the wing loading value. From this we get:

EQUATION 95 - THRUST TO WEIGHT RATIO

Where,

1. q = dynamic velocity =

32

Roskam, equation 3.16, Part 1, Chapter 3, Page 113 33

Roskam, equation 3.15, Part 1, Chapter 3, Page 113 34

Roskam, equation 3.1, Part 1, Chapter 3, Page 90 35

Roskam, Example of FAR 25 Landing Distance Sizing, part 1, chapter 3,3.3.4, page 113 36

Roskam, 3.6.4, Part 1, Chapter 3, Page 167 37

Roskam, equation 3.60, Part 1, Chapter 3, Page 167.

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2. CDO = profile drag from drag estimates section 3. AR = Aspect ratio = 9.45 (B737-900NG) 4. E = Oswald efficiency factor = 0.85

Stall

Stall was found simply by using a range of maximum lift coefficients with the corresponding stall velocity of the B737-900NG. Using the basic lift equation we get:

Rearranging this to get wing loading we get:

Using a lift coefficient range from 1.2 to 1.8 allowed the development of the takeoff slopes.

Climb

Assuming climb velocity is 15% more than the stall velocity then we can calculate the climb wing loading. Using a rearranged lift equation to cater for the modification yields:

Using a lift coefficient range from 1.2 to 1.8 allowed the development of the takeoff slopes.

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8.2 STRUCTURAL DRAG ESTIMATES

In order to determine the structural drag estimates the Reynolds numbers for each component must be determined. The components considered for this are the following:

1. Fuselage = 41.9 m 2. Nacelle = 2.51 m 3. Wing = 3.38 m (MAC) 4. Horizontal Stabilizer = 2.31 m (MAC) 5. Vertical Stabilizer = 3.72 (MAC)

Flight Conditions

1. Cruise Velocity = 228.61 m/s 2. Dynamic Viscosity = 1.39E-5 3. Density at 41000 ft = 0.30 4. Wing Reference Area = S = 108.26 m2

Using the Reynolds number equation below we can determine the Reynolds number for each component.

EQUATION 96 - REYNOLDS NUMBER EQUATION

Where

Re = Reynolds Number

= Density (kg/m3)

L = Component reference length (m)

= dynamic viscosity (Pa.s)

TABLE 58 - STRUCTURAL DIMENSIONS

Item Reference Length (m)

Velocity (m/s2)

Dynamic Viscosity

Density at 41000 ft

Reynolds Number

Fuselage 41.9 228.61 1.39E-05 0.30 2.06E+08

Nacelle 2.51 228.61 1.39E-05 0.30 1.24E+07

Wing 3.38 228.61 1.39E-05 0.30 1.66E+07

Horizontal Stabilizer

2.31 228.61 1.39E-05 0.30 1.14E+07

Vertical Stabilizer

3.72 228.61 1.39E-05 0.30 1.83E+07

Drag Estimates

Fuselage

Diameter = 3.88 m

Length = 41.9 m

Reynolds Number = 2.06E8

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Wetted Area

Coefficient of Friction

EQUATION 97 - CO-EFFICIENT OF FRICTION

F(L/D)

CDO

Nacelle

The same calculations will be performed for the nacelle except for the f(L/D) value which will be 1. This is because the airflow does not necessarily flow over the structure but through it.

Diameter = 1.5 m

Length = 2.51 m

Reynolds Number = 1.24E7

Wetted Area

Coefficient of Friction

CDO

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Wing

Thickness = 0.3 m

Mean Aerodynamic Chord = 3.38 m

Coefficient of friction is equal to the value calculated for the fuselage.

F(t/c)

Wetted Area

The wetted area is simply the top surface and the bottom surface of the wings. Therefore, the wetted area is:

CDO

Horizontal Stabilizer

Thickness = 0.2 m

Mean Aerodynamic Chord = 2.3068 m

Horizontal Stabilizer Area = SHS = 32.78 m2

Coefficient of friction is equal to the value calculated for the fuselage.

F(t/c)

Wetted Area

The wetted area is simply the top surface and the bottom surface of the wings. Therefore, the wetted area is:

CDO

Vertical Stabilizer

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Thickness = 0.2 m

Mean Aerodynamic Chord = 3.72 m

Horizontal Stabilizer Area = SHS = 26.44 m2

Coefficient of friction is equal to the value calculated for the fuselage.

F(t/c)

Wetted Area

The wetted area is simply the top surface and the bottom surface of the wings. Therefore, the wetted area is:

CDO

Total Drag

TABLE 59 - STRUCTURAL DRAG ESTIMATE

Item CDO

Fuselage 0.00782

Nacelle 0.00063

Wing 0.0048

Horizontal Stabilizer 0.00072

Vertical Stabilizer 0.00108

TOTAL 0.0151

8.2.1 CL vs. CL/CD

Using the equation below performance of the projected aircraft over a range of CL can be plotted.

Where:

Oswald Factor = e = 0.85

Aspect Ratio = AR = 9.45

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Two sets of results were produced; the aircraft without drag reducing technologies and the aircraft with drag reducing technologies (Winglet and Wing Waggle – total of 27% reduction). Results are as follows.

TABLE 60 - CL VS CL/CD VALUES WITHOUT WING TECHNOLOGY

CL CD CL/CD

0.1 0.0154 6.47

0.2 0.0166 12.02

0.3 0.0186 16.11

0.4 0.0214 18.70

0.5 0.0250 20.03

0.6 0.0293 20.46 (max)

0.7 0.0345 20.31

0.8 0.0404 19.79

0.9 0.0472 19.09

1 0.0547 18.29

1.1 0.0630 17.46

1.2 0.0721 16.64

1.3 0.0820 15.85

1.4 0.0927 15.10

1.5 0.1042 14.39

TABLE 61 - CL VS CL/CD WITH WING TECHNOLOGY

Without Tech.

With Tech. (27% saving)

CL CD CD CL/CD

0.1 0.0154 0.0097 10.27

0.2 0.0166 0.0105 19.08

0.3 0.0186 0.0117 25.57

0.4 0.0214 0.0135 29.68

0.5 0.0250 0.0157 31.80

0.6 0.0293 0.0185 32.48

0.7 0.0345 0.0217 32.23

0.8 0.0404 0.0255 31.42

0.9 0.0472 0.0297 30.30

1 0.0547 0.0344 29.03

1.1 0.0630 0.0397 27.71

1.2 0.0721 0.0454 26.41

1.3 0.0820 0.0517 25.16

1.4 0.0927 0.0584 23.97

1.5 0.1042 0.0657 22.85

FIGURE 113 – CL/CD COMPARISON

0.00

5.00

10.00

15.00

20.00

25.00

30.00

35.00

0 0.5 1 1.5

CL/

CD

CL

CL/CD vs. CL

Projected Aircraft Without Technologies

Projected Aircraft With Technologies

B737-900NG

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TABLE 62 - LIFT-TO-DRAG COMPARISON

B737-900NG

Projected Aircraft (Without Technology)

Projected Aircraft (with Technologies)

Mass [kg] 85130 73750

Weight [N] 835125.

3 723487.5

Cruise Velocity [m/s] 228.61 228.61

Density At Cruise 41000 ft [kg/m3]

0.3 0.3

Wing Area [m2] 125 108.26

Coefficient of lift 0.854 0.85

Oswald Factor 0.85 0.85

Aspect Ratio 9.45 9.45

Profile Drag 0.03 0.0151

Coefficient of drag 0.059 0.04 0.03

Lift-To-Drag 14.499 19.44 30.86

Improvement - 25.42% 53.01%

Lift-To-Drag Improvement

RFP Requirement: Without the use of lift-to-drag technologies and based solely on the weight reduction the projected aircraft was capable of a 25.42% improvement over the B737-900NG and a 53.01% improvement through the use of drag reducing technologies. This comparison was performed for a cruise condition at 41000 ft. RFP requirement states that a 25% improvement is desired and that need was fulfilled.

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8.3 CRITICAL FLIGHT CONDITIONS

The critical flight conditions must be analyzed in order to have a rough idea if it can meet the RFP design requirements as well as FAR 25 Rules and Regulations.

8.3.1 Takeoff

FIGURE 114 - TAKEOFF MISSION PROFILE Takeoff can be broken down into 4 headings: ground run distance, rotation, airborne phase and takeoff time.

8.3.2 Ground Run

1. Wing Height above runway, (h) = 8 m (B737-900NG Specifications) 2. Aircraft Wing span, (b) = 32.32 m (Wing sizing calculations) 3. Rolling Resistance38, (μR) = 0.02 4. Oswald Factor, (e) = 0.85 (Assumption) 5. Aspect Ratio, (AR) = 9.45 (B737-900NG) 6. Profile Drag, (CDO) = 0.015 (Drag Estimates) 7. Maximum Thrust, (T) = 102309.08 N (PW1000G Specifications) 8. Aircraft Weight, (W) = 723487.5 N (Weight Estimate) 9. Gravity, (g) = 9.81 m/s2 10. Lift-curve slope, (a) = 5 per rad (Assumption) 11. Density at Sea-Level, (ρSL) = 1.225 kg/m3 12. Wing Area, (S) = 108.26 m2 (Wing Sizing)

First constant that needs to be found is kL

Now we can calculate the optimum co-efficient of lift.

38

Roskam, Airplane Flight Dynamics and Automatic Flight Controls, Part 1, Chapter 4, Page 290.

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Using this we can determine the co-efficient of drag

Total Thrust (TS) is the thrust from the PW1000G multiplied by two engines. Therefore:

Constant A

Constant B

Stall Velocity

39

Rotational Velocity

Ground Run Distance

Rotation Distance

39

Roskam, Table 3.Typical Values For Maximum Lift Co-efficient, part 1, chapter 3, page 91,

(Transport aircraft - averaged value)

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Rotation time is usually around 1 to 3 seconds. For this analysis we will assume a 3 second rotation. Therefore:

The horizontal distance, (SR) is calculated using a simple distance equation.

Airborne Phase

Obstacle height, (ho) is based on FAR requirements and states that it is 35 feet (10.668 meters). The following calculation determines the distance travelled horizontally as the aircraft clears the obstacle height.

1. Obstacle Height, (ho) = 35 ft = 10.668 m 2. Aircraft Weight, (W) = 723487.5 N 3. Wing Area, (S) = 108.26 m2 4. Airborne Velocity is usually 15% more than the stall velocity

The average CL must be calculated using the maximum co-efficient of lift for takeoff. Therefore:

40

Now the lift can be calculated using these two values for sea level conditions.

Load Factor

Radius of Pull-Arc

Angle of Arc

40

Roskam, Table 3.1 Typical Values For Maximum Lift Co-efficient, part 1, chapter 3, page 91,

(Transport aircraft - averaged value)

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Horizontal Airborne Distance

Total Takeoff Distance

RFP Requirement

The RFP requirement of maximum 8200 feet for total take-off distance has been achieved.

Takeoff Time

Constants A and B will once again be utilized.

Horizontal Airborne Speed at Takeoff

Time of pull-up

Total Takeoff Time

Maximum Rate of Climb

Maximum Rate of Climb occurs when the following equation is satisfied:

By rearranging the formula you can find the co-efficient of lift value which will give the maximum rate of climb.

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1. Thrust, (T) = 204618.17 N (PW1000G Specifications) 2. Weight, (W) = 723487.5 N (Weight Estimate) 3. Oswald Factor, (e) = 0.85 (Assumption) 4. Aspect Ratio, (AR) = 9.45 (B737-900NG) 5. Profile Drag, (CDO) = 0.015 (Drag Estimates) 6. Density at Sea-level, (ρSL) = 1.225 kg/m3 7. Wing Area, (S) = 108.26 m2 (Wing sizing)

K Value

Maximum Rate of Climb formula which will give you co-efficient of lift

Velocity

Drag

Maximum Rate of Climb

Climb Angle

Maximum Angle for Climb

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Maximum climb angle occurs when CDi is equal to CDO. Therefore:

Co-efficient of lift for Maximum Climb angle

Velocity

Co-efficient of Drag

Minimum Drag

Maximum Angle

Rate of Climb for Maximum Angle

8.3.3 Cruise Conditions

1. Weight, (W) = 723487.5 N (Weight Estimate) 2. Density at 41000ft, (ρ41000) = 0.30 (Calculated – See Appendix) 3. Cruise Velocity, (VCR) = 228.61 m/s (B737-900NG) 4. Wing Area, (S) = 108.26 m2 (Wing Sizing) 5. Profile Drag, (CDO) = 0.015 (Drag Estimates) 6. Oswald Factor, (e) = 0.85 (Assumption) 7. Aspect Ratio, (AR) = 9.45 (B737-900NG) 8. K Value, (k) = 0.04 (From above Maximum rate of climb)

Co-efficient of Lift

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Co-Efficient of Drag

Drag

Lift

Lift-To-Drag Ratio

This value will no doubt increase as new technologies in terms of laminar flow technology and drag reducing devices have not yet been taken into consideration.

8.3.4 Landing

FIGURE 115 - LANDING MISSION PROFILE

Landing will be assessed based on the FAR requirements for landing.

1. Density at 41000ft, (ρ41000) = 0.30 (Calculated – See Appendix) 2. Wing Area, (S) = 108.26 m2 (Wing Sizing) 3. Profile Drag, (CDO) = 0.015 (Drag Estimates) 4. K Value, (k) = 0.04 (From above Maximum rate of climb)

FAR states that an aircraft under FAR 25 regulations must be able to land in 5000 feet. Therefore:

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Approach Velocity41, VA

EQUATION 98 - APPROACH VELOCITY

Landing Stall Velocity42, VSL

EQUATION 99 - LANDING STALL VELOCITY

This is the landing speed of our aircraft. RFP requirements state that 140 KCAS (72.02 m/s at sea level conditions) is the maximum allowable. We have indeed met that requirement.

Using the maximum lift co-efficient for landing from Roskam43

Coefficient of Drag Estimate

Drag Estimate

This drag estimate will also be the thrust required.

41

Roskam, equation 3.16, Part 1, Chapter 3, Page 111 42

Roskam, equation 3.15, Part 1, Chapter 3, Page 111 43

Roskam, Table 3.Typical Values For Maximum Lift Co-efficient, part 1, chapter 3, page 91,

(Transport aircraft - averaged value)

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8.3.5 RFP Performance Requirements

TABLE 63 - RFP REQUIREMENTS

Item RFP

Requirement Projected Aircraft

Condition

Takeoff Distance <8200ft 6757.06ft YES

Landing Speed <140KCAS 99.31KCAS YES

Cruise Speed MACH0.8 MACH0.78 NO

Maximum Cruise Altitude

41000ft 41000ft YES

Maximum Range 3500nm 3500nm YES

As can seen from Table 63 above, all performance requirements have been met except the cruise speed requirement. The reason for this is that the cruise speed used for the projected aircraft is solely based on the B737-900NG and not calculated based on the projected aircrafts own merits. Meaning that, the projected aircrafts cruise speed will actually be higher than MACH0.78 because of the more efficient engines and the lighter body.

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8.4 WEIGHT AND BALANCE

1. Root Chord, (cRoot) = 7.877 m (B737-900NG) 2. Mean Aerodynamic Chord, (c) = 3.38 m (Wing Sizing) 3. Distance from the nose (reference point is the nose of the aircraft going backwards

towards the tail) to centre of gravity, (hc) = 22.50 m (Assumption – approximately half of the fuselage length)

4. Distance from mean aerodynamic chord to the centreline, (Y) = 6.06 m 5. Fuselage Length, (L) = 41.90 m (Fuselage Sizing) 6. Fuselage Diameter, (D) = 3.76 m (Fuselage Sizing) 7. Main Wing Sweep Back Angle, (Λ) = 25 degrees (B737-900NG) 8. Distance from Tail Centre of pressure to fuselage end = 1 meter (Assumption)

The distance from the nose to the centre of gravity, (hc) is an adjusted value. Weight and balance was performed once before using exactly half of the fuselage length, i.e. 20.95 meters. This showed a very small distance between the centre of pressure and the centre of gravity (1.03 meters). Therefore the distance was increased to 22.50 meters to allow for a larger distance. This method is aloud for this stage in design in order to comply with early longitudinal stability calculations. This will be analysed much more accurately in the detailed design phase.

The following drawing shows the break-up of unknowns to determine the aircrafts centre of pressure. Note: This location of centre of pressure, (CoP) and centre of gravity, (CG) will change due to factors such as longitudinal stability analysis which will in turn change the geometry of the empennage section, centre of pressure location will change when the new airfoil is applied and the wing is resized.

FIGURE 116 - CENTRE OF PRESSURE AND CENTRE OF GRAVITY LOCATIONS

The distance from the tail centre of pressure to the centre of gravity, (lT) can be calculated using the following equation.

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The distance from the datum to the centre of pressure, (hoc) can be calculated using simple geometry. The distance from the datum to the beginning of the root chord is approximately 16 meters. This value was obtained using the B737-900NG planform scaled drawing (See Appendix 8). The “x” value as seen on Figure 116 can be found using trigonometry. Half of the fuselage diameter must be subtracted from the distance from the mean aerodynamic chord to the centreline. i.e.

X-Value

Centre of Pressure Location

The Distance from the tail centre of pressure to the centre of pressure, (l) can be found using simple geometry.

The distance between the CoP and the CG, (hc-hoc) is the most important value as it dictates the aircrafts stability. There are three conditions of stability:

1. Positive - means that the CoP is ahead of the CG and that the aircraft is stable. 2. Negative - it means that the CoP is behind the CG and that it is unstable. 3. Zero – this means that the CoP and the CG lie on the same longitudinal location and that

the aircraft is neutrally stable.

Therefore the aircraft is longitudinally stable based solely on the fact that the CoP is ahead of the CG. Longitudinal stability will be discussed further in the stability section as well as in the detail design phase.

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8.5 STABILITY – LONGITUDINAL, LATERAL AND DIRECTIONAL

As mentioned earlier the empennage section is te sted for stability.

TABLE 64 - SPECIFICATIONS

PARAMETER MAGNITUDE UNIT SYMBOL

Weight of the Aircraft 723488 N W

Wing Area 108.26 m2 S

Tailplane Area (B737-900NG) 32.78 m2 ST

AR 9.45

Elevator Area (B737-900NG) 6.55 m2 SE

Wing Mean Aerodynamic Chord 3.42 m c-

Elevator Chord 0.9 m c-E

Tailplane Setting Angle -2 degree iT

Stick Gearing 2.5 rad/m GE

Wing-body Pitching Moment Coefficient -0.04 per

radian CM

Wing-body Lift Curve Slope 5.7 per

radian aWB

Tailplane Lift Curve Slope 5 per

radian aT1

Elevator Lift Curve Slope 2 per

radian aT2

Elevator Tab Lift Curve Slope 0.5 per

radian aT3

Tailplane Hinge Moment Coefficient Curve Slope -0.05 per

radian bT1

Elevator Hinge Moment Coefficient Curve Slope -0.15 per

radian bT2

Elevator Tab Hinge Moment Coefficient Curve Slope -0.2 per

radian bT3

Distance From Centre of Gravity To Tailplane Aerodynamic Centre

18.41 m lT

Location of Centre of Gravity Arf of The Wing-body Aerodynamic Centre

2.59 m (h-h0).c-

Distance from tail CoP to Wing CoP 21 m l

Distance from nose to CoP at centreline from main wing 19.92 m hoc

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Longitudinal Stability (Stick fixed)

Downwash angle

( )

Tail Volume Ratio

Aircraft Lift-Curve Slope

Stick Fixed

Note : Moment derivative with respect to CL , stick fixed, for neutral stability we must have ∂C ∂C T CK F XED

Stick fixed Neutral Point

The neutral point, hn is a longitudinal location for the aircraft which determines its longitudinal stability in respect to the location of the centre of gravity. To determine its location we use the following equation:

FIGURE 117 - AERODYNAMIC GEOMETRY

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Therefore, the distance between the centre of gravity and the neutral point relative to the centre of gravity, also known as the static margin Kn is as follows:

By definition, if:

0,

0,

0,

Therefore the projected aircraft is longitudinally stable for the stick fixed condition.

Longitudinal Stability (Stick Free)

Aircraft lift-curve slope

To find the neutral point

. Using the following equation we can determine its

location.

Therefore, the distance between the centre of gravity and the neutral point relative to the centre of gravity, also known as the static margin Kn is as follows:

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By definition, if:

0,

0,

0,

Therefore the projected aircraft is longitudinally stable for the stick fixed condition.

Lateral Stability

Lateral Stability and Control is related to moments about the roll axis.

TABLE 65 - SPECIFICATIONS

PARAMETER MAGNITUDE UNIT SYMBOL

Aileron Deflection – Right Wing 10 Degrees

Aileron Deflection – Left Wing -10 Degrees

Density 1.225 kg/m3 Velocity 228.62 m/s V

Wing Lift-Curve Slope 5.4 Per radian a1

Aileron Lift-Curve Slope 3.2 Per radian a2

Wing Area 108.26 m2 S

Wing Span 32.32 m b Semi Wing Span 16.16 m s

Distance to aileron from wing root 10 m yA

Wing Mean Aerodynamic Chord 3.42 m

Rolling Moment

Dimensional Derivative

Non-Dimensional

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Rolling Moment Co-Efficient due to aileron deflection

Aircraft moment of inertia about longitudinal axis

For a wing span of 32.32 m the following assumption was made:

Roll Acceleration

Roll Velocity

Roll Displacement

Lift on the opposite wing due to rolling moment due to p

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Rolling Moment Coefficient

Notice that the moment due to rolling ( Nm) is almost 3 times but opposite the moment due to aileron If the ailerons ( Nm) are deflected, the roll rate builds up until there is a balance between the two moments. That is

To determine

we use the following equations

Therefore, to counteract the rolling moment in a one side of the moment will create in the opiate wing ailerons at a rate of until there is a balance between the two moments.

Directional Stability and Control

TABLE 66 - DETAILS

PARAMETER MAGNITUDE UNIT SYMBOL

Side slip angle 7 (Assumed) Degrees β

Rudder lift curve slope 2(Assumed) per radian av1

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The following criteria determines lateral stability

∂N/∂β Nβ, is the dimensional derivative of rolling moment due to sideslip. (yawing moment N)

The yawing moment can be non-dimensionalized as:

Non dimensional derivative ∂CN/∂β CNβ is the main contribution to the directional stability.

The vertical tail volume coefficient is used to measure this as

An approximate expression for CNβ can be derived by considering the effect of the vertical tail

This value is very small, therefore it is stable. \An approximate expression for CNβ can be derived by considering the effect of the vertical tail.

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9 SECTION 4 – PROJECTED AIRCRAFT SPECIFICATIONS

FIGURE 118 - FRONT VIEW

FIGURE 119 - TOP VIEW

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FIGURE 120 - TOP VIEW (WITH SEATING ARRANGEMENT)

FIGURE 121 - SIDE VIEW

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FIGURE 122 - CONCEPT ARTWORK

FIGURE 123 – FUSELAGE (SEATING ARRANGEMENT)

FIGURE 124 - FUSELAGE CROSS SECTION

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FIGURE 125 - CATIA MODEL (VIEW 1)

FIGURE 126 - CATIA MODEL (VIEW 2)

FIGURE 127 - CATIA MODEL (VIEW 3)

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9.1 COMPARISONS

TABLE 67 – B737-900NG VS. PROJECTED AIRCRAFT

Comparison - B737-900NG vs. Projected Aircraft

B737-900NG

Projected Aircraft

Condition

Flight Crew 2 2 -

Maximum Passengers 215 200 7% Reduction

Nominal Range (Nautical Miles)

- - -

Maximum Range (Nautical Miles)

3200 3500 8.6% Improvement

Weight kg lbs kg lbs

Maximum Take-off Weight 85130 187679.

47 73750

162591

13.4% Weight Reduction

Empty Weight 44676 98493.6

9 35444.5

7 78141.

9 20.7% Weight

Reduction

Landing Weight 66361 146300.

92 62686.8

1 13820

1 5.5% Weight

Reduction

Payload Capability - 19050.8

6 42000 -

Maximum Fuel Capacity (Litres)

29660 22180.15 25% Reduction

Dimensions (Lengths) m ft m ft

Fuselage Length 42.1 138.12 41.9 137.47 0.5% Reduction

Wing Span 35.7 117.13 34.32 112.58 4.8% Reduction

Aircraft Height 12.5 41.01 12.26 40.22 1.9% Reduction

Fuselage Width 3.76 12.34 3.66 12.01 2.7% Reduction

Fuselage Height 4.01 13.16 3.88 12.73 3.2% Reduction

Cabin Width 3.54 11.61 3.4 11.15 4% Reduction

Cabin Height 2.2 7.22 2.2 7.22 -

Dimensions (Area) m2 ft2 m2 ft2

Wing Area 125 410.1 108.26 355.18 13.4% Reduction

Dimensions (Volume) m3 ft3 m3 ft3

Cargo Capacity 52.5 1854.02 42.68 1507.2

6 18.7% Reduction

Performance (Velocities) Mach Mach

Cruise Speed 0.78 0.78 -

Maximum Operating Speed 0.82 0.82 -

Altitude m ft m ft

Initial Cruise Altitude - - -

Maximum Cruise Altitude 12496.

8 41000 12496.8 41000 -

Fuel Type Jet Fuel Biojet Fuel -

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TABLE 68 – A320-200 VS. PROJECTED AIRCRAFT

Comparison - A320-200 vs. Projected Aircraft

A320-200

Projected Aircraft

Condition

Flight Crew 2 2 -

Maximum Passengers 180 200 10% Improvement

Nominal Range (Nautical Miles)

- - -

Maximum Range (Nautical Miles)

3200 3500 8.6% Improvement

Weight kg lbs kg lbs

Maximum Take-off Weight 78000 171960.

52 73750

162591

5.4% Weight Reduction

Empty Weight 42600 93916.9 35444.5

7 78141.

9 16.8% Weight

Reduction

Landing Weight 64500 142198.

12 62686.8

1 13820

1 2.8% Weight

Reduction

Payload Capability - 19050.8

6 42000 -

Maximum Fuel Capacity (Litres)

24050 22180.15 7.8% Improvement

Dimensions (Lengths) m ft m ft

Fuselage Length 37.57 123.26 41.9 137.47 10.3% Increase

Wing Span 34.1 111.88 34.32 112.58 0.3% Reduction

Aircraft Height 11.8 38.71 12.26 40.22 8.5% Increase

Fuselage Width 3.95 12.96 3.66 12.01 2.7% Reduction

Fuselage Height -

3.88 12.73 -

Cabin Width 3.7 12.14 3.4 11.15 8.1% Reduction

Cabin Height - 2.2 7.22 -

Dimensions (Area) m2 ft2 m2 ft2

Wing Area 122.6 402.23 108.26 355.18 11.4% Reduction

Dimensions (Volume) m3 ft3 m3 ft3

Cargo Capacity 37.41 1321.12 42.68 1507.2

6 12.3% Increase

Performance (Velocities) Mach Mach

Cruise Speed 0.78 0.78 -

Maximum Operating Speed 0.82 0.82 -

Altitude m ft m ft

Initial Cruise Altitude - - -

Maximum Cruise Altitude 11887.

2 39000 12496.8 41000 4.9% Increase

Fuel Type Jet Fuel Biojet Fuel -

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TABLE 69 – PROJECTED AIRCRAFT VS. RFP REQUIREMENTS

Comparison - Projected Aircraft vs. RFP Requirements

Projected Aircraft

RFP Requirements

Condition Requireme

nt

Flight Crew 2 2 - YES

Maximum Passengers 200 Minimum 175 12.5%

Increase YES

Nominal Range (Nautical Miles)

N/A 1200

-

Maximum Range (Nautical Miles)

3500 3500 - YES

Weight kg lbs lbs

Payload Capability 19050.

86 42000 37000

11.9% Improvement

YES

Dimensions (Lengths) m ft ft

Wing Span 34.32 112.58 - - -

Aircraft Height 12.26 40.22 - - -

Fuselage Width 3.66 12.01 - - -

Fuselage Height 3.88 12.73 - - -

Cabin Width 3.4 11.15 - - -

Cabin Height 2.2 7.22 - - -

Dimensions (Area) m2 ft2 ft2

Wing Area 108.26 355.18 - - -

Dimensions (Volume) m3 ft3 ft3

Cargo Capacity 42.68 1507.26 1240 17.7%

Increase YES

Performance (Velocities) Mach Mach

Cruise Speed 0.78 0.8 2.5%

Reduction NO

Maximum Operating Speed

0.82 0.83 1.2%

Reduction NO

Altitude m ft ft

Initial Cruise Altitude - 35000 - -

Maximum Cruise Altitude

12496.8

41000 41000 - YES

Fuel Type Biojet Fuel Alternative

Fuels - YES

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TABLE 70 - RFP REQUIREMENTS

Item Projected Aircraft

Requirement Unit Condition

Safety and Airworthiness Regulations

FAR25 FAR25 - Yes

Crew 2 2 - Yes

Passengers 200 175 (1 class) Yes

Seating Pitch 32 32 inches Yes

Seating Width 17.2 17.2 inches Yes

Cabin Width 11.15 12.5 feet No

Cabin Height 7.22 7.25 feet No

Cargo Capacity 1507.26 1240 feet3 Yes

Take-off distance 6757.06 8200 feet Yes

Landing Speed 129.10 140 KCAS Yes

Maximum Weight 162589.25 Maximum Zero Fuel Weight plus

reserves for maximum range Yes

Cruise Speed 0.78 0.8 MACH No

Maximum Operating Speed

0.82 0.83 MACH No

Initial Cruise Altitude 41000 35000 feet Yes

Nominal Range - 1200 nm Yes

Maximum Range 3500 3500 nm Yes

Payload Capability 42000.00 37000 lbs Yes

Alternate Fuels Biojet Fuel Biofuels - Yes

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10 COMPARISON DISCUSSION

B737-900NG vs. Projected Aircraft

The projected aircraft matched and/or exceeded all specifications except the cargo capacity.

A320-200 vs. Projected Aircraft

The projected aircraft matched and/or exceeded all specifications.

RFP requirements vs. Projected Aircraft.

The projected aircraft matched and/or exceeded all specifications except for cabin width, cabin height, cruise and maximum airspeed. Seeing as how the cabin dimensions are short only by a few percent, it can be stated that the RFP requirement for cabin dimensions are met. In terms of velocity the initial aircraft velocity was based on the B737-900NG. Therefore, the newly designed aircraft has not been analyzed incorporating the improved structure and reduced weight as well as through the utilization of more efficient engines and reduced overall drag. It is then safe to say that the aircraft is indeed capable of the cruise and maximum speed requirements set by the RFP.

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11 SECTION 5 – TOTAL ENVIRONMENTAL IMPACT

TABLE 71 - NOISE REDUCTION TECHNOLOGIES

Item Value

Pratt and Whitney PW1000G Turbofan 15dB

Integrated OGV 3.8dB

Pylon-based noise suppressors for engine attachments 2.1dB

Actuated panels through the use of sensors for damping noise damping through the fuselage

15dB

Fixed Winglets 6.50%

Single Slotted Flap 23%

TABLE 72 - FUEL CONSUMPTION AND EMISSION SAVING TECHNOLOGY

Item Saving

Pratt and Whitney PW1000G Turbofan Fuel Saving 15%

NextGen Navigational Equipment Fuel Saving 10%

Fixed Winglets - Climb Thrust 3%

Fixed Winglets - Cruise Thrust 6%

Wing Waggle - Drag 20%

Wing Waggle - Fuel 20%

No-Bleed System 3%

Magnetic Braking Fuel Saving of US$2.4-

3.4M/year

Weight reduction through the use of composites and smarter materials

Approximately 7.5-10% Fuel Saving

Hydrogen Power Ground Support Vehicles

Hydrogen fuel cells for airport management vehicles will produce zero emissions during operation and will eventually replace the existing ground support vehicles thus completely eliminating ground emissions.

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12 CONCLUSION

As mentioned at the beginning of this investigation, there is a strong desire for all aircraft designers to design and create an aircraft which is more fuel efficient and produced less emissions than existing aircrafts. The projected aircraft was designed with the emphasis of environmental friendliness in all aspects of its creation, manufacture, operation and disposal. From this viewpoint the design indeed succeeded.

The aircraft now has an improved lift-to-drag ratio in comparison with the B737-900NG and also exceeded the RFP requirement of 25%. The use of the newly designed Pratt and Whitney PW1000G engine is much more fuel efficient than pre-existing engines, has an increased by-pass ratio as per requested by the RFP and more importantly can utilize the alternate fuel source, biojet fuel, leading to further reduced emissions.

Advances in structural technology has also been incorporated through the use of advanced composite material. This new material in specified and carefully selected areas has lead to a 10% reduction in empty weight as well as improved the structural integrity of specific structures of the aircraft.

Fuel burn, noise and emissions in all areas including the airport were analyzed and allowed not just an improvement in the aircraft itself, but all sectors which are involved with its daily routine. The use of advanced navigational and aircraft systems, accompanied with the new fleet management system and hydrogen based ground support vehicles has substantially reduced the airport emissions. The movement and control of the aircrafts whereabouts and flight patterns has never been more efficient.

Technologies, systems and materials selected are all within the 2020 design date parameter. Performance has been improved and all necessary FAR25 and RFP requirements have been met.

In comparison to both the Boeing 737-900NG and the Airbus A320-200 aircrafts, the projected aircraft is indeed years ahead in terms of technology, system utilization, structure and overall management. All general requirements stated in the RFP have been fulfilled.

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15 APPENDIX 1 - DETAILED DESIGN TIMELINE

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16 APPENDIX 2 - DETAILED DESIGN WORK BREAKDOWN STRUCTURE

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17 APPENDIX 3 – AIRCRAFT SPECIFICATIONS – A320 FAMILY

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18 APPENDIX 4 – AIRCRAFT SPECIFICATIONS – B737 FAMILY (737-100 TO NEXT GENERATION)

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19 APPENDIX 5 - AIRCRAFT SPECIFICATIONS – B737 FAMILY (NEXT GENERATION FAMILY)

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20 APPENDIX 6 – AIRCRAFT DIMENSIONS – SCALE DRAWING

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21 APPENDIX 7 – NEEDS ANALYSIS

Background

1. Date of aircraft operation: 2020

2. Improvement in energy efficiency for long term national goals in aerospace

3. To reduce the environmental impact of the industry in order to continue aviation

research

4. Reduction in fuel consumption and emissions due to ever increasing fuel costs and a

possible carbon footprint tax in the future.

5. Aircraft systems, structures and performance

a. Performance improvement

b. Use of new technologies internal and external of the aircraft

c. Impact of alternative fuels on aircraft design as well as airline infrastructure

d. Carbon footprint reduction

e. Noise reduction

f. Enhanced laminar flow

g. Active controls for aircraft and engine design

h. Integration of high by-pass ratio for lower fuel consumptions

i. Advances in materials and structures for an increase in strength to weight ratios

j. 25% improvement of lift-to-drag through the utilisation of multidiscipline

configuration and laminar flow technology

k. Weight fractions improvement

l. Produce an optimised design

6. The aircraft must be economically feasible in terms of purchase price. A slight increase

in costs over the A320 and B737 is acceptable on the premise that it is much more cost

effective for the life of the proposed aircraft. However, it will be targeted to be cheaper

and lower maintenance than these aircrafts.

7. Incorporation of new technologies, operational procedures and alternate fuels.

8. Design will need to target improvement in accordance with the National Aeronautics

Research and Development challenges, goals and objectives

9. The design will need to be considered a 737NG/A320 replacement aircraft

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Structural and Performance Requirements

Item Requirement Note

1 (

Un

fix

ed

)

2

3

4

5 (

Fix

ed

)

Safety and Airworthiness Regulations FAR N/A

x

Crew 2 N/A

x

Passengers 175 (1 class)

x

Seating Pitch 32 inches

x

Seating Width 17.2 inches

x

Cabin Width 12.5 feet

x

Cabin Height 7.25 feet

x

Cargo capacity 1240 feet3

x

Take-off distance 8200 feet

x

Landing Speed 140 KCAS

x

Maximum Weight Maximum Zero Fuel Weight plus reserves for maximum

range

x

Cruise Speed 0.8 MACH

x

Maximum Operating Speed 0.83 MACH

x

340 KCAS

Initial Cruise Altitude 35000 feet

x

Nominal Range 1200 nm

x

Maximum Range 3500 nm

x

Payload Capability 37000 lb

x

Alternate Fuels Biofuels

x

Aspect ratio N/A N/A

x

Fuselage Width N/A N/A

x

Fuselage Height N/A N/A

x

Fuselage Length N/A N/A

x

Max Landing Weight N/A N/A

x

Max Fuel Capacity N/A N/A

x

Cruise Thrust N/A N/A

x

Engine length N/A N/A

x

Engine Clearance N/A N/A

x

Wingspan N/A N/A

x

Sweepback N/A N/A

x

Tail Height N/A N/A

x

Empty Weight N/A N/A

x

MTOW N/A N/A

x

Fuel Tank Size N/A N/A

x

Number of Engines N/A N/A

x

Max Thrust N/A N/A

x

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Projected Aircraft, Engine Technology and Assumptions

1. An increase in airfoil performance is needed and therefore airfoil technology will be

used to delay the transition by 20% for a 30o swept wing.

2. A reduction in overall weight is required to reduce fuel consumption and the size of the

engine required. Using the structural and materials property data of composites from

the B787 we can assess its level of use.

3. A need for an improvement in ground support and support systems in order to reduce

costs, emissions, increase performance and productivity.

4. The need for an improved engine design which will be cheaper to run, easier to

maintain, efficient, low fuel consumption and lower noise production.

5. An estimation of operation and maintenance costs using current in-service aircraft.

6. Assess the projected environmental impacts which include carbon footprint, acquisition

of alternate fuels as well as changes in airline infrastructure.

7. Assess the impact of alternate fuel use on airlines.

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22 APPENDIX 8 – FEASIBILITY REPORT

A I R C R A F T S T R U C T U R A L C O N F I G U RA T I O N

Configuration 1 - Conventional Aircraft With Engines Attached To Wings

General:

Our configuration will be very similar to the Boeing 737-900NG. It will be a low wing commercial transport aircraft. The wings will be swept-back and dihedral for aerodynamic stability purposes. It will be a conventional design with two engines, one on the underside of each wing and basic empennage configuration.

Reasons for Use:

The reason we are using an existing configuration is because it is a reliable design and it meets most of the criteria set out by our customer. Because the B737 has been in service for such a long period, reasonable amounts of knowledge in regards to its performance and capabilities have been acquired. Therefore, we can use this information for our calculations and modifications. This is important because it reduces the design, calculation and certification aspects of our proposed modified Boeing 737-900NG design.

Risks:

The possible risks associated with using a current design, is if the structural configuration can accommodate the desired improvements or new engines. This is of vital importance as we hope to incorporate the use of composites in certain areas as well as new avionic systems.

Advantages

Using an existing design is always cheaper. Designers, manufacturers, investors and airlines will have prior experience with the aircraft and will therefore trust in its capabilities.

The use of a conventional design also provides control and stability benefits. The use of dihedral, sweptback and conventional tail design all provide stability improving characteristics. Also, as the engines are mounted on the underside of the wing they much more easily accessible than that of an tail engine mounted design.

Viability

This configuration will be used over the blended wing body because it is a trusted medium sized commercial aircraft configuration. It is a reliable design and it meets almost all the design requirements set out by our customers. The aircraft can be upgraded and therefore it is possible for us to choose which modifications to make for our purpose. It is also the optimised design for commercial aircrafts at the present time.

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Configuration 2 – T-Tail Configuration With Engines Attached To The Empennage Section

General:

Our proposed configuration will be very similar to the Boeing 737-900NG with a minimum of double engine mounted on the wing section but in this configuration only a single engine can be used and mounted on to the empennage section. T-tail is optimum for SHORT TAKE-OFF AND

LANDING, so T-tail arrangement is also commonly found on smaller airliners with rear mounted engines which are designed for very steep approaches.

Risks:

One of the risks we may face when using this type of configuration, is that it has too many structural problems. The engine being placed on the empennage section creates high stress due to excess load and becomes a source of primary failure. This means it will lead to an additional cost for structure and inspection maintenance.

Advantages:

T-tail is designed for a smoother ride because the tailplane surfaces are kept well out of the airflow behind the wing, giving smoother flow, more predictable design characteristics, and better pitch control. T-tail configuration also allows high performance aerodynamics and excellent glide ratio as the empennage is not affected by wing slipstream. The effective distance between wing and tailplane can be increased without a significant increase in the weight of the aircraft. The horizontal stabilizer is kept farther away from the ground, which helps reduce damage to it by objects on the ground when taking off or landing.

Disadvantages:

T-tail aircrafts will tend to be much more prone to a dangerous deep stall condition due to blanking of the airflow over the tailplane and elevators which could lead to a total loss of pitch control. The vertical stabilizer must be made considerably stronger and stiffer to support the forces generated by the tail-plane. The T-tail configuration can cause several maintenance concerns and can be difficult to casually inspect from the ground. The lack of airflow over the elevator from a mounted low speed engine is reduced and low speed operation is more difficult for aircraft not designed for low speed operation.

Viability

This configuration cannot be used because the horizontal stabilizer is much bulkier than a conventional design. That means all of that surface area being moved up there will cause excess stress on the tail of the aircraft. In conventional designs most of these forces will be placed directly into the airframe. In a T-tail design many of these strong forces are placed upon the empennage which creates more leverage as it transfers its forces to the rest of the airframe. Another reason is because this type of configuration is designed for aircraft with low speed.

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W I N G C O NF I G U R A T I O N

Conventional Wing Configuration

Low-Wing Configuration:

Advantages:

One of the main advantages with low wing configuration in regards to commercial aircraft is its ability to allow for more internal room for payload as the wing structure penetrates below the cabin area. It does this by allowing the main spar to bypass the cabin area and run below the feet of the passengers.

It allows for storage space for the main undercarriage for a convention landing gear design and allows the gears to be attached to the main spar. This combination with the landing gear also allows for easy access, maintenance and refuelling. Upon landing the low wing configuration allows the wheels to be spaced further apart which increases stability.

Disadvantages:

Disadvantages include low ground clearance which is a problem when landing. Since the wing is closer to the ground it is more susceptible to foreign impact damage than that of a high wing design.

Swept-Back Wing:

Advantages:

This design reduces drag during transitions between different Mach levels as well as delay the formation of shock waves. It can also provide more manoeuvrability and agility in the air. Since the air resistance increases rapidly well before the speed of sound, this type of configuration is a solution for aircrafts that want to travel at the speed of sound or just below the speed of sound.

Disadvantages:

When an aircraft with a sweep back configuration travels at high speeds the airflow has little time to react and as a result the airflow flows straight over the wing. At low velocities the air does not have time to react and causes the airflow to flow to the leading edge towards the tip. This is known as spanwise flow and is a problematic as it reduces the overall lift of the wing.

Another concern is the torque transmitted by the wing to the fuselage. Because of the sweep back the centre of lift lies behind the wing root and causes the fuselage to twist when lift is produced.

Dihedral:

Advantages:

The use of a dihedral configuration adds lateral stability which is needed for a low wing passenger airliner. The trick with using a dihedral is by making it slightly unstable to allow for a quick change in direction by the pilot. Using an anhedral design provides too much stability and

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makes it hard to manoeuvre. The dihedral also allows the engines to be raised higher allowing for more clearance to the tarmac.

Disadvantages:

Dihedral is actually aerodynamically less efficient during banking. Since the lift vectors each wing compete with each other and cause more drag.

Viability:

This is the most viable design for a commercial transport aircraft. It has proven with many years of service that it is reliable and trustworthy in design, production and operation. It is also cost effective for our design as all it requires are some avionic or structural modifications for composites or new engine attachments.

Blended Wing Body:

General:

The blended wing body design is comparable to the Airbus A380 payload and range. It is therefore a large passenger commercial transport and has 32% lower fuel burn per seat. It is a new and upcoming design.

Advantages:

An advantage of the blended wing body is that the whole aircraft can produce lift and minimise drag. Therefore it increases fuel efficiency and consumes 20% less fuel than current commercial aircrafts during cruise. It also has larger storage space for passengers and cargo. This type of design exhibits high lift, low drag, reduced ground noise due to top-mounted engines, increased carrying capacity and improved structural weight.

The primary structure will be made of composites which make it lighter and more efficient that aircrafts within its class. It also uses 3 engines rather than 4 like the A380 reducing its overall fuel usage.

Disadvantages:

Modifications are needed for high speed aircraft as tests performed at low speeds. The main drawback of the blended wing body is that it is less structurally stable for internal pressurisation. The blended wing body has many significant advantages over the conventional aircraft design. But as revolutionary as it is, it will require a large and expensive engineering effort to become a reality.

Viability:

The blended wing body is constructed mainly by composites which are very expensive compared with an aluminium aircraft design. It does however have some desirable advantages such as fuel efficiency and noise reduction. It cannot however be used as a design because there is a lack of knowledge when it comes to higher speeds. Aesthetically, it may also be a concern as customers may not understand the safety factor on this particular design as it looks very different to existing transports. This may affect ticket sales.

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E N G I N E S E L E C T I O N

PW1000

Reason for use:

The Pratt & Whitney PW1000 which is to be certified in late 2011 is among the new line of “Pure-Power Engines”. It is a geared turbofan which can provide the necessary thrust requirements for our design. It operates simply by separating the engine fan from the compressor and turbine which allows the whole system to operate at its optimum speeds. This design allows the fan to rotate at a different velocity than the rest of the engine allowing for increased engine efficiency, less noise, less emissions and significant fuel reduction.

Operational/Maintenance Costs:

Compare to the current engine use in the B737-900NG, CFM International CFM56, Pratt & Whitney PW1000 offers a fuel burn improvement of 15%. Used on an aircraft such as the A320 or B737-800, this fuel burn advantage would save the airline US$750,000 per aircraft per year. As airlines usually operate more than one of these possible aircrafts, the savings are substantial over the number of aircrafts and the number of years of service.

Furthermore, PW1000 has a maintenance benefit. This is because this turbofan has less life-limited disks as well as 1500 less airfoil blades. Each disk is approximately US$100,000. These life-limited parts plus lower airfoil blade count will enable the PW1000 to have greater than 20% engine maintenance cost advantage over the CFM56.

In addition, PW1000 has lower noise production, reduced pollution and significant CO2 reduction. This improvement helps a further 2-3% reduction in airline operating costs for avoiding airport noise fee and the CO2 emission penalties.

Risks:

However, one of the drawbacks of choosing the PW1000 is that there is not enough information to predict the future development and improvement for this engine. In the future, there may be a better engine that will be released such as the CFM International Leap-X engine. But seeing as how this aircraft will be released 5 years earlier than the Leap-X, it is a much safer investment as research and experiments will be conducted much earlier. It also gives more time for us to research alternative engines in the event that it does not match the requirements of the RFP or it falls below the predicted performances stated during its design.

Viability:

In terms of noise, carbon emission, thrust capacity and other improvements, the PW1000 matches the requirements of the RFP. It shows significant improvement over existing engines and is soon to be available. This will be our primary engine and the Leap-X will be the secondary engine choice.

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Leap-X

Reason for use:

The LEAP-X engine is made by the same company as the CFM-56’s, the current engines of the B737-900NG. It is a new fully developed variant with implementation of newer technology and greater savings in all the key objectives of our RFP. At this moment the company claims it has possibilities to use alternative fuel sources. As time progresses, test and experiments will be conducted to determine actual types of fuel sources that can be used.

Based on the CFM56 engine performance the new Leap-X engine will out-perform it in every aspect. The range and thrust will be greater then the current engines and will still give better fuel consumption, lower noise and lighter weight. The range will depend on the amount of fuel that can be carried on the aircraft. However based on fuel consumption, the engines run 16% less fuel then the best CFM56 Engine model on the current B737-900NG aircrafts.

Availability:

The engine is already under endurance testing in France as of mid 2009. It will need to be tested for 7 years to be certified under the certification regulations. This will mean the engine is available for sale and fully certified, hopefully by 2016. Although this is the case, the first full demonstrator engine is scheduled to run in 2012 for real time testing, ensuring that the engine does meet the design parameters and performance requirements. Cost for the engine is still not determined. As production of the engine is not underway and in testing phase it would be estimated as similar cost as the current CFM56 engines, which is $2-3USD million dollars each.

The first major test was performed in mid 2009. The results were welcoming in terms of performance as all key criteria were met. The next major test is scheduled on 2012 on its first full run test demonstrator.

Risk:

Until the second major test is underway, the engine may fail to deliver the required performance and specifications by its release date. These may include the correct weight per engine, specific thrust, fuel consumption and emissions. However as the engine is set to be fully certified by 2016, there will be leeway to choose an alternative engine before the 2020 deadline.These risks will not substantially affect our design. As this engine is the alternative option, if the engine fails to meet our key criteria then we will lose our backup option. This may affect our design.

Advantages:

The key advantages compared to the current CFM56 engines are: Less fuel consumption, noise reduction, substantial weight reduction and infinite fan blade life cycle.

Viability:

This will be our secondary engine choice. The reason is because the primary engine is already in production and will be certified by 2011 giving us plenty of time for tests and contingency options in the event that the primary engine cannot deliver. However both engines have similar performance and specifications.

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M A T E R I A L S

Composites

General:

Modern day designed passenger aircrafts are pushing the limits with high-strength, light weight composite materials. Highly researched carbon fibre composite materials are replacing more traditional aluminium alloy materials, especially for the frame structures and skins. This results in the structure possessing more rigidity, strength and less overall weight. Carbon fibre reinforced epoxy composites are the most commonly used composite material in modern airliner passenger jets.

Reason for use:

Typically, the weight saving achieved by replacing an aluminium vertical fin box & inboard aileron with a carbon/epoxy composite, results in a weight reduction of around 28%. Also, by substituting the aluminium parts with carbon/epoxy composites, there will be fewer assemblies, parts & fasteners required. Also, composite materials can easily be built up into exotic shapes and thicknesses, whereas for aluminium, it is hard to accomplish this without several aluminium fittings and one or more skin sections which need to be fastened together.

Composite materials structural properties can be tailored during the manufacturing phase of the material whereas more traditional aluminium metal fittings would require difficult machining and manufacturing techniques. One of the most impressive benefits of using composites in passenger airliners is their outstanding resistance to seawater & chemical corrosion. Aircraft are used most frequently over sea water than land. Therefore using composites will keep the aircraft flying longer and more safely without corrosion problems.

Material Density (g/cm3)

Tensile Strength (GPa)

Tensile Modulus (GPa)

Carbon/Epoxy Composite

1.6 0.93 213

Aluminium Alloy 2.8 0.46 72

Mild Steel 7.8 0.99 207

Titanium Alloy 4.5 0.93 110

Table 2 - Comparison of Material Properties

Carbon Footprint:

During the manufacturing of composites there are toxins and emissions which are released into the air which come from the resin. Unfortunately, any time that uncured resin is exposed to the air, some monomer will evaporate. But once the curing stage is complete, all of the monomers have essentially reacted, and there will be no more evaporation. During flight composite materials used throughout an aircraft will not release any toxins into the atmosphere. This is due to the composite materials all being permanently set into the various forms required by the structure.

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A major problem with composite materials is the disposal of the retired or unwanted parts. There are various ways to dispose of composites, but most solutions currently used are not environmentally friendly. These include burning of the composite, which in turn releases huge amounts of toxins and greenhouse gases, or burying unwanted composites in the earth which can take hundreds of years to break down due to their high integrity and toughness. Recycling technology which is substantial enough to remove all unwanted composite is currently unavailable. Therefore disposal of composite is a future problem.

Availability:

Composite materials for skin sections of wings, fuselage, structural beams and spars are available with the current technology and research for passenger aircraft. Taking the highly developed Boeing 787 as an example it can be seen that a large quantity of passenger aircraft can be made from composites. The 787 is made up of 50% composites by weight resulting in an overall weight saving of around 20% when compared to its replacement aircraft, the B777.

Risks:

By using large amounts of composite materials in a passenger aircraft, there are moderate risks involved. This is due to the latest developments in composites being so new, which leaves the company with minimal information on composites over a long period of time when compared to all the readily available information on aluminium alloys used in aircraft. This of course does not mean that composites aren’t safe to use for passenger aircraft. Extremely rigorous and highly extensive research has been conducted on all composite components before they go anywhere near an aircraft.

Advantages:

If composite materials are used in both secondary and structural sections, the result would be an aircraft with a total weight reduction of around 20%. This would, in turn improve the aircrafts fuel efficiency which would result in less overall emissions and greenhouse gas being released into the atmosphere. This can be seen in the new, highly developed Boeing 787 and future Airbus A350 passenger airliners.

This possible weight reduction and use of composites will lead to lower emissions, increased performance, longer life, smaller maintenance intervals and increased overall fuel saving. Cost saving will also occur due to the high quality properties and corrosion resistance of composites.

Disadvantages:

There are a few disadvantages in using composites in our design. This includes an initial increase in manufacturing cost. The cost of composites used in aerospace applications is higher than for steel and aluminium components due to the higher cost of raw materials and higher production costs. Though, in the long run more money will be saved due to the fuel-saving and reduced through-life maintenance costs. Also if any of the composite components are damaged, for example by bird-strike or other foreign object damage, then the repair costs are higher than repairing aluminium. Composites also exhibit extremely low tolerances for lightning strike. This continues to be a problem.

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Application:

After researching the passenger aircraft which our new, environmentally friendly passenger aircraft is going to replace, including the Boeing 737 NG and Airbus A320, it was found that very few components on the aircraft were made from composite material. This can be put down to a few reasons. Firstly when these aircraft were designed and built the technology and research on composite materials were minimal and not really an option. Also these two airliners are moderately small when compared to jumbo jets like the Boeing 787 and Airbus A350. Therefore there was no great need to use large amounts of composites in those aircraft due to the fact that the overall weight saving produced wouldn’t be worth the overall research and cost needed, since it wouldn’t save an adequate amount of fuel.

In our design we will use various composite components for our aircraft including; parts of the wings and wing control surfaces, fin box and rudder, tailplane and elevator, engine cowlings and nose section. The rest of the aircraft will be made primarily from aluminium alloys.

Viability:

In terms of a single aisle aircraft, the benefits of using a ~50% composite design is just not feasible. This is because the benefits of using high levels of composites do not outweigh the increased manufacturing and maintenance costs. Therefore the decision to use a select amount of composite materials, around 15% weight in our design is justifiable. This will provide enough weight saving to increase efficiency whilst still ensuring a safe reliable passenger.

Aluminium

General:

Aluminium alloys are widely used in the construction of aircraft structures such as wings and fuselages due to their superior metal fatigue resistance and their light weight (when compared to most other metallic materials). The strength-to-weight ratio of aluminium alloys is proficient enough to be capable of supporting the aircraft under all loading conditions. With its high strength and excellent fatigue resistance, it is used to advantage on structures and components on the aircraft with the knowledge that they won’t fail during flight.

Reason for use:

In cases where aluminium alloy components are damaged during flight or when on the ground, the repair and replacement of the parts is relatively cheap and easy. Most basic aircraft maintenance workshops would have the necessary equipment and tools needed to achieve a successful repair.

Aluminium alloys have been widely used in the aerospace industry since the early 1900’s, adequate amounts of information on their structural properties and loading cases are readily available. With this information it is safe to say that aluminium alloyed aircraft can fly safely and sufficiently for passenger aircraft.

Carbon footprint:

The effect of aluminium and its alloys on the environment is reduced when compared to composite materials. This is due to the fact that aluminium can be easily recycled. This advantage over composite materials is a big positive since there are many older aircraft which are retiring and in need of disposal. Recycling large quantities of these aircraft is great for the environment since there is a large reduction in waste.

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Though, during the production and manufacturing of aluminium in foundries and recycling plants, there are large amounts of emissions and greenhouse gases produced. This can’t be helped since the aluminium needs to be heated to very high temperatures to achieve necessary levels of quality.

Risks:

The risks involved with using aluminium alloys in aerospace structures are fairly minimal due to the wide availability and use in the current aerospace industry. There are small risks including the production of components not being manufactured up to the approved standards, or complexity of components rendering more fasteners and aluminium parts, which over the design stage increases the overall weight.

Availability:

Aluminium alloyed structures and components are readily available and in constant production around the world. This is due to aluminium being the most widely used material for the production of aircraft especially for passenger airliners. The price of manufacturing the aluminium then producing specific parts is relatively cheap, especially when compared to the price of composite materials. Since aluminium is widely available and recyclable, it is much easier to produce a good component in a shorter amount of time.

Advantages:

Aluminium alloys have many advantages when using them for aerospace applications. In terms of cost, they are one of the cheapest aerospace materials. Due to their high strength-to-weight ratio and fatigue resistance, aluminium alloys will be continued to be used throughout aircraft structures. When disposing of unused or retired aluminium alloyed aerospace structures, it is easily done by recycling the components by melting and re-using the aluminium and alloys for other cheap products, like aluminium cans.

During flight, when the aircraft is subjected to poor weather conditions like thunderstorms, with the use of aluminium alloys the aircraft will be protected from lightning strikes. Unlike composite materials, the aluminium alloys will dampen the lightning strike leaving minimal damage.

In the case where damage does occur, the repair of aluminium components is routine.

Disadvantages:

The biggest downfall of using aluminium alloys is their susceptibility to corrosion. Since aircraft fly in all weather conditions, including rain and freezing temperature, it is only a matter of time before the aluminium components and structures begin to corrode and weaken. To reduce the effect and cost of corrosion, aircraft are maintained and inspected regularly and always have a good coat of aerospace grade paint. However, this results in more maintenance events than composites.

Even though aluminium alloyed structures are lightweight, they are still not as light as composite materials. When comparing the strength-to-weight ratios of composites and aluminium alloys, composite materials are far more superior.

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Viability:

The use of aluminium will remain the majority of the aircraft structure. This is because the use of extensive composites is not beneficial for a single aisle aircraft. By using slightly lower levels of aluminium than that of the Boeing 737-900 NG and replacing the rest with composites, we can reduce the overall costs of production as well as still achieve some weight saving through the use of composites.

A L T E R N A T I V E F U E L S O U R C ES

Biodisel

Reason For Use:

Biodiesel is a type of biofuel that is manufactured from renewable energy sources such as plant crops, oils or animal fats. Jet fuel blended with biodiesel has already been tested on the Boeing 737-800 and has proved successful for a two hour test flight. The 737-800 model uses the same model engine as the 737-900 and can therefore be utilised on our base design. Its energy content is also higher than petroleum based aviation fuels which show that it has the potential to replace it in the future. The National Renewable Energy Laboratory has run tests which show that biodiesels substantially reduce carbon dioxide emissions. This is however dependent on the type of crop used to produce the fuel. Biodiesel can be transported, delivered and stored in very much the same way as diesel fuel whilst using the same equipment making it easily controlled. It has approximately the same weight, volume and performance as Jet-A Fuel and therefore it is relatively easy to use and no design changes are necessary for the intended aircraft.

Availability:

The use of biofuel in aerospace is still under development. With current economic climate, crop production, oil production and extraction process factors all in mind, it is clear that its success is dependent on its cost-effectiveness. Even though it is a renewable resource, its rate of production compared with that of Jet-A fuel is poor. However, seeing as we will most likely be utilising a 50-50 mix between biodiesel and kerosene, the resultant product, bio-kerosene, will most likely be cheaper than Jet-A fuel and pure biofuel.

Risks:

The price of biofuel depends on the ingredients, availability of necessary resources and the currency between countries. Therefore it will vary from time to time. Biofuel is still regarded as a new product and therefore little is known about it compared with that of Jet-A fuel.

Disadvantages:

The problem with producing biofuel is that an extremely large land mass is required to produce a sufficient supply. Therefore deforestation might occur and in turn increase the level of CO2

emissions. However, this problem can be solved depending on the types of crops used. Another issue with deforestation is that the crops being manufactured for biofuel could compete with existing food production which could ultimately increase the food price for the future.

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Biofuel does produce fewer emissions during operation than that of existing fuel sources; however it still produces emissions during production, acquisition and transportation. Overall, the possible emissions saved by use of biofuel are slightly reduced.

Biofuel also have issues in regards to its freezing points. When used for jet aircraft it must be tailored to solve the freezing problem. Research shows that mixing biofuel with Jet-A fuel can resolve this problem.

Storage is not really an issue as the handling of biofuel is very similar to the practices used for Jet-A fuel and therefore the similar techniques and existing equipment can be utilised.

Hydrogen

Hydrogen is a potential renewable resource and the most considerable long term renewable aviation fuel. Pure hydrogen can be generated using electrolysis and biomass which is a simple chemical process. This process can be supported by renewable energy such as wind turbine generators, wave and tidal stream energy or nuclear energy.

Reasons for use:

Hydrogen is an alternative fuel source possibility. It produces fewer emissions than kerosene during production and also has a considerable weight reduction compared with current jet fuel. The method for storing or delivering hydrogen is similar to the practices for natural gas transportation and utilisation. Therefore, new storage technology is not required.

Availability

Hydrogen has been used in the past and is still under research as an alternative fuel source because of its potential on commercial aircraft. The reason for this is that the energy content is similar to kerosene with lower densities. In early tests it has shown that hydrogen can be used on current engines with some slight modifications on the combustion section. Positioning of storage tanks and structural support has to unfortunately be redesigned. This is due to its chemical and physical properties. When used in airports there are a number of issues associated with its use. Pressure vessels must be built for compressed or liquid hydrogen and the transportation of these fuels is limited.

The cost of hydrogen production in terms of biomass and electrolysis are quite different. Biomass technique reveals a similar price compared with kerosene. However, electrolysis is approximately one and half times more expensive.

Risks

Due to hydrogen’s chemical properties, it has low vapour and ignition temperatures which exhibit difficulties when it comes to storing, transporting or maintaining it. In order achieve reasonable efficiency for jet engine use, the hydrogen must first be condensed to a liquid or compressed gas stage. Hence, it must be stored in a pressure vessel. Also, the production of hydrogen vapour is apparent and must be addressed during transportation due to its low ignition point. The design of the aircraft structure must also be able to reduce and control the level of hydrogen vapour. This problem is somewhat similar to problems that kerosene experiences and therefore there is the chance of improvement.

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Disadvantages:

Problems associated with this fuel are the distribution of resources, compatibility issues with some current aircrafts and pollutant levels. Its methods of production have with them geographic limitations. Wind turbine generation is only available in a few countries and the use of nuclear power is not favoured by many countries. Biomass technique is only provided in some countries.

Emission levels during production and operation are also an issue. The combustion process of hydrogen only discharges water. However, oxides of nitrogen are secondary emissions caused by the engines. The water vapour released is also a kind of greenhouse gas and varies with altitude but has a low residence time which is a benefit. Lastly, the overall nitrogen oxides released is relatively low on liquid hydrogen for jet engines.

The last drawback is that this type of fuel requires the aircraft structure and propulsion systems must be modified. This is an issue as some aircraft just cannot support these kinds of modifications. Also, ground support equipment must be built which increases overall costs.

Liquid Methane

Liquid methane is a type of cryogenic fuel. It is a gas at room temperature and has been cooled to its boiling and stored as a low temperature liquid.

Reason for use:

Methane is a type of natural gas that is stored within the earth and could potentially offer a fuel source for hundreds of years. Is it considered as a renewable energy source and can be produced using renewable energy methods. Biomass technique is used to produce and as mentioned earlier can be supported by wind turbine generators, wave and tidal stream energy or nuclear power. With liquefied natural gas the CO2 emission levels are approximately 25% lower than kerosene levels. With the use of bio-methane (methane acquired using biomass technique) the potential savings could be much more.

Availability:

There are vast quantities of methane trapped within the earth in the form of methane hydrates. It can possible be available in the future.

Risks:

Although there is the chance of it being available in the future in regards to its current progress. The question is when. This sort of risk is just not viable when it comes to designing our proposed aircraft.

Methane itself is a greenhouse gas. Producing it is very complex as well as costly. Throughout its fuel cycle acquisition and operation it undergoes the following procedures: extraction, production, distribution and combustion. All of which produce emissions.

Disadvantages:

The use of methane is new and complex at all stages of its utilisation. Ground transportation, storage and vent capture systems are very difficult to control and manage. Also, aircrafts which utilise jet-A fuel are not readily compatible with liquefied methane.

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Advantages of Biodiesel, Hydrogen and Liquid Methane:

Table 1 and figure 1 show the advantages of these alternate fuels compared with that of Jet-A fuel .It can be clearly seen that they have substantial advantages in some areas.

Fuel Specific Energy

(MJ/kg) Density at 15˚C

(kg/100 MJ)

Jet A 43.2 0.808

Bio-kerosene (Unknown percentage of Biodiesel developed by Boeing)

48 (approx.) Unknown (lighter

than Jet A)

Biodiesel (Typical) 38.9 0.87

Liquid Methane 50 0.424

Hydrogen 120 0.071 Table 1 – Specific Energy and Density levels

Figure 1 – Relative CO2 emissions of alternative fuel resources compared to Jet fuel

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Fuel Choice:

Liquid Methane:

From figure 1, we can see that use of methanol from natural gas can substantially reduce CO2 emission. However, methane is a powerful greenhouse gas and leakage of methane is unavoidable when using it. Liquid methane is not chosen for our project due to its performance, availability and its greenhouse gas emission.

Hydrogen:

As shown on the table and figures above, hydrogen is the best alternative by its properties. It does not emit any CO2, has the highest energy content and lowest density in the three alternative fuels chosen. However, the use of hydrogen and liquid methane requires modifications and redesign of all existing aircrafts. The technology is not mature and it is unlikely to be used in the near future as too many changes such as storage and transportation in airports and aircrafts will have to be made.

Biodiesel:

On the other hand, biodiesel seems like a more suitable alternative to Jet-A fuel because it has a similar energy content and density with fewer emissions. Boeing also claims that it’s developing bio-jet fuels that perform better than the conventional Jet-A fuel. Its emission is much lower and can be used in aircraft with conventional design. It is also being tested and likely to be extensively used in the aviation industry in the near future.

Therefore biofuel will be chosen to be our priority fuel source. It will work on most engines and requires no modifications on the aircraft, which implies that our design will work with the conventional Jet-A fuel as a contingency plan if in the future biodiesel becomes unviable. There are several types of biofuel and all appear to be compatible with Jet A fuel engines. Therefore we could choose the most suitable biofuel to be used with our project in the future. The biofuel that could potentially replace Jet-A fuel at the moment is the biodiesel and kerosene blended fuel Boeing is developing.

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G R O U N D S U P P O R T A N D I N F R A S T R U C T U R E

General:

There are a number of solutions that improve the overall efficiency that are not related directly to the aircrafts design. Improvements in the supports of an aircraft can improve the overall cost and maintenance of the aircraft. These solutions can be in implementation through a more efficient management vehicle system, thus reducing the ground time of the aircraft. Another option could be to implement hydrogen fuel cells to the management vehicles.

Risks:

The major risks when implementing new ground support systems are related to cost and down time. While the new system will improve overall efficiency and thus save money in the long term, the set up cost may be too high for an airport to handle in a short period of time. Another problem comes from the down time of sections of the support systems as they are upgraded putting more work onto the operating sections. This may cause problems/delays while the new system is being introduced.

Advantages:

The introduction of a more efficient ground support system will improve the overall efficiency of the aircraft as its ground time will be reduced. Also the implementation of cleaner fuel sources, such as hydrogen in the management vehicles will reduces the emissions of the airport thus leading to a greener solution for the management of the aircraft. If the overall operational costs of the aircraft are reduced the ticket price can also be reduced thus encouraging more people to fly on the aircraft and thus increasing sales.

Disadvantages:

The disadvantage of the implementing of a new system comes from the use of un-proven technology. There will be revisions to the logistics and procedures to remove problems associated with the new system. This will lead to some sectors being unable to do required duties as they are currently undergoing upgrades. The introduction of a new system can also be a costly venture.

What options will be advised to be implemented with the aircraft.

There are two systems the will be advised with the implementation of this aircraft. These are that the management vehicles are upgraded to a hydrogen fuel cell. This will dramatically reduce the carbon-footprint of the ground support system. The systems are the “Zebra Enterprise” and “Airport Visualiser” software. This system will allow for a more efficient use of the ground equipment. The system will reduce idle time for the ground vehicles and storage facilities. Thus improving productivity and reducing operations costs.

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O P E R A T I O N A N D M A I N T E N A N C E C O S T S

Cost Reduction Methods:

If biofuel is used then production in a vicinity closer to the airport can reduce transportation costs as well as emissions levels. A similar production method may be used for hydrogen. As we are hoping to change management vehicles within the airport to utilize alternate fuels, we can reduce not only transportation costs but also operational ones within the airline.

With new technologies available in terms of maintenance and design of an aircraft, we hope to reduce maintenance time intervals by increasing wear and tear resistance of the aircrafts. In the long term, increasing maintenance intervals even by only one year will exponentially reduce costs.

With the introduction of new engines, maintenance is now easier to perform and servicing is much more cost effective. These factors mentioned above will all reduce operation and maintenance costs within the airline and in the manufacturing sense, which will in turn hopefully reduce airfares.

As aerospace technology and production methods increase in quality, we hope to reduce production time by outsourcing components to other manufacturers around the world. Reasons for this include cost saving and an incentive for these countries involved to purchase the proposed aircraft.

Purchase Price:

At the present time we are assuming the price of our aircraft will be around the same as that of the existing Boeing 737-900NG and A320-200 models which is between US$51.5-87M. Even if our aircraft is slightly more expensive, the amount of potential savings which come from manufacturing reduction time, fuel efficiency and low maintenance will cover this slight increase in price.

Operational/Maintenance Costs

We are assuming that operational and maintenance costs are similar to existing aircrafts within the same class. In terms of maintenance, a twin engine narrow-body aircraft is US$515/hour and a twin engine wide-body is US$787/hour. These values are based on the Bureau of Transportation Statistics 1999-1998. Delays cost approximately $10000 per hour and operational costs, based on the Boeing 737-800 is US$1665 per hour.

Conclusion:

Using current technologies, engines, materials and engineering practices we are assuming that this aircraft will be lower maintenance and much more cost-effective throughout its life than the aircraft it intends to replace.

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23 APPENDIX 9 – COMPLIANCE REPORT – B737-900NG

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24 APPENDIX 10 – COMPLIANCE REPORT - A320-200

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25 APPENDIX 11 – DENSITY CALCULATION

Density Calculation

Sea-Level Conditions

ρ 1.225 kg/m3

P 101325 Pascals

T 288 Kelvin

Cruise Altitude

Feet Meters

41000 12496.8 (h)eight

a -0.0065 k/m

R 287.4 J/kg-k

At Cruise Altitude

ρ 0.30 kg/m3

P 17784.18 Pascals

T 206.77 Kelvin

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26 APPENDIX 14 - AIRFOIL GEOMETRY

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27 APPENDIX 15 - MATCHING CHART CALCULATIONS

Take-off Distance - (FAR 25 Requirement)

CL-MAX-To (Trasport Jets Table 3.1)

W/S 1.6 1.8 2 2.2

0 0 0 0 0

50 0.234 0.208 0.188 0.170

100 0.469 0.417 0.375 0.341

150 0.703 0.625 0.563 0.511

200 0.938 0.833 0.750 0.682

250 1.172 1.042 0.938 0.852

300 1.406 1.250 1.125 1.023

Landing Distance - (FAR 25 Requirement)

CL-MAX-L (Trasport Jets Table 3.1)

W/S 1.8 2 2.2 2.4 2.6 2.8

0 0 0 0 0 0 0

40 70.758 78.62 86.482 94.344 102.206 110.068

80 70.758 78.62 86.482 94.344 102.206 110.068

120 70.758 78.62 86.482 94.344 102.206 110.068

160 70.758 78.62 86.482 94.344 102.206 110.068

200 70.758 78.62 86.482 94.344 102.206 110.068

240 70.758 78.62 86.482 94.344 102.206 110.068

Stall

1.2 1.4 1.6 1.8

Wing Loading 228.522 266.609 304.696 342.783

Climb

1.2 1.4 1.6 1.8

Wing Loading 302.220 352.590 402.960 453.330

Thrust/Weight 2.531 2.169 1.898 1.687

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Cruise

W/S

CDO Density at Cruise

(lbs/ft3) Cruise Velocity

(kts) Dynamic Pressure

Oswald Factor

Aspect

Ratio T/W

40 0.0376 0.02 542.4 2754.868 0.85 9.45 2.590

80 0.0376 0.02 542.4 2754.868 0.85 9.45 1.296

120

0.0376 0.02 542.4 2754.868 0.85 9.45 0.865

160

0.0376 0.02 542.4 2754.868 0.85 9.45 0.650

200

0.0376 0.02 542.4 2754.868 0.85 9.45 0.521

240

0.0376 0.02 542.4 2754.868 0.85 9.45 0.435