3 Seat Light-Sport Aircraft Development | Alon Contents: Report Attachments

POLYTECHNIC UNIVERSITYOF CATALONIA

AEROSPACE ENGINEERING

Engineering Projects

3 seat Light-Sport Aircraft development | AlOn Contents: Report Attachments

Group: G06-AE-2018/19-Q1

Delivery date: 20/12/2018

Students: BERNAD SERRA, P. CARRILLO CÓRCOLES, X. FERNÁNDEZ MARTÍNEZ, A. GAGO CARRILLO, E. GÓMEZ ESCANDELL, E. KALINA CAPDEVILA, A. Supervisor MARIN DE YZAGUIRRE, M. Lluís Manuel Pérez Llera MÉNDEZ GÁLVEZ, C. MEDINA RODRÍGUEZ, C. NADAL VILA, P. PÉREZ RICARDO, C. RODRIGUEZ POZO, D. SANS MOGULLÓ, A. UGARTEMENDIA RODRÍGUEZ, I. G06-AlOn LSA 3 seats | Project report

RA 2 Contents

1 Aerodynamics 15 1.1 Wing ...... 15 1.1.1 Airfoil selection ...... 15 1.1.1.1 Polar distribution ...... 16 1.1.1.2 Efficiency versus α ...... 16 1.1.1.3 Momentum versus α ...... 17 1.1.2 General Wing Plant forms ...... 17 1.1.2.1 Rectangular ...... 18 1.1.2.2 Elliptical ...... 18 1.1.2.3 Trapezoidal ...... 18 1.1.2.4 Delta ...... 18 1.1.3 Plant form definition ...... 18 1.1.3.1 Initial efficiency analysis ...... 19 1.1.3.2 Airfoil definition result ...... 19 1.1.3.3 Validation of the initial wing parameters configuration . . . . . 20 1.1.3.4 Wing Efficiency ...... 22 1.1.3.4.1 Aerodynamic Torsion ...... 22 1.1.3.4.2 Wingtips ...... 22 1.1.3.5 Stall Behaviour ...... 23 1.1.3.6 High-lift Device ...... 24 1.1.4 Wing configuration criteria ...... 25 1.1.4.1 Unweighted Average Method ...... 26 1.1.4.2 Ordered Weighted Average Method ...... 26 1.2 Tail ...... 26 1.2.1 Definition of tail ...... 26 1.2.2 Utility of a tail ...... 26 1.2.3 General tail designs ...... 27 1.2.3.1 Conventional tail design ...... 27 1.2.3.2 T-tail design ...... 28 1.2.3.3 Cruciform tail design ...... 28 1.2.3.4 V-tail design ...... 29 1.2.3.5 Twin-tail design ...... 29 1.2.3.6 Boom-tail design ...... 29 1.2.3.7 Dual tail design ...... 29 1.2.3.8 Triple tail ...... 29 1.2.3.9 Other tail configurations ...... 29 1.2.4 Tail configuration selection ...... 29 1.2.4.1 Simple Hierarchy Method ...... 30 1.2.4.2 Unweighted Average Method ...... 30 1.2.4.3 Ordered Weighted Average Method ...... 31 1.2.5 Decision making ...... 31 1.2.6 Tail design ...... 31 1.2.6.1 Parameters determination ...... 31

RA 3 G06-AlOn LSA 3 seats | Project report 1.2.6.1.1 Horizontal tail ...... 32 1.2.6.1.2 Vertical tail ...... 33 1.3 Fuselage ...... 33 1.3.1 Fuselage selection criteria ...... 35 1.3.1.1 Unweigthed Average Method ...... 36 1.3.1.2 Ordered Weighted Average Method ...... 36 1.4 Control Surfaces ...... 36 1.4.1 Elevator ...... 37 1.4.2 Ailerons ...... 38 1.4.3 Rudder ...... 39 1.4.4 Parasite Drag ...... 40 1.4.4.1 Wheels parasite drag ...... 41 1.4.4.2 Tube parasite drag ...... 42 1.4.4.3 Drag correction ...... 42 1.5 Final plane conﬁguration ...... 42 1.5.1 Design parameters ...... 42 1.5.2 Range study ...... 43 1.6 Final Plane Analysis ...... 45 1.6.1 Efﬁciency ...... 45 1.6.2 Static Stability ...... 45 1.6.3 Dynamic Stability ...... 46 1.6.3.1 Longitudinal Mode 1 ...... 46 1.6.3.2 Lateral Mode 1 ...... 46 1.6.3.3 Lateral Mode 2 ...... 47 1.7 Flight envelope ...... 47

2 Structures 51 2.1 Materials ...... 51 2.1.1 Fuselage and wings internal structure ...... 51 2.1.2 Skin of the aircraft ...... 53 2.1.3 Landing gear ...... 54 2.1.4 Transparent surfaces ...... 55 2.1.4.1 Windshield ...... 55 2.1.4.2 Windows ...... 55 2.1.5 Additional material considerations ...... 55 2.1.5.1 Welding ...... 55 2.1.5.2 Corrosion Prevention ...... 55 2.2 Landing Gear ...... 56 2.2.1 Analysis of the landing gear regulations for an LSA ...... 56 2.2.1.1 Landing gear options ...... 57 2.2.1.2 Atec 322 Faeta ...... 57 2.2.1.3 TL-2000 Sting S4 ...... 57 2.2.1.4 TL-3000 Sirius ...... 57 2.2.1.5 Pipistrel Taurus M ...... 58 2.2.1.6 Alexander Schleicher ASG 29 E ...... 58 2.2.1.7 ONE Aircraft ...... 58 2.2.1.8 Sling 4 ...... 58 2.2.2 Landing gear calculations ...... 59 2.2.2.1 Static analysis ...... 59 2.3 Wing ...... 64 2.3.1 Initial analysis ...... 64 2.3.1.1 Lift approximation ...... 65 2.3.1.2 Weight approximation ...... 65 2.3.1.3 Results ...... 65

RA 4 G06-AlOn LSA 3 seats | Project report 2.3.2 Sizing of the beam ...... 66 2.3.2.1 I section ...... 67 2.3.2.2 Square section ...... 68 2.3.2.3 Results ...... 69 2.3.3 Ribs ...... 70 2.3.4 Final Result ...... 71 2.4 Fuselage ...... 72 2.4.1 Analysis ...... 74 2.4.1.1 Performed analysis ...... 75 2.4.1.2 Results of the analysis ...... 75 2.5 Tail ...... 79 2.5.1 Elevator beam ...... 79 2.5.2 Fin beam ...... 81 2.5.3 Ribs ...... 83 2.5.4 Final Result ...... 83

3 Power plant 85 3.1 Motor ...... 85 3.1.1 OWA ...... 87 3.1.2 PRESS ...... 87 3.2 Motor Mount ...... 90 3.3 Auxiliary motor systems ...... 91 3.3.1 Gear box ...... 91 3.3.2 Exhaust system ...... 91 3.4 Propeller ...... 92 3.4.1 Custom propeller design ...... 93 3.5 Overall Powerplant evaluation ...... 95

4 Systems and avionics 97 4.1 Flight and Navigation Instruments ...... 97 4.1.1 Analogic Flight and Navigation Instruments ...... 98 4.1.2 Electronic Flight and Navigation Instruments ...... 102 4.2 Powerplant Instruments ...... 109 4.3 Miscellaneous Equipment ...... 117 4.3.1 Aircraft Lights ...... 117 4.3.2 Batteries ...... 118 4.3.3 Fire extinguishers ...... 122 4.3.4 Aircraft’s parachute ...... 122 4.3.5 Flight Recoder ...... 123 4.4 Safety Belts and Harnesses ...... 125 4.5 Installation ...... 126 4.6 Controls ...... 126 4.6.1 Controls for wing ailerons ...... 127 4.6.2 Controls for tail elevator ...... 127 4.6.3 Controls for tail vertical estabilitzer ...... 127 4.6.4 Control Cables and Accessories ...... 128

5 Overall Design 129 5.1 3D design ...... 129 5.1.1 Exterior ...... 129 5.1.2 Interior ...... 130 5.1.2.1 Structure ...... 130 5.1.2.2 Power plant ...... 130 5.1.2.3 Instruments and equipment ...... 131

RA 5 G06-AlOn LSA 3 seats | Project report 5.1.2.4 Seats and passengers ...... 131 5.2 Blueprints ...... 133

6 Business 135 6.1 Manufacturing costs ...... 135 6.1.1 Cost of facilities ...... 135 6.1.2 Cost of human resources ...... 137 6.1.3 Additional costs ...... 138 6.1.4 Marketing costs ...... 138 6.1.4.1 Initial marketing campaign costs ...... 138 6.1.4.2 Study of marketing costs over time ...... 139 6.1.4.3 Study of future marketing campaigns ...... 140 6.2 Marketing campaign ...... 141 6.2.1 SWOT analysis ...... 141 6.2.1.1 Research on Market Opportunities ...... 141 6.2.1.2 Research on Market Threats ...... 141 6.2.1.3 Research on Aircraft Strengths ...... 142 6.2.1.4 Research on Aircraft Weaknesses ...... 142 6.2.2 Study of potential customers ...... 142 6.2.3 Set marketing goal ...... 143 6.2.4 Study of advertisement ...... 144 6.2.5 Design of marketing campaign ...... 144 6.3 Initial investment ...... 146 6.3.1 Research on possible investors ...... 146 6.4 Payback Analysis ...... 146 6.4.1 Study of proﬁtability margins ...... 146

7 Organization, planning and scheduling. 149 7.1 Gantt Diagram ...... 149

8 Minutes of the Meeting 151

Bibliography 179

A Code 183 A.1 Weigth-Range Diagram ...... 183 A.2 Gust-Airspeed envelope ...... 184 A.2.1 Second grade polynomial ﬁtting ...... 186 A.2.2 First grade polynomial ﬁtting ...... 186 A.3 Propeller Design ...... 186 A.3.1 Variable declaration ...... 187 A.3.2 Core function ...... 187 A.3.3 importAirfoil ...... 187 A.3.4 discretization ...... 188 A.3.5 nonDymensionalization ...... 188 A.3.6 getLambdaInduced ...... 188 A.3.7 getSpecs ...... 190 A.3.8 Energy ...... 191 A.3.9 Other secondary functions ...... 191 A.4 Gantt diagram ...... 193

RA 6 List of Figures

1.1 Polar distribution of the selected airfoils ...... 16 1.2 Cl versus α of the selected airfoils ...... 16 Cd 1.3 Momentum coefficient of the airfoils selected ...... 17 1.4 Efficiency curve for the N2305 and the B24 airfoils ...... 20 1.5 NACA 23015 ...... 20 1.6 Trapezoidal Wing, S=7,5 m2 ...... 21 1.7 Trapezoidal Wing, S=12 m2 ...... 21 1.8 Efficiency comparison of trapezoidal wings with different surface ...... 21 1.9 Efficiency comparison between wing with and without aerodynamic torsion . . 22 1.10 Winglets ...... 23 1.11 Elliptic Tip ...... 23 1.12 Efficiency comparison between elliptic wingtips, winglets and without wingtips 24 1.13 Cl comparison between torsion and non-torsion wing ...... 24 1.14 Efficiency comparison between torsion and non-torsion wing ...... 25 1.15 Tail configurations ...... 28 1.16 Seats configuration ...... 34 1.17 Efficiency in both configuration ...... 35 1.18 Cm in both configurations ...... 35 1.19 Elevator Airfoil ...... 37 1.20 Hmom Elevator ...... 37 1.21 Cl versus α for the elevator at its maximum deflection angle ...... 38 1.22 Aileron Airfoil ...... 38 1.23 Hmom Ailerons ...... 39 1.24 Rudder Airfoil ...... 39 1.25 Hmom Rudder ...... 40 1.26 Cl versus α fortherudderatitsmaximumdeflectionangle ...... 40 1.27 CD experimental data ; Source: researchgate.net ...... 41 1.28 Comparison between both behaviours ...... 42 1.29 Weight-Range diagram ...... 44 1.30 Efficiency of the final configuration ...... 45 1.31 Static stability of the final configuration ...... 45 1.32 Longitudinal Mode 1 ...... 46 1.33 Lateral Mode 1 ...... 47 1.34 Lateral Mode 2 ...... 47 1.35 Airspeed envelope diagram ...... 49 1.36 Gust-Airspeed envelope diagram ...... 50

2.1 CentrAl conﬁguration ...... 52 2.2 First variant of the TL-3000 Sirius’ landing gear ...... 57 2.3 Second variant of the TL-3000 Sirius’ landing gear ...... 58 2.4 ASG 29 E ﬂying ...... 58 2.5 Front leg of the landing gear...... 59 2.6 Half of the rear landing gear...... 59

RA 7 G06-AlOn LSA 3 seats | Project report 2.7 Stress analysis of the front landing gear...... 61 2.8 Displacements of the front landing gear...... 62 2.9 Stress analysis of half of the rear landing gear...... 62 2.10 Displacements of half of the rear landing gear...... 63 2.11 Tension results of the dropping test...... 63 2.12 Detail of the point where the tension gets its maximum value...... 63 2.13 Displacement results of the dropping test...... 64 2.14 Loads distribution and approximations ...... 64 2.15 Moment diagram for n=1 ...... 65 2.16 Moment diagram for n=4 ...... 66 2.17 Moment diagram for n=-2 ...... 66 2.18 Stress distribution for composite beam ...... 67 2.19 I section parameters ...... 67 2.20 Stress distribution ...... 68 2.21 Square section parameters ...... 68 2.22 Stress analysis ...... 69 2.23 Displacement results ...... 70 2.24 Sandwich ribs ...... 70 2.25 Section approximation ...... 71 2.26 Beam and ribs ...... 72 2.27 Fuselage with a beam thickness of 1 cm...... 73 2.28 Fuselage with a beam thickness of 0.5 cm...... 73 2.29 Fuselage with a beam thickness of 0.5 cm and a reinforcement...... 73 2.30 Fuselage with a U beam...... 74 2.31 Fuselage with a rectangle beam...... 74 2.32 10mm thick I:Results of the stress analysis...... 76 2.33 10mm thick I:Results for the displacement...... 77 2.34 5mm thick I:Results of the stress analysis...... 77 2.35 5mm thick I:Results for the displacement...... 77 2.36 5mm thick I with reinforcement:Results of the stress analysis...... 78 2.37 5mm thick I with reinforcement:Results for the displacement...... 78 2.38 3,5mm thick U:Results of the stress analysis...... 78 2.39 3,5mm thick U:Results for the displacement...... 79 2.40 16mm Rectangle: Results of the stress analysis...... 79 2.41 16mm Rectangle:Results for the displacement...... 79 2.42 Elevator momentum diagram ...... 80 2.43 Elevator FEM analysis ...... 81 2.44 Fin moment diagram ...... 82 2.45 Fin FEM analysis ...... 82 2.46 Beams and ribs ...... 83

3.1 Wankel motor IAE50R – AA ...... 85 3.2 IAE50R – AA dimensions ...... 89 3.3 IAE50R – AA motor mount points ...... 90 3.4 Standard frame mount ...... 90 3.5 Gear Box ...... 91 3.6 HAUTECLAIRE Propeller ...... 93 3.7 Results of computation ...... 94

4.1 Winter 7 FMS 5 Air Speed Indicator ...... 98 4.2 WINTER EBH Altimeter ...... 98 4.3 WINTER Vanetype Variometer 5 STV 5 ...... 99 4.4 Kelly MFG RCA22 Artiﬁcial Horizon ...... 99 4.5 Kelly RCA15BK-1 Directiornal Gyro ...... 100

RA 8 G06-AlOn LSA 3 seats | Project report 4.6 WINTER Turn Coordinator QM II ...... 101 4.7 WINTER Hours Counter Analog ...... 101 4.8 SV-D600/B 7" SkyView SE Display ...... 102 4.9 Dynon Avionics Skyview Display Harness ...... 102 4.10 SV-ADAHRS-200 Air Data/Attitude/Heading Reference System ...... 103 4.11 Location of SV-ADAHRS-200 Air Data/Attitude/Heading Reference System . 103 4.12 SV-BAT-320 SkyView Backup Battery ...... 104 4.13 SV-GPS-2020 GPS Receiver/Antenna ...... 104 4.14 SV-XPNDR-261 Mode-S Transponder ...... 105 4.15 SV-KNOB-PANEL/H SkyView Knob Control Panel ...... 105 4.16 SV-COM-C25/H SkyView VHF Com Radio 25 kHz ...... 106 4.17 SV-INTERCOM-2S Two-Place Stereo Intercom ...... 107 4.18 SV-NET-3CC Network Cable ...... 107 4.19 AOA - Pitot Probes ...... 108 4.20 AOA - Pitot Mount Bracket ...... 108 4.21 J-3 8-Gal. LH Wing Tank ...... 109 4.22 J-3 8-Gal. LH Wing Tank ...... 109 4.23 Westach Dual Level Fuel Level Gauge Model: 2DA4 ...... 110 4.24 Aviasport ROTAX 503-582 Ducati Ignition Tachometer ...... 111 4.25 IM-510 Thermometer Coolant for ROTAX 912/914 ...... 112 4.26 IM-554 EGT Gauge for Rotax 914 ...... 112 4.27 IM-560 Oil Temperature Indicator for Rotax 912S ...... 113 4.28 IM-584 Fuel Pressure Gauge for Rotax 912 BAR ...... 113 4.29 IM-543 Oil Pressure Gauge for Rotax 912/914 BAR ...... 114 4.30 IM-561 Voltimeter ...... 115 4.31 Ultra Galactica™ Series lights ...... 117 4.32 Disposal of Ultra Galactica™ Series lights ...... 117 4.33 Aveo Hercules 30 Landing/Taxi/Wigwag Light Module ...... 118 4.34 Aerovoltz 16 cell battery ...... 119 4.35 H3R Aviation Model A344T - Halon 1211 Fire Extinguisher ...... 122 4.36 Sequence of BRS Ballistic Parachute System for LSA ...... 123 4.37 BRS Ballistic Parachute System for LSA ...... 123 4.38 BRS 7 LSA Canister Parachute System ...... 123 4.39 Simpson 5-point Harness System ...... 125 4.40 Poliester straps to restrain the baggage ...... 125 4.41 7x7 Strand Stainless Steel by Amstrong Cables...... 128 4.42 Stick Control Grip in ...... 128

5.1 Exterior model ...... 129 5.2 Full structure ...... 130 5.3 Power plant model ...... 130 5.4 Instruments displayed, Cockpit ...... 131 5.5 Passenger model ...... 131 5.6 Seats and prices of Air-Tech Inc.[34] ...... 132 5.7 Measures, technical sheet of the Ultralight seat [34] ...... 132 5.8 Initial sizing ...... 133

6.1 Specifications of Office 1 ...... 136 6.2 Specifications of Office 2 ...... 136 6.3 Specifications of Office 3 ...... 136 6.4 Specifications of Office 4 ...... 137 6.5 Specific weigh t by category ...... 138 6.6 Scheme of a SWOT analysis...... 141 6.7 Airplane Shipments Worldwide (1995–2017) ...... 142

RA 9 G06-AlOn LSA 3 seats | Project report 6.8 Different logo proposals ...... 145 6.9 Net Present Value (years = 7, k = 11,4%) for different selling prices ...... 147 6.10 Pay-Back Time for different selling prices ...... 147 6.11 Internal Rate of Return for different selling prices ...... 147

RA 10 List of Tables

1.1 Wing parameters of the most similar airplanes and gliders ...... 19 1.2 AlphaOne initial main parameters configuration ...... 19 1.3 Analysis configuration ...... 19 1.4 Initial wing definition ...... 21 1.5 Initial wing definition ...... 21 1.6 Final wing definition ...... 25 1.7 Wing Unweighted Average Method ...... 26 1.8 Wing Ordered Weigthed Average method ...... 26 1.9 Simple Hierarchy Method application ...... 30 1.10 Unweighted Average Method application ...... 30 1.11 Ordered Weighted Average Method application ...... 31 1.12 Unweighted Average Method application(fuselage) ...... 36 1.13 Ordered Weighted Average Method application(fuselage) ...... 36 1.14 Final wing configuration data ...... 43 1.15 Final fin configuration data ...... 43 1.16 Final elevator configuration data ...... 43 1.17 Control surfaces main sizes ...... 43

2.1 Global properties of aluminium ...... 52 2.2 Properties of components of CentrAl ...... 52 2.3 Properties of CentrAl ...... 53 2.4 Properties of T300 ...... 54 2.5 Properties of T400H ...... 54 2.6 Properties of T1000G ...... 54 2.7 Results of the static analysis of the different alternatives of the landing gear. . . 60 2.8 Results of the application of OWA method ...... 60 2.9 Properties of the steel 4130 normalized at 870 ºC...... 61 2.10 Load distribution for n=1 ...... 65 2.11 Lift distribution for each load factor ...... 65 2.12 Beam configurations and results ...... 68 2.13 Beam configurations and results ...... 68 2.14 Results of buckling analysis ...... 71 2.15 Principal results of the different fuselage alternatives analysis ...... 76 2.16 Results of OWA method ...... 76 2.17 Total load and distribution ...... 80 2.18 Dimensions for elevator beam. Referred to Fig. 2.21 ...... 80 2.19 Total load and distribution ...... 81 2.20 Dimensions for fin beam. Referred to Fig. 2.21 ...... 82 2.21 Buckling analysis for fin ...... 83 2.22 Buckling analysis for elevator ...... 83

3.1 Rotax 912 Vs AE50R comparison ...... 86 3.2 OWA importance table ...... 87

RA 11 G06-AlOn LSA 3 seats | Project report 3.3 OWA results ...... 87 3.4 Relative weight table ...... 87 3.5 Graded criteria table ...... 87 3.6 Valuation matrix table ...... 88 3.7 Domination matrix table ...... 88 3.8 PRESS results ...... 88 3.9 OWA and PRESS analysis results ...... 88 3.10 Motor characteristics ...... 89 3.11 Motor frame characteristics ...... 90 3.12 Motor frame characteristics ...... 91 3.13 Custom propeller characteristics ...... 92 3.14 HAUTECLAIRE propeller characteristics ...... 92 3.15 Custom propeller cruising specs ...... 95

4.1 Specifications of WINTER 7 FMS 5 Aispeed Indicator ...... 98 4.2 Specifications of WINTER EBH Altimeter ...... 99 4.3 Specifications of WINTER Vanetype Variometer 5 STV 5 ...... 99 4.4 Specifications of Kelly MFG RCA22 Artificial Horizon ...... 100 4.5 Specifications of Kelly RCA15BK-1 Directiornal Gyro ...... 100 4.6 Specifications of WINTER Turn Coordinator QM II ...... 101 4.7 Specifications of WINTER Hours Counting Analog ...... 101 4.8 Specifications of SV-D600/B 7" SkyView SE Display ...... 102 4.9 Specifications of SV-ADAHRS-200 Air Data ...... 103 4.10 Specifications of SV-BAT-320 SkyView Backup Battery ...... 104 4.11 Specifications of SV-GPS-2020 GPS Receiver/Antenna ...... 104 4.12 Specifications of SV-XPNDR-261 Mode-S Transponder ...... 105 4.13 Specifications of SV-KNOB-PANEL/H SkyView Knob Control Panel ...... 106 4.14 Specifications of SV-COM-C25/H SkyView VHF Com Radio 25 kHz ...... 106 4.15 Specifications of SV-INTERCOM-2S Two-Place Stereo Intercom ...... 107 4.16 Specifications of AOA - Pitot Probes ...... 108 4.17 Specifications of J-3 8-Gal. LH Wing Tank, FAA/PMA’d and Compulsory Extra Elements ...... 110 4.18 Specifications of Westach Dual Level Fuel Level Gauge Model: 2DA4 ...... 110 4.19 Specifications of Aviasport ROTAX 503-582 Ducati Ignition Tachometer . . . . . 111 4.20 Temperature limits for the engine IAE50R-AA ...... 111 4.21 IM-510 Thermometer Coolant for ROTAX 912/914 ...... 112 4.22 Specifications of IM-554 EGT Gauge for ROTAX 914 ...... 112 4.23 Specifications of IM-560 Oil Temperature Indicator for Rotax 912S ...... 113 4.24 Pressure limits for the engine IAE50R-AA ...... 113 4.25 Specifications of IM-584 Fuel Pressure Gauge for Rotax 912 BAR ...... 114 4.26 Specifications of IM-543 Oil Pressure Gauge for Rotax 912/914 BAR ...... 114 4.27 Specifications of IM-561 Voltimeter ...... 115 4.28 Total of Flight-Navigation and Powerplant Instruments ...... 116 4.29 Specifications of Aveo Ultra Galactica™ Series lights ...... 117 4.30 Specifications of Aveo Hercules 30 Landing/Taxi/Wigwag Light Module . . . . 118 4.31 Specifications of Aerovoltz 16 cell battery ...... 119 4.32 Table 1 for choosing battery using OWA method ...... 120 4.33 Table 2 for choosing battery using OWA method ...... 120 4.34 Table 1 for choosing battery using PRESS method ...... 121 4.35 Table 2 for choosing battery using PRESS method ...... 121 4.36 Specifications of H3R Aviation Model A344T - Halon 1211 Fire Extinguisher . . 122 4.37 Specifications of BRS 7 LSA Canister Parachute System ...... 123 4.38 Total costs of Safety and Miscellanous Components and Controls ...... 126

RA 12 G06-AlOn LSA 3 seats | Project report 6.1 Comparative of the 4 possibilities ...... 137 6.2 Basic human resources cost ...... 137 6.3 Marketing costs by category ...... 139 6.4 Evolution of Airplane Shipments worldwide between 1995 and 2017 ...... 143

RA 13 G06-AlOn LSA 3 seats | Project report

RA 14 Chapter 1

Aerodynamics

Due to the complexity that suppose the design of a three seat aircraft which can only weight up to 600Kg, the aerodynamic design is going to be a decisive part on this project. The main idea about the design of this aircraft is to create a hybrid between a traditional LSA (Light- Sport Aircraft) and a glider. The aim is to combine the efﬁciency of a glider and the main characteristics of the LSA in order to develop an aircraft light enough to be able to transport 3 passengers, including the pilot, without exceeding 600 Kg and having an acceptable performance. Taking into consideration the previous aspects, the aerodynamics study will be divided in the following sections:

1. Wing

2. Tail

3. Fuselage

4. Combined study

1.1 Wing

1.1.1 Airfoil selection The airfoil selection began searching for information about airplanes and gliders that meet the requirements of our conﬁguration or performance characteristics. The possible airfoils that could ﬁt in our model are:

• Davis B-24

• NACA 23015 [1]

• NACA 4412 [2]

• GA-30415 [3]

• FXS0-2196 [2]

The airfoils mentioned above were the ones that could give the best performance to our airplane. In order to validate our decision, an airfoil analysis with XFLR5 has been done for a range of Reynolds numbers and α.

It could not be found the distribution of the GA-30415 airfoil, which in particular was a characteristic one because it was a modiﬁcation of a 4415 and had a very good performance in terms of efﬁciency and momentum. The airfoil was found in the book [3] but there was no clue of it neither on the internet nor in some well-known airfoil databases.

RA 15 G06-AlOn LSA 3 seats | Project report 1.1.1.1 Polar distribution

Figure 1.1: Polar distribution of the selected airfoils

As shown in Fig. 1.1 the airfoils can be separated in two groups: 1. NACA 23015 and 4415 2. Davis B-24 and FXS0-2196

The ﬁrst group has wider range of Cl, the NACAs almost get a Cl = 1, 75 for a Cd = 0, 02; while the second group gets its maximum Cl practically when Cd = 0, both values are really close. As can be seen, the most signiﬁcant differences are the way the curves evolve with the Cd and the maximum values.

1.1.1.2 Efﬁciency versus α

Figure 1.2: Cl versus α of the selected airfoils Cd

This graphic is a decisive one because it defines one of the most significant parameters, the efficiency. As can be seen in Fig. 1.2 there is quite a big efficiency difference amongst the four,

RA 16 G06-AlOn LSA 3 seats | Project report so as a ﬁrst conclusion the best airfoil is thought to be the FXS0.

However, there is also another parameter that should be taken into account, the momentum coefﬁcient, which could cause further problems trying to make up for the wing momentum with the horizontal tail.

So not necessarily the best efﬁciency airfoil will provide the best performance. On the next section, a balance will be made in order to decide the optimum airfoil.

1.1.1.3 Momentum versus α Another important parameter that could impact on the efﬁciency of our aircraft is the momentum generated by the wing. In order to design a stable wing it is important to select an airfoil that can provide us an easy way to correct the momentum generated by the wing, which means a low momentum, in terms of the absolute value.

Figure 1.3: Momentum coefﬁcient of the airfoils selected

As seen in Fig. 1.3, the most efficient airfoil analyzed (FXS0) has a huge Cm at α = 5º, which is precisely the angle of attack where it’s efficiency is maximum. A Cm = −0.1 would difficult the tail design, trying to compensate the huge momentum the wing could cause. In order to simplify the tail design and the position of the tail and wing within the fuselage, a wing efficiency analysis has been made using the David B-24 and the NACA 23015. These two airfoils provide the balance between Cm/E desired.

1.1.2 General Wing Plant forms The wing plant form is a very important aspect in design. It affects the velocity, efficiency, bending moment, fuel tanks, cost... and so on. There are different types of wing plant forms and each one is used in a specific application.They can be classified in 4 major groups used in airplanes:

• Rectangular

• Trapezoidal

• Elliptic

• Delta

RA 17 G06-AlOn LSA 3 seats | Project report 1.1.2.1 Rectangular

Is the simplest wing plant form and consequently the less expensive to build. However, it provides an undesired aerodynamic and structural behaviour. Nevertheless, is widely used in small general aviation because the aerodynamic demands are not very strict and the cost of fabrication is very low in comparison to wings with other plant forms.

1.1.2.2 Elliptical

It is the wing plant form with less induced drag. It has a constant Cl distribution which, with an elliptic chord distribution, results in elliptical lift distribution. This behaviour is suitable for aerodynamics and structures. However, it is not very used today because of its price, which is the highest, and also because it has also an undesired stall behaviour due to its constant Cl distribution, which means that at a certain angle of attack the whole wing stalls at once. To solve this problem swept was implemented in WWII emblematic planes, such as the Spitﬁre.

1.1.2.3 Trapezoidal

Is the most used in heavy transport planes such as commercial planes. A good trapezoidal design can achieve almost the same behaviour as the elliptic wing but in a shorter range of angles of attack, where the cruise angle of attack should be included. In addition, they are easier to produce and, consequently, less expensive than elliptic wings. However, the stall behaviour is not suitable, due to the fact that for very narrow trapezoidal wings the stall starts at the tip where the ailerons are often located. In commercial planes trapezoidal wings are designed with swept to increase the cruise speed and with some torsion to counter the stall non-desired behaviour.

1.1.2.4 Delta

Delta Wing plant form is the trapezoidal concept brought to the limit where the tip chord is 0. It is used in ﬁghter jets in addition to swept because of its ability to increase the airplane speed in exchange for efﬁciency.

1.1.3 Plant form deﬁnition

The deﬁnition of the main parameters of the wing plant form is based on a combination of the data obtained from other similar aircrafts and a wing efﬁciency analysis using the airfoils mentioned in the previous section.

The next table contains the wingspan and area of the airplanes analyzed in order to do an initial sizing of the wing. As can be seen, the wing area is similar for all aircrafts; while the wingspan is divided in two groups:

• LSA < 10 m

• Gliders > 10 m

These values make sense considering the nature of each aircraft. As mentioned above, the plane designed in this project is neither an LSA nor a glider, but it is a combination of both. Due to this reason, the initial wing conﬁguration is the following:

RA 18 G06-AlOn LSA 3 seats | Project report Wing span (m) Wing area (m2) Atec 322 Faeta 9,6 10,1 TL-2000 Sting S4 9,12 11,1 TL-3000 Sirius 9,4 11,26 Pipistrel Taurus M 14,97 12,33 Alexander Schleicher ASG 29 E 18 10,5

Table 1.1: Wing parameters of the most similar airplanes and gliders

Wing span (m) Wing area (m2) AlphaOne 14 11

Table 1.2: AlphaOne initial main parameters conﬁguration

1.1.3.1 Initial efficiency analysis In order to validate the airfoils and the wing parameters selected, an efficiency analysis has been done. The objectives of this first analysis are:

1. Compare and deﬁne the best airfoil

2. Validate the wing parameters selected

3. Reﬁnement and analysis of better possibilities.

1.1.3.2 Airfoil definition result In order to compare the two previous airfoils, an analysis with XFLR51 has been done to determine the maximum efficiency and the momentum coefficient Cm once the wing is set.

The analysis had the configuration shown in Table 1.3. As a result, shown in Fig. 1.4, the B24 has a higher maximum efficiency than the N23015, it is not a huge difference but could be a good improvement for the wing. But as regards performance of the airplane, it is also important to consider the flight mechanics and the different flight configurations it will have.

α -5º < α < 20º Weight 600 kg Gravity center (0,0) Polar type Fixed Lift Method Horshoe vortex (viscous corrections) kg ρ 1,225 m3 −5 m2 v 1, 5 ∗ 10 s Table 1.3: Analysis conﬁguration

Taking into account the ﬂight mechanics and considering that both top efﬁciency values are quite high, it would be a better choice to have a wider operative range of angle of attack α

1is an analysis tool for airfoils, wings and planes operating at low Reynolds Numbers.

RA 19 G06-AlOn LSA 3 seats | Project report than a smaller one with a bigger maximum efﬁciency.

A wider angle of attack will give the airplane a better manoeuvrability and will make possible bigger α0s for those ﬂight conﬁgurations that could require them.

Figure 1.4: Efﬁciency curve for the N2305 and the B24 airfoils

Thus, the NACA 23015 is the airfoil chosen to deﬁne the aircraft’s wing. This airfoil can be seen in Fig. 1.5.

Figure 1.5: NACA 23015

1.1.3.3 Validation of the initial wing parameters configuration In order to validate and check if the values we estimated are good enough, an analysis for the parameters defined in Table 1.2 has been developed using the previous analysis configuration.

The first objective was to obtain the surface needed to make a horizontal flight with the maximum efficiency. It was also important to consider that the stall speed in landing configuration should be 22,5m/s or less. First, a smaller surface, in comparison with similar planes, has been studied. The first configuration was a trapezoidal plant form with 7,5 m2:

RA 20 G06-AlOn LSA 3 seats | Project report

Wing span (m) 15 Root chord (m) 0,6 Tip chord (m) 0,4 Surface (m2) 7,5 Airfoil 23015 Aspect ratio 30 Maximum efﬁciency 51,1 Stall speed (m/s) 34,0

Figure 1.6: Trapezoidal Wing, S=7,5 m2 Table 1.4: Initial wing deﬁnition

This configuration is able to fly fast and has good efficiency but the stall speed is too high. In order to reduce the stall speed, an increase in the surface is needed. It was decided not to increase the wingspan in order to make the structure and the storing of the plane simpler. For this reason, the following configuration has a chord increase.

Wing span (m) 15 Root chord (m) 1 Tip chord (m) 0,6 Surface (m2) 12 Airfoil 23015 Aspect ratio 18,75 Maximum efﬁciency 44,7 Stall speed (m/s) 27,5

Table 1.5: Initial wing deﬁnition Figure 1.7: Trapezoidal Wing, S=12 m2

With this wider configuration, the stall speed is also higher than the maximum value desired. At this point, it is easy to see that a high-lift device is needed in order to achieve the desired stall speed without losing too much efficiency, because from the first to the second configuration a reduction of efficiency has already occurred, produced by the enlargement of the surface. In the next figure, both configurations are compared.

Figure 1.8: Efﬁciency comparison of trapezoidal wings with different surface

The ﬁnal wing will have these major parameters. From now on, some implementations will be done to improve the aerodynamic characteristics. Speciﬁcally, three main aspects need to be revised:

RA 21 G06-AlOn LSA 3 seats | Project report 1.1.3.4 Wing Efficiency The efficiency is a very important parameter in the design. An increment of this value allows producing the same lift with lower drag and, consequently, lower thrust. This makes possible the implementation of a smaller engine with smaller consumption, which involves using smaller fuel tanks. Overall, an increment of efficiency will end in a reduction of weight, which is a hard restriction in this project. In order to improve efficiency, two techniques have been used:

• Aerodynamic Torsion

• Wingtips

1.1.3.4.1 Aerodynamic Torsion The variation of the airfoil along the wingspan can also result in an improvement of efficiency. Since the tip of the wing will have smaller structural stress compared with the root, it can be thinner. This allows the airfoil at the tip to be a NACA 23012 with a linear evolution from the NACA 23015 of the root. In figure 1.9 is shown a comparison in efficiency between the trapezoidal wing and without aerodynamic torsion.

Figure 1.9: Efﬁciency comparison between wing with and without aerodynamic torsion

As can be seen, the configuration with different airfoils is more efficient, 45,7 in front of 44,7. After these results, the final configuration will implement aerodynamic torsion.

1.1.3.4.2 Wingtips The wingtips are able to reduce the induced drag by making a softer transition between the higher pressure Intrados and the lower pressure Extrados at the tip of the wing. Two types of wingtips have been studied:

• Winglets: These devices have a complex geometry because an almost 90 degrees soft turn is needed with strong narrowing. After several iterations the best results were obtained with the following geometry:

RA 22 G06-AlOn LSA 3 seats | Project report

Figure 1.10: Winglets

Because of its complex geometry, winglets are difﬁcult to study and expensive to build.

• Elliptic wingtips: These devices consist in an elliptic chord distribution at the tip of the wing. The geometry is much simpler than winglets’ as it can be seen:

Figure 1.11: Elliptic Tip

The following graph shows the efficiency of the wing with elliptic wingtips, winglets and without wingtips: As can be seen, the wingtips improve the efficiency of the wing. Winglets and Elliptic wingtips show similar results being the second one a little bit better. As a consequence of the lower cost of production and the performance of the elliptic wingtips, the final wing configuration will incorporate them.

1.1.3.5 Stall Behaviour When the plane reaches the stall speed it is not desired that the whole wing enters the stall at the same time; neither the tip, because the ailerons control would be lost; nor the root,

RA 23 G06-AlOn LSA 3 seats | Project report

Figure 1.12: Efficiency comparison between elliptic wingtips, winglets and without wingtips because the stall wake could affect the tail. To solve this problem, +3º of geometrical torsion has been given to the central part of the wing, at 3,75m from the root. This implementation makes that zone to have more angle of attack and, consequently, more Cl. That results to be the first section to stall. The next figure shows it, in blue the torsion configuration and in red the non-torsion one:

Figure 1.13: Cl comparison between torsion and non-torsion wing

An efficiency study has also been done to know the repercussion of this implementation. It can be seen that it has moved the maximum efficiency angle but the value did not change significantly: This technique allows having a soft stall behaviour without losing efficiency and control of the plane. For these reasons, the final configuration will incorporate geometrical torsion.

1.1.3.6 High-lift Device The stall speed in landing configuration should not exceed 22,5m/s (45 Knots) as the regulation specifies. In table Table ?? it can be seen that the stall speed is above the maximum. An increase in the surface would fix it, but then the plane would be less efficient. To solve this

RA 24 G06-AlOn LSA 3 seats | Project report

Figure 1.14: Efﬁciency comparison between torsion and non-torsion wing problem a high-lift device can be implemented.

The configuration will implement a trailing edge simple flap due to its simplicity. The design has a 0,25 chord flap with a maximum deflection of 20º. The span of the flap is 4m and it is located at the root. This modification makes the stall speed go just under 22,5m/s in landing configuration.

The ﬁnal wing conﬁguration has these properties:

Wing span (m) 16 Root chord (m) 1 Tip chord (m) 0,6 Surface (m2) 12,5 Root airfoil 23015 Tip airfoil 23012 Aspect ratio 20,6 Maximum efﬁciency 46,7 Stall speed (m/s) 22,3

Table 1.6: Final wing deﬁnition

1.1.4 Wing conﬁguration criteria Once the alternatives have been studied, the next step is to draw a comparison among them to choose the one that provides better performance and efﬁciency for lower weight.

The comparison is based in four main criteria:

• Efficiency: it measures ratio between the functionality provided by the wing and the increase of drag produced. It gives a general value of how beneficial it is. A higher grade in this criterion means a higher efficiency.

• Weight: it considers the complexity of the structure required in each conﬁguration and, consequently, the weight of the wing. A higher value in this criterion indicates a lighter structure.

• Cost: it is related with the difﬁculty to build each wing, which is also related with the complexity and the amount of material needed. A higher grade in this criterion involves a lower cost.

RA 25 G06-AlOn LSA 3 seats | Project report • Stall performance: it deﬁnes how good the stall performance is in each wing conﬁg- uration, when and how it starts and how it evolves. A higher grade in this criterion indicates a better performance. For the selection of the most adequate tail, different methods have been applied in order to see if the best alternative does not depend on the procedure of the method. Besides, the same criteria have been used in all methods and the grades are proportionally calculated.

1.1.4.1 Unweighted Average Method

Unweigthed Average Method CRITERIA Trap Ellip Rec Trap +Elip ﬁnal Trap + Wing Efﬁciency 3 3 2 5 5 Weight(struc) 4 3 4 3 2 Cost 4 2 5 4 3 Stall perf 5 2 2 5 5 SUM 16 10 13 17 15 Max 20 UA 2 0,5 0,65 0,85 0,75

Table 1.7: Wing Unweighted Average Method

1.1.4.2 Ordered Weighted Average Method

Ordered Weigthed Average Method Trapeziodal Elliptic Rectangular Trapeziodal + wingtip Trapeziodal + winglet CRITERIA P PxG P PxG P PxG P PxG P PxG Efﬁciency 3 12 3 12 2 8 5 20 5 20 Weight(struc) 4 12 3 9 4 12 3 9 2 6 Cost 4 4 2 2 5 5 4 4 3 3 Stall perf 5 10 2 4 2 4 5 10 5 10 SUM(PXG) 16 38 10 27 13 29 17 43 15 39 Max(PXG) 20 50 50 50 50 50 OWA 2 0,76 0,5 0,54 0,65 0,58 0,85 0,86 0,75 0,78

Table 1.8: Wing Ordered Weigthed Average method

As it can be seen in both tables the best option is the trapezoidal plus a wingtip at the end. Also it can be noticed that the simple trapezoidal is better than the trapezoidal with wingltes when using the OWA. So the ﬁnal wing conﬁguration is the trapezoidal plus the wingtips.

1.2 Tail

1.2.1 Definition of tail A tail or empennage is a structure located usually at the rear of an airplane which is in charge of providing aerodynamic stability during flight. It can also contain control surfaces used to deliberately change the yaw and pitching moment of the aircraft in order to adjust its trajectory. It is commonly composed of a horizontal stabilizer (or tailplane), a vertical stabilizer (or vertical fin) and two control surfaces: an elevator and a rudder.

1.2.2 Utility of a tail In conventional aircraft, tails are generally designed to fulﬁll a set of concrete roles. These roles are:

RA 26 G06-AlOn LSA 3 seats | Project report • Trim (longitudinal and directional)

• Stability (longitudinal and directional)

• Control (longitudinal and directional)

The ﬁrst function, trimming, consists in maintaining balance, both longitudinal and directional, during unsymmetrical ﬂight conditions. Its major purpose is to eliminate the need for the pilot to keep constantly focusing on the control surfaces in order to maintain the equilibrium personally. The trimming function is usually done by trim tab elements or by adjustable stabilizers.

The second function of a tail is to stabilize the plane throughout the flight. Airplane stability is defined as a tendency to recover from undesired transitory perturbations such as, primarily, atmosphere phenomena. The tailplane is in charge of longitudinal stability while the vertical fin is responsible to preserve the directional stability.

The third and most important responsibility of tails is the in-ﬂight control of the aircraft, so as to the pilot can consciously modify its trajectory. The elevator is used to change the pitch moment of the plane and therefore adjust its longitudinal behavior. The rudder is used to change the yaw moment, thus leading to a directional adjustment of its behavior. Combining the rudder and the ailerons (usually located somewhere in the wing surface), the pilot is able to perform a coordinated turn of the plane.

1.2.3 General tail designs The ﬁrst step to design the tail of our airplane is to decide the conﬁguration to be used. In this case, it is convenient to compare different alternatives in a theoretical way evaluating the advantages and drawbacks for each option.

The most common tail designs used in civil aviation are the following (see ﬁgure 1.15):

Conventional tail T tail Cruciform tail Dual tail Triple tail V tail Inverted V tail Inverted Y tail Twin tail Boom tail High boom tail Multiple-plane tail

The main function of the tail of an airplane is to provide both stability and control in pitch and yaw. Taking into account the fact that our airplane is a hybrid between an LSA (Light-Sport Airplane) and a glider, the main aspects to consider to select the optimum tail configuration are not only the stabilization and control provided but also the weight and complexity of the structure required. From all the alternatives mentioned above, it is easy to see that some of them will not provide the desired performance. Moreover, the most common tail configurations for both LSAs and gliders are only conventional tail and T tail; the other configurations are rarely used and have become obsolete. In the following, every possible tail configuration will be analyzed.

1.2.3.1 Conventional tail design The conventional tail configuration is an aft tail design (rear-mounted tail) which is considered as the simplest tail configuration. The horizontal stabilizer does not produce any load on the vertical stabilizer, consequently, weight and complexity can be reduced. Its overall performance (trim, stability and control) is presumably acceptable. It is also one of the lightest configurations. That’s why it is the most common configuration, especially in LSAs. The weak point of this design is the fact that the horizontal stabilizer is located in the wing wake, causing a loss of efficiency.

RA 27 G06-AlOn LSA 3 seats | Project report

Figure 1.15: Tail conﬁgurations

1.2.3.2 T-tail design

Another usual tail configuration in General Aviation airplanes is the T-tail, which is an aft tail configuration with the horizontal stabilizer located on top of the vertical fin. Its main advantage over the conventional design is that during cruise flight the tailplane remains above of the regions disturbed by the wing wake, downwash, wing vortex and engine prop wash; thus, it provides a higher aerodynamic efficiency. The lower effects from the wings and engine also lead to a diminution in tail vibrations, greatly reducing its fatigue and increasing its durability. Nonetheless, the bending moment created by the elevator is also higher, so the vertical tail structure becomes heavier. Besides, its main disadvantage appears when flying at a high angle of attack, when the turbulent flow separated from the wings might incise upon the elevator and might result in a complete loss of the aircraft’s longitudinal control, situation known as deep stall.

1.2.3.3 Cruciform tail design

This tail design is a hybrid variation of the two previous designs. The horizontal empennage is located higher than in the conventional tail, so that it is away from the wing wake and the propeller flow, but it is not as high as in T tail configuration. With this design, the lower part of the stabilizer and the rudder receive undisturbed airflow due to the lifting force of the horizontal stabilizer. It is important to have undisturbed airflow on the rudder, especially to recover from spins. Although this configuration does not improve significantly the main strengths of its predecessors; it significantly reduces its major drawbacks, especially the deep stall distress.

RA 28 G06-AlOn LSA 3 seats | Project report 1.2.3.4 V-tail design In V-tail configuration instead of three surfaces (horizontal and vertical empennages with different variants), there are only two, which intend to serve the same function. The purpose of this tail design is to reduce the total tail area, as both parts of the tail act as horizontal and vertical stabilizers. Despite the fact that this configuration is quite suitable for trimming the aircraft, it is highly inefficient at maintaining its stability, letting the airplane really susceptible to perturbations. It is mainly employed in unmanned reconnaissance aircraft.

1.2.3.5 Twin-tail design This tail conﬁguration is commonly used when a big tail is required but it would be too heavy or unfeasible to make a single regular tail. The big horizontal tail improves hugely the longitudinal performance of the airplane. In addition, all control surfaces remain clear of the fuselage wake region, thus increasing overall controllability, at a cost of a signiﬁcant weight.

1.2.3.6 Boom-tail design The boom-mounted tail conﬁguration is often used when the aircraft has a prop-driven engine installed at the rear of the aircraft, so as to the interference between the prop wash and the tail is reduced.

1.2.3.7 Dual tail design The placement of two vertical stabilizers at the ends of the horizontal one provide a better directional control in low-speed operations and allows a smaller and more aerodynamically efﬁcient horizontal stabilizer. Although the size and consequently the weight of the horizontal empennage are reduced, the total weight of the tail is higher due to the addition of two vertical stabilizers instead of one as in other conﬁgurations.

1.2.3.8 Triple tail This conﬁguration includes has two vertical stabilizers at the ends of the horizontal one and one more on the fuselage. This fact allows reducing the height of the tail but increases the total weight.

1.2.3.9 Other tail configurations As mentioned above, there are many different tail configurations, but the most common are the ones previously analyzed. Configurations such as inverted V tail, inverted Y tail or multiple-plane tail are rarely used and have been dismissed from the study because they are heavier and less efficient than the most common ones.

1.2.4 Tail conﬁguration selection Once the alternatives have been studied, the next step is to draw a comparison among them to choose the one that provides better performance and efﬁciency for lower weight. The comparison is based on four main criteria:

• Efficiency: it measures the ratio between the functionality provided by the tail and the increase of drag produced. It gives a general value of how beneficial it is. A higher grade in this criterion means a higher efficiency.

• Weight: it considers the complexity of the structure required in each conﬁguration and, consequently, the weight of the tail. A higher value in this criterion indicates a lighter structure.

RA 29 G06-AlOn LSA 3 seats | Project report • Cost: it is related to the difﬁculty to build each tail, which is also related to the complexity and the amount of material needed. A higher grade in this criterion involves a lower cost.

• Stability and control: it deﬁnes the capability of the tail to develop its main function, and it is related to the performance of the plane. It also considers the facility of the tail to be immersed in the turbulent wake coming from other parts of the plane, which the control surfaces less effect. A higher grade in this criterion indicates a better performance.

For the selection of the most adequate tail, different methods have been applied in order to see if the best alternative does not depend on the procedure of the method. Besides, the same criteria have been used in all methods and the grades are proportionally calculated.

1.2.4.1 Simple Hierarchy Method The scale used to evaluate each criterion goes from 1 to 10, and the threshold to reject inade- quate alternatives is 6. Then, the grades and results can be seen in the following table:

ALTERNATIVES CRITERIA Conventional Tail T-Tail Cruciform Tail Efﬁciency 4 10 8 Weight 10 6 8 Cost 8 6 4 Stability and control 6 8 6

Table 1.9: Simple Hierarchy Method application

The threshold set in 6 means that the alternatives that have one or more grades under this number are rejected. Analyzing the efﬁciency, the conventional tail is discarded; considering the cost of manufacturing, the cruciform is also discarded. In this way, the best alternative is the T tail.

1.2.4.2 Unweighted Average Method The new scale applied goes from 1 to 5, but the grades are proportionally calculated from the previous method in order to follow the same procedure for the different methods. The results obtained are shown in the next table:

ALTERNATIVES CRITERIA Conventional Tail T-Tail Cruciform Tail Efﬁciency 2 5 4 Weight 5 3 4 Cost 4 3 2 Stability and control 3 4 3 SUM 14 15 13 UA 0,7 0,75 0,65

Table 1.10: Unweighted Average Method application

This method considers the sum of all grades for each conﬁguration to calculate an unweighted average; this means that the four criteria are considered equally important. This procedure indicates that the best alternative is the T tail.

RA 30 G06-AlOn LSA 3 seats | Project report 1.2.4.3 Ordered Weighted Average Method It is applied the same scale as in the previous method, from 1 to 5. The main difference is the consideration of a speciﬁc weight for each criterion, which indicates the importance or weight that will have in the ﬁnal result. The results obtained can be seen in the following table:

ALTERNATIVES Conventional Tail T-Tail Cruciform Tail WEIGHT CRITERIA P PxG P PxG P PxG G Efﬁciency 2 8 5 20 4 16 4 Weight 5 15 3 9 4 12 3 Cost 4 8 3 6 2 4 2 Stability and control 3 3 4 4 3 3 1 SUM (PxG) 34 39 35 OWA 0,68 0,78 0,7

Table 1.11: Ordered Weighted Average Method application

This method is possibly the most accurate among the ones used, due to the fact that each criterion is weighted in order to consider its importance. According to this method, the best alternative is again the T tail.

1.2.5 Decision making In this case, the best alternative is the same for all methods, so undoubtedly the ﬁnal decision is the T tail conﬁguration. The next step is to determine the main parameters of the tail in order to provide the necessary stability and control with the minimum increment of drag.

1.2.6 Tail design There are two main aspects to take into account from the definition and sizing of the tail. Firstly, it is important to remember its main functions: trim, stability and control. The movable surfaces of the tail, typically the rudder and the elevator, provide control; and the horizontal and vertical tails, stability, which is what is going to be studied first. The desired tendency of the plane is to automatically recover the initial state after a transitory perturbation, which is, in fact, the definition of a stable aircraft. However, the more stable a plane is, the less manoeuvrable. Due to this correlation, to correctly define the size of the tail, a midpoint should be found in order to design a stable plane but also controllable and easy to handle.

The second aspect to consider, especially in T-tail conﬁguration, is the location of the horizontal stabilizer. As explained in a previous section, there is a risk of deep stall, which consists in the loss of control of the tail due to its immersion in the wing wake when the plane is ﬂying at high angles of attack, causing the complete loss of the longitudinal control. On the one hand, to avoid this critical situation, the horizontal tail should be high enough to ensure it is not affected by the turbulent wake at high angles of attack. But on the other hand, an increase in the distance between the horizontal tail and the longitudinal plane axis involves a higher bending moment and, consequently, a heavier vertical tail structure is needed. Once again, the optimum design point should be found to avoid both drawbacks.

1.2.6.1 Parameters determination The procedure to determine the main parameters of the tail begins with a preliminary sizing having as a reference similar airplanes, which in this case are both LSAs and gliders. The

RA 31 G06-AlOn LSA 3 seats | Project report main reference has been the glider Pipistrel Taurus, due to its similar weight, fuselage length and wing.

Considering the reference data, an average or an approximated value can be used initially to design the tail and study the plane’s behaviour with XFLR5. Depending on the results, some parameters should be reﬁned to achieve the optimum performance.

Moreover, a generic body has been created in order to study the relative position between the wing and the tail, which is a decisive parameter because it inﬂuences the longitudinal stability of the aircraft.

1.2.6.1.1 Horizontal tail

• Aspect ratio The main influences of the aspect ratio are the slope of the lift curve (CLα), the induced drag coefficient (CDi) and the structural weight. Considering the fact that the lift and drag produced by the horizontal tail are significantly smaller than those produced by the wing, it is more important to focus on the reduction of weight rather than on the aerodynamic forces to determine the aspect ratio.

When the aspect ratio decreases, the structural weight does it too. Consequently, the aspect ratio of the tail should be lower than that of the wing. As for the wing the value is about 20, for the tail this parameter is cut approximately by half, ending in a value about 10.

• Taper ratio The taper ratio affects the induced drag, the structural weight and the facility to build the wing and, consequently, the cost. A low taper ratio (λ = 0.3 - 0.5) involves low induced drag and low structural weight, but it is easier to build an untapered wing (λ = 1).

The equilibrium point between a good performance (low induced drag and weight) and an affordable cost is a value of 0.5.

• Sweep angle The sweep angle apparently has no direct advantages and, in fact, it has an adverse effect, which results in a reduction of lift (lower CLα and a lower CLmax), an increment of induced drag (higher CDi) and also a higher structural weight. Swept wings are used to increase the drag divergence Mach number, but in this case, it is not necessary, because our plane ﬂights in low Mach numbers. Hence, the sweep angle of the horizontal tail is only about 5º.

• Airfoil selection

Both horizontal and vertical tails have a symmetric airfoil, because the movable surfaces can deﬂect on both sides of the undeﬂected position.

For low Mach number ﬂights, one of the most common airfoils for the horizontal tail is the NACA 0008, which provides the necessary aerodynamic forces.

RA 32 G06-AlOn LSA 3 seats | Project report • Incidence angle The incidence angle of the horizontal tail is the angle between the fuselage reference axis and the reference chord of the horizontal tail. The tail incidence is such that during cruise ﬂight the lift required from the tail to make the pitching moment zero is produced without elevator deﬂection.

To determine the incidence angle is necessary to know the angle of attack of the plane in cruise ﬂight and the downwash of the tail. This last value is difﬁcult to measure properly. Hence, the tail is built with zero incidence angle and the performance of the airplane will be controlled with the corresponding control surface.

1.2.6.1.2 Vertical tail • Aspect ratio

The inﬂuence of the aspect ratio in vertical tails is the same as in horizontal tails, so the points mentioned previously are applicable. Besides, there are additional inﬂuences: increment of the height of vertical tail and, consequently, the plane is higher; increment of moment of inertia about the longitudinal axis, which involves lower lateral control; and better directional control due to the increment of the moment arm.

However, the main aspect to consider in a T-tail is the risk of deep stall, so it is important to place the horizontal tail above the wing wake when ﬂying at high angles of attack. This is the criterion applied for our plane to determine the height of the vertical tail.

• Taper ratio

To determine the taper ratio of the vertical tail, the same criterion as in horizontal tail has been applied. As a result, the value selected is 0.67.

• Sweep angle

The adverse effects of sweep explained for the horizontal tail are also applicable for vertical tail. In addition, a swept vertical tail has a higher moment arm.

Considering these effects, a low sweep angle is commonly applied, in this case it is about 7º.

• Airfoil selection

As for the horizontal tail, a symmetric airfoil is commonly used in vertical tails. In this case, the NACA 0008 is selected again, due to its aerodynamic characteristics.

1.3 Fuselage

The main problem about the fuselage configuration is the unusual number of passangers and also the maximum take-off weight. A configuration of three people does not allow the typical LSA or glider two seat configuration. So the casuistry is the next:

RA 33 G06-AlOn LSA 3 seats | Project report • 3 seat conﬁguration

• Relative position of the wing and the tail

• Maximum efﬁciency

• Costs

• Weight and structural facilities

Analyzing the options, the two different dispositions were studied in order to deﬁne which was the best. Fig. 1.16 shows the 2 different conﬁgurations.

(a) One plus two (b) In line

Figure 1.16: Seats conﬁguration

Firstly an analysis has been done in order to see which conﬁguration has better behaviour. The main performance graphics are going the be analyzed next.

• Efficiency It is expected that the three seats in line configuration has a better aerodynamic performance because of it’s geometry. The results obtained can be seen in the following graphic, which show both efficiency curves. As said before, Fig. 1.17 shows both graphics are almost the same. There is not a big difference between the maximum values. So, in this case, the efficiency is not a decisive point.

• Momentum coefficient As can be seen in Fig. 1.18 the difference is not remarkable enough to decide which configuration is better. Although the one plus two model has the Cm = 0 best situated than the in line one, the maximum efficiency is more centered in the wider configuration. But in both cases the plane shows the same maneuverability, as both lines are parallel.

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Figure 1.17: Efﬁciency in both conﬁguration

Figure 1.18: Cm in both conﬁgurations

Another important aspect to consider is the union between the wing and the fuselage. Be- cause of its wider conﬁguration, the one plus two design offers a better union and allows a better ﬁt performance.

In order to evaluate all the variables and choose the one with the best performance, the Un- weighted Average Method (UA) and the Ordered weighted Average Method (OWA) have been applied.

1.3.1 Fuselage selection criteria For the application of this method, it is necessary to deﬁne the main variables to be evaluated for each conﬁguration. The comparison criteria will be based on:

• Efficiency: it measures ratio between the functionality provided by the fuselage and the increase of drag produced. It gives a general value of how beneficial it is. A higher grade in this criterion means a higher efficiency.

• Weight: it considers the complexity of the structure required in each conﬁguration and, consequently, the weight of the fuselage. A higher value in this criterion indicates a lighter structure.

• Cost: it is related with the difﬁculty to build each fuselage, which is also related with the complexity and the amount of material needed. A higher grade in this criterion involves a lower cost.

RA 35 G06-AlOn LSA 3 seats | Project report • Wing position: it deﬁnes how big is the area where the union between the wing and the fuselage is going to be make. A Higher value in this criterion means a higher area, thus a better union.

1.3.1.1 Unweigthed Average Method The scale applied goes from 1 to 5, where a score of 5 in any variable shows the best possible value and a score of 1 shows the worst. This method does not take into account the weight of the criteria evaluated in relation to the other conﬁguration. The obtained results are shown in Table 1.12:

Unweigthed Average Method Criteria Seats in line One plus two Efﬁciency 5 4 Weight 4 5 Cost 3 4 Wing position 4 5 SUM 16 18 Max 20 20 UA 0,8 0,9

Table 1.12: Unweighted Average Method application(fuselage)

As can be seen in Table 1.12, the best option in this case is the One plus two conﬁguration, which offers better wing position, lower weight and better structural performance, although the efﬁciency is worse.

1.3.1.2 Ordered Weighted Average Method It is applied the same scale as in the previous method, from 1 to 5. The main difference is the consideration of a speciﬁc weight for each criteria, which indicates the importance or weight that will have in the ﬁnal result. The results obtained can be seen in the following table:

Ordered Weighted Average Method Seats in line One plus two Weight Criteria P PxG P PXG G Efﬁciency 5 10 4 8 2 Weight 4 16 5 20 4 Cost 3 3 4 4 1 Wing position 4 12 5 15 3 SUM(PxG) 41 47 Max(PxG) 50 50 OWA 0,82 0,94

Table 1.13: Ordered Weighted Average Method application(fuselage)

As can be seen in Table 1.13, the best option is again the one plus two configuration. Hence, the final fuselage configuration will be the one plus two, which is shown in Fig. 1.16(a).

1.4 Control Surfaces

The control surfaces allow the pilot to govern the plane, this means that angles of yaw, pitch and roll can be controlled. The three control surfaces are:

• Elevator: its task is to control the pitch.

RA 36 G06-AlOn LSA 3 seats | Project report • Ailerons: their task is to control the roll, however, it induces a moment that also affects the yaw due to the difference of drag produced by the difference of lift that both wings have. This effect is called adverse yaw.

• Rudder: its task is to control the yaw, however, it induces a moment that also affects the roll because the rudder is higher than the center of mass, the rotation point. This effect is called adverse roll.

In order to size the different control surfaces a similarity study has been done. The plane chosen was the glider Taurus from Pipistrel because of its similar characteristics.

1.4.1 Elevator

From the similarity study it can bee seen that the elevator should be a third of the horizontal tail’s surface. This means that Alpha-One should have a 0.5m2 elevator. In order to make it easier to construct, the elevator will have the same span that the horizontal empennage, 4m; so the chord will be 0.125m to fulfill the two conditions. The hinge moment was calculated assuming that the elevator was a flap of 0,33 chord with 15º of deflection:

Figure 1.19: Elevator Airfoil

Assuming this, XFLR-5 is capable to compute the non-dimensional hinge moment, which is the following:

Figure 1.20: Hmom Elevator

RA 37 G06-AlOn LSA 3 seats | Project report

To compute the maximum hinge moment the vne should be used and a Hmom of 0.025:

2 Moment = 0.5 · ρ · vNE · Se · ce · Hmom (1.1)

Moment = 0.5 · 1.225 · 82.32 · 0, 5 · 0, 125 · 0.025 = 6.48Nm (1.2) In order to determine the influence of the elevator, the lift coefficient at its maximum deflection angle has been obtained for different values of angle of attack, which can be seen in Fig. 1.21. For further structural calculations, it is necessary to obtain the maximum lift coefficient provided by the elevator, which is approximately 1.2.

Figure 1.21: Cl versus α for the elevator at its maximum deﬂection angle

1.4.2 Ailerons In the Taurus from Pipistrel, each aileron has a surface of one tenth of the wing. In the designed plane this equals to 0.6m2. The span of the aileron will be 4m with 0,15m chord and will be located at the end of the wing.

Figure 1.22: Aileron Airfoil

To obtain the hinge moment the same method will be used. The next figure shows the aileron was a flap of 0.25 chord with 15º of deflection:

RA 38 G06-AlOn LSA 3 seats | Project report

Figure 1.23: Hmom Ailerons

To compute the maximum hinge moment for each aileron, the vne should be used and a Hmom of 0,010: 2 Moment = 0.5 · ρ · vNE · Sa · ca · Hmom (1.3)

Moment = 0.5 · 1.225 · 82.32 · 0.6 · 0.15 · 0.010 = 3.73Nm (1.4)

1.4.3 Rudder In the similar plane, the rudder has a surface of one half of the ﬁn. In the designed plane this equals to 0.4m2. The span of the rudder will be the whole ﬁn, 1.2m, and 0.4m chord.

Figure 1.24: Rudder Airfoil

In order to calculate the hinge moment, the same method will be used. Here we can see the airfoil with a ﬂap of 0.5 chord with 15º of deﬂection:

RA 39 G06-AlOn LSA 3 seats | Project report

Figure 1.25: Hmom Rudder

To compute the maximum hinge moment the vne should be used and a Hmom of 0,065:

2 Moment = 0.5 · ρ · vNE · Sr · cr · Hmom (1.5)

Moment = 0.5 · 1.225 · 82.32 · 0.4 · 0, 4 · 0.065 = 43.15Nm (1.6) In addition, to determine the influence of the rudder to the aerodynamic forces, the lift coefficient at a maximum angle of deflection as a function of the angle of attack has been obtained using XFLR5. The results can be shown in Fig. 1.26. From this graphic, the maximum lift coefficient can be obtained, which is approximately 1.2.

Figure 1.26: Cl versus α fortherudderatitsmaximumdeflectionangle

1.4.4 Parasite Drag In order to get a more realistic performance of the aircraft, the drag coefﬁcients of the landing gear have been estimated. For the estimation of the wheels CD it has been used the following method:

• Deﬁne the reference distances.

RA 40 G06-AlOn LSA 3 seats | Project report

• Get the CD value from a experimental graphics.

• Deﬁne the calculated CD on XFLR5 to improve the veracity of the simulations.

And for the tube that connects the wheels it has been used the next following system:

• Deﬁne the reference distances.

• Create an equivalent wing using a symmetric air foil.

• Make a simulation in order to get equivalent CD.

1.4.4.1 Wheels parasite drag The reference distance used for the wheel estimation have been: the following ones:

• Radius = 0.4 m

• Width = 0.1 m

CD are not as usual as it might seem, there are not to much forms analyzed and, a disk through the longitudinal face is one of the ﬁgures which have not been analyzed.

The ﬁgures that might be an approximation to the disk in volume or in some referent distance are:

• Cylinder

• Sphere

• Bullet

Because of its geometry, the bullet its the most realistic way to simulate the wheel, because it has an smooth curve at the beginning and, but there is not as much rounded as the sphere.

The following ﬁgure shows the CD distribution through the Re. The Re used has been the one for cruise ﬂight conditions, which is:

Re = 1.7M (1.7)

Thus, using the following ﬁgure, the CD for each wheel have been estimated.

Figure 1.27: CD experimental data ; Source: researchgate.net

l Using the bullet body and a d ratio of 1, to be conservative. The extra drag for each wheel is :

−1 CD = 0.85wheel (1.8)

RA 41 G06-AlOn LSA 3 seats | Project report 1.4.4.2 Tube parasite drag As it has already been said, the tube has been estimated as a wing, a rectangular one. It can be noticed there is a huge difference between a wing and a tube, but taking into consideration that the tube is nor a cylinder neither a wing, it has been decided to estimates as a rectangular wing using a huge symmetric airfoil.

The geometry used to deﬁne the wing tube is:

• lcenter = 2m

• lside = 0.5m (each diagonal longitude).

Using the listed geometry the extra CD obtained is:

CD0 = 0.05 (1.9)

1.4.4.3 Drag correction Applying the drag correction to the ﬁnal plane conﬁguration on XFLR5, have been obtained the following results:

Figure 1.28: Comparison between both behaviours

As it can be on the Fig. 1.28 the maximum efficiency decreases due to the extra parasite drag caused by the landing gear. It was expected result, even though the addition of the landing gear drag, the efficiency performance of the airplane is still really good, with a maximum efficiency of approximately 35.

1.5 Final plane conﬁguration

1.5.1 Design parameters Once everything has been correctly calculated, the main parameters have been summarized in the following tables:

RA 42 G06-AlOn LSA 3 seats | Project report Wing span (m) 16 Wing area (m2) 12,424 Wing load (kg/m2) 48,295 Croot (m) 1 Ctip (m) 0,6 Wing span (m) 4 Aspect ratio 20,606 Wing area (m2) 1,5 Taper Ratio 10 Wing span (m) 1,2 Croot 0,5 2 Efﬁciency 35.42 Wing area (m ) 0,8 Ctip 0,25 Tilt angle (º) 2 Croot 0,75 Aspect ratio 10,67 v(m/s) 43,6 Ctip 0,6 Taper ratio 2

Table 1.14: Final wing configuration Table 1.15: Final fin configu-Table 1.16: Final elevator con- data ration data figuration data

Fin Elevator Aileron Sw 0,8 1,5 6,212 Croot 0,375 0,5 1 S(control surface) 0,4 0,5 0,6212 C(control surface) 0,375 0,15 0.25

Table 1.17: Control surfaces main sizes

1.5.2 Range study Also in order to estimate the different possible ranges, a Weight-range diagram has been obtained using the Breguet’s equation.The code developed is attached in the annexes.

The next ﬂight conditions are considered in order to apply the Breguet equation in cruise ﬂight conditions: L = W (1.10) T = D (1.11) Both equation result into the next relationship:

C T = W D (1.12) CL The range can be calculated as:

Z x f Z t f R = dx = dt · v (1.13) xi ti It is also known the relationship between time an weight, which relates the fuel intake with the mass ﬂow:

dW c · dt = − (1.14) g Where: Tv c = c · P = c (1.15) j j η Also applying (1.10) it is obtained the next result: s 2W v = (1.16) CL · Sw · ρ

RA 43 G06-AlOn LSA 3 seats | Project report Now, applying (1.15),(1.14), (1.16) and (1.12) into (1.13) the next result it is obtained: s η 2 Z Wf dW R = − · E 1/2 (1.17) cjg CL · Sw · ρ Wi W Integrating the last equation it is obtain. s 2η 2 1/2 1/2 R = · E · (Wi − Wf ) (1.18) cjg CL · Sw · ρ

Where CL is the optimum one to cruise mode. Applying (1.17) the W-R it is obtained:

Figure 1.29: Weight-Range diagram

The code designed to obtain the Fig. 1.29 can be found in the annexes of the report.

RA 44 G06-AlOn LSA 3 seats | Project report 1.6 Final Plane Analysis

1.6.1 Efﬁciency

The efﬁciency of the ﬁnal plane is shown in the following graph:

Figure 1.30: Efficiency of the final configuration

As it can bee seen in Fig. 1.30 the maximum efﬁciency is 34,5 at 1,5º of angle of attack.

1.6.2 Static Stability

The static stability is the immediate response of the plane in front of a perturbation. A static plane is the one that in front of a sudden increase of angle of attack, a restorer moment appears that decreases the perturbation. This means that the variation of the pitching moment coefficient as function of alpha should have a negative slope. Furthermore, the alpha with zero pitching moment coefficient should be roughly the same that the angle of attack of maximum efficiency. This condition would make the plane to have the angle of attack with the best glide ratio without the pilot’s help.

The following figure shows the final plane pitching moment coefficient, Cm, as function of alpha:

Figure 1.31: Static stability of the ﬁnal conﬁguration

RA 45 G06-AlOn LSA 3 seats | Project report The slope of the pitching moment coefﬁcient is negative which will make the plane statically stable. The alpha with no Cm is 0,7º, the alpha of maximum efﬁciency is 1,5º as it is shown in Fig. 1.30. It’s important to notice that the position of the center of gravity has a huge impact in stability analysis. In order to achieve the behaviour described in Fig. 1.31 the center of gravity should be at 0,225m behind the leading edge of the wing at the root.

1.6.3 Dynamic Stability The dynamic stability consists in the behaviour of the plane in front of a perturbation that changes its ﬂight conditions. The desired behaviour is the one that tries to counter the effect of the perturbation and damp the amplitude of it.

In order to evaluate the dynamic stability of the plane a modal analysis has been done. If the Modes have a damped evolution in time, the plane is dynamically stable. The three more important modes are the followings:

• Longitudinal Mode 1

• Lateral Mode 1

• Lateral Mode 2

1.6.3.1 Longitudinal Mode 1 This mode is related to the pitch angle. The next ﬁgure shows the angular velocity related to this mode:

Figure 1.32: Longitudinal Mode 1

As it can bee seen, the angular velocity decreases in time so it’s a stable second order response to the ﬁrst longitudinal mode. It takes about 2s to dissipate the perturbation.

1.6.3.2 Lateral Mode 1 This mode is related to the roll angle. The next ﬁgure shows the angular velocity related to this mode:

RA 46 G06-AlOn LSA 3 seats | Project report

Figure 1.33: Lateral Mode 1

As it shows the ﬁgure, the angular velocity decreases rapidly in time so it’s a very stable response. This is a ﬁrst order behaviour with a time constant of τ = 0.1s. This means that in 0.4s the perturbation is almost gone.

1.6.3.3 Lateral Mode 2 This mode is related to the yaw angle. The next ﬁgure shows the angular velocity related to this mode:

Figure 1.34: Lateral Mode 2

The ﬁgure shows that is a stable mode but is the slowest one. It takes about 8s to dissipate the perturbation and is more oscillating than the other two studied modes.

1.7 Flight envelope

In order to define the flight envelope where the LSA is going to flight, an airspeed-gust envelope has been done at sea level. To develop the diagram, the rules and requirements set by the governmental authorities have been taken into account.

RA 47 G06-AlOn LSA 3 seats | Project report The [4] has been followed in order to make sure to meet all the requirements and to calculate the necessary velocities. Following the structural requirement established by the ASTM F225, the load factors needed are:

• n1 = 4 : Positive maneuvering load factor

• n2 = −2. Negative maneuvering load factor

• nF1 = 2 :Positive manoeuvring load factor ﬂaps extended

Firstly, the vs is needed in both cases, with and without ﬂaps. Those velocities have been determined by simulations using XFLR5, which already gives the stall velocity for a Cl. In this case the maximum cl is required. The following velocities will allow computing the other ones in order to be able to create the airspeed diagram:

vs = 45.40kts; (1.19)

vs0 = 43.13kts < 44kts −→ EASA requirement; (1.20) Now the needed velocities can be calculated.

Design manoeuvring speed va: √ vA = vs n1 = 90.8kts (1.21)

Flaps maximum operating speed vF:

Should be bigger than 1.4ss and 2vs0. Doing this operation the result is:

vF = 85kts; (1.22)

Design cruising speed vc:

Needs to be bigger than: r wMTOW vCmin = 4.77 = 103kts (1.23) Sw And smaller than: 1.4VH = 121.21kts (1.24)

So, approximately vC can be deﬁned as:

vC = 110kts (1.25)

And ﬁnallly 4.7 Design dive speed VD:

vD > 1.4vCmin = 145.34 ≈ 160kts because needs to be higher (1.26) Using the same methodology for the ﬂap and for the negative load factor the resultant airspeed diagram is the following:

RA 48 G06-AlOn LSA 3 seats | Project report

Figure 1.35: Airspeed envelope diagram

The next step is to calculate the gust lines for 7.5 and 15 m/s for both vC and vD.

In order to get the load factor increment due to an instant gust the ASTM F2245 has been used. Using the aircraft data and the graphics that can be found in [4], the increments of the load factor for the different cases can be determined.

Using the increments previously calculated from the ASTM F2245, the following diagram has been developed:

RA 49 G06-AlOn LSA 3 seats | Project report Gust-Airspeed envelope diagram Figure 1.36:

RA 50 Chapter 2

Structures

In this chapter, design and analysis of the different structures that conform the aircraft will be presented . Main consideration that will be taken into account is weight.Since LSAs are thought to fly with a maximum of 2 passengers, adding a third one forces the structure to be ligther than convetional LSAs. For this reason, the first that will be done is a materials study in order to find the most suitable options. Another important item in the design is the resistance of the parts, whose analysis will be done during the second part of this chapter, where the design and the analysis if the parts will be presented.

2.1 Materials

As previously introduced, the ﬁrst thing that needs to be done is an analysis of the materials. In order to do this, it will be presented a choice of materials for: the main structure (fuselage and wings), the skin, the landing gear and the transparent surfaces.

2.1.1 Fuselage and wings internal structure

For both ribs and beams, a light but strong material must be chosen. The most used materials for these applications are aluminium-based composites and, among all of the different possible composites, two of them have been compared: Regular aluminium alloy and Central Reinforced Aluminium (CentrAl).

There are a lot of great advantages when using Aluminium alloys: they are an exception- ally light material with a relatively high strength and with an acceptable electric and thermal conductivity. They are neither magnetic nor toxic while at the same time they reﬂect light (it means it allows to have a lower heat accumulation), they are water-resistant and they ease the manufacturing process due to their ductility and malleability.

They have really important environmental advantages as well. Current aluminium alloys are completely recyclable and as they reduce the aircraft weight, a reduction of fuel consumption is fulﬁlled due to the reduction of the needed thrust to propel itself.

The main properties of the chosen aluminium alloy [5] are presented in Table 2.1.

RA 51 G06-AlOn LSA 3 seats | Project report Property Value Density 2.698,4 kg/m3 Fusion point 933,47 K Ce 900 J/K·kg Electric conductivity 37,7.106 S/m Thermal conductivity 237 W/K.m Young modulus 66,6 Gpa Traction resistance 230-570 MPa Mechanical resistance 690 MPa Elastic limit 215-505 MPa Elongation 10-25

Table 2.1: Global properties of aluminium

The second aluminium-based material is called Central Reinforced Aluminium which is a composite consisting of several layers of aluminium and glass ﬁbers glued with epoxy-based adhesives. Between aluminium and the glass ﬁber composite are layers of a proprietary resin- rich material called “BondPreg” by the developers. The main developer and manufacturer of CentrAl is Alcoa and the structure of the material they distribute is presented in Fig. 2.1.

Figure 2.1: CentrAl conﬁguration

The different materials used to manufacture the CentrAl [6] have diverse properties and when they are glued together, some new values are achieved. This data is presented in Ta- ble 2.2 and Table 2.3.

Unidirectional Lamina Property Aluminum 2024-T3 Unidirectional BondPreg® (S2-Glass Fiber and FM94K Adhesive) UTS (L (material 441,26 MPa 2.213,22 MPa 1.103,16 MPa rolling direction)) UTS (LT (perpendicular to 434,37 MPa 22,06 MPa 21,99 MPa material rolling direction)) Y (L) 324,05 MPa 355,76 MPa 177,88 MPa Y (LT) 289,58 MPa 51,71 MPa 29,03 MPa E (L) 72,39 GPa 53 GPa 27,61 GPa E (LT) - 5,31 GPa 3,12 MPa G12 27,57 GPa 2 GPa 1,17 GPa nu 12 0,33 0,27 0,30 nu 21 - 0,027 0,034 alpha 12 - 0,00000161 F−1 0,00000337 F−1 alpha 21 - 0,0000224 F−1 0,00003904 F−1

Table 2.2: Properties of components of CentrAl

RA 52 G06-AlOn LSA 3 seats | Project report Property Value Density 2.591 kg/m3 Young Modulus 65,2 GPa Poisson 0,33 Elastic Limit 812,2 MPa Y (Yield Strength) 383,8 MPa Deformation failure 0,04761

Table 2.3: Properties of CentrAl

CentrAl provides nearly 25 percent more tensile strength than high-strength aluminum alloys, is extremely resistant to metal fatigue and is highly damage-tolerant. Taking all those properties into account it is possible to consider that CentrAl, without further analysis, seems a better material for this purpose. The main reasons are:

1. 25 percent more tensile strength than high-strength aluminum alloys

2. Extremely resistant to metal fatigue

3. Highly damage-tolerant

4. So light that a transport-aircraft wing made from a combination of CentrAl and aluminum – which is better than CentrAl at resisting the compression strains on surfaces such as upper wing-skins – would not only be much stronger than a wing completely made of aluminium, but also could be 20 percent lighter.

So the main conclusion would be that the best material to build the structures is the CentrAl, despite it would have to be combined with aluminium alloys in the structures that have to support a higher compression strain. Nevertheless, it must be mentioned that other possibilities must not be directly discarded because even though CentrAl presents excellent properties its density is not specially low and the main limitation for this aircraft is the weight. Due to this problem ﬁber and honeycomb composites will be also studied for the wings composition. These materials are also light [7] but enough resistant to some parts of the aircraft due to the combination.

2.1.2 Skin of the aircraft

In order to build a light and effective aircraft skin the materials considered will be different types of carbon fibers. There are a lot of different types of carbon fiber depending on its composition (fiber type, fiber amount, fiber orientation, etc). Some of them are specifically used in aerospace applications, but all of them share certain interesting properties such as high longitudinal tensile strengths, low density, low coefficient of thermal expansion and low thermal and electric conductivity. A properties comparison between different carbon fibers used in aerospace applications is shown in Table 2.4, Table 2.5 and Table 2.6.

RA 53 G06-AlOn LSA 3 seats | Project report T300 (2,7%Epoxy /93% Carbon) Property Value Density 1.760 kg/m3 Tensile Modulus 140 GPa Tensile Strength 1.820 MPa -0,41 alpha Coefﬁcient of thermal expansion ·10-6 ºC−1 Speciﬁc Heat 0,777 J/g ºC Thermal Conductivity 0,105 J/cm s ºC Electric Resistivity 1,7 · 10−3 Ω · cm

Table 2.4: Properties of T300

T400H (1,6%Epoxy /94% Carbon) Property Value Density 1.800 kg/m3 Tensile Modulus 145 GPa Tensile Strength 2.250 MPa -0,45 alpha Coefﬁcient of thermal expansion ·10-6 ºC−1 Speciﬁc Heat 0,18 J/g ºC Thermal Conductivity 0,0252 J/cm s ºC Electric Resistivity 1,6 · 10−3 Ω · cm

Table 2.5: Properties of T400H

T1000G (0,7%Epoxy /95% Carbon) Property Value Density 1.800 kg/m3 Tensile Modulus 145 GPa Tensile Strength 2.250 MPa -0,45 alpha Coefﬁcient of thermal expansion ·10-6 ºC−1 Speciﬁc Heat 0,18 J/g ºC Thermal Conductivity 0,0252 J/cm s ºC Electric Resistivity 1,6 · 10−3 Ω · cm

Table 2.6: Properties of T1000G

The main conclusion would be that Carbon Fiber is such an interesting material to be used in the aircraft skin. The ﬁber conﬁguration chosen to be used in the aircraft is the T300 because it is able to hold enough tensile strength and has a very low density so the weight of the plane would not specially increased.

2.1.3 Landing gear To choose the materials for the landing gear the main requirements are the weight and the resistance of the materials. Because of their tensile strength the materials that will be evaluated are titanium alloys and also various types of steel.

RA 54 G06-AlOn LSA 3 seats | Project report Initially, titanium [8] was selected for these reasons: 1. Its tensile strength

2. Its low density, which it is considerably lower that any of the steels. But the main problem it presents is its price. Unlike, the steels present a higher density, have a similar tensile strength and its cost are considerably lower. Based on this considerations, when the structure is built, an analysis will be performed to see which material matches better the requirements.

2.1.4 Transparent surfaces 2.1.4.1 Windshield Aircraft windshields [9] may be purchased either from the original aircraft manufacturer or from any of several FAA-PMA sources. These windshields are formed to the exact shape required, but can be slightly larger than necessary so they may be trimmed to the exact size.

2.1.4.2 Windows Are mainly made with two different thermoplastics [9]: 1. Acrylic plastics. They are the current substitute to Cellulose acetates and they are commonly known with the name of Plexiglas. Two main kinds: MIL-P-5425 for military regular speciﬁcations and MIL-P-8184 for military craze-resistant.

2. Cellulose acetate plastics. It was vastly used in the past but was dimensionally unstable and turned yellow after being installed for a long time. It is important to emphasize that transparent thermoplastic sheets soften and deform when they are heated, they must be used and stored where the temperature will never become excessive. Store them in a cool, dry location away from heating coils, radiators, or steam pipes.

2.1.5 Additional material considerations In this subsection there will be added some things that have to be taken into consideration for the manufacturing and the conservation of the aircraft materials [9].

2.1.5.1 Welding Before trying to weld any aluminium alloy it must be tested to determine if it is able to heat treatment or not. In many cases materials are marked to distinguish between heat and non- heat treatable alloys but if in any case it is not marked, there are some ways to identify the materials. First way would be immersing a sample of the material in a 10 % solution of caustic soda. Materials with a high copper content will turn black and would make an evidence of heat tractability. However, if it does not turn black, it is not an evidence of non-heat-treatable so these materials could be tested by chemical or spectre-analysis.

2.1.5.2 Corrosion Prevention There are 3 levels to categorize the severity of corrosion damage (IATA): 1. The damage does not require structural reinforcement or replacement of parts. Also corrosion occurring between inspections that exceeds allowable limits but is local and can be attributed to an event not typical of operator usage of other aircraft of the same ﬂeet.

RA 55 G06-AlOn LSA 3 seats | Project report 2. Corrosion occurring between inspections that requires a single rework/blend -out, which exceeds allowable limits, requiring a repair/reinforcement or complete/partial replacement of applicable structure.

3. Corrosion found during ﬁrst inspection, which is determined to be an urgent airwor- thiness concern requiring expeditious action.

Corrosion prevention [10] involves cleaning the surface and providing a surface ﬁnish through layers of coasting. For the surfaces of aluminium alloys, the coating of a corrosion-inhibiting primer is the ﬁrst coat. Primer is a base protecting used on the outer surface of a substrate before painting. They shield the area and ensure better grid of color to the outer surface. Primers increase the color strength and prevent attack on the substrate by air, water or other elements. Urethane primer is formulated as a high-performance corrosion inhibitor for all types of aircraft. Urethane consists of 2 parts: the base and certain binders with catalysts. Urethane advantages:

• Good inter-coat adhesion

• High durability

• Good chemical resistance

• Abrasion resistant

• High gloss ﬁnish

There are some other primers such as polyurethane or epoxy primers. Corrosion-Resistant Alloy can also be used. They are alloys consisting of metals such as chrome, stainless steel, cobalt, nickel, iron, titanium and molybdenum. When combined, such metals can promote corrosion resistance and can offer reliable protection from corrosion, eradicating the need for maintenance and repair. This will make a compulsory point to check if the chosen materials are corrosion-resistant or if they will need an additional concrete product to prevent material from corrosion.

2.2 Landing Gear

2.2.1 Analysis of the landing gear regulations for an LSA The regulations on the landing gear for LSA are not so speciﬁc [11], so, basically, it will be the most convenient. The only points that must be considered are the following ones:

1. Depending on the type of the LSA we finally select our landing gear will be approximately conditioned. If it is finally chosen a conventional LSA, the landing gear must be obligatorily fixed. Conversely, if it is decided a glider LSA the landing gear can be fixed or retractile. In order to avoid possible conflicts in a further point of the project it will be chosen a fixed landing gear, which would fit in any case.

2. Depending on the landing gear chosen, it will need structural considerations. If it affects to the wing, the ground loads must be justiﬁed in the wing structure.

3. The manufacturer must certiﬁcate the landing gear, with its corresponding wheels, brakes and tires, and install all the landing gear system components.

Based on these previous premises, we can conclude that with its required by the norms, the best option for us is to choose an existing ﬁxed landing gear that does not affect to the wings of our plane.

RA 56 G06-AlOn LSA 3 seats | Project report 2.2.1.1 Landing gear options

The evaluated existing landing gears are the following ones:

2.2.1.2 Atec 322 Faeta

The landing gear is a ﬁxed tricycle [12] undercarriage with a steerable nose wheel. The main gear is constructed as a pair of leaf springs of composites. The front leg is made of composites and metal tube suspended with rubber spring. The main gear is a pair of composite springs. Electronic main wheels size is 350x120 mm, front wheel size is 300x100 mm. The main wheels are equipped with hydraulic disc brakes. Fairings are installed on all wheels. Landing Gear (tricycle with front wheel) speciﬁcations:

1. Wheel spacing: 6,234 ft / 1,900 m

2. Wheel base: 4,742 ft / 1,445m

3. Front tire: 400x400 / 102 x 102 mm

4. Main tire dimensions: 600x400 / 152 x 102 mm

5. Tire pressure: 23,2 psi / 160 KPa

2.2.1.3 TL-2000 Sting S4

The landing gear is convention a ﬁxed [13], tricycle type with a steerable nose gear and two main landing gears. Hydraulically-actuated brakes are attached on each main landing gear wheel.

2.2.1.4 TL-3000 Sirius

TL-3000 Sirius [14] is fitted with a fixed tricycle undercarriage with main wheels fuselage mounted on composite cantilever spring legs; all of them have brakes. The main gear legs made of carbon and glass fibre composite, which dampens impact and makes the aircraft forbearing in the event of hard landings, are attached to the undercarriage bulkhead trestle located under the pilot seats. The gear is equipped with a shock absorber and attached to the firewall; all wheels are almost completely enclosed in spats. There are two variants of the undercarriage which differ from undercarriage wheels depending on the version of the airplane. First variant has its nose gear and main gear wheels with dimension 400 x 100 mm. The nose gear is non-steerable, and it is equipped with shimmy dumper.

Figure 2.2: First variant of the TL-3000 Sirius’ landing gear

RA 57 G06-AlOn LSA 3 seats | Project report

Figure 2.3: Second variant of the TL-3000 Sirius’ landing gear

Both undercarriage options have different type of laminate wheel covers. It is not possible to combine these two undercarriage variants. First, it is necessary to recognize the undercarriage version of your aircraft to follow all instructions related to your type of undercarriage that are described in this article. Besides, the company has also developed a ﬂoat installation for the TL-3000 to allow water operations.

2.2.1.5 Pipistrel Taurus M Double retractable main landing gear [15]. However, it is retractable, and it has been decided not to use a retractable landing gear for the project.

2.2.1.6 Alexander Schleicher ASG 29 E There’s no speciﬁc information about its landing gear [16]. Nevertheless, regarding to its glider nature it is assumed a one wheel retractable and a rear wheel non-retractable; as it is observed in any of its pictures.

Figure 2.4: ASG 29 E ﬂying

2.2.1.7 ONE Aircraft Fixed tricycle landing gear and a single engine in tractor conﬁguration.

2.2.1.8 Sling 4 The landing gear [17], which is made from composites, is a tricycle landing gear with a steerable nose wheel. The main landing gear uses a single continuous composite spring section. Landing gear speciﬁcations:

1. Wheel track: 1,95 m (6,4 ft).

2. Wheel base: 1,68 m (5,51 ft).

3. Brakes: Hydraulic.

4. Main gear tyres: 15x6.00-6, 6-ply

RA 58 G06-AlOn LSA 3 seats | Project report 5. (2,5 bar (36,26 psi) pressure).

6. Nose gear tyres: 5.00-5, 6-ply

7. (1,8 bar (26,11 psi) pressure).

2.2.2 Landing gear calculations

2.2.2.1 Static analysis This ﬁrst analysis will be determining for the landing gear because it will be used to choose the material for the landing gear. The ﬁrst step to proceed with this section has been modeling the whole landing gear with an appropriate software. The results will be presented in the report attachments but for the reader to have a basic idea of what is going to be analyzed, the main parts are shown in Fig. 2.5 and Fig. 2.20.

Figure 2.5: Front leg of the landing gear.

Figure 2.6: Half of the rear landing gear.

Once the models are obtained, the static analysis are performed with the same software. All the analysis were performed over the models meshed with tetrahedral elements and with a quadratic quality. The meshes are created with four Jacobean points and the elements have a size of 6mm.

RA 59 G06-AlOn LSA 3 seats | Project report In order to determinate which was the most appropriate design for the landing gear it has been done a static analysis of the front landing gear and a posterior analysis of the results to make a decision. The front landing gear has two main designs, based on the same structure but one with a solid tube and the other with a void tube to see if the weight reduction compensates the reduction in the resistance. Therefore, the same static analysis has been done over the same front landing gear in six different conﬁgurations:

1. Solid tube in a titanium alloy.

2. Solid tube in a steel 4340.

3. Solid tube in a steel 4130.

4. Void tube in a titanium alloy.

5. Void tube in a steel 4340.

6. Void tube in a steel 4130.

Even though in the materials sections titanium is chosen as the best option, different materials are evaluated because of its price.

This analysis is performed with a ﬁnite element software that builds structured meshes automatically with customized parameters. The parameters chosen for this analysis are: 4 Ja- cobean Points, an element quality of quadratic order, tetrahedral elements and an element size of 10.73mm.

After performing the same analysis, which has consisted in the application of a stress over the wheel axis equal to the weight of the whole aircraft ( 600kg ) - even thought it must be taken into account that in the end this part of the landing gear must only bear with between 8 to 15 per cent of the aircraft gross weight) the results have been those shown in Table 2.7.

Weight (g) Max. Deformation (mm) Material cost ($) Titanium solid 24.364,49 0,1360 901,48 (37$/kg approx) Steel 4340 solid 39.680,76 0.0814 119,04 (3000$/tn approx) Steel 4130 solid 39.680,76 0,0816 79,36(2000/tn approx) Titanium void 8.646,90 0,2837 319,94$ Steel 4340 void 14.082,60 0,1441 42,25$ Steel 4130 void 14.082,60 0,1446 28,16$

Table 2.7: Results of the static analysis of the different alternatives of the landing gear.

Since is not possible to make a decision with only this data, an OWA decision method will be applied to determine which material and structure is more convenient for the case.Results are shown in Table 2.8.

Weight Titanium solid Steel 4340 solid Steel 4130 solid Titanium void Steel 4340 void Steel 4130 void p p*w p p*w p p*w p p*w p p*w p p*w Weight (g) 6 2,97 17,82 1 6 1 6 5 30 4,29 25,74 4,29 25,74 Max. Deformation (mm) 2 3,91 7,82 5 10 4,99 9,98 1 2 3,75 7,5 3,74 7,48 Material cost ($) 1 1 1 4,58 4,58 4,76 4,76 3,66 3,66 4,93 4,93 5 5 Total sum of p*g 45 26,64 20,58 20,74 35,66 38,17 38,22 OWA score 1 0,592 0,457 0,461 0,792 0,848 0,849

Table 2.8: Results of the application of OWA method

RA 60 G06-AlOn LSA 3 seats | Project report As one can see, the model that presents the best result is the void bar model in steel 4130, therefore, this will be the used material.

Hence, the following analysis in both, the front and the rear landing gear will be performed with the parameters from Table 2.9, corresponding to a steel 4130.

Young Modulus 205 GPa Poisson coefﬁcient 0,285 Shear Modulus 80 GPa Density 7.850 kg/m3 Traction limit 731 MPa Elastic limit 460 MPa Thermal conductivity 42,7 W/(m·K) Speciﬁc heat 477 J/(kg·K)

Table 2.9: Properties of the steel 4130 normalized at 870 ºC.

And the result of these analysis are the following ones:

First, for the front landing gear, with the parameters mentioned above, results are shown in Fig. 2.7 and Fig. 2.8.

Figure 2.7: Stress analysis of the front landing gear.

RA 61 G06-AlOn LSA 3 seats | Project report

Figure 2.8: Displacements of the front landing gear.

And must be added that it was also performed a buckling test which granted that there would be no buckling on the structure.

And then for the rear landing gear, the analysis is performed with a solid mesh. This mesh is created with four Jacobean points and tetrahedral elements with quadratic quality. The size of these elements is of 24,9mm which makes a total number of elements of 21.439. With this parameters, the results are shown in Fig. 2.9 and Fig. 2.10.

Figure 2.9: Stress analysis of half of the rear landing gear.

RA 62 G06-AlOn LSA 3 seats | Project report

Figure 2.10: Displacements of half of the rear landing gear.

In this case it was also granted that there would be no buckling in the main bar. And ﬁnally, the last analysis that was performed was a dropping test. This test gave excellent results because it can be appreciated that the drop does not a have a signiﬁcant impact on the landing gear. The results are shown in Fig. 2.11 and Fig. 2.13.

Figure 2.11: Tension results of the dropping test.

As the last image Fig. 2.11 is not very clear, detail Fig. 2.12 is added to give more clarity to the reader.

Figure 2.12: Detail of the point where the tension gets its maximum value.

RA 63 G06-AlOn LSA 3 seats | Project report

Figure 2.13: Displacement results of the dropping test.

2.3 Wing

2.3.1 Initial analysis In order to design the structure of the wing, it is necessary to ﬁx design criteria and know the loads that it will deal with.

In this case, analysis will be focused on stresses produced by bending moment and beam will be designed considering only this moment. After the beam is designed, qualitative twist analysis will be performed in order to see that beam resists.

As speciﬁed in F2245[18], the aircraft must be able to operate with a maximum load factor of 4 and a minimum of -2. Thus, these will be the design criteria.

Before calculating momentum along the wing, load distribution is needed. In order to simplify this ﬁrst analysis, there will be some hypothesis: 1. Cantilever beam. Semi-wing study 2. Constant weight distribution: 750N per wing (structure+fuel). 3. Elliptical lift distribution approximated as four uniform distributions. Also, from requirements and dimensions: 1. MTOW is 600 kg, total lift for n=1 will be approximately 6000 N 2. Wingspan is 16 m, then beam length will be 8 m.

Figure 2.14: Loads distribution and approximations

RA 64 G06-AlOn LSA 3 seats | Project report 2.3.1.1 Lift approximation

Beam will divided in four regions of 2 m length. Total lift will be divided by a multiple of four and a proportional part will be given to each region. Since each region is a uniform distribution, the value applied will be the force divided by the region length. For n different from 1, each distribution will be multiplied by n. For n=1 it was distributed as show in Table 2.10.

R1 R2 R3 R4 Total Fraction 7/16 5/16 3/16 1/16 16/16 Force [N] 1.312,50 937,50 562,50 187,50 3.000 Distribution [N/m] 656,25 468,75 281,25 93,75 -

Table 2.10: Load distribution for n=1

2.3.1.2 Weight approximation

The distribution must sum a force of 750 N. Since it is uniform, dividing by beam length we could ﬁnd its value of 750/8=93,75 N/m

2.3.1.3 Results

With this information, moment along the beam was calculated with STRIAN[19].The distribution for load factors of interest are shown in Table 2.11 and results are shown in Fig. 2.15, Fig. 2.16 and Fig. 2.17.

Distribution [N/m} R1 R2 R3 R4 n=1 656,25 468,75 281,25 93,75 n=4 2.625 1.875 1.125 375 n=-2 -1.312,5 -937,5 -562,5 -187,5

Table 2.11: Lift distribution for each load factor

Figure 2.15: Moment diagram for n=1

RA 65 G06-AlOn LSA 3 seats | Project report

Figure 2.16: Moment diagram for n=4

Figure 2.17: Moment diagram for n=-2

2.3.2 Sizing of the beam

With the moment diagram, maximum requirements for the beam are determined and its section can be sized. From n=4, maximum moment at the embedding is 30 kN·m. This will be the requirement used to choose beam’s section. Also, from chord at the embedding of 1.000 mm, maximum thickness of 15 per cent (see Table 1.6) is 150 mm, which determines how big the beam can be. Despite the wing has variable chord, the most critical moment is found in the embedding. Thus, as decision parameter, only embedding sizing will be considered. In order to fit in the wing, beam will have variable size section. After choosing the final section from the embedding parameter, the full beam will be analyzed with FEM methods and if critical stresses appear in any regions, it will be reinforced. From F2245[4], CentrAl maximum stress can be half of its rupture tensile strength of 812 MPA. Then, maximum stress for aluminum can be 406 MPA.For carbon fibers, the maximum is 400 MPA. In order to calculate maximum stress for each material, maximum axial stress produced as result of beam’s bending will be calculated. If this stress is lower than the maximum allowed, the beam can be accepted for the structure.

Also, it will be considered that buckling won’t appear thanks to the ribs. More information about them will be given in Section 2.3.3. Two possibilities are proposed for the beam:

1. Wagner CentrAl beam (I section) with carbon ﬁber reinforcements

2. Square carbon ﬁber section with extra thickness in superior and inferior faces.

RA 66 G06-AlOn LSA 3 seats | Project report 2.3.2.1 I section I section is useful for bending stresses but requires a second beam to support torsion. In this analysis only bending will be considered. If the result has a similar weight to the other option, this will be rejected because of the extra weight added by the second beam. In order to reinforce the beam, 1 mm carbon ﬁber layers will be applied to the superior and inferior faces of the beam. This layer reduces the maximum thickness of the beam. It will be set in 144 mm. The procedure applied to calculate axial stresses is that for ideal composite beams [20]. Beam will be idealized with a young modulus:

∗ E = E0 (2.1)

Where E0 is Young modulus of one of the materials.Then proportional factor will be deﬁned:

E n = i (2.2) i E∗

Where subindex determines the material(consequently, n0=1).Then inertia can be calculated as: ∗ I = ∑ ni Ii (2.3) Finally axial stress: Moment σ∗(y) = (2.4) I∗ Thus, axial stress in each material, shown in Fig. 2.18, will be:

∗ σ(y) = niσ (y) (2.5)

Figure 2.18: Stress distribution for composite beam

In Table 2.12, different dimensions, referred to those shown in Fig. 2.19, for the section and their maximum stress and weight are shown. In all the cases, web will be a 0.5 mm thick plate.

Figure 2.19: I section parameters

RA 67 G06-AlOn LSA 3 seats | Project report E [mm] Y [mm] B [mm] H [mm] Al Stress [MPa] Fiber Stress [MPa] Weight [kg] 20 72 80 4 435,9 441,9 46,0 30 72 80 4 435,4 441,4 52,9 20 72 90 4 387,5 392,9 50,0 20 72 100 4 348,8 353,7 54,0 20 72 90 3 457,0 463,3 38,8 20 72 100 3 411,4 417,1 41,9

Table 2.12: Beam conﬁgurations and results

2.3.2.2 Square section Square section might be heavier for same material beams because of its vertical walls but, since it acts as a torsion box, it won’t need a second beam. In this case, axial stress calculations are simpliﬁed to: Moment σ(y) = y (2.6) Inertia

Figure 2.20: Stress distribution

Like in the previous case, different dimensions, referred to Fig. 2.21, and its results are shown in Table 2.13.

Figure 2.21: Square section parameters

H[mm] B[mm] Y[mm] T[mm] Stress[MPa] W[kg] 4 80 72 4 421,3 50,5 4 90 72 4 391,4 52,7 4 100 72 4 365,4 55,0 3 90 72 4 430,2 44,6 3 100 72 4 399,0 46,9 3 110 72 4 372,1 49,1

Table 2.13: Beam conﬁgurations and results

RA 68 G06-AlOn LSA 3 seats | Project report 2.3.2.3 Results

Both possibilities show similar results for similar dimensions but ﬁrst option is slightly heavier and needs a second beam, which makes it even heavier. Since weight requirement is really restrictive, it is very important to reduce weight as possible. Thus, the rectangular section has been chosen. Dimensions chosen will be those of the lightest beam within regulations. From Table 2.13, these are:

H = 3mm; B = 100mm; T = 4mm (2.7)

Since the wing tip isn’t as required as the embedding, beam’s height will be adapted to wing’s thickness but B will be reduced to 80 mm. This way, section is reduced to a 80x80 mm square beam and weight is decreased.

Beam is studied with SolidWorks FEM analysis, as speciﬁed in Section 2.2.2.1, in order to see that it can be used. Loads are those for load factor of n=4. As shown in Fig. 2.22, stress never exceed the maximum allowed of 400 MPa. The maximum stress is found near the embedding as expected. In Fig. 2.23, the maximum displacement appears at the wing tip. Given that these are results for a load factor of 4, a displacement of 1.100 mm is acceptable. Using trigonometrical relations, angle of displacement is calculated and results in approximately 10 degrees, which can be considered a minor displacement.

(a) Results

(b) Embedding detail

Figure 2.22: Stress analysis

RA 69 G06-AlOn LSA 3 seats | Project report

Figure 2.23: Displacement results

If this displacement appeared for cruise operation conditions, instead, the beam should be redesigned. Since real shape of the wing would be completely different from the one used for aerodynamic calculations, the beam must be changed in order to reduce displacement to a value that could be considered negligible.

However, operation with a load factor of 4 is limited to few particular maneuvers. Then, this maximum displacement would only appear momentarily.

It is interesting how FEM analysis show lower maximum stress than the initial one. The reason for this difference might be that 3D model used for computational analysis is different from geometry used for the initial approach. To ensure that beam ﬁts in the wing, the beam has been modelled completely tangent to the skin. Then, thickness of the walls of the beam is measured from the nearest point to the center. This way, walls are slightly thicker than previous geometry and, consequently, inertia is higher and stress is lower.

Despite it is out of the scope of this project, it is important to mention that further analysis should be done in order to study if aeroelastic effects could take the wing to failure.

2.3.3 Ribs

Ribs are used to avoid buckling and shape the wing. They will be a sandwich composite with a central 2 mm layer made of aramid honeycomb and external 1 mm carbon ﬁber layers, as shown in Fig. 2.24. With its 48 kg/m3 density, honeycomb support great compression stresses with really low weight.

Figure 2.24: Sandwich ribs

To calculate distance between ribs, an approximation of the critical buckling length will be calculated for the most critical section of the wing. Since higher inertia helps to avoid buckling, tip’s section, with the lowest inertia, will be the most critical one.

RA 70 G06-AlOn LSA 3 seats | Project report With the purpose of simplifying calculations, the skin will be approximated as a rectangular thin airfoil beam with length equal to the chord, height equal to the maximum thickness of the section airfoil and thickness equal to skin’s thickness, as shown in Fig. 2.25. Given that thickness of the skin is out of the scope of the project, a mean value of 1 mm will be taken. Further analysis of the skin should be taken for more precise results.

Figure 2.25: Section approximation

With this considerations critical length can be calculated from:

π2EI = Pcrit 2 (2.8) Lcrit s π2EI Lcrit = (2.9) Pcrit E is Young modulus of the material. This value for carbon ﬁber can be found in Section 2.1. Pcrit can be calculated from critical stress and section’s surface:

Pcrit = σcritS (2.10) Since regulations stipulate a maximum stress of 400 MPa for carbon ﬁber, this value will be taken as the critical one. Skin can’t receive higher stresses, then it has to be ensured that buckling does not appear before this limit.

Section’s surface can be calculated with its perimeter and thickness:

S = pt = (2c + 2h)t (2.11) Inertia for this rectangular section is:

th3 ct3 h I = 2 + 2( + ct( )2) (2.12) 12 12 2 For tip’s section of 600 mm of chord and maximum thickness of 12 per cent (see Table 1.6) the results are shown at Table 2.14.

2 4 c [mm] h [mm] p [mm] S [mm ] I[mm ] Pcrit [N] Lcrit [mm] 600 72 1.344 1.344 1.617.508 537.600 1.442

Table 2.14: Results of buckling analysis

Since critical length is the maximum distance between ribs that avoid buckling, a lower distance will be used to ensure buckling doesn’t appear despite the approximations. Distance chosen will be 1.000 mm.

2.3.4 Final Result The ﬁnal result is a square section beam made of carbon ﬁber with composite ribs. An overall image is shown in Fig. 2.26

RA 71 G06-AlOn LSA 3 seats | Project report

Figure 2.26: Beam and ribs

2.4 Fuselage

In order to proceed to design the fuselage a series of models with different beam proﬁles and different conﬁgurations are going to be analyzed.

The first section that was going to be analyzed is a beam with an I shape. This shape is chosen because it is the standard beam shape and it could give an initial approach to the problem solution. In order to choose the ideal profile there will also be analyzed two normalized profiles with different shapes and a more compact section.

Once the sections were chosen, the following step was deﬁning the number of stringers and frames the fuselage would be formed for. Usually, aircraft fuselages are formed by a large number of both of the elements [21], but as the main requirement of this concrete aircraft was a strictly limited weight, hence the main objective will be to use the minimum number of structural components at a minimum thickness.

Therefore, there will be used four main stringers that will go from leading edge to trailing edge. It could have been considered the possibility of using only three stringers (one for the part above, one for the part below and one on one side) but this would have generated an horizontal displacement of the gravity center, which is not convenient. Then, it must also be taken into consideration that the stringer that crosses the windshield will be divided in two parts. For the frames, two main frames will be considered, located the first one where the windshield begins and the other one, formed by only half arch; where it ends, to maintain the structural stability. It will not be considered the addition of more frames because even though the skin is thin, it will have a clamping function. With this distribution, three possible solutions can be raised: one with this exact configuration and a thickness of the profile of 1 cm Fig. 2.27, another with the same structure but a thickness of 0.5 cm Fig. 2.28, and a last one as the previous one but with the two frames with their structure complete, in case reducing the thickness could affect to this specially empty part Fig. 2.29. For the other two cases, there will be used two main stringers that will go from the leading edge to the trailing edge. To complement these ones, two shorter stringers will be added along the cabin. However, to compensate the structural loss of reducing the length of the stringers three additional frames are added. Those two models have a total of seven frames, five along the cabin and two more frames along the tail. These options raise two more possible solutions Fig. 2.30 Fig. 2.31.

The ﬁve options can be seen in the following images:

RA 72 G06-AlOn LSA 3 seats | Project report

Figure 2.27: Fuselage with a beam thickness of 1 cm.

Figure 2.28: Fuselage with a beam thickness of 0.5 cm.

Figure 2.29: Fuselage with a beam thickness of 0.5 cm and a reinforcement.

RA 73 G06-AlOn LSA 3 seats | Project report

Figure 2.30: Fuselage with a U beam.

Figure 2.31: Fuselage with a rectangle beam.

To choose the best one between this alternatives, a series of analysis will be performed.

2.4.1 Analysis

For the ﬁve options, three main possible problems have to be considered. The principal problem is that the stringers will be working at traction, considering that the thrust will be acting pulling the airplane forward while the drag will be acting pulling the aircraft backwards. Given these forces, another problem that has to be taken into account is the possibility of buckling in the stringers. At the same time, the last problem that can happen is the deformation of the frames.

As it has been chosen a concrete profile to avoid bending moments and buckling the main problem that may concern to the aircraft structure is that the stringers will be working at traction and has to be checked that none of the configurations might generate a configuration of tensions that surpasses the elastic limit of the material.

However, it must be taken into account that due to the choice of a composite material for the fabrication of the fuselage, the Young’s modulus will be different depending on the direction that is analyzed. This could have meant that ﬁnite element methods are not valid for this analysis but given that the analysis will be performed with the material working at traction and with the forces applied in a longitudinal direction to the stringers; it can be used the longitudinal Young’s modulus with a negligible error. A slightly higher security factor will be taken although to ensure the security of the structure because the cross Young’s modulus is lower.

RA 74 G06-AlOn LSA 3 seats | Project report Then, provided that the analysis will be acceptable to the case, it can be performed to decide which conﬁguration will be used, and the thickness of the proﬁles, will be used the following criteria:

1. Weight of the fuselage structure.

2. Maximum tension.

3. Security factor (Elastic limit of 8, 122e + 8).

4. Maximum displacement.

2.4.1.1 Performed analysis To evaluate this, the five analysis have been performed with a finite element method under the same conditions: two distributed loads applied in each edge that will represent the thrust and the drag. The thrust has been estimated based on the engine power and the aircraft characteristic velocity and the drag has been simply calculated for a horizontal straight flight. This way, the applied loads are:

Enginepower Thrust = (2.13) characteristicvelocity

59.656W Thrust = = 1.368, 2569N (2.14) 43, 6m/s and, for a case where the thrust is not equal to the drag (a ﬂight with acceleration) and from equalling the maximum aircraft weight to lift:

1 MTOW = ρv2SC (2.15) 2 L From this equation, using one more time the characteristic velocity, it can be obtained the value of the CL, and furthermore, the value of the CD:

CD = 0, 01016 (2.16)

1 D = ρv2SC = 146, 984N (2.17) 2 D These loads will be applied in a perpendicular direction to the aircraft axis. Aside, it has to be said that Lift and Weight loads will not be considered because they do not affect directly to the fuselage structure. The last thing that must be speciﬁed are the mesh parameters. It is used a structured mesh with four Jacobean points. The quality of the tetrahedral elements is quadratic and its size is of 30,6mm.

2.4.1.2 Results of the analysis These results are presented graphically in ﬁgures from Fig. 2.32 to Fig. 2.41, and the choice between the alternatives has been taken using an OWA method.

In order to establish the weights, it will be taken into account that the weight is the most restricting item. Then, the second one will be the security factor because it is recommended to be of 2. So, overdimensionated structures will be directly related to extra weight. Finally, the maximum deformation will also have a point, but not so important because the skin will also retain the fuselage structure.

RA 75 G06-AlOn LSA 3 seats | Project report The results are presented in the following table:

Weight Max. tension Security factor Max. displacement 1cm thick I 215,5 kg 1,123e+7 Pa 72,32 7,39 mm 0,5 cm thick I 72,5 kg 2,487e+8 Pa 3,27 18,25 mm 0,5 cm thick I 94,6 kg 3,567e+7 22,77 3,173 mm with reinforcement 0,35 cm thick U 41,49 kg 2,181e+7 37,24 2,01 mm 0,6 cm thick rectangle 72,54 kg 8,06e+6 100,77 1,21 mm

Table 2.15: Principal results of the different fuselage alternatives analysis

Then, in order to choose the better option, an ordered weighted averaging method was used, obtaining the following results.

Weight (kg) Max. Deformation (mm) Security factor Total p*g Score OWA Weight 6 2 4 70 p 1 3,54 2,17 1cm thickness I p*w 6 6,9 8,68 21,58 0,31 p 3,86 1 5 0.5cm thickness I p*w 23,16 2 20 45,16 0,65 p 3,42 4,54 4,20 0,5cm thickness reinforced I p*w 20,52 9,08 16,8 46,4 0,66 p 5 4,81 3,61 0,35cm thickness U p*w 30 9,62 14,56 54,18 0,77 p 4,76 5 1 1,6cm thickness rectangle p*w 28,56 10 4 42,56 0,61

Table 2.16: Results of OWA method

Which led to the conclusion that most adequate alternative was the fuselage with a U proﬁle beam.

Figure 2.32: 10mm thick I:Results of the stress analysis.

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Figure 2.33: 10mm thick I:Results for the displacement.

Figure 2.34: 5mm thick I:Results of the stress analysis.

Figure 2.35: 5mm thick I:Results for the displacement.

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Figure 2.36: 5mm thick I with reinforcement:Results of the stress analysis.

Figure 2.37: 5mm thick I with reinforcement:Results for the displacement.

Figure 2.38: 3,5mm thick U:Results of the stress analysis.

RA 78 G06-AlOn LSA 3 seats | Project report

Figure 2.39: 3,5mm thick U:Results for the displacement.

Figure 2.40: 16mm Rectangle: Results of the stress analysis.

Figure 2.41: 16mm Rectangle:Results for the displacement.

2.5 Tail

2.5.1 Elevator beam

For beam analysis, same procedure applied to the wing will be applied. This time, force distribution will be approximated as uniform. The value for the total lift is calculated with the

RA 79 G06-AlOn LSA 3 seats | Project report maximum lift coefﬁcient from Fig. 1.21.Then:

1 L = ρV S CL = 7.500N (2.18) 2 ne e max Using uniform distribution, the value of the loads are shown in Table 2.17.

Length [mm] Total Force [N] Distribution[N/m] 4.000 7.500 1.875

Table 2.17: Total load and distribution

The results, calculated with STRIAN [19] are shown in Fig. 2.42.

Figure 2.42: Elevator momentum diagram

Then, maximum moment at the embedment is 3.750 N·m. The results for different rectangular beams are shown in Table 2.18.

H [mm] B [mm] Y [mm] T [mm] I [mm4] Stress[MPa] W [kg] 2 45 30 2 223.382 503,6 3,0 3 45 30 2 259.384 433,7 3,8 2 45 30 3 291.316 386,2 3,6 3 45 30 3 327.321 343,7 4,4 3 40 30 3 302.953 371,3 4,2 1 55 30 3 304.047 370,0 3,2 1 50 30 3 284.553 395,4 3,0

Table 2.18: Dimensions for elevator beam. Referred to Fig. 2.21

After some iteration, the lightest beam is chosen and its dimensions are:

H = 1mm; B = 50mm; T = 3mm (2.19) Finally, FEM analysis is shown in Fig. 2.43.

As happened with the wing, stress from the analysis is below the stress previously calculated. The reason might be the same, 3D model used is slightly different from ideal geometry used in initial analysis.

From trigonometrical relations,a displacement of 330mm results in about 10 degrees. As considered in the wing, it can be accepted.

RA 80 G06-AlOn LSA 3 seats | Project report

(a) Stress

(b) Displacement

Figure 2.43: Elevator FEM analysis

2.5.2 Fin beam The procedure and approximations for the ﬁn are the same applied to the elevator. The value for the total lift is calculated with the maximum lift coefﬁcient of 1,2 from Fig. 1.26.Then:

1 L = ρV S CL = 4.000N (2.20) 2 ne e max Now, assuming an uniform distribution, the value of the loads are shown in Table 2.19.

Length [mm] Total Force [N] Distribution[N/m] 1.200 4.000 3.335

Table 2.19: Total load and distribution

The results, calculated with STRIAN[19] are shown in Fig. 2.44.

Then, maximum moment at the embedment is 2.400 N·m. The results of some iterations for beams are shown in Table 2.20.

This time, weight barely changes between options compared to MTOW of 600 kg. Then, decision criteria will be two: weight must be lower than 1 kg and maximum stress on the

RA 81 G06-AlOn LSA 3 seats | Project report

Figure 2.44: Fin moment diagram

H [mm] B [mm] Y [mm] T [mm] I [mm4] Stress[MPa] W [kg] 1 45 45 2 187.381 384,2 0,6 2 45 45 2 223.383 322,3 0,9 1 45 45 1 114.323 629,8 0,4 1 40 45 2 170.561 422,1 0,6 2 40 45 2 206.563 348,6 0,8 2 35 45 2 189.743 379,5 0,8 1 30 45 3 182.210 395,1 0,6

Table 2.20: Dimensions for ﬁn beam. Referred to Fig. 2.21 embedment will be the minimum between different options. Thus, beam chosen and its dimensions are:

H = 2mm; B = 45mm; T = 2mm (2.21)

Finally, FEM analysis is shown in Fig. 2.45.

(a) Stress (b) Displacement

Figure 2.45: Fin FEM analysis

Results are within regulations. Same conclusions as in elevator can be taken.

RA 82 G06-AlOn LSA 3 seats | Project report 2.5.3 Ribs Using the same methodology that is used to calculate distance between ribs in the wing, results for ﬁn and elevator are shown in Table 2.21 and Table 2.22. For ﬁn, chord is 500 mm and maximum thickness is 12 per cent. For elevator, chord is 250 mm and maximum thickness is 12 per cent.

2 4 c [mm] h [mm] p [mm] S [mm ] I [mm ] Pcrit [N] Lcrit [mm] 500 60 1.120 1.120 936.083 448.000 1.201

Table 2.21: Buckling analysis for ﬁn

2 4 c [mm] h [mm] p [mm] S [mm ] I [mm ] Pcrit [N] Lcrit [mm] 250 30 560 560 117.042 224.000 601

Table 2.22: Buckling analysis for elevator

As explained in Section 2.3.3, lower distances are used to ensure that ribs work. Thus, length chosen for the vertical distance is 1.000 mm and 500 mm for the horizontal.

2.5.4 Final Result The ﬁnal result are two rectangular carbon ﬁber beams with the same composite ribs used in the wing. An overall image is shown in Fig. 2.46.

Figure 2.46: Beams and ribs

RA 83 G06-AlOn LSA 3 seats | Project report

RA 84 Chapter 3

Power plant

Figure 3.1: Wankel motor IAE50R – AA

Due to the limited weight design conﬁgurations, the motor choice and its systems will be critical. The objective is to arrange a light motor with notable power-to-weight ratio. Also the infeasibility to carry to much fuel weight will be crucial in order to ﬁnd of a suitable engine. The design of the propellers will also follow the requisites mentioned above.

3.1 Motor

The election of the motor is restricted for its weight and power. After a market research of current engines manufactured by major brands like Rotax, which has a worldwide reputation in building motors for small recreational aircraft, it was found that the majority of conventional small motors were too heavy and powerful for a gilder like aircraft characteristics. Among all the Rotax motors, the Rotax 912 is clearly the most applicable engine for a three seater LSA[22]. Another engine suitable for the powerplant requirements is the Wankel AE50R-AA motor, which is manufactured by the Austrian company Austro Engine founded in 2007. Despite the novelty of the company, this motor has an unique set characteristics that are really convenient:

"The AE50R series is a single stage rotary engine developing 41 kW and is the only rotary engine worldwide that is certiﬁed according to EASA Part 22 Subpart H on today‘s market. The remarkable power-weight ratio (2 hp : 1 kg) makes it the ideal engine for light or

RA 85 G06-AlOn LSA 3 seats | Project report unmanned vehicles. With more than 2000 engines produced, the AE50R has proven its relia- bility in both, manned and unmanned applications[23]."

The table below shows its features compared with the Rotax motor:

Name Rotax 912 AE50R Type: 4- engine stroke Wankel Power(hp): 80 55 Mass(Kg): 55,4 28 Aprox extra Weight(Kg): 11 6 Coolant: Air Water and Glycol Fuel pump: Yes Yes Gear box: included not included Alternator: 12V/40A 14V/18A RPM: 5800 7750 PWR (max) (hp/Kg): 1,2 1,7 Torque(Nm): 103 52,5 Fuel: AVGas100 RON 95 AVGas 100LL RON 95 Fuel consumtion: 15,8L/h 12L/h Certiﬁcation: FAR22 JAR33 ESA Price: —– —–

Table 3.1: Rotax 912 Vs AE50R comparison

Even though the Rotax 912 is considerably heavier than the AE50R, both engines might be feasible for the power plant conﬁguration of our LSA like aircraft. In order to decide which engine would be the most suitable according to the needs and limitations of the project, both engines are going to be evaluated using the OWA and PRESS methods, which are going to assess the principal characteristics of the engines. Those characteristics and its importance regarding the requirements of the project are listed below:

• Power (hp): The horse power provided by the engine.

• Mass (Kg): The net mass of the engine itself, excluding the mass of extra or auxiliary elements such as cooling system and ﬂuids.

• Coolant: Capability of the engine of avoiding overheating using a cooling system. Lack- ing of a system will get the lowest grade whereas cooling systems will be graded according to the complexity of the system.

• Torque (N·m): Rotating force generated by the engine

• Accessories: Availability of engine extra parts such as the gear box or the engine bench.

Numeric criteria have been graded according to the quotient between the value provided by each engine and the difference between the maximum and minimum value:

Criteria value (3.1) Max value-Min value

RA 86 G06-AlOn LSA 3 seats | Project report 3.1.1 OWA Starting for the OWA method:

Criteria Rotax 912 AE50R Importance Power 3,2 2,2 6 Mass 0,5 0,99 10 Coolant 1 5 4 Torque 0,49 0,96 4 Accessories included 5 1 2

Table 3.2: OWA importance table

Using the importance table, the OWA method can then be applied to sort out which alternative is probably going to be the best one:

Weight Rotax 912 AE50R Criteria G P PXG P PXG Power 6 3,2 19,2 2,2 13,2 Mass 10 0,5 5 0,99 9,9 Coolant 4 1 4 5 20 Torque 4 0,49 1,96 0,96 3,84 Accessories included 2 5 10 1 2 SUM(pxg) 26 40,16 48,94 OWA 0,309 0,376

Table 3.3: OWA results

3.1.2 PRESS Otherwise, if the alternatives are to be evaluated using the PRESS method, ﬁrst of all a relative weight table is needed:

Criteria Power Mass Coolant Torque Accessories included Weight 6 10 4 4 2 Relative weight 0,231 0,385 0,154 0,154 0,077

Table 3.4: Relative weight table

Secondly, using the relative weight table and the grades each alternative has got in each criteria, another table can be written:

Grades Alternatives Power Mass Coolant Torque Accessories included Rotax 912 3,2 0,5 1 0,49 5 AE50R 2,2 0,99 5 0,96 1 Pmax 3,2 0,99 5 0,96 5

Table 3.5: Graded criteria table

RA 87 G06-AlOn LSA 3 seats | Project report Using the table above, the valuation matrix table can be obtained. It is to be reminded that the PRESS method is being followed:

0,231 0,194 0,031 0,079 0,077 0,159 0,385 0,154 0,154 0,015

Table 3.6: Valuation matrix table

From the the valuation matrix table, the domination matrix can be done:

0 0,134 0,134 0,389 0 0,389 0,389 0,134

Table 3.7: Domination matrix table

Finally, the PRESS method score for each alternative is attained:

Alternative Rotax 912 AE50R PRESS 0,344 2,909

Table 3.8: PRESS results

Hence, employing the OWA and PRESS methods we get the following results:

Rotax 912 AE50R OWA 0,309 0,376 PRESS 0,344 3,909

Table 3.9: OWA and PRESS analysis results

Therefore, the AE50R engine has turned out to be the most suitable engine regarding the requirements of the project.

The dimensions of the motor can be seen in the ﬁgure 3.2. Is is notably small and can easily be installed in the front of the airplane.

RA 88 G06-AlOn LSA 3 seats | Project report

Figure 3.2: IAE50R – AA dimensions

AE50R-AA Wankel motor Weight: 28g Approximate price: approx: 9000 eur

Table 3.10: Motor characteristics

RA 89 G06-AlOn LSA 3 seats | Project report 3.2 Motor Mount

The Austro Engine company doesn’t manufacture speciﬁc motor mounts for the AE50R-AA engine and there are no companies that manufacture this model due to its small market fraction. Therefore, the motor mount will need to be adapted from an existing one. The mount will be readjusted from a Rotax 912 motor as this motor has a similar mount attachment and dimensions.

The ﬁgure 3.3 shows the attachment points designed for this function. During the mount design it has to be taken into account that there are side parts of the motor mount that might be in the path of the main subjection structure of the aircraft.

Figure 3.3: IAE50R – AA motor mount points

The ﬁxation system will be bought from a third party provider and adapted for the Wankel engine. The similarities in the design will simplify the readjustments and minimize the need for re-welding parts[24]. The ﬁgure 3.4 illustrates the engine mount parts and design.

Figure 3.4: Standard frame mount

Motor mount frame Weight: approx 4,0kg Price: 981,27 eur

Table 3.11: Motor frame characteristics

RA 90 G06-AlOn LSA 3 seats | Project report 3.3 Auxiliary motor systems

For the proper functionality and optimal behaviour of the motor some auxiliary components are required to guarantee it’s best performance.

3.3.1 Gear box

The Gear box assures that the optimal RPM is transmitted from the engine to the propeller, providing maximal thrust according to the propeller performance. The Gear box selected is a Helical Gear from the Aeromomentum Company. It is made from CNC 6061-T6 billet case for high strength and light weight and heat treated CrMo4130 steel gears for a prolonged life. Despite being designed for 4 cylinder engines, up to 225hp, it will also be convenient for the Wankel motor. This gear box will convert the maximum RPM from 7750rpm to 2994rpm[25].

Figure 3.5: Gear Box

Aeromomentum Gear Box Gear ratio: 2,588:1 Weight: approx 3kg Price: 1412 eur

Table 3.12: Motor frame characteristics

3.3.2 Exhaust system

The Wankel motor requires an exhaust system composed of pipes to be attached to redirect the combustion gases produced by the engine. All different parts will be bought and assem- bled to redirect the ﬂux under the chassis once the engine is installed and mounted inside the plane[26]. The different components will be:

• Micro louver air shield: It will allow air to ﬂow through the metal shield. The movement of the vehicle creates dynamic air ﬂow,which has a cooling effect by natural convection. The textured surface created by the micro air louvers creates more area for heat dissipation.

• Exhaust clamps: This clamp is used on the exhaust system. In case that the spot weld fails, the clamp will avoid the release of high-temperature gas inside the engine compartment, which could possibly cause in-ﬂight ﬁre.

RA 91 G06-AlOn LSA 3 seats | Project report • Flexible ceramic heat shield: It can be bent and manipulated to suit different geometries, without damaging the thermal barrier. Used to protect sensitive components from heat.

• Mufﬂer Shroud PA-18

• Stainless Steel 90º bend

• Stub exhaust Stacks: An existing exhaust Stack that will be adapted from a similar motor exhaust system

3.4 Propeller

In the designing process of the propeller of the plane, two options have been taken into consideration. The ﬁrst one would be purchasing a regular standardized propeller for 50hp motors which had already been designed and was being manufactured by a third company. The second option would be manufacturing a custom propeller according to the requirements of the project. In order to choose between these two options, an early prototype of a custom blade has been designed (the designing process its thoroughly explained in subsection 3.4.1) and its main characteristics have been compared with those of an already existent LSA propeller:

• Custom propeller:

Custom propeller Number of blades: 2 Diameter: 144 cm Total mass: 1,1 kg RPM: 2000 rpm Price: >2500 EUR

Table 3.13: Custom propeller characteristics

• HAUTECLAIRE 3 propeller: Manufactured by the French E-PROP company, it is a light carbon propeller with ground adjustable pitch and a leading edge protection based on nanostrength composites[27].

HAUTECLAIRE propeller Number of blades: 3 Diameter: 155cm Total mass: 1,7Kg Moment of inertia: 1.400 kg·cm2 RPM: 3000rpm Price: 724 EUR

Table 3.14: HAUTECLAIRE propeller characteristics

RA 92 G06-AlOn LSA 3 seats | Project report

Figure 3.6: HAUTECLAIRE Propeller

Although the custom propeller might be more convenient to the power plant of the project, both propellers have overall similar characteristics. Therefore, it was ﬁnally preferred to purchase the Hauteclaire propeller, as it had already been certiﬁed and would always be cheaper than manufacturing a custom one.

3.4.1 Custom propeller design

In order to maximize efficiency for the aircraft it would definitely be better to completely design a propeller specifically fit for the LSA like aircraft.

A Matlab program has been created to study possible design shapes for the propeller. Using several physical models undertaking certain assumptions, the program outputs the thrust, useful power and engine power as a function of a plethora of geometric variables such as: chord and torsion distribution, aerodynamic proﬁle, radius of the propeller and the central cone, etc. Flight variables are also contemplated: average speed, altitude. Since a DNS or otherwise computationally intensive analysis would be too expensive (especially for an iterative process such as optimization), a simpler, less reliable but more affordable method was chosen. In this case it was a Blade Element Method, wherein each blade is discretized in thin slices or blade elements. This way airﬂow is almost two-dimensional. Drag and Lift are calculated from each blade element, projected onto and integrated along the blade.

The propeller’s most important design values were iterated so as to optimize its performance. Those values were varied by until a good performance scheme was obtained. The results of the calculations in static ﬂight at cruising speed are presented in Figure 3.7:

RA 93 G06-AlOn LSA 3 seats | Project report

Figure 3.7: Results of computation

The plots are limited at 1920 rpm, bellow which the propeller acts as an aerobrake, and at 2225 rpm, where power consumed reaches the engine maximum at 41kW (55 hp). It is to take into account, however, that this limits are dependent on various factors, such as altitude and speed. For instance, it is only necessary to ﬂy at 1100 rpm to overcome stall speed.

These graphs are therefore the instantaneous response to a change of thrust when cruising. The rather lineal nature in the thrust plot implies that the engine will be easy to control, because an increase of RPM leads to a very similar increase of thrust regardless of the ﬂight regime.

RA 94 G06-AlOn LSA 3 seats | Project report

In static ﬂight at cruising speed (calculated in section 2.5), the results are shown in Table 3.15:

Custom propeller results Power consumed: 14.01 kW Power delivered: 8.33 kW Efﬁciency: 59% RPM: 2021 rpm Thrust: 147 N

Table 3.15: Custom propeller cruising specs

3.5 Overall Powerplant evaluation

The overall evaluation of the power plant can conclude that all the specifications have been acquired. The motor is lightweight enough for our LSA like design, and the auxiliary systems and the motor mount doesn’t add much weight. The propeller designed is a two blade body with good power efficiency and enough thrust to power the airplane successfully. Also the motor is very efficient and doesn’t consume much fuel for hour meaning that the fuel tanks can be relative small and save some weight.

RA 95 G06-AlOn LSA 3 seats | Project report

RA 96 Chapter 4

Systems and avionics

In this chapter will be presented the different instruments and systems that an aircraft needs according to [4] and [28]. These instruments pretend to provide information to the pilot about ﬂight, engine and systems conditions; as to allow the pilot to control the aircraft through levers, buttons and pedals.

The main considerations that will be taken into account are that the equipment chosen must obey the laws established in [4] and [28] ( all of them have to be LSA approved or certiﬁed). Without forgetting that weight is an important factor for design limitation, and also price has to be taken into account. And other important factor is the compatibility between equipment in order to achieve it has been selected instruments from the same company or group to ensure that fact. And the equipment has been carefully chose to read the speciﬁc range in which the airplane or engine parameters work.

In the section 8 Required Equipment of [4], an aircraft shall be designed with a mininum instrumentation and equipment. These equipment can be classiﬁed into the following groups:

• Flight and Navigation Instruments

• Powerplant Instruments

• Miscellanous Equipment

• Safety Belts and Harnesses

4.1 Flight and Navigation Instruments

Flight and Navigation instruments are the instruments in the cockpit of an aircraft that provide the pilot with information about the ﬂight situation of the aircraft, such as altitude, airspeed and direction. In [4] it is established that an aircraft must have:

• Airspeed indicator

• Altimeter

While in [28], rules and instrumentation for VFR, night VFR and IFR are more speciﬁcied.

For visual flight (VFR) an airspeed indicator, an altimeter, as ASTM requirements [4] settle; and a compass or a suitable magnetic direction indicator are required. On the other hand, instrument flight rules (IFR) additionally require a gyroscopic pitch-bank (artificial horizon), direction (directional gyro) and rate of turn indicator, plus a slip-skid indicator, adjustable altimeter, and a clock.

RA 97 G06-AlOn LSA 3 seats | Project report Light-sport aircrafts (LSA) must be able to fly following VFR rules. Nevertheless, the aircraft will also have IFR instrumentation being able to fly following instrumental flight rules if it is necessary. On the other hand, safety is improved by giving more useful information to the pilot.

4.1.1 Analogic Flight and Navigation Instruments A summary of the Flight and Navigation instruments selected is featured below.

1. Airspeed indicator Instrument used in an aircraft to display the aircraft’s airspeed, typically in knots, to the pilot. It measures the difference in pressure between the air around the aircraft and the increased pressure caused by propulsion.

Figure 4.1: Winter 7 FMS 5 Air Speed Indicator

WINTER 7 FMS 5 Airspeed Indicator Dimensions: 60 W x 60 H x 62,5 D [mm] Weight: 0,075 kg Price: 453,02 e

Table 4.1: Speciﬁcations of WINTER 7 FMS 5 Aispeed Indicator

2. Altimeter Instrument used to measure the altitude of an object above a ﬁxed level. Altitude can be determined based on the measurement of atmospheric pressure. It can be measure in QNH (Q code indicating the atmospheric pressure adjusted to mean sea level) and QFE (Indicating the height from the level of the airport or aerodrome).

Figure 4.2: WINTER EBH Altimeter

RA 98 G06-AlOn LSA 3 seats | Project report WINTER EBH Altimeter Dimensions: 57 ∅ x 57 D [mm] Weight: 0,105 kg Price: 377,52 e

Table 4.2: Speciﬁcations of WINTER EBH Altimeter

3. Variometer Flight instrument used to inform the pilot of the rate of descent or climb. Variometers measure the rate of change of altitude by detecting the change in air pressure (static pressure) as altitude changes.

Figure 4.3: WINTER Vanetype Variometer 5 STV 5

WINTER Vanetype Variometer 5 STV 5 Dimensions: 63 Ø x 290 D [mm] Weight: 0,176kg Price: 530,00 e

Table 4.3: Speciﬁcations of WINTER Vanetype Variometer 5 STV 5

4. Artificial horizon The artificial horizon is a flight instrument that informs the pilot of the aircraft orientation relative to Earth’s horizon providing the pilot an immediate indication of the smallest orientation change.

Figure 4.4: Kelly MFG RCA22 Artiﬁcial Horizon

RA 99 G06-AlOn LSA 3 seats | Project report Kelly MFG RCA22 Artiﬁcial Horizon Dimensions: 85,72 W x 85,72 L x 62,5 D [mm] Weight: 1,247kg Voltage: 14-24 DC Current: 0,62 A Price: 1033,58 e

Table 4.4: Speciﬁcations of Kelly MFG RCA22 Artiﬁcial Horizon

5. Compass / Directional gyro A compass is an instrument used for navigation and orientation. It shows direction relative to the geographic cardinal directions. While, the heading indicator or directional gyro is a ﬂight instrument used in an aircraft to inform the pilot of the aircraft’s heading.

Figure 4.5: Kelly RCA15BK-1 Directiornal Gyro

Kelly RCA15BK-1 Directiornal Gyro Dimensions: 80 L x 80 W x 170 D [mm] Weight: 1,043 kg Voltage: 14-24 DC Current: 1,7 A Price: 969,06 e

Table 4.5: Speciﬁcations of Kelly RCA15BK-1 Directiornal Gyro

6. Rate of turn indicator and slip-skid indicator The turn indicator indicates the rate of turn or the rate of change in the aircraft’s heading.

The slip-skid indicator is an inclinometer that at rest displays the angle of the aircraft’s transverse axis with respect to horizontal, and in motion displays this angle as modiﬁed by the acceleration of the aircraft.

RA 100 G06-AlOn LSA 3 seats | Project report

Figure 4.6: WINTER Turn Coordinator QM II

WINTER Turn Coordinator QM II Dimensions: 65 W x 36 L x 13 D [mm] Weight: 240 g Price: 56,00 e

Table 4.6: Speciﬁcations of WINTER Turn Coordinator QM II

7. Clock A clock is an instrument used to measure, keep and indicate time.

Figure 4.7: WINTER Hours Counter Analog

WINTER Hours Counter Analog Dimensions: 60mm W x 60mm H x 108mm D Weight: 0,150 kg Voltage: 12-24 V DC Current: 20 mA Price: 373,00 e

Table 4.7: Speciﬁcations of WINTER Hours Counting Analog

RA 101 G06-AlOn LSA 3 seats | Project report 4.1.2 Electronic Flight and Navigation Instruments The aircraft will provide to the pilot two displays of each instrument in case one of them fails. One display is shown in a screen and the other will be the reading in the instrument directly (explained in 4.1.1 Analogic Flight and Navigation Instruments).

For the digital display, all components are from Dynon Avionics.

1. SV-D600/B 7" SkyView SE Display, with Harness (for Simple VFR) Dynon’s SkyView SE system offers simpliﬁed experience for pilots of simple VFR aircraft, it provides the most intuitive ﬂight and engine instruments on the market.

Figure 4.8: SV-D600/B 7" SkyView SE Display

SV-D600/B 7" SkyView SE Display Dimensions: 194 W x 140 H x 60 D [mm] Weight: 1,1 kg Power: 2,1A - 12V DC Resolution: High Resol. 480 H x 800 W pixels Price: 1850,00 US dollars

Table 4.8: Speciﬁcations of SV-D600/B 7" SkyView SE Display

The price includes a harness to connect the SV-ADAHRS-200 with the SV-D600/B 7" SkyView SE Display.

Figure 4.9: Dynon Avionics Skyview Display Harness

2. SV-ADAHRS-200 for Air Data, Attitude and Heading ADAHRS provides data for Air, Attitude and Heading Reference System, the data necessary to drive the Artiﬁcial Horizon/Synthetic Vision, Airspeed, Altitude, Vertical Speed, Slip, Turn Rate, Angle of Attack, and Gyro-Stabilized Heading. Additionally, Outside Air Temperature is obtained from the SV-OAT-340 Sensor which is connected to the ADHARS Module.

RA 102 G06-AlOn LSA 3 seats | Project report In the price there are included: SV-ADAHRS-200 SV-OAT-340, 3 Pneumatic ports for Pitot, Static, and Angle of Attack, 9-Pin DIN Connector for the SkyView Network, 2-Pin Connector for SV-OAT-340 Sensor.

Figure 4.10: SV-ADAHRS-200 Air Data/Attitude/Heading Reference System

SV-ADAHRS-200 Air Data Dimensions: 120 W x 31 H x 66 D [mm] Weight: 0,2 kg Power: 0,15A - 12V DC Price: 1.200,00 US dollars

Table 4.9: Speciﬁcations of SV-ADAHRS-200 Air Data

The ADAHRS internal sensors measure rotational motion and acceleration about the center of gravity (CG) of the aircraft. In order to maintain accuracy, the ADAHRS should be mounted within six feet laterally (side-to-side) and twelve feet longitudinally (front-to-back) of the aircraft CG. Quite often the best place to mount the ADAHRS is in the fuselage behind the cockpit area. But other locations can work (see Fig. 4.11). Furthemore, the ADAHRS Module includes a Magnetometer (Remote Compass). This means that it needs to be located in magnetically benign location, away from high current wires and moving aircraft components that contain a ferrous material.

Figure 4.11: Location of SV-ADAHRS-200 Air Data/Attitude/Heading Reference System

3. SV-BAT-320 SkyView Backup Battery SV-BAT-320 provides at least an hour of power to a single SkyView display and all connected SkyView Network Modules including the ADAHRS and EMS.

RA 103 G06-AlOn LSA 3 seats | Project report

Figure 4.12: SV-BAT-320 SkyView Backup Battery

SV-BAT-320 SkyView Backup Battery Dimensions: 84 W x 53 H x 99 D [mm] Weight: 0,37kg Power: 1,0 A - 12V DC Price: 590,00 US dollars

Table 4.10: Speciﬁcations of SV-BAT-320 SkyView Backup Battery

4. SV-GPS-2020 GPS Receiver/Antenna GPS latitude and longitude data is necessary for the SkyView to locate its position to display the Moving Map and to calibrate the Remote Compass or Magnetometer.

Figure 4.13: SV-GPS-2020 GPS Receiver/Antenna

SV-GPS-2020 GPS Receiver/Antenna Dimensions: 56 W x 19 H x 87 D [mm] Weight: 0,2 kg Update Rate: 4 Hz Price: 200,00 US dollars

Table 4.11: Speciﬁcations of SV-GPS-2020 GPS Receiver/Antenna

When combined with Dynon’s SV-XPNDR-261 Mode-S transponder, SkyView and Ad- vanced Flight Systems, accomplised all ADS-B Out requirements.

5. SV-XPNDR-261 Mode-S Transponder The SV-XPNDR-261 is a Class 1 transponder. For US aircraft, the FAA 2020 ADS-B Out mandate requires a Class 1 transponder.

RA 104 G06-AlOn LSA 3 seats | Project report A transponder is an electronic device that produces a response when it receives a radio- frequency interrogation. Aircraft have transponders to assist in identifying them on air trafﬁc control radar.

Figure 4.14: SV-XPNDR-261 Mode-S Transponder

SV-XPNDR-261 Mode-S Transponder Dimensions: 62 W x 45 H x 141 D [mm] Weight: 400 g Supply Voltage: 9 - 33 V Power Consumption: 6 Watts (typical) [idle: 0.15 A, active: 0.34 A at14 V] Operating Temperature: -40ªC to + 70ºC Cooling Requirement: No fan required Certiﬁcation: ETSO C88a, 2C112b, C166 and TSO C88b, C112c, C166a, approved for IFR and VFR ﬂight Price: 2200,00 US dollars

Table 4.12: Speciﬁcations of SV-XPNDR-261 Mode-S Transponder

6. SV-KNOB-PANEL/H SkyView Knob Control Panel (OPTIONAL) The SV-KNOB-PANEL offers comfort with three knobs for the most used functions: Altimeter setting (BARO), Heading bug and the Altitude bug. The SV-KNOB-PANEL installs easily with simple plug-n-play integration to the SkyView network - no additional aircraft wiring needed.

Figure 4.15: SV-KNOB-PANEL/H SkyView Knob Control Panel

RA 105 G06-AlOn LSA 3 seats | Project report SV-KNOB-PANEL/H SkyView Knob Control Panel Dimensions: 89.7 W x 45.7 L x 32.4 D [mm] Weight: 0,18 kg Power Input: Via SkyView Network Price: 250,00 US dollars

Table 4.13: Speciﬁcations of SV-KNOB-PANEL/H SkyView Knob Control Panel

7. SV-COM-C25/H SkyView VHF Com Radio 25 kHz The SkyView COM Radio tunes frequencies by airport and station type at the touch of a button. The radio can send frequencies directly from SkyView. There are two versions available: vertical and horizontal.

Figure 4.16: SV-COM-C25/H SkyView VHF Com Radio 25 kHz

SV-COM-C25/H SkyView VHF Com Radio 25 kHz Control panel: Dimensions: 89,7 W x 45,7 H x 32,4 D [mm] Weight: 159 g SV-COM-C25 Transceiver Module: Dimensions: 186,6 L x 60,45 W x 38,23 H [mm] Weight: 390 g Frequencies: 118.000 to 136.975 MHz Spacing between frequencies: 25 kHz Price: 1295,00 US dollars

Table 4.14: Speciﬁcations of SV-COM-C25/H SkyView VHF Com Radio 25 kHz

There is the possibility to add a new Radio with a more precise spacing of 8,33 kHz. SV-COM-X83 costs 2195.00 US dollars. There is the possibility of even substitute the standard radio with a spacing of 25 kHz (SV-COM-C25/H SkyView VHF Com Radio 25 kHz) to one more precise of 8.33 kHz (SV-COM-C25/H SkyView VHF Com Radio 8.33 kHz).

8. SV-INTERCOM-2S Two-Place Stereo Intercom In order to listen the messages and communications that the Radio collects, a Stereo is needed.

RA 106 G06-AlOn LSA 3 seats | Project report SV-INTERCOM-2S Two-Place Stereo Intercom includes Horizontal and Vertical face- plates, Stereo Headset Panel Jacks and Pilot and Copilot Jack set.

Figure 4.17: SV-INTERCOM-2S Two-Place Stereo Intercom

SV-INTERCOM-2S Two-Place Stereo Intercom Dimensions: 89,71 W x 45.72 L x 106,17 D [mm] Weight: 204 g Input Voltage: 10-30V DC Power Usage: 0,1A at 14V Muting Inputs: - 1x stereo, differential, noise-rejecting music input, with panel-mounted music jack override. - 1x mono input for additional muting source. Non-muting Inputs: - 1x stereo, differential, noise-rejecting EFIS input. - 3x mono inputs for radios and other avionics. One of these is fail-safed to the pilot headset. Headsets: - 2x stereo headsets supported. Radio Outputs: - 2x com radio outputs. External PTT selector required. Price: 1295,00 US dollars

Table 4.15: Speciﬁcations of SV-INTERCOM-2S Two-Place Stereo Intercom

In order to connect the SV-COM-C25/H SkyView VHF Com Radio 25 kHz or 8.55 kHz with the SV-INTERCOM-2S Two-Place Stereo Intercom, a cable is needed. SV-NET-3CC Network Cable costs 40 US dollars.

Figure 4.18: SV-NET-3CC Network Cable

RA 107 G06-AlOn LSA 3 seats | Project report 9. AOA - Pitot Probes Pitot probe is a flow measurement device used to measure fluid flow velocity. It is used to determine the airspeed of an aircraft. The pitot probe is unheated.

Figure 4.19: AOA - Pitot Probes

AOA - Pitot Probes Dimensions: 75 W x 790 L x 250 D [mm] Weight: 0,18 kg Price: 200,00 US dollars

Table 4.16: Speciﬁcations of AOA - Pitot Probes

The Pitot Probe has to be located under the wing, it has L-shape and requires a separate mounting bracket (AOA/Pitot Mount Bracket). The mounting bracket costs 95.00 US dollars.

Figure 4.20: AOA - Pitot Mount Bracket

RA 108 G06-AlOn LSA 3 seats | Project report 4.2 Powerplant Instruments

Powerplant instruments are the instruments in the cockpit of an aircraft that provide the pilot with information about fuel and oil temperatures and pressures, revolutions of the engine, voltage of the system, amount of fuel of the tank and so forth.

The section 8. Required Equipment of [4] establishes for Powerplant instruments:

• Fuel quantity indicator

• Tachometer (RPM)

• Engine “kill” switch

• Engine instruments as required by the engine manufacturer.

The aircraft motor is IAE50R – AA from AustroEngine. This model has integrated an elec- tronically controlled (dual) ignition and injection system.

The instruments featured below are selected carefully to read the parameters in the speciﬁc range of the engine works.

1. Fuel quantity indicator Instrument that displays the amount of fuel left in the tank. First of all, the fuel tanks are selected. The fuel tanks will have to host a mininum 50 litres of fuel. A disposal with the fuel tanks in the wings has been chosen, this disposal is favorable for ideal position of the centre of gravity than a fuel tank behind or in front of the cockpit area.

Figure 4.22: J-3 8-Gal. LH Wing Tank

The speciﬁcations for the left or right wing fuel tank are: The wing fuel tanks ﬁt perfectly in the wing, taking into account that the root chord is 1 meter, the thickness of the airfoil is 15 (per cent) of the chord, and the semi-span is 7 meters. The two wing fuel tanks can host 60.57 litres of fuel. They satisfy the demand of the minimum of litres of fuel.

RA 109 G06-AlOn LSA 3 seats | Project report J-3 8-Gal. Wing Tank, FAA/PMA’d Dimensions: 9,4 L x 26,6 W x 120 D [cm] Volume: 30,28 l (8 gal) Weight: 4,309 kg Price: 544,00 US dollars Mounting plate G-749-100 Price: 13,50 US dollars 1/4" Brass Ball Shutoff Valves Price: 94,50 US dollars Fuel Strainer G-775-000 Price: 6,75 US dollars Total price: 658,75 US dollars

Table 4.17: Speciﬁcations of J-3 8-Gal. LH Wing Tank, FAA/PMA’d and Compulsory Extra Elements

For measuring the quantity of fuel left in the tank, it has been chosen: Westach Dual Level Fuel Level Gauge Model: 2DA4.

Figure 4.23: Westach Dual Level Fuel Level Gauge Model: 2DA4

Westach Dual Level Fuel Level Gauge Model: 2DA4 Dimensions: 57 L x 57 W x 52 D [mm] Weight: 250 g Supply Voltage: DC 8-30 V Power Current: 0,1 A Max. Price: 133,00 US dollars

Table 4.18: Speciﬁcations of Westach Dual Level Fuel Level Gauge Model: 2DA4

RA 110 G06-AlOn LSA 3 seats | Project report 2. Tachometer (RPM) Instrument used to measure the rotation speed of the motor. A tachometer displays the revolutions per minute (RPM) of the engine. The maximum speed of the engine IAE50R - AA is 8000 rpm.

• Maximum Continuous Speed: 7100 rpm • Minimum Speed: 2500 rpm • Engine Speed (Take-off): 7750 rpm, max. 3 min. • Maximum Engine Over-Speed (20 sec.): (20 sec.): 8000 rpm

The tachometer chosen is able to display starting at 0 up to 8000 rpm.

Figure 4.24: Aviasport ROTAX 503-582 Ducati Ignition Tachometer

Aviasport ROTAX 503-582 Ducati Ignition Tachometer Dimensions: 63 Ø x 50 D [mm] Range: 0-8000rpm Weight: 185 g Price: 125,75 US dollars

Table 4.19: Speciﬁcations of Aviasport ROTAX 503-582 Ducati Ignition Tachometer

3. Temperature readings In [29], temperature limits for IAE50R-AA engine are:

Temperature limits for IAE50R-AA Cooling Fluid, take off, min.: 60 oC Cooling Fluid, take off, max.: 90 oC Cooling Fluid, continued operation, max.: 100 oC Exhaust Gas Temperature, max.: 970 oC Fan Cooling Air at Cooling Air Exhaust, max.: 120 oC, max. 3 min. 110 oC Ambient Temperature for Starting, min.: -10 oC Ambient Temperature, continued operation, max.: 55 oC

Table 4.20: Temperature limits for the engine IAE50R-AA

For the reading of Cooling Fluid temperature (taking into account that Cooling Fluid Temp. min.: 60 oC and Cooling Fluid Temp. max.: 90 oC ):

RA 111 G06-AlOn LSA 3 seats | Project report

Figure 4.25: IM-510 Thermometer Coolant for ROTAX 912/914

IM-510 Thermometer Coolant for ROTAX 912/914 Dimensions: 56 w x 56 L x 52 D [mm] Weight: 200g Power Supply: DC 10-30V Power Current: 0,1 A Max. Scale: 50-150º C Subdivision scale: 1ºC Price: 101,5 US dollars

Table 4.21: IM-510 Thermometer Coolant for ROTAX 912/914

In Table Table 4.20, the maximum Exhaust Gas Temparature is 970oC, therefore for the instrument selected for reading of Exhaust Gas temperature is:

Figure 4.26: IM-554 EGT Gauge for Rotax 914

IM-554 EGT Gauge for ROTAX 914 Dimensions: 80 w x 80 L x 75 D [mm] Weight: 200g Power Supply: DC 10-30V Power Current: 0,1 A Max. Scale: 300-1000 ºC Subdivision scale: 10ºC Price: 185,00 e

Table 4.22: Speciﬁcations of IM-554 EGT Gauge for ROTAX 914

RA 112 G06-AlOn LSA 3 seats | Project report Finally, for the reading of the temperature of oil:

Figure 4.27: IM-560 Oil Temperature Indicator for Rotax 912S

IM-554 EGT Gauge for ROTAX 914 Dimensions: 56 Ø x 52 D [mm] Weight: 200g Power Supply: DC 8-35V Power Current: 0,1 A Max. Scale: 50-150 ºC Subdivision scale: 1ºC (central zone) Price: 70,00 e

Table 4.23: Speciﬁcations of IM-560 Oil Temperature Indicator for Rotax 912S

4. Pressure Limits: In Table Table 4.24 there are included the pressures of the engine recommended by the engine supplier. This data is taken from [29] and [30].

Pressure limits for IAE50R-AA Fuel Pressure – Injection System: 3,0 bar ± 0,2 bar Fuel Pressure at Carburettor Inlet: 0,276 - 0,414 bar

Table 4.24: Pressure limits for the engine IAE50R-AA

In Section 6.3.6 - Fuel Pressure of [30], the nominal fuel pressure is 3 ± 0.2 bar (43.5 ± 2.9 psi). A Powerplant instrument to measure the fuel pressure is needed.

Figure 4.28: IM-584 Fuel Pressure Gauge for Rotax 912 BAR

IM-554 Fuel Pressure Indicator for Rotax 912 BAR measures only the tolererance of fuel pressure, if the needle is in the left red zone, the pressure is lower than 2.8 bar; if the needle is on the right, the pressure is higher than 3.2 bar.

RA 113 G06-AlOn LSA 3 seats | Project report IM-584 Fuel Pressure Gauge for Rotax 912 BAR Dimensions: 56 Ø x 52 D [mm] Weight: 200g Power Supply: DC 10-30V Power Current: 0,1 A Max Range: 0 a 0,6 Bar Subdivision scale: 0,01 Bar Price: 72,00 e

Table 4.25: Speciﬁcations of IM-584 Fuel Pressure Gauge for Rotax 912 BAR

As in the reading of fuel pressure, for oil pressure the gauge IM-543 Oil Pressure Gauge for Rotax 912/914 BAR only measures the relative pressure:

Figure 4.29: IM-543 Oil Pressure Gauge for Rotax 912/914 BAR

IM-543 Oil Pressure Gauge for Rotax 912/914 BAR Dimensions: 56 Ø x 52 D [mm] Weight: 220g Power Supply: DC 10-30V Power Current: 0,1 A Max Range: 0 - 8 Bar Subdivision scale: 0,02 Bar Price: 70,00 e

Table 4.26: Speciﬁcations of IM-543 Oil Pressure Gauge for Rotax 912/914 BAR

5. Voltimeter Most of the instruments can work in a range of 10-14 V; even some of them can work in bigger range between 10-30 V in DC. The work tension selected is 14 V, this tension suits perfectly for every instrument op- erational range of power supply. Therefore, the Voltimeter IM-561 has the following speciﬁcations:

RA 114 G06-AlOn LSA 3 seats | Project report

Figure 4.30: IM-561 Voltimeter

IM-561 Voltimeter Dimensions: 52 Ø x 52 D [mm] Weight: 220g Range: 8 - 16 V Subdivision scale: 0,1 V Price: 71,00 e

Table 4.27: Speciﬁcations of IM-561 Voltimeter

6. Engine electrical components IAE50R – AA engine has an Electrical Control Unit (ECU) and a starter that need a power supply to run. In [29],there is speciﬁed that the starter is an electric starter which operates with 12 V and 50 A. It is of a Bendix Type, engaging the gear when operated. While in [30], it is established that the ECU needs 14.8 V and 2 A to work properly for a condition of 3000 rpm.

In Table 4.38 all Flight and Navigation, Powerplant and other instruments are recopilated in a table.

The total price of Flight, Navigation and Powerplant instruments is:

12,101.75 e(13,791.09 US dollars) 14,244.61 e(16,237.29 US dollars) with optionals

RA 115 G06-AlOn LSA 3 seats | Project report ------40,00 95,00 72,00 70,00 70,00 71,00 590,00 200,00 295.00 200,00 658,75 133,00 101,50 185,00 1850,00 1200,00 2200,00 1295,00 (250,00) 9.326,25 (2195,00) (11.771,25) Price(US dollars) ) e ------56,00 453,02 377,52 969,06 373,00 125,75 1033,58 - 530,00 3.917,93 Price( (3.917,93) ------6 1 1 12 14 1,8 1.4 1,4 1,4 1.4 1,4 600 8,68 23,8 0,24 25,2 29,6 730,32 (730,32) Power(W) ------14 14 12 12 12 12 14 14 14 14 10 10 12 0,1 0,1 0,1 14,8 Voltage(v) ------1 1 2 14 50 14 14 1,7 2,1 0.1 0,1 0,1 0,1 0,62 0,02 0,15 0,34 Intensity (A) - - - - 1,1 0.2 0,2 0,4 0,37 0,18 0,075 0,105 0,176 1,247 1,043 0,240 0,150 0,549 0.204 8,618 0,250 0,185 0,200 0,200 0,200 0,220 0,200 0,220 (0,18) 16,352 (0,549) (17,081) Weight(kg) Powerplant Instruments Total of Flight-Navigation and Powerplant Instruments Analogic Flight and Navigation Instruments Electronic Flight and Navigation Instruments Table 4.28: Total: Instruments IM-561 Voltimeter Total /w optionals: AOA - Pitot Probes WINTER EBH Altimeter AOA - Pitot Mount Bracket IAE50R – AA Engine Starter SV-NET-3CC Network Cable WINTER Hours Counter Analog WINTER Turn Coordinator QM II SV-D600/B 7" SkyView SE Display IM-554 EGT Gauge for ROTAX 914 Electric Control Unit of IAE50R-AA Kelly RCA15BK-1 Directiornal Gyro WINTER 7 FMS 5 Aispeed Indicator SV-GPS-2020 GPS Receiver/Antenna Kelly MFG RCA22 Artiﬁcial Horizon SV-BAT-320 SkyView Backup Battery SV-XPNDR-261 Mode-S Transponder J-3 8-Gal. Wing Tank + Extra Elements WINTER Vanetype Variometer 5 STV 5 IM-560 Oil Temp Indicator for ROTAX 912S Westach Dual Fuel Level Gauge Model: 2DA4 SV-INTERCOM-2S Two-Place Stereo Intercom IM-584 Fuel Pressure Gauge for Rotax 912 BAR SV-COM-X83 SkyView VHF Com Radio 8,33 kHz IM-510 Thermometer Coolant for ROTAX 912/914 IM-543 Oil Pressure Gauge for Rotax 912/914 BAR SV-COM-C25/H SkyView VHF Com Radio 25 kHz SV-KNOB-PANEL/H SkyView Knob Control Panel Aviasport ROTAX 503-582 Ducati Ignition Tachometer SV-ADAHRS-200 Air Data/Attitude/Heading Ref. System

RA 116 G06-AlOn LSA 3 seats | Project report 4.3 Miscellaneous Equipment

4.3.1 Aircraft Lights As speciﬁed in 91.209 Aircraft lights of [28], no person may operate an aircraft unless it has position lights during the period from sunset to sunrise; no person may park or move an aircraft in the area of an airport during night unless the aircraft is clearly illuminated, has lighted position lights and it is an area that is marked by obstruction lights, no person may anchor an aircraft unless the aircraft has lighted anchor lights.

For the signalling of the wings, a pair of strobe lights from Aveo Engineeting Group S.R.O. has been chosen. Aveo Ultra Galactica™ Series Lights are TSO certiﬁed by the EASA.

Figure 4.32: Disposal of Ultra Galactica™ Series lights

Aveo Ultra Galactica™ Series lights Dimensions: 100,3 L x 46 W x 34 D [mm] Weight: 127g Operating Voltage Range: 9 – 36 VDC Input Current – Position - green: 0,25A @36V / 0.3A @28V / 0.62A @14V / 0.9A @9V Input Current – Position - red: 0,2A @36V / 0,25A @28V / 0,48A @14V / 0,8A @9V Input Current – Strobe: 1,1A @36V / 1,4A @28V / 2,3A @14V / 2,9A @9V Input Power - Position - green: 9 W @36V / 8,4 W @28V / 8,7 W @14V / 8,1 W @9V Input Power - Position - red: 7 W @36V / 7 W @28V / 6,7 W @14V / 7,2 W @9V Input Power - Strobe: 39,6 W @36V / 39,2 W @28V / 32,2 W @14V / 26,1 W @9V Input Power - Strobe: 39,6 W @36V / 39,2 W @28V / 32,2 W @14V / 26,1 W @9V Repetition Rate of Strobe: 50 cycles per minute Ambient Temperature: from -55oC to +8oC Price: 1.199,00 e

Table 4.29: Speciﬁcations of Aveo Ultra Galactica™ Series lights

Aveo Ultra Galactica™ Series lights can work with a voltage of 14 V DC and the power at this voltage is 9 W + 32,2 W (for the green one) and 7 W + 32,2 W (for the red one). The Input Power total is 80,4 W.

Finally, the landing, taxi and wigwag lights:

RA 117 G06-AlOn LSA 3 seats | Project report

Figure 4.33: Aveo Hercules 30 Landing/Taxi/Wigwag Light Module

Aveo Hercules 30 Landing/Taxi/Wigwag Light Module Dimensions: 110 Ø x 45,7 D [mm] Material: Housing/Heatsink: Aluminum Alloy, natural anodizing Lens: Clear PMMA Weight: less than 490 g with Aveo bracket Number of LED: 30 Input voltage: DC 9-36 V Current: Taxi -> max. 4,2A (12V ) Landing -> max. 6,3A (12V ) Landing + Taxi -> max. 10,5A (12V ) Output power [Hi/Low]: Taxi -> 45/21W Landing -> 68/33W (12V ) Landing + Taxi -> 113W/54W Ambient temperature: from -55oC to +85oC Useful life: not less than 30.000 aircraft ﬂight hours Price: 625,00 e

Table 4.30: Speciﬁcations of Aveo Hercules 30 Landing/Taxi/Wigwag Light Module

4.3.2 Batteries In Tables 4.32 and 4.33, OWA (Ordered Weight Average) method has been used to select the best battery considering seven different models that cover the basic speciﬁcations but have different characteristics. The factors have been assigned a weight value in order to value more some speciﬁcations than others such as the weight in g and the power pulse. Each category is grades from 1 to 5 being 1 the lowest grade and 5 the highest.

With the OWA method the better batteries are Aerovoltz 16 and the Anigravity 16 having not a signiﬁcative difference between both.

While in Tables 4.34 and 4.35, PRESS method has been used.

RA 118 G06-AlOn LSA 3 seats | Project report The obtained results with the PRESS method are that the better batteries are Aerovoltz 16 and the Anigravity 16 and as the OWA results, there is not a signiﬁcative difference so between the two options the chosen one is Aerovoltz 16 cell because the is more obtained information and the Power Pulse is higher.

Figure 4.34: Aerovoltz 16 cell battery

Aerovoltz 16 cell battery Dimensions: 114 L x 114 W x 112 H [mm] Negative Terminal Location: Left Weight: 1611 grams Voltage (Charged): 13,6V Lead Acid Equivalent Amperage: 28 A/H @ 10 A/H Rating Pulse Cranking Amps: 500 amps Operating Environment: -18oC to 60oC Price: 333,95 US dollars

Table 4.31: Speciﬁcations of Aerovoltz 16 cell battery

RA 119 G06-AlOn LSA 3 seats | Project report 9 35 450 5400 0,575 102,75 11,377 31,164 19,991 8164,66 106,534 3686315 Powersonic Powersonic 480 1450 6528 0,824 260,99 21,353 40,131 23,272 34,286 33,4657 152,509 1258446 Anigravity 16 Anigravity 16 5 7 35 45 500 6315 0,568 14000 13,206 7257,47 105,206 5238617 MCI True Blue MCI True Blue 7 9 25 45 400 135 1620 0,652 147,95 34,796 370800 120,796 Aerovoltz 4 Aerovoltz 4 768 275 3740 0,725 188,75 23,323 43,293 17,739 15,164 34,612 778848 134,134 Aerovoltz 8 Aerovoltz 8 410 1106 5576 0,792 249,95 21,959 41,726 28,095 20,503 34,336 146,622 1110816 Aerovoltz 12 Aerovoltz 12 Table 1 for choosing battery using OWA method Table 2 for choosing battery using OWA method 35 Table 4.32: Table 4.33: 500 1611 6800 0,826 333,95 20,543 39,385 24,063 33,957 152,949 1455552 Aerovoltz 16 cell Aerovoltz 16 cell 5 9 7 9 7 5 9 7 9 7 37 Weight category Weight category ] ]: 3 3 mm mm Sum: OWA: Name Name Weight [g]: Weight [g]: Power Pulse: Power Pulse: Volume [ Volume [ Pulse Amperage: Pulse Amperage: Price (US dollars): Price (US dollars):

RA 120 G06-AlOn LSA 3 seats | Project report 7 7 1 5 0,189 4,851 4,905 4,944 4,970 4,898 1,071 0,926 0,839 0,862 1,187 0,999 Price (US dollars) 0,257 0,237 0,250 0,268 0,179 0,252 1,443 0,352 9 9 1 5 5 0,243 2,673 2,278 1,684 2,585 0 0,013 0,012 0,030 0,049 0,126 0,233 4,294 Power Pulse 0 0,371 0,394 0,411 0,430 0,381 2,173 0,546 7 7 5 1 5 5 0 0,189 4,013 2,534 4,780 0,233 0,176 0,091 0,346 0,220 1,258 0,685 Pulse Amperage 0 Grades 0,141 0,085 0,019 0,255 0,129 0,730 1,150 9 9 5 5 0 0,243 4,376 4,636 4,810 1,467 4,459 0,057 0,017 0,037 0,170 0,006 0,306 3,024 Weight [g] ] 0 Table 1 for choosing battery using PRESS method Table 2 for choosing battery using PRESS method 3 0,022 0,040 0,059 0,113 0,010 0,250 4,284 mm 5 5 5 1 5 0,135 4,108 4,391 4,664 4,270 Table 4.34: Table 4.35: Volume [ PRESS: Aerovoltz 8: Aerovoltz 4: Aerovoltz 12: Anigravity 16: MCI True Blue: Aerovoltz 16 cell: Name P MAX: Weight [g]: Aerovoltz 8: Aerovoltz 4: Aerovoltz 12: Anigravity 16: MCI True Blue: Weight average: Weight category: Aerovoltz 16 cell:

RA 121 G06-AlOn LSA 3 seats | Project report 4.3.3 Fire extinguishers As specified in [31] the onboard systems designed to extinguish fires which occur either in the air or on the ground. In LSA, at least one fire extinguisher must be installed.

Fire can occur due to several factors: energized electrical equipment; combustibles which fires in cloth, paper, rubber, and many plastics, fires in flammable liquids, oils... The fire extinguishers have to be portable to extinguish the fire in any place where the fire is produced.

Halon 1211 extinguishers are the most used nowadays. Halons are ﬁre extinguishing agents which are gaseous when discharged in the aircraft environment. Halon 1211 is made of Bro- moChlorodiFluoromethane (CBrClF2) also known as ’BCF’.

Figure 4.35: H3R Aviation Model A344T - Halon 1211 Fire Extinguisher

H3R Aviation Model A344T - Halon 1211 Fire Extinguisher Agent Weight: 0,567 kg Gross Weight: 1,021 kg Discharge range: 2,7-3,7 m Discharge time: 10 sec Height: 25,7 cm Width: 9,7 cm Cylinder diameter: 6,6 cm Price: 173,75 US dollars

Table 4.36: Speciﬁcations of H3R Aviation Model A344T - Halon 1211 Fire Extinguisher

The aircraft will dispose of two H3R Aviation Model A344T - Halon 1211 Fire Extinguishers. The total cost of them is 347.5 US dollars, and the total weight is 2.042 kg. Model A344T is the smallest Halon 1211 ﬁre extinguisher in H3R Aviation. It is recommended for a 1-4 person aircraft, including the pilot, and it is also included Inspection and Maintenance Requirements Manual.

4.3.4 Aircraft’s parachute The purpose of the parachute system is to descent safely an entire light aircraft to the ground in the event of loss of control, failure of the aircraft structure, or other in-ﬂight emergencies.

In [32], ASTM settles the requirements of the parachute system. The Cirrus Airframe Parachute System (CAPS) is a whole-plane ballistic parachute recovery system designed. The design is

RA 122 G06-AlOn LSA 3 seats | Project report certiﬁed by the FAA.

Even though, an aircraft’s parachute is not compulsory. For Light Sport Aircrafts, a BRS parachutes could save the lives of the passengers in a possible incident in mid air. BRS parachutes have saved 395 lives up to now.

The aircraft should have installed a BRS parachute.

Figure 4.37: BRS Ballistic Parachute System for LSA

Figure 4.38: BRS 7 LSA Canister Parachute System

BRS 7 LSA Canister Parachute System Weight: 13 kg Maximum Weight: and Deployment Speed: 603,73kg at 222 km/h in 10 sec Dimensions: 8 Ø x 12,5 D cm Repack Cycle: 6 yrs Price: 4872,00 US dollars

Table 4.37: Speciﬁcations of BRS 7 LSA Canister Parachute System

4.3.5 Flight Recoder Flight Data Recorder (FDR) device is used to record speciﬁc aircraft performance parameters. The purpose of an FDR is to collect and record data from a variety of aircraft sensors onto a medium designed to survive an accident.

RA 123 G06-AlOn LSA 3 seats | Project report In [link Faa regulations], it is settled that Cockpit Voice Recorders are required for:

• Large 4-engined turbine-powered planes

• Large 4-engined pressurized planes

• Multiengined turbine-powered planes seating 10+ passengers

• Multiengined turbine-powered rotorcraft seating 20+ passengers

• Multiengined turbine-powered planes and rotorcraft seating 6+ passengers and requiring two pilots

While, Flight Data Recorders are required for:

• Multiengine turbine-powered airplanes or rotorcraft seating 10+ passengers

• Multiengine turbine-powered airplanes or rotorcraft seating 6+ passengers and requiring two pilots

• Large airplanes meant for use either over 25,000 feet

• Large turbine-powered airplanes

• Turbine-powered transport category planes

Nevertheless, the LSA aicraft does not satisfy the previous requirements for Cockpit Voice Recorders neither for Flight Data Recorders. Because the aircraft is a mono engine airplane for 3 passengers, only one pilot is needed, their dimensions are not large, and the airplane is not meant to ﬂy over 25,000 feet obviously.

Therefore, a Flight Data Recorder is not compulsory needed.

RA 124 G06-AlOn LSA 3 seats | Project report 4.4 Safety Belts and Harnesses

As established in [4] there must be a seat belt and harness for each occupant and adequate means to restrain the baggage.

Due to their years experience and motorsport safety for the harness of each occupant the ones from the brand Simpson have been chosen.

Figure 4.39: Simpson 5-point Harness System

Lightweight aluminum adjusters are the smoothest in the industry and provide exceptional locking action. Simpson harnesses are premium polyester webbing only. Polyester minimizes elongation and sunlight degradation. Polyester also absorbs energy better in a crash. Simp- son aviation harnesses are engineered with proven racing technology to offer you top of the line performance and safety.

The Simpson 5-point Harness System costs is 320.00 US dollars. Therefore, the price of 3 harnesses is 960.00 US dollars.

Figure 4.40: Poliester straps to restrain the baggage

A pair of straps are needed to restrain and ﬁx the baggage. The ones that have been chosen are from Leroy Merlin and are made of poliester with 25 mm wide and 2.5 m long. The cost of the strap is 9.95 e. Therefore, three straps to ensure the ﬁxation of the baggage will be enough.

RA 125 G06-AlOn LSA 3 seats | Project report In Table 4.38, a recopilation with weights and prices of Safety Systems, Aircraft Lights and Controls can be found.

Instruments Weight (kg) Price (e) Price (US dollars) Safety Systems Simpson 5-point Harnesses System - - 960.00 H3R Aviation Model A344T - Halon 1211 Fire Ext. 2,042 - 374,50 Poliester Straps - 29,85 - BRS 7 LSA Canister Parachute System 13 - 4872,00 Aircrafs Lights Aveo Ultra Galactica™ Series lights 0,127 1.199,00 - Aveo Hercules 30 Landing Light Module 0,490 625,00 - Aerovoltz 16 cell battery 1,611 - 333,95 Total cost: 17,27 7.443,84 8.485,98

Table 4.38: Total costs of Safety and Miscellanous Components and Controls

4.5 Installation

Each instrument and system mentioned in 5.Systems and Avionics has to be connected correctly, has to be located in a speciﬁc position in the cockpit area or has to follow some requirements settled in [4].

One of the main priorities for the selection of the equipment was to have its own manual easy to ﬁnd, so the steps to do a correct installation and maintenance can be found.

On other hand, most of the instruments have been chosen wisely from the same company or brand to ensure the compatibility between them, and to maintain the aesthetics in the cockpit. A design of the cockipit with the instruments mentioned before can be found in (6.1.2.3 In- struments and equipment).

Another part of the installation is the cabling. To size it, is necessary to evaluate the power that must support. The maximum power is 700W but the higher intensity the 50 amperes of the starter so it must be considered. The required cable length to cover at least the compulsory lights is 16m each wing lights and 7,22m to cover the tail lights. To do a logical installation, the starter cable

4.6 Controls

This section will show how the control system has been sized according to the pilot force established in [4]. The torque value of the control surfaces is written in (1.4 Control Surfaces). The controls are done with mechanical transmission directly from the control surfaces because is the easiest system and much more lighter than the ﬂy-by-wire system.

The parameters selected to design the control stick are chosen in order to ensure the pilot’s comfort and to bring a perfect sensitivity about any plane movement.

The stick has a length of 14 inches, the allowed movement is 10 inches in the pilot direction, 5 inches in front direction and 10 inches in left and right direction. The pedal movement is from 0 to 10 inches. The lever relation between surface control and the stick lever must be

RA 126 G06-AlOn LSA 3 seats | Project report accurate to ensure the pilot force parameters.

4.6.1 Controls for wing ailerons The maximum pilot force in roll is 22 lbf and the horizontal displacement is 10 inches, the maximum angle of deﬂection is 15 degrees and the produced torque is 3,73 N.m.

In order to respect all the maximum values there is a compulsory relation between the actuation point of the stick and the actuation point of the control surface, this relation is 1/4 and to maintain constructive values the actuation point in the stick is located 1cm from the centre of rotation and the control surface actuation point is 4 cm from the empennage.

Additionally, to maintain the force sensation to the pilot a torque spring must be installed and for not passing the maximum force value the K must be 33,5 N.m/rad or lowest but the recommended one is 30 N.m/rad because is the one that suits better with the highest force value and allows the pilot to have the best and most accurate sensation between force and aircraft movement.

Lastly in the recommended design the maximum force done by the pilot to control the roll is 19,55 lbf.

4.6.2 Controls for tail elevator The maximum pilot force in pitch is 45 lbff and the horizontal displacement is 10 inches, the maximum angle of deﬂection of the control surface is 15 degree and the produced torque of it is 6,48 N.m.

In order to respect all the maximum values there is a compulsory relation between the actuation point of the stick and the actuation point of the control surface, this relation is 1/3 and to maintain constructive values the actuation point in the stick is located 1cm from the centre of rotation and the control surface actuation point is 3 cm from the emppenage.

Additionally, to maintain the force sensation to the pilot a torque spring must be installed and for not passing the maximum force value the K must be 69 N.m/rad or lowest but the recommended one is 65 N.m/rad because is the one that suits better with the highest force value and allows the pilot to have the best and most accurate sensation between force and aircraft movement.

Lastly in the recommended design the maximum force done by the pilot to control the pitch is 42,45 lbf.

4.6.3 Controls for tail vertical estabilitzer The maximum pilot force in yaw is 90 lbf because the tail rudder is controlled by pedals. As is controlled with the legs the force done by the pilot can be higher than the one done with the arms. The displacement of the pedal is 10 inches and is based with a 14 inch lever , the maximum angle of deﬂection of the control surface is 15 degree and the produced torque of it is 43,15 N.m.

In order to respect all the maximum values there is a compulsory relation between the actuation point of the pedal lever and the actuation point of the control surface, this relation is 1/3 and to maintain constructive values the actuation point in the pedal lever is located 1cm from

RA 127 G06-AlOn LSA 3 seats | Project report the centre of rotation and the control surface actuation point is 3 cm from the emppenage.

Additionally, to maintain the force sensation to the pilot a torque spring must be installed and for not passing the maximum force value the K must be 128 N.m/rad or lowest but the recommended one is 125 N.m/rad because is the one that suits better with the highest force value and allows the pilot to have the best and most accurate sensation between force and aircraft movement.

Lastly in the recommended design the maximum force done by the pilot legs to control the yaw is 88.11 lbf.

4.6.4 Control Cables and Accessories According to the aircraft dimensions the cables had been selected to ensure a safety ﬂy in any operation. The selected cable is a 7x7 Strand Stainless Steel Fabricated by Amstrong Cables.

Figure 4.41: 7x7 Strand Stainless Steel by Amstrong Cables.

The cable characteristics are: 1/16” of diameter, 7,5 lbff for each 1000 feet and a breaking strength 480lbff. The needed cable length is 1693 in and this cover all the control surfaces. The global weight of the cabling system is 1 lbs and the total price of the cabling is 41 USD.

There is also a control stick grip chosen in order to ensure the comfortably of the stick in the pilot hands. The sellected grip is Ray Allen G305 Stick Grip, the unity price is 154 USD.

Figure 4.42: Stick Control Grip in

RA 128 Chapter 5

Overall Design

The principal reason to have a CAD model is to make it easier to draw the blueprints needed for the technical sheets. However, since CAD software offers much more possibilities than just drawing, the model has been used as a support for other departments. In this case, the software used have been Catia v5 and SolidWorks. Furthermore, renders and a mock-up will be done.

5.1 3D design

Once the materials have been chosen in section 3.1, CAD software gives the center of gravity of the model, which is necessary to know for stability purposes. This model have some exterior parts, which will be used for the 3d printed model, an interior parts, which are only used to ﬁnd the c.g position. This is why in some cases, they aren’t a real representation, they are solid blocks positioned as required with volume and density of these parts.

5.1.1 Exterior Exterior includes fuselage, wing, tail and landing gear as shown in Fig. ??. Fuselage and surfaces have been imported from XFLR5 aerodynamic analysis. Landing gear was chosen as explained in Section 2.2.2.

Figure 5.1: Exterior model

RA 129 G06-AlOn LSA 3 seats | Project report 5.1.2 Interior Interior includes everything which must be inside the skin: structure, seats arrangement, belts, baggage, parachute, ﬁre extinguishers, instruments and cockpit.

Regarding the seats, it has been decided a 1:2 conﬁguration: 1 in front of the cockpit after the lever and 2 behind. Five-point belts have been placed in each seat and explained in 5.4. The luggage will be placed under the seats, weight depending on the number of passengers. Two ﬁre extinguishers have been placed in the posterior part and a parachute is allocated also in the posterior part.

Furthermore, the cockpit is in the front of the ﬁrst seat where the pilot maneuver every movement of the LSA. A lever is 127 mm in front of the cockpit, separated 254 mm from the pilot between their legs. Finally, it has been placed a throttle lever next to the pilot and pedals under the cockpit.

5.1.2.1 Structure

Used to calculate weights and surfaces. Weight is used as a decision criteria for beams and surface are used to calculate the amount of material needed. After selection, all models were simulated in SolidWorks

Figure 5.2: Full structure

5.1.2.2 Power plant

It has been approximated as a box of 40 kg and dimensions 500x500x420 mm. As propeller modelling only has aesthetical purpose, it was downloaded from GrabCad [33].

Figure 5.3: Power plant model

RA 130 G06-AlOn LSA 3 seats | Project report 5.1.2.3 Instruments and equipment Regarding the instruments and equipment in the inside of our LSA, it has been used belts in each seat and additional ones in the posterior part for the baggage of the passengers.Besides, it has been displayed the different instruments used and explained above in Fig. 5.4

Figure 5.4: Instruments displayed, Cockpit

Which its size is 500x400 mm so it ﬁts in the cabin. The deepest instrument is 141 mm, so it is made all along the cabin covering every instrument and leaving space under it for the pedals placement.

5.1.2.4 Seats and passengers It has been estimated a block of 87 kg, which includes seat, belts, passenger and their luggage. Its dimensions were 1.200mm height and 500mm wide. The model is shown in ﬁgure 5.5

Figure 5.5: Passenger model

The seats used in our LSA have been selected after a research of different seats in other LSAs. These seats must be decided according to the sizing of the inside part and an average person and their minimum comfort. The three seats will be the same.

In the following, it is showed two companies which provide seats for sports aircraft: Air-Tech Inc. and Sport Aircraft Seats. The first one gives us a general overview of the dimensions of a seat in this type of airplanes. This company offers, focused on what it is needed, a fiberglass seat structure and a cover seat in black for that structure, besides its technical drawing is given in figures 5.6b and 5.7 [34].

RA 131 G06-AlOn LSA 3 seats | Project report

Figure 5.6: Seats and prices of Air-Tech Inc.[34]

Figure 5.7: Measures, technical sheet of the Ultralight seat [34]

The second company Sport Aircraft Seats, provides a whole seat in “one-piece” between fabric and leather[35]. Therefore, it has been chosen a fabric seat instead of a leather one because of its breath-ability as well as how it behaves in cold weather. In this case, it is possible to choose the P110 structure-seat, so it ﬁts for the passenger and our designed aircraft.

RA 132 G06-AlOn LSA 3 seats | Project report 5.2 Blueprints

On early stages of the project, a first document with general dimensions of the aircraft was given to Structures department in order to start sizing of the structure. This overview was prepared with the dimensions used by Aerodynamics department during their analysis and weren’t definitive, so there might be minor changes between figure 5.8 and the final aircraft. Final blueprints are those shown at the technical sheets.

Figure 5.8: Initial sizing

RA 133 G06-AlOn LSA 3 seats | Project report

RA 134 Chapter 6

Business

In this chapter there will be an analysis, study and planning of all the economical and business points regarding the aircraft itself and the enterprise. A further analysis of the latter will be carried out as well.

6.1 Manufacturing costs

Taking into account all the costs of the project and the fact that there will be a development of an enterprise behind the aircraft, in order to make proﬁt of this product (sold by the company), one must consider all the costs the company will face. All these manufacturing costs have been split into four different categories: • Materials • Facilities • Human resources • Additional costs Some of this categories will be discussed in the next sections.

6.1.1 Cost of facilities To be able to choose an office, some previous calculations must be done to have an order of magnitude of how big the work area must be. Assuming one person takes up approximately 3.5m2 of personal work room, where they have their desk and personal objects. The total space required for one person to work will be evaluated as 14m2 [36]. As there are 14 people working the required room will be approximated as ≈ 14 · 14 = 196m2. Now that an order of magnitude has been set of how big the total room of the office must be, these four different cases can be studied in order to make the final choice. 1. Office 1: C/ Almogàvers 119, Edifici ECOURBAN, 08018 Barcelona. Total room available: 431 - 909 m2. Cost: 16e/m2.

RA 135 G06-AlOn LSA 3 seats | Project report

(a) Floor plant (b) Location

Figure 6.1: Speciﬁcations of Ofﬁce 1

2. Ofﬁce 2: C/ Calabria 169, 08015 Barcelona. Total room available: 170 - 343 - 418 m2. Cost: 11e/m2.

(a) Floor plant (b) Location

Figure 6.2: Speciﬁcations of Ofﬁce 2

3. Ofﬁce 3: Av. Diagonal, 309, 08013 Barcelona. Total room available: 113 m2. Cost: 10, 00e/m2.

(a) Floor plant (b) Location

Figure 6.3: Speciﬁcations of Ofﬁce 3

4. Ofﬁce 4: C/ Bailén, 3, 08010 Barcelona. Total room available: 188 m2. Cost: 11e/m2

With these four options considered, a decision will be taken on the best ofﬁce on terms of room available and location. Taking into account the room needed is approximately 200m2,

RA 136 G06-AlOn LSA 3 seats | Project report

(a) Floor plant (b) Location

Figure 6.4: Speciﬁcations of Ofﬁce 4

Ofﬁce 3 falls behind as far as room is concerned. Then, considering the total amount it would cost to rent each ofﬁce, one gets:

Price per m2 [e] Room [m2] Total rent cost [e] Office 1 16 431 6.896 Office 2 11 170 1.870 Office 3 10 113 1.130 Office 4 11 188 2.068

Table 6.1: Comparative of the 4 possibilities

Now that the total amounts have been computed, an objective and thoughtful choice can be taken. The best option would be Ofﬁce 4, as it provides enough room (it just lacks 8m2), it is in a very good spot in Barcelona, and the price per squared meter is low enough to keep the budget within the margins. So, ﬁnally, one gets a total cost of facilities of 2.068 eper month.

6.1.2 Cost of human resources In the human resources category it will be assessed how much money the company must spend on the people that will work in it. Such quantity is mainly composed of the different salaries of the workers (engineers and coordinators). In order to get the total sum of salaries, the workers wil be divided into two separate categories: nine engineers and four coordinators. The total amount of hours per week will be estimated to be eight hours for the two categories. With this, the salary will be determined by the total amount of hours required to do the project, and the cost per hour of an engineer/coordinator. The remaining team member is the Project Manager, who belongs to a different category. Assuming the project takes a total amount of 11 weeks to be ﬁnished, so the total amount of hours worked is:

Workers Weeks Hours/week/worker Salary/hour Total amount [e] Engineers 9 13 8 25 23.400 Coordinators 4 13 8 35 14.560 Project Manager 1 13 8 50 5.200 Total 14 13 8 - 43.160

Table 6.2: Basic human resources cost

With this ﬁrst estimation one gets the total amount that must be paid to the workers as gross salary. Until this point, any extra hours nor external personnel have been considered. In order to get a better approximation of the total cost of the human resources, one must add a percentage of the gross amount, in case the workers have to work extra-hours. Such percentage will be estimated as the 10% of the total. So, in conclusion, the total cost of human resources will be estimated at 43160 e+ 4316 e= 47.476 e. This ﬁnal result consists only of the salaries of internal personnel of the company. It does not

RA 137 G06-AlOn LSA 3 seats | Project report contain any additional personnel costs, as it will be considered in the Additional costs subsection.

6.1.3 Additional costs

Certification of the aircraft: The certification of the aircraft is a process that must be carried out while all the aircraft is being designed. EASA gives the companies the opportunity to certificate its own aircraft model, delivering all the needed documents once it is finished, or delivering each document separately once that specific part is finished. This system works very well because it gives room for error, so if an specific part of the aircraft does not satisfy EASA’s regulations (Part21A.701 (15)) [37], we can easily go back and correct it before the aircraft is completely finished. So, in order to compute the cost of the certification of the aircraft, one can take the Total Budget of Human Resources (43160 e) and add a 10%, as it is approximately the time an engineer will have to dedicate to check if the part they are working on satisfies the regulations. The total amount of the certification is then: 0.1 · 43.160e = 4.316e.

6.1.4 Marketing costs

One of the main costs of the company is the marketing cost, not in terms of total volume, but in terms that it has to be well deﬁned, as a high percentage of the sells will depend on how good the marketing campaign is. In this section, the cost of the marketing campaign will be studied. The marketing campaign in itself will be thoroughly studied on section 7.2 Marketing campaign.

6.1.4.1 Initial marketing campaign costs

For new companies, the cost of the marketing campaign is set to be ≈ 8 − 12% of the total budget. This amount, though, is lower in this project, since the marketing campaigns of prod- ucts like planes are far more centered in the potential costumers. The total amount invested in marketing is, thus, 8.650 e. With 8.650edesignated to the budget of the marketing campaign, this total amount will be split into the marketing subsections that can be seen on Table 6.3: So the total amount spent on marketing is $ 8.650,00, what has been designated. The weight distribution of each subcategory can be seen on Figure 6.5.

Figure 6.5: Speciﬁc weigh t by category

RA 138 G06-AlOn LSA 3 seats | Project report CATEGORY QUANTITY UNIT COST SUBTOTAL

National Marketing 250,00 e Banner Ads 1 250,00 e 250,00 e

Local Marketing 1.300,00 e In-Store Marketing 2 400,00e 800,00 e POP 1 00,00 e 500,00 e

Public Relations 700,00 e Public Events 1 500,00e 500,00e Press Releases 2 100,00e 200,00e

Online 400,00e Blog 1 400,00e 400,00e

Advertising 2.350,00e Online 1 1.500,00e 1.500,00e Print 1 850,00e 850,00e

Sales Campaigns 800,00e Campaign A 1 800,00e 800,00e

Market Research 350,00e Surveys 2 175,00e 4.800,00e

Content Marketing 600,00e Sponsored Content 3 100,00e 300,00e Landing Page 2 150,00e 300,00e

Web 750,00e Development 1 750e 750,00e

Other 900,00e Premiums 2 200,00e 400,00e Corporate Branding 1 250,00e 250,00e Business Cards 10 25,00e 250,00e

GRAND TOTAL 8.650,00e

Table 6.3: Marketing costs by category

6.1.4.2 Study of marketing costs over time

Being able to sustain the marketing costs implies the success on sales of the company. Taking into account the high amount that a 1-year marketing campaign costs 8.650 e, and considering the capital assigned to marketing is approximated to be a ≈ 6 − 10% of the company’s benefits, the benefits of the company on the first year should be around $ 86.500. If this initial marketing campaign is sustained over time (see subsection 7.1.5.3 Study of future marketing campaigns), and thus having benefits of over 86.500 e, the company will carry on having this marketing cost. The impact of the initial investment on the annual benefits will be studied in the sections 7.3 Initial Investment and 7.4 Payback Analysis.

RA 139 G06-AlOn LSA 3 seats | Project report 6.1.4.3 Study of future marketing campaigns Once the first exercise is completed (i.e. the first year), the future marketing campaigns may vary from the first one. This differences may reside on the approach, on the scope or on the platforms used in the new marketing campaign. In this prediction, two different situations can occur:

• Success of sales. Assuming that at the end of first exercise, the benefits are high enough (so the sales have been successful, and so the marketing campaign), there would not be any need for a wider marketing campaign. So the marketing campaign for exercise 2 would be equal, or even a reduction of the first one. Considering this situation, the total amount of capital needed for marketing issues on the future can be reduced to a 7-10% of the benefits of the company. This is the best scenario for the company.

• Lack of interest/ low number of sales. Assuming now, that at the end of the first exercise the sales have not been high enough, so not enough profit has been made by the company, there should be some changes on the marketing campaign of exercise 2. Firstly, the amount of capital destined to marketing would rise to 15% of the benefits of the company in exercise 1. Secondly the scope of the campaign could be broadened in order to reach more potential customers and achieve the goal in exercise 2.

RA 140 G06-AlOn LSA 3 seats | Project report 6.2 Marketing campaign

This section is focused on the marketing issues of the project, from the possible customers, to the design of the marketing campaign.

6.2.1 SWOT analysis

Here we explore the Strengths and Weaknesses of our product and the Opportunities and Threats to our company; to better approach decision-making and to more adequately focus our marketing efforts. This kind of analysis is called a SWOT analysis (See Figure 6.6), and it provides a simpliﬁed and structured vision of the whole Business project. [38] [39]

Figure 6.6: Scheme of a SWOT analysis.

Source: https://research-methodology.net/theory/strategy/swot-analysis/

6.2.1.1 Research on Market Opportunities

The location in Barcelona gives the company access to a large specialized labour pool. Ties and contacts within Catalonia’s Polytechnic University can yield us cheap interns as well as further access to aerospace engineers, both recent graduates and veterans in the ﬁeld. Lack of airplane manufacturers in the area only reinforces this access to the labour pool, allowing us to pay lower salaries. Low cost-of-living in Spain vs. the US, Canada and the rest of the EU (where most our competitors are based) further reduces our labour costs relative to the competition.

6.2.1.2 Research on Market Threats

The aforementioned lack of plane manufacturers in the Area also means a less experienced labour pool. Well-established companies in the field could undercut our prices because they benefit from economy of scale. Since the product sold is a sport aircraft without LSA attributions (because of its higher than three seat capacity), it is taken as competitors manufacturers both that sell actual LSA and those that sell three-seated sport aircraft. The main competitors for the company are Flight Design USA with its model CTLS and CTLS Lite,then Czech Sport Aircraftwith the models CLub, Tourer and Professional,and finally CubCrafters with the models Carbon Cub SS and Sport Cub S2. These three companies are known as the leaders of LSA’s sellers. Also One Aircraft has developed a model of LSA with three-seats called ONE 2+1 that will be in the market in 2020.Since this product is basically the same type of LSA model as Alpha One , this could be the main threat for the company, so Alpha One’s aircraft should be in the market before ONE 2+1 Another issue is that volatility of oil prices and EUR-USD exchange rates can affect customers’ demand and harm the company’s long term plans.

RA 141 G06-AlOn LSA 3 seats | Project report 6.2.1.3 Research on Aircraft Strengths

The aircraft offers high fuel efﬁciency unmatched by any other aircraft capable of carrying three passengers. The aircraft being manufactured out of mostly custom-made parts secures us future funding in the form of repairs and maintenance. Furthermore, the possibility of the airplane entering the category of LSA in a future, makes it even more appealing to Sport Pilots, given they will be able to carry more passengers (and thus more weight) without the need of another license.

6.2.1.4 Research on Aircraft Weaknesses

Limited ﬂying range makes the product unsuitable for the American and Canadian markets, as well as alienating any potential customers who seek medium to long range sport aircraft. The aircraft being manufactured out of mostly custom-made parts makes us more liable to litigation due to malfunction-related accidents. Also the main weakness for any starting company is the lack of experience in the aviation industry, and not having any reputation makes it harder to get into the labour market.

6.2.2 Study of potential customers The main customer market to which the company can center its sells is the one that is already buying conventional two-seats LSAs. With this in mind, the attractive this project’s LSA (AlOn) has over other conventional LSAs is its bigger capacity. Although the total weight is the same, as it is regulated by the OACI (and thus cannot be exceeded), it provides extra room for one person, or to carry more payload if necessary. The market of aviation grew approximately a 6,5% between 2016 and 2017 (see Image 6.7), so a new idea of LSA has enough room in the market to start its own client portfolio. [40] In the next table one can see the Gen- eral Aviation Airplane Shipments of Piston-Engine planes between 1995 and 2017. The great decay from 2007 to 2010 is believed to have been caused by the economical crisis Europe and the US suffered on this period of time. [41]

Figure 6.7: Airplane Shipments Worldwide (1995–2017)

[42]

RA 142 G06-AlOn LSA 3 seats | Project report Airplane Shipments Worldwide (1995–2017) Year Total Single-Engine Piston Multi-Engine Piston Change 1995 666 605 61 - 1996 801 731 70 20,27% 1997 1123 1043 80 40,20% 1998 1606 1508 98 43,01% 1999 1801 1689 112 12,14% 2000 1980 1877 103 9,94% 2001 1792 1645 147 -9,49% 2002 1721 1591 130 -3,96% 2003 1896 1825 71 10,17% 2004 2051 1999 52 8,18% 2005 2465 2326 139 20,19% 2006 2755 2513 242 11,76% 2007 2675 2417 258 -2,90% 2008 2119 1943 176 -20,79% 2009 963 893 70 -54,55% 2010 889 781 108 -7,68% 2011 898 761 137 1,01% 2012 908 817 91 1,11% 2013 1030 908 122 13,44% 2014 1129 986 143 9,61% 2015 1056 946 110 -6,47% 2016 1019 890 129 -3,50% 2017 1085 936 149 6,48%

Table 6.4: Evolution of Airplane Shipments worldwide between 1995 and 2017

6.2.3 Set marketing goal As a start-up, the main goal is to make the company known. Putting itself in the public eye is, however, not a must. The company needs only make itself known to its potential customers, a very specific subset of the general population. Furthermore, an effort to appear professional and trustworthy must be undertaken, since the lack of an established reputation will make clients wary of acquiring the product, fearing to be scammed or given a sub-par purchase. A policy of openness and transparency about the way business is conducted needs be approached, whilst avoiding to harm the company’s own interests. Because of the potential for a profitable after-sales relationship with the customer, the company must strive for a good relationship with them, presenting itself as friendly and close, but without sacrificing the aforementioned professionalism.

Moreover as the marketing goal is set, other concepts need to be deﬁned such as: • Mission: this concept answers the questions: What is our service at the moment? Which is our competitive advantage? A good mission allows a company to reach their vision. In our case our mission is to provide a brand new product, an LSA suitable for an additional passenger, to all the aviation fanatics and LSA lovers. • Vision: this refers to what do we want to become in the future as a company. Since we have come up to the market with a new product, our vision is to expand the company, become the best sellers of this type of light sport aircraft and maybe design a variant of the product. • Values: this is one of the most important aspects to consider. The values of a company

RA 143 G06-AlOn LSA 3 seats | Project report refer to the work philosophy that all the workers of it share such as respect, cooperation, discipline and perseverance.

6.2.4 Study of advertisement To develop the advertising plan that will be developed,it will be done a general study of the best options to invest in advertising, which media is the best one and possible companies that can provide the company this service.

The main media used for advertising in aviation is print media, magazines basically and digital media. Although nowadays digital media seems to be the most powerful way to reach the maximum possible costumers, Alpha One’s target is set on people that are used to print media, so we will focus on that.

Following it is seen the different companies that could provide the company the advertisement desired:

• Air Charter Guide : it is the most-read authoritative guide to the air charter industry. This company can provide the company visibility by advertising in the guide pub- lished twice a year, and as long as we commit to advertise the company will promote Alpha One on their website with no extra charge. The Air Charter Guide has 171.738 subscribers all over 148 countries so we would have a wide visibility. The price for a quarter page ad costs around 1.165 $.[43]

• IATA (International Air Transport Association : this large association has half million read- ers worldwide and the website has over 2,3 million views per month. Although the association is more focused in air transport and commercial aviation, more feasible for airline advertising, we could consider advertising with them since they provide ﬂexi- bility to pick and choose our target audience.[44]

• Flying Magazine by Bonnier’s Marine & Aviation Group: it has the largest paid subscription for an aviation magazine, that positions it as the world’s most widely read aviation magazine.[45] Bonnier’s Group offers the possibility to advertise in either the monthly printed publication in any size desired or in their website. Moreover the website can promote our product via digital media not just on their website, also in their social network and encourages the company to have an account for networks like Instagram, Facebook...[46]

• Perfect Landing: this last company is an aviation marketing agency that it would taken into account as an auxiliary help for designing the website or if needed to help us ﬁnd possible advertisers as they are specialized in this ﬁeld.[47]

The company has decided to work with all the companies mentioned above. Since all of them can provide us good visibility, we will advertise by printed media and also via online websites with IATA, Air Charter Guide and Flying Magazine with the periodicity set by each company.We will hire the service of Perfect Landing as well to help us develop our website and supervise our marketing plan. Another digital platform that we could use as a way of broaden the possible costumers is using social network as Instagram and Facebook since the both of them are free and would not be an additional cost, a part from designing the Alpha One website.

6.2.5 Design of marketing campaign Basic marketing campaigns are based on a few requirements such as the analysis of the actual market, determining a marketing goal and decide which will be the actions done to set that

RA 144 G06-AlOn LSA 3 seats | Project report goal.

First of all, the analysis of the actual market was made when the SWOT analysis was done, and it was checked who the main competitors in the industry were, companies that sold a similar product to ours and had more experience and fame than our start-up.

On the other hand, the marketing goal and advertisement development was set in previous sections. So in this section one can see the logotypes designed for the brand . Here-under we will show the logotypes designed and the ﬁnal choice: So ﬁnally the last logo was chosen because it was thought that it was simple and serious,

(a) Logo proposal 1

(b) Logo proposal 2

Figure 6.8: Different logo proposals although the other logos are really modern and could be more suitable for start-up. Since our target is mainly business men and not young people, we thought that the simplicity and the colour black would show more seriousness and could be more suitable for our project.

RA 145 G06-AlOn LSA 3 seats | Project report 6.3 Initial investment

6.3.1 Research on possible investors This is a key section of our project. Since the Initial Investment is far larger than what it takes to sustain the company once it has started producing, the project members must seek some investor to be able to make the project a reality. The origin of the money is a very important aspect when it comes to the economical feasibility, as it has a big impact on the factor k, that measures the proﬁtability of the whole project within a period of time. The origin of the money could be: • A bank

• A group of people that invests money (angel investor)

• The members of the project Since we have already discarded the last option, as the Initial Investment is beyond the group’s range, we will focus on the other two options: a bank, or an angel investor. Investing in a brand new project implies risking the money you have invested, as one can not have 100% certainty that the project will be successful. Given this, having a bank investing money in a project like this is usually quite difficult, as they tend not to have a risk factor over ≈ 2%. On the other hand, angel investors demand a high IRR (Internal Rate of Return) in order to earn money in a short period of time. This IRR is usually between 20-30% in the first 5 years of the project. [48][49] The project can find angel investors online, for example at https://angel.co/europe/investors. Last year, in Spain, these angel investors invested a total amount of 21.2 M e. This total amount was a 78% larger than last year. This increase demonstrates the fact that every year the angel investors are more relevant worldwide.[50]

6.4 Payback Analysis

Once the project is all set up, one must carry out an analysis of economic feasibility of the project, regarding if this project is worth developing or not. This economical feasibility analysis is thoroughly carried out in the Chapter 6 Economical Feasibility of the main report. In this section, we will analyze how the price of the plane (set as 160.000,00 e), can vary in order to make it more attractive to the clients or to make it more proﬁtable for the project. This analysis is called the Study of proﬁtability margins

6.4.1 Study of profitability margins As commented previously, in this section the price of the plane will be varied and analyzed in a range from 100.000 to 200.000, computing for each price the NPV, the PBT, and the IRR. This values will be later compared with the values obtained for our project, in the Economical Feasibility analysis, carried out in the Report, Chapter 6. On the figures 6.9, 6.10 and 6.11 one can see the graphics of this computations. It has been highlighted with an orange line the selling price set for our project, in order to have a better visual on where AlphaOne lays in each graphic. As figure ?? shows, the company will not make any profit (in 7 years, with rate of decay k = 11,4%) until the selling price per aircraft is more than 130.000 e. After this price is surpassed, the company can increase its revenues by increasing the price. The intersection of the two curves gives us the NPV calculated for a selling price of 160.000,00 ewith a total value of 938.113,37 e.

On ﬁgure 6.10 one can see the evolution of the Pay-Back Time. As it can be seen, it decreases with the price, as expected, tending to 0 as the price increases. The PBT for prices per unit

RA 146 G06-AlOn LSA 3 seats | Project report

Figure 6.9: Net Present Value (years = 7, k = 11,4%) for different selling prices

Figure 6.10: Pay-Back Time for different selling prices

Figure 6.11: Internal Rate of Return for different selling prices less than 120.000,00 eincreases significantly as it tends to ∞. Once more, the intersection of the two curves gives the PBT of AlphaOne, 3,61 years. Finally, on figure 6.11, one can see the evolution of the Internal Rate of Return with the selling prices, like in the two previous figures. This latter figure shows that for prices less than ≈ 130.000,00 e, the IRR turns out to be negative. This indicates us the non-feasibility of the project, without taking into account the decay rate k. If one sets the IRR to be higher than k, the project starts being feasible once the price reaches ≈ 140.000,00 e, and then gains feasibility with higher prices. Two things could be done once we have this results:

• Lowering the selling price: Lowering the selling price might lead us to more sells, but not in the ﬁrst years, as the company has to adapt itself to the new market (as commented on the Report). Even if the company sold one more unit per year, a price of 150.000 ethe PBT would not be any higher (3,74 years), and would have an IRR of 32 %, lower than the current IRR of 33 %.

RA 147 G06-AlOn LSA 3 seats | Project report • Increasing the selling price: Increasing the selling price would turn out to be proﬁtable once the company has enough fame, and a customer wallet big enough to be able to do so without losing sells. Increasing the selling price to 170.000, and considering one unit sold less every year, one gets a PBT of 3,7 years, and an IRR of 32 %, both with economical disadvantages to the current selling price.

With these economical evaluations on how the main economical aspects vary with the price and units sold, it can be concluded that a selling price of 160.000,00 elays in a range where it gives the company a high enough NPV with, an acceptably low PBT, and an IRR higher than k.

RA 148 Chapter 7

Organization, planning and scheduling.

7.1 Gantt Diagram

Due to the complexity of the diagram and in order to keep the quality of the image, the Gantt diagram has been attached on the annexes in its original size. There it can be found the Gantt Annex A.4 that has been followed during the development of the project.

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RA 150 Chapter 8

Minutes of the Meeting

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MINUTES OF THE MEETING 01

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 20th September 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT

AGENDA:  Presentation of the group members and selection of a coordinator and a secretary.  Description of the Project Charter and its structure  Presentation of the object of the project

ITEMS ON THE AGENDA: 4th October 2018: Delivery of the Project Charter with the following content:  Aim of the project  Scope of the project  Requirements of the project  Justification 1

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 State of art  Group Organization o Organization Structure o Roles and responsibilities  Planning o Tasks identification from Work Breakdown Structure (WBS) o Brief tasks description o Interdependency relationship among tasks o Human resources and level of effort (hours) to develop each task  Budget (initial estimation)

DECISIONS MADE – ASSIGNMENTS:  Presentation of the group members and selection of the main roles: o Coordinator: Ariadna Fernández, who will deal with the management tasks and will be responsible for the communications between the customer and the other group members. o Secretary: Itziar Ugartemendia, who will write up the minutes of the meetings and will be in charge of keeping BSCW updated and organized.  The working environment will be BSCW. This platform allows organizing files in different folders and subfolders to classify the uploaded documents in the corresponding category. All the members of the group have access and can upload, download or modify the documents.  The communication between the customer and the group members will be though ATENEA, where the final documents will be delivered.  The communication tool among the group members will be Slack. This application allows having different communication groups. There will be a general group with all the members to discuss the global and most relevant aspects of the project. In addition, there will be a specific group for each department, where only the corresponding members will take part.  The aim of the project is to design a Light Sport Aircraft (LSA) for 4 passengers, taking the European regulation into consideration.  The next meeting will be the 27th September 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Definition of the aim and scope of the project o Planning of the project and tasks identification o Group organization in different departments

SIGNATURES Team coordinator Supervisor

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MINUTES OF THE MEETING 02

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 27th September 2018, from 9:00 to 12:00 h Place: Seminary 6 – ESEIAAT

AGENDA:  Definition of the aim, scope and justification of the project  Analysis and determination of the requirements of the project  Group organization in different departments  Discussion of the planning and tasks to be done by each department  Schedule and time distribution of the tasks  Revision of the Project Charter and determination of the remaining contents to be finished

ITEMS ON THE AGENDA: 4th October 2018: Delivery of the Project Charter.

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DECISIONS MADE – ASSIGNMENTS:  The project management application called Trello will be used to have a real-time organization of the tasks. It allows us to create different blocs, one for each department, and write down the tasks and subtasks to be done by each department in a list called “To do”. When a task is started, it is moved to the list called “In progress” until it is finished and moved to the list called “Verification”, where it is revised by two or three people (that haven’t worked in that task) and moved to the list called “Done”. In this way it is easy to see the evolution of the project and the remaining tasks to be done.  Revision of different aspects of the Project Charter. There are some aspects to change: o The first step is the scope definition; afterwards, the tasks determination and the organization in departments; and finally the time estimation and schedule. o The scope defines the amount of work and the level of determination of the project, but it does not include the specific tasks in each department. The department organization is created afterwards according to the scope and the WBS. o The project must meet all the requirements of an LSA, except the number of passengers (three instead of two). Consequently, the aeroplane developed can’t be called LSA. o The State of Art contains the study of planes under 600Kg and planes for 3 passengers, but does not include the planes of two passengers.  Discussion of new suggestions and resolution of questions, doubts or problems. The conclusions taken are the following: o The project will be focused on the development of a hybrid between a glider and an LSA in order to reduce the structural weight and increase the efficiency. o The technology used in the plane cannot be determined yet, but it will be considered whether to install Fly-by-wire in order to use electrical actuators instead of mechanical. The weight and the requirements of each configuration will be studied to determine the best option.  Development of the BWS and the scope of the project: Definition of tasks and subtasks with a numerical code to organize them clearly.  The team organization in departments will be the following:

Departments Members 1. Pol Bernad 2. Edgar Gago Aerodynamics team 3. Pau Nadal 4. Itziar Ugartemendia 1. Eduard Gómez Power plant team 2. Marcel Marín Technical 3. Pau Nadal 1. Xavi Carrillo Design team 2. Alejandro Sans 1. Pol Bernad 2. Ariadna Fernández Structures team 3. Carlos Méndez 4. Pau Nadal

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5. Alejandro Sans 1. Ariadna Fernández Systems team 2. Carlos Medina 3. Carlos Pérez 1. Eduard Gómez Economics team 2. David Rodríguez Non-technical 1. Alexandra Kalina Team management 2. David Rodríguez

It is important to remark that the distribution of the team members could be modified if necessary due to the amount of work or the difficulty of the tasks.  There are also the following department coordinators, who are responsible for the coordination and communication among departments: o Aerodynamics coordinator: Pol Bernad o Structures coordinator: Carlos Méndez o Technical coordinator: Pau Nadal o Non-technical coordinator: David Rodríguez Finally, two more important roles, already defined in the first meeting, are: o Main coordinator: Ariadna Fernández o Secretary: Itziar Ugartemendia  The next meeting will be the 1th October 2018, from 14:00 to 15:00 h, in a free classroom to be determined (ESEIAAT). Over these days, the following tasks will be carried out: o Correction of the aspects of the Project Charter previously described o Development of the last contents of the Project Charter in order to revise it together during the meeting.

SIGNATURES Team coordinator Supervisor

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MINUTES OF THE MEETING 03

Project: Design of a LSA for 3 passengers Participants: Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 1st October 2018, from 14:00 to 15:00 h Place: TR5-3.2 – ESEIAAT

AGENDA:  Revision of the WBS  Discussion of the remaining contents of the Project Charter  Schedule and time distribution of the tasks

ITEMS ON THE AGENDA: 4th October 2018: Delivery of the Project Charter.

DECISIONS MADE – ASSIGNMENTS:  Once the WBS is finished and the list of tasks for each department is completed, it is time to assign the resources required to develop each task. It is quite difficult to determine how many hours each task will take. For this reason the method used is the following: the measure used instead of hours is the “Story Point” (SP), which considers both time and difficulty. One Story Point is equivalent to the work of a person in a morning/afternoon (that

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is, 2 or 3 hours). Each task can be defined by a multiple of 0,5*SP. In this way it is easy to count the number of Story Points available in order to set the deadlines for each task and create the Gantt Diagram. It will also help to distribute the tasks in an equitable and fair way among the group members.  From the revision of the Project Charter the conclusions to be drawn are the following: o The remaining contents are: description of tasks; resources, time and preceding tasks assignment; estimation of the budget; references and bibliography. o The deadline set to conclude these contents is the 2nd October. The following day the Project Charter will be revised and ready to be delivered.  The next meeting will be the 4th October 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Development of the Gantt Diagram and other remaining tasks. o Final revision and delivery of the Project Charter o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.

SIGNATURES Team coordinator Supervisor

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MINUTES OF THE MEETING 04

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 4th October 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT

AGENDA:  Analysis of the progress done and resolution of the problems  Delivery of the provisional Project Charter  Revision of the Project Charter with the supervisor and analysis of things to be changed  Last corrections of the Project Charter

ITEMS ON THE AGENDA: 4th October 2018: Delivery of the Project Charter Specific tasks for each department according to the Gantt Diagram

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DECISIONS MADE – ASSIGNMENTS:  Delivery of the provisional Project Charter in BSCW according to the deadline.  Revision of the Project Charter with the supervisor and discussion of the following aspects to be modified: o The State of Art: . It should not include specific aspects of our project . It should appear after the requirements, not after the aim . There are some changes that not appear in the last version o Scope: . There are some generic words that should be changed in order to specify more what is going to be done, for example, the words “design”, “analysis”, “study” should describe concretely what is included. Also, when it is said “main parts” it should be specified which parts. o Stakeholders: . It does not give much information in a Project Charter o WBS: . It would be better to turn the page in order to see it clearer . It is also possible to present it in a table o Gantt Diagram: . It must appear at the end of the document, not after the WBS, following the logic organization . It can be included a specific Gantt Diagram for each department to see the tasks clearer . The tasks of the point 7 are not included in the description of tasks o Definition of tasks: . It is very specific but that’s fine if we can deal with it o Design Budget: . We cannot include the annual licenses of Catia and Ansys, only partially because we are working only 4 months in this project . The computers could be include partially or consider that every team member uses his/her own one. . The necessary materials should also appear  These aspects have been corrected and modified during the meeting.  The next meeting will be the 11th October 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram. o Evaluate how is the group organization working and determine if it is necessary to make changes in any department

SIGNATURES Team coordinator Supervisor

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MINUTES OF THE MEETING 05

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 11th October 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT

AGENDA:  Revise the work done and discuss problems or doubts  Update the Gantt Diagram and organize the following tasks to be done

ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram

DECISIONS MADE – ASSIGNMENTS:  Develop the corresponding tasks for each department according to the schedule. 

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 Each department has explained the progress done and the most important decisions made in order to have an overall view of the project. A brief summary of the tasks developed by each department is given next: o Aerodynamics: . Discussion of possible airfoils and analysis of different studies of airfoils and wings in order to define the basic configuration of the wing, which at first are: airfoil 230XX (the thickness could change according to future analysis); wing span, 16m; root chord, 0.8m; and tip chord, 0.6m. . Study of different tail configurations and selection of two possibilities: Conventional tail or T-tail. The final decision will depend on further analysis of the whole aircraft. . While developing the tasks according to the Gantt Diagram a problem has come up: two blocks of tasks (wing and tail) can’t be done at the same time by different members, as it was at first planned. The department has accorded to work together in both blocks and design all of them in order to study the aerodynamic forces and moments and the stability. o Structures: . Study and sizing of the landing gear according to the most commonly used in similar aircrafts. . Study of the possible materials for the construction of the aircraft and the regulations in this field. o Power Plant: . Study of the different engines and propellers available in the current market, estimation of the power required by the plane and selection of a few possible engines according to the results obtained. . Revision of the Certification Specifications for an LSA by ASTM and the specific regulations for EASA and FAA o Systems and Avionics: . Research about compulsory and complementary systems. . Research on possible control systems. o Economics: . Estimation of costs, including materials, facilities, human resources and other additional costs o Team management: . Coordination and supervision of the progress done.  The next meeting will be the 18th October 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.

SIGNATURES Team coordinator Supervisor

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MINUTES OF THE MEETING 06

Project: Design of a LSA 3 for passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 18th October 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT

AGENDA:  Revise the work done and discuss problems or doubts  Update the Gantt Diagram and organize the following tasks to be done  Develop the corresponding tasks for each department according to the schedule

ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 20th December: Project Delivery and Final Presentation

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DECISIONS MADE – ASSIGNMENTS:  Each department has continued developing the corresponding tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . Last analysis of plane designs in Xflr5 in order to close the design as soon as possible and start the CAD design. The main wing and tail parameters are already fixed. o Structures: . Final decision of the materials that will be used. . Tests with the landing gear to decide the optimum configuration. o Power Plant: . The engine is already selected and the current work is based in the engine bench. . Development of a Matlab program to analyse the propeller in order to design it and evaluate its viability compared to the other propellers selected. o Systems and Avionics: . The work done by the each member has been joined. It has been discussed what is missing to finish the Flight and Navigation instruments part. . Contact with the Power Plant Department in order to start selecting the power plant instruments for next week. o Economics: . The estimation of the facilities’ and the human resources’ costs is already finished. . Estimation of the marketing campaign costs. . The SWOT analysis has been started. o Team management: . Coordination and supervision of the progress done.  The next meeting will be the 18th October 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.

SIGNATURES Team coordinator Supervisor

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MINUTES OF THE MEETING 07

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 25th October 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT

ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 20th December: Project Delivery and Final Presentation

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DECISIONS MADE – ASSIGNMENTS:  Each department has continued developing the corresponding tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . There are two completed designs of airplane without control surfaces, which are the two possible options for the final plane: the first one with the three passengers located in a row, and the second one with two passengers in a line behind the pilot. In order to select one, the decision methods explained in class will be applied. . Once the final airplane design is selected, the definitive parameters and graphics will be exported. . The current work is focused on the control surfaces (sizing and location), the weight-range diagram and the parasite drag determination. o Structures: . The analysis of the landing gear has already been finished. . Start sizing the structure of the wings and fuselage o Power Plant: . Last steps in the study of the engine bench . Progress in the propeller design and selection of alternative propellers from the market, just in case that it is necessary. . Study of the exhaust fumes and the correct escapement o Systems and Avionics: . All the power plant instruments have been finished . The costs, weight and power of each instrument has been recorded in a table that will be useful for future tasks . The 2D design of the dashboard of the cockpit has been started o CAD design: . Design of the two alternatives for the plane. o Economics: . Finish the SWOT analysis. . Make progress in the estimation of the marketing campaign costs. o Team management: . Coordination and supervision of the progress done.  The next meeting will be the 8th November 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.

SIGNATURES Team coordinator Supervisor

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MINUTES OF THE MEETING 08

Project: Design of a LSA for 3 passengers Participants: Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 8th November 2018, from 9:00 to 12:00 h Place: Seminary 6 – ESEIAAT

ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 20th December: Project Delivery and Final Presentation

DECISIONS MADE – ASSIGNMENTS:  The report has been updated to include all the contents that should appear. This document contains all the necessary information about the design of the LSA: the procedures developed and the results obtained from the different analysis.

RA 167 G06-AlOn LSA 3 seats | Project report

 Every department is writing its corresponding part of the report while developing the tasks in order to record all the procedures, analysis, conclusions and decisions made.  Each department has continued developing the corresponding tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . From the final design it has been necessary to size the tail again and make it bigger to improve its performance. . The current work is focused on the sizing of the control surfaces, which involves the study of the stability and control of the airplane. Besides, some parameters from equivalent airplanes will be used as a first reference. o Structures: . The landing gear design has been finished. . For the fuselage sizing, the momentum equilibrium has been calculated. The same procedure will be applied for the wing and the tail. o Power Plant: . Progress in the design of the propeller in Matlab. . End of the tasks related to the engine mount and the exhaust fumes system. o Systems and Avionics: . All the systems are defined. The task remaining is the determination of the parameters of the controls of the pilot. Some data related to the control surfaces, which the current work of aerodynamics department is required, but it will be available in two or three days. o CAD design: . Make the necessary modifications to the design previously done according to the changes in some aspects of the structure and aerodynamics. o Economics: . Conclude the SWOT analysis . Close the estimation of the marketing campaign costs and the study of the future campaigns. . Begin the study of potential customers and the marketing study. o Team management: . Coordination and supervision of the progress done.  The next meeting will be the 15th November 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.

SIGNATURES Team coordinator Supervisor

RA 168 G06-AlOn LSA 3 seats | Project report

MINUTES OF THE MEETING 09

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 15th November 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT

ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 29th November: Delivery of a draft of the report. 20th December: Project Delivery and Final Presentation

RA 169 G06-AlOn LSA 3 seats | Project report

DECISIONS MADE – ASSIGNMENTS:  Considering the deadline for the delivery of the draft of the report, one of the main tasks in each department is write all the information required in the document, adding graphs, figures and conclusions.  Each department has made progress in their tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . The current work is focused on sizing the control surfaces, including flaps, ailerons and rudder. o Structures: . Structural analysis of the wing and tail separately. An assembly analysis will also be done. o Power Plant: . Selection of the engine and the engine bench. . The current work is based in the design of the propeller and the exhaust fumes system. o Systems and Avionics: . Selection of the most appropriate batteries. . Study of security components, such as fire extinguishers, and selection of the seat belts. . The next tasks are lighting, fuel tank and pilot joysticks for control surfaces. o CAD design: . Selection and design of the seats. . The files of the external part of the plane are ready to print and do the renders. The internal part is in process. . Calculation of the centre of gravity, for which the total mass of the instruments is needed and will be provided by the Systems Department. o Economics: . End the study of potential customers and the advertisement study. . Research on possible investors. . Design of a marketing campaign. . Start the payback analysis. o Team management: . Coordination and supervision of the progress done.  The next meeting will be the 22th November 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram.

SIGNATURES Team coordinator Supervisor

RA 170 G06-AlOn LSA 3 seats | Project report

MINUTES OF THE MEETING 10

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 22th November 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT

ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 29th November: Delivery of a draft of the report. 20th December: Project Delivery and Final Presentation

RA 171 G06-AlOn LSA 3 seats | Project report

DECISIONS MADE – ASSIGNMENTS:  General aspects of the oral presentation have been discussed.  Each department has made progress in their tasks according to the Gantt Diagram. A summary of the present situation in each area of the project is given next: o Aerodynamics: . Design of the control surfaces. . Describe the final results and conclusions in the report. o Structures: . Building of the fuselage structure . Analysis of the wing and fuselage and joint analysis of the whole aircraft. o Power Plant: . Continue the propeller design and exhaust fumes system. . It remains an overall performance evaluation, but previously the propeller design must be finished. o Systems and Avionics: . The selection of batteries and the sizing of the electric system are finished. . The tasks related to the fuel tank are completed. . The compulsory illumination is also finished; and the night flight illumination has been studied and it is not possible to incorporate. . It remains to study of the pilot joysticks, but some data from the control surfaces is needed. o CAD design: . Refinement of the designed aircraft. o Economics: . Payback analysis, income vs. Time analysis o Team management: . Coordination and supervision of the progress done.  The next meeting will be the 29th November 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Make progress in the tasks assigned to each department according to the schedule and the Gantt Diagram. o Try to finish the tasks related to the design and analysis of the aircraft in order to complete the draft of the report. o Start planning and developing the deliverable documents

SIGNATURES Team coordinator Supervisor

RA 172 G06-AlOn LSA 3 seats | Project report

MINUTES OF THE MEETING 11

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 29th November 2018, from 9:00 to 12:00 h Place: Seminary 6 – ESEIAAT

AGENDA:  Revise the work done and discuss problems or doubts  Delivery of the draft of the report in BSCW  Update the Gantt Diagram and organize the following tasks to be done  Develop the corresponding tasks for each department according to the schedule

ITEMS ON THE AGENDA: Specific tasks for each department according to the Gantt Diagram 20th December: Project Delivery and Final Presentation

RA 173 G06-AlOn LSA 3 seats | Project report

DECISIONS MADE – ASSIGNMENTS:  Some aspects of the contents of the report have been discussed. The Attachments will contain the detailed explanation of every part of the airplane, procedure, graphs, analysis, etc. The Report will be a summary of the necessary information considering that the reader should be able to understand the project without consulting the attachments. Hence, the report will mention every aspect of the final plane configuration (decisions made and method used) and will develop these ideas. The report has initially two main sections: Main alternatives and selection of the best one and Development and design of the chosen solution, but they will probably be merged in only one if necessary. Another important aspect is the fact that the report is thought to be read by engineers, so it should be technical and include de corresponding graphs and calculations.  The last tasks have been finished in order to include them in the drafts of the report.  Presentation of the drafts of the report by uploading both documents (Technical Report and Report Attachments) in BSCW to be revised by the supervisor.  General aspects of the oral presentation have been discussed. Only three or four people will do the presentation, but one person of each department (5 people) will be especially devoted to organise the presentation. Meanwhile, the other team members will be more focused on the deliverable documents.  The next meeting will be the 13th December 2018, from 10:00 to 12:00 h, in the Seminary 6 (ESEIAAT). Over the week, the following tasks will be carried out: o Refinement and correction of the corresponding aspects in the report. o Development of the deliverable documents. o Planning and creation of the project Presentation.

SIGNATURES Team coordinator Supervisor

RA 174 G06-AlOn LSA 3 seats | Project report

MINUTES OF THE MEETING 12

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Marcel Marín De Yzaguirre (MMD) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Alejandro Sans Monguiló (ASM) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 13th December 2018, from 10:00 to 12:00 h Place: Seminary 6 – ESEIAAT

AGENDA:  Revise and correct the Report and the Report Attachments.  Make progress in the deliverable documents.  Organise the presentation.

ITEMS ON THE AGENDA: 20th December: Project Delivery and Final Presentation

RA 175 G06-AlOn LSA 3 seats | Project report

DECISIONS MADE – ASSIGNMENTS:

 Discussion with the project supervisor about some aspects of the conclusions of the project. The content should be not also general conclusions of the project (such as the difficulty of the project and how we have managed to achieve it, how satisfied we are about the results obtained...) but also opinions and comments about the organization, the new competences acquired, the experience of developing such type of project... Another important aspect to comment, although the coordinator of the subject already knows, is the fact that the theory classes are taught simultaneously to the development of the project, so sometimes we have manage aspects that we haven’t seen theoretically in class, such as the project charter or the decision making.  During this week, every member of the group will evaluate the others and himself/herself in an Excel document sent by the supervisor. There are different criteria to evaluate with a mark from 1 (minimum) to 4 (maximum).  In this session, the 4 members that will do the presentation have continued preparing it, while the others are focused on the deliverable documents and the revision of the report and other documents that are nearly finished.  The next meeting will be the 18th December 2018, from 12:00 to 14:00 h, in the Conference Room (ESEIAAT). Over these days, the following tasks will be carried out: o Refinement and correction of all the documents. Everyone will read and revise the whole documents to ensure that all fits. o Correction of the style of the documents in order to standardize the format and style of graphs, images, references, etc. o Finish the presentation and start the rehearsals.

SIGNATURES Team coordinator Supervisor

RA 176 G06-AlOn LSA 3 seats | Project report

MINUTES OF THE MEETING 13

Project: Design of a LSA for 3 passengers Participants: Lluís Manuel Pérez Llera (LMPL) Department of Project and Construction Engineering Pol Bernad Serra (PBS) G06-AE-2018/19-Q1 Xavier Carrillo Córcoles (XCC) G06-AE-2018/19-Q1 Ariadna Fernández Martínez (AFM) G06-AE-2018/19-Q1 Edgar Gago Carrillo (EGC) G06-AE-2018/19-Q1 Eduard Gómez Escandell (EGE) G06-AE-2018/19-Q1 Alexandra Kalina Capdevila (AKC) G06-AE-2018/19-Q1 Carlos Medina Rodríguez (CMR) G06-AE-2018/19-Q1 Carlos Méndez Gálvez (CMG) G06-AE-2018/19-Q1 Pau Nadal Vila (PNV) G06-AE-2018/19-Q1 Carlos Pérez Ricardo (CPR) G06-AE-2018/19-Q1 David Rodríguez Pozo (DRP) G06-AE-2018/19-Q1 Itziar Ugartemendia Rodríguez (IUR) G06-AE-2018/19-Q1 Date and time: 19th December 2018, from 8:00 to 10:00 h Place: Classroom 3.5 (TR5) – ESEIAAT

AGENDA:  Revision of all documents  Rehearsal of the presentation

ITEMS ON THE AGENDA: 20th December: Project Delivery and Final Presentation

DECISIONS MADE – ASSIGNMENTS:  General revision of documents and last changes, if necessary.  Rehearsal of the presentation with all the team in a classroom with the same projector in order to prove that all slides are properly visible.

RA 177 G06-AlOn LSA 3 seats | Project report

 The full presentation has been done and some important aspects have been discussed: general ways to improve, aspects that should (not) be commented in order to adjust the time of the presentation, recommendations for the people that present, etc.  Preparation for possible doubts and questions that could be asked after the presentation.

SIGNATURES Team coordinator Supervisor

RA 178 Bibliography

[1] Michaels Selig, James J Guglielmo, Andy P Broeren, and Philippe Giguere. “Summary of Low-Speed Airfoil Data Summary of Low-Speed Airfoil Data”. In: 1 (1995), p. 315. ISSN: 1742-6596. DOI: 10.1115/1.1793208. URL: https://m-selig.ae.illinois.edu/ uiuc{\_}lsat/Low-Speed-Airfoil-Data-V5.pdf. [2] David Lednicer. The Incomplete Guide to Airfoil Usage. 2010. URL: https://m-selig.ae. illinois.edu/ads/aircraft.html (visited on 10/12/2018). [3] Harrey Riblett. Ga-airfoils. Sixth Edif. Harry C. Riblett, 1996, p. 142. [4] ASTM International. Standard Specification for Design and Performance of a Light Sport Air- plane. Tech. rep. ASTM International, 2006. URL: https://www.astm.org/Standards/ F2245.htm. [5] Propiedades del Aluminio — Universidad de Cádiz. URL: http://tablaperiodica.uca.es/ Tabla/elementos/Aluminio/Grupo1/Prop.Al (visited on 12/13/2018). [6] Adnene Tlili and Sofiene Bouhjar. “Performance Study of a Metal Matrix Composite Alloy for Aircraft Industry Use”. In: June (2015). DOI: 10.13140/RG.2.1.2292.3360. [7] Nomex aramid honeycomb 2mm. URL: http : / / www . fibermaxcomposites . com / shop / nomex-aramid-honeycombbrthickness-mmbrcell-size-32-mm-p-962.html (visited on 12/15/2018). [8] CÁTALOGO DE PERFILES DE ALUMINIO NORMALIZADOS ALUMINIUM STAN- DARD PROFILES CATALOGUE. Tech. rep. URL: http://www.extrusax.com/imagenes/ descargas/es/12/STANDARDPROFILES-PERFILESNORMALIZADOS.pdf. [9] Advisory Circular TITLE 14 OF THE CODE OF FEDERAL REGULATIONS (14 CFR) GUIDANCE MATERIAL. Tech. rep. URL: https://www.faa.gov/documentLibrary/ media/Advisory{\_}Circular/AC{\_}43.13-1B{\_}w-chg1.pdf. [10] Aviation Coatings for Corrosion Prevention. URL: https://www.corrosionpedia.com/2/ 1823/industries/transportation/aviation-coatings-for-corrosion-prevention (visited on 12/13/2018). [11] Powered Sailplanes. “Standard Specification for Design and Performance of a Light Sport Airplane 1”. In: (). [12] Libice Nad Cidlinou. “ATEC 322 FAETA Flight and Operations Manual”. In: March (2013), pp. 1–54. URL: http://www.atecaircraft.be/dossiers/Manuals/flight- manual-atec-322-faeta.pdf. [13] Ultralight. “TL-2000 Sting S4 PILOT´S OPERATING HANDBOOK”. [14] Aircraft Maintenance Manual. “TL-3000 Sirius”. In: (). [15] Pipistrel Glider Taurus M l Gliding - Pipistrel Andorra. 2003. [16] Alexander Schleichers. “ASG 29 E. Sneak into the Design Process”. [17] Strength Tests of the plane - TomarkAero | Production of UL ultralight / LSA Light sport aircraft Viper SD4 and Skyper GT9. URL: http://www.tomarkaero.com/en/manufacturing/ strength-tests.html (visited on 10/13/2018).

RA 179 G06-AlOn LSA 3 seats | Project report [18] ASTM International. “F2245-15: Standard Specification for Design and Performance of a Light Sport Airplane”. In: (2016), pp. 1–30. DOI: 10.1520/F2245-12D.2. URL: http: //compass.astm.org.proxy.library.carleton.ca/download/F2245.1538.pdf. [19] STRIAN—Structural analysis. URL: http : / / structural - analyser . com/ (visited on 12/15/2018). [20] Chun-yun. Niu. Composite airframe structures : practical design information and data. 1st Ed. Conmilit Press, 1992, p. 664. ISBN: 9627128066. [21] Michael Chun-Yung Niu. Airframe : stress analysis and sizing / Michael Chun-Yung Niu. Dragon Terrance, North Point : Hong Kong Conmilt Press, 1999. ISBN: 9627128082. [22] Rotax. ENGINE TYPE 912 | 80 hp. URL: https://www.flyrotax.com/files/Bilder/ Produkte20Rotax / Datasheets / Produktdatenblatt _ 912 _ 80hp _ rev . BRP - Rotax _ 20160823.pdf. [23] Austro Engine. AE59R-AA Engine Manual. 2011. URL: http : / / austroengine . at / uploads/pdf/EME10105r6IAE50RAA.pdf. [24] Aircraft Spruce. Rotax engine mount. URL: https://generalaviationnews.com/2009/ 07/13/rotax-engine-mount-at-aircraft-spruce/ (visited on 12/19/2018). [25] Aeromomentum. Aircraft engine, Gearbox. URL: http://aeromomentum.com/gearbox. html (visited on 12/19/2018). [26] Aircraft Spruce. Exhaust System & Components. URL: https://www.aircraftspruce. com/menus/ep/exhaustcomponents.html (visited on 12/19/2018). [27] E-PROPS. HAUTECLAIRE carbon propellers for ultralights. URL: http://www.e-props. fr/16/hauteclaireA.php (visited on 12/19/2018). [28] Cornell Law School. 14 CFR Part 91, Subpart C - Equipment, Instrument, and Certificate Requirements- US Law. 1989. URL: https://www.law.cornell.edu/cfr/text/14/part- 91/subpart-C (visited on 11/13/2018). [29] European Aviation Safety Agency (EASA). EASA Type-Certificate Data Sheet for Austro Engine AE50R series engines. Tech. rep. 2011. URL: https : / / www . easa . europa . eu / sites/default/files/dfu/EASA- TCDS- E.085{\_}Austro{\_}Engine{\_}GmbH{\_ }EA50R{\_}series{\_}engines-01-04042011.pdf. [30] Engine Manual. “IAE50R – AA”. In: (2014). [31] SKYbrary. Aircraft Fire Extinguishing Systems - SKYbrary Aviation Safety. URL: https:// www.skybrary.aero/index.php/Aircraft{\_}Fire{\_}Extinguishing{\_}Systems (visited on 11/13/2018). [32] ASTM F2316 - 12(2014) Standard Specification for Airframe Emergency Parachutes. 2014. URL: https://www.astm.org/Standards/F2316.htm (visited on 11/22/2018). [33] Stavros Salampoukidis. 3 Blade propeller. 2016. URL: https://grabcad.com/library/3- blade-propeller-2. [34] Seats & Seat Covers - Air-Tech Inc. URL: https://air-techinc.com/topic{\_}std{\_ }prods.php?catid=196{\&}pmid=18 (visited on 11/22/2018). [35] Sport Aircraft Seat Company. URL: http://www.sportaircraftseats.com/sportaircraftseats/ Home.html (visited on 11/22/2018). [36] La oficina ideal: 14m 2 por empleado | Pyme | Cinco Días. URL: https : / / cincodias . elpais.com/cincodias/2014/10/28/pyme/1414500383{\_}553511.html (visited on 12/11/2018). [37] Easa - Rps. AMC and GM to Part 21 Acceptable Means of Compliance and Guidance Material. Tech. rep. 2012. URL: https://www.easa.europa.eu/sites/default/files/dfu/ AnnexItoEDDecision2012-020-R.pdf.

RA 180 G06-AlOn LSA 3 seats | Project report [38] Susan E. Jackson, Aparna Joshi, and Niclas L. Erhardt. “Recent Research on Team and Organizational Diversity: SWOT Analysis and Implications”. In: Journal of Management 29.6 (2003), pp. 801–830. ISSN: 0149-2063. DOI: 10.1016/S0149-2063(03)00080-1. URL: https://www.sciencedirect.com/science/article/pii/S0149206303000801. [39] Seyed J. Sadjadi, Maryam Oroujee, and Mir. B. Aryanezhad. “Optimal Production and Marketing Planning”. In: Computational Optimization and Applications 30.2 (2005), pp. 195– 203. ISSN: 0926-6003. DOI: 10.1007/s10589-005-4564-8. URL: http://link.springer. com/10.1007/s10589-005-4564-8. [40] 2017 ANNUAL REPORT General Aviation Manufacturers Association. Tech. rep. URL: www. wfscorp.com. [41] General-Aviation-Manufacturers Association. 2017 ANNUAL REPORT General Aviation Manufacturers Association. Tech. rep. URL: www.wfscorp.com. [42] EASA Anual Safety Review 2017. Tech. rep. URL: https://www.easa.europa.eu/sites/ default/files/dfu/209735_EASA_ASR_MAIN_REPORT_3.0.pdf. [43] Air Charter Guide. URL: http://aircharterguide.com/Product.aspx?Prod=acg (visited on 11/15/2018). [44] IATA. IATA - Advertising. URL: https://www.iata.org/services/advertising/Pages/ index.aspx (visited on 11/15/2018). [45] Flying Magazine. Advertising Specs FLYING MAGAZINE. Tech. rep. URL: http://www. google.com. [46] “Print Speciﬁcaions for Flying Magazine”. In: (). URL: http://www.bonniermarinegroup. com/files/{\_}attachments/media{\_}kit/flying{\_}2018{\_}print{\_}specs. pdf. [47] Perfect Landing Media, LLC. - Aviation Marketing. URL: https://www.perfectlandingmedia. com/our-services/aviation-marketing (visited on 11/15/2018). [48] The Importance of Angel Investing in Financing the Growth of Entrepreneurial Ventures. Sba.gov. URL: http://www.sba.gov/advo/research/rs331tot.pdf. [49] A Guide to Angel Investors. Entrepreneur. URL: http://www.entrepreneur.com/article/ 52742. [50] 14 business angels in Spain that startups CEOs can’t miss. URL: https://startupxplore. com/en/blog/business-angels-in-spain/ (visited on 11/22/2018).

RA 181 G06-AlOn LSA 3 seats | Project report

RA 182 Appendix A

Code

The following code has been use to develop several parts of the project.

A.1 Weigth-Range Diagram

1 % Data in order to create the weight-range diagram 2 clear all ; 3 clc; 4 5 %Plane data 6 TF = 46 ;%Kg trip fuel 7 RF = 4;% Kg --> reserve fuel 8 FW = TF+RF;% FUEL WEIGHT 9 E = 35 ;%Efficiency for alpha0 10 Sw = 12.424;%[m^2] 11 CL = 0.6 ; 12 rho = 1.225 ;% needs to be re-calculated 13 PL_max = 70 *3; 14 Pl_min = 70; 15 Pl_2 = 70*2; 16 OEW = 340; 17 18 MZFW = OEW + PL_max; 19 ZFW = OEW + Pl_min; 20 ZFW2 = OEW + Pl_2; 21 22 MZFW_R = MZFW + RF; 23 ZFW_R = ZFW + RF; 24 ZFW_2R = ZFW2 + RF; 25 26 %Engine data 27 Rend= 0.85;% average 28 Ce = 75e-8;% micro-grams/J 29 C= 12;%l\h 30 g = 9.81; 31 32 %Configurations 33 MTOW = OEW + PL_max + FW; 34 MLW = OEW + PL_max + RF; 35 36 TOW = OEW + Pl_2 + FW; 37 LW2 = OEW + Pl_min + + RF; 38 LW = OEW + 2*70 + RF; 39 40 %K of breguet 41 K = 2*(Rend * E) / (Ce * g) * sqrt(2 /(CL*Sw*rho)); 42 43 R_MPL = K*log(MTOW/MLW);

RA 183 G06-AlOn LSA 3 seats | Project report

44 R_max = K*log(TOW/LW2); 45 R_MTOW = K*log(MTOW/LW); 46 47 y = [MZFW_R MTOW MTOW TOW]; 48 x = [0 R_MPL R_MTOW R_max]; 49 x=x./1000; 50 OEW_V = [OEW OEW OEW OEW]; 51 MZFW_V = [MZFW MZFW ZFW2 ZFW]; 52 MZFW_RV = [MZFW_R MZFW_R ZFW_2R ZFW_R]; 53 54 55 plot(x,y,'r',x,OEW_V,'g',x,MZFW_V,'b-',x,MZFW_RV,'c'); 56 title('Weight-Range Diagram','Interpreter','Latex'); 57 ylabel('Weight(kg)','Interpreter','Latex'); 58 xlabel('Range(km)','Interpreter','Latex'); 59 %axis([0 95000 600]); 60 legend('MTOW','OEW','MZFW','MZFW+RF','Location','SouthEast');

A.2 Gust-Airspeed envelope

1 clc; 2 clear; 3 close all; 4 5 %% Flight envelope 6 7 %Factores de carga por normativa(LSA) 8 n1 = 4; 9 n2 = -2; 10 11 %% velocidades[kts] para factor de carga4 12 Vs1 = 45.40 ; 13 Va1 = 90.8; 14 Vc = 110; 15 Vd = 160; 16 17 %Ajuste de par bolas 18 X1 = [0 Vs1 Va1]; 19 Y1 = [0 1 4]; 20 [Fit1, gof1] = CurveFitting(X1, Y1); 21 22 %Rectas 23 Rx1=[Va1 Vc Vd]; 24 Ry1=[4 4 4]; 25 26 Vx1=[Vd Vd]; 27 Vy1=[0 4]; 28 29 %% Velocidades para factor de carga -2; 30 Vs2 = 54; 31 Va2 = 76.45; 32 33 %Ajuste de par bolas 34 X2 = [0 Vs2 Va2]; 35 Y2 = [0 -1 -2]; 36 [Fit2, gof2] = CurveFitting(X2, Y2); 37 38 %Rectas 39 Rx2=[Va2 Vd ]; 40 Ry2=[-2 -2]; 41 42 Dx2=[Vd Vd];

RA 184 G06-AlOn LSA 3 seats | Project report

43 Dy2=[0 -2]; 44 %[Line2, gof]= LineFitting(Dx2, Dy2); 45 46 %% Flapn= 2; 47 Vaf = 61.8; 48 Vdf =85; 49 50 Xf = [0 42 Vaf]; 51 Yf = [0 1 2]; 52 [Fitf, goff] = CurveFitting(Xf, Yf); 53 54 %Rectas 55 Rxf=[Vaf Vdf]; 56 Rx3f=[0 Vdf]; 57 Ryf=[2 2]; 58 Ry3f=[0 0]; 59 60 Vxf=[Vdf Vdf]; 61 Vyf=[0 2]; 62 63 %% Gust envelope 64 n3=[1 5.4]; 65 n4=[1 -3]; 66 n3_1=[5.4 4]; 67 n4_1=[-3 -2]; 68 n5=[1 4]; 69 n6=[1 -2]; 70 x=[0 Vc]; 71 x2=[0 Vd]; 72 x1=[Vc Vd]; 73 74 %%Reference lines 75 vs1_v=[Vs1 Vs1]; 76 vs2_v=[Vs2 Vs2]; 77 vc_v=[Vc Vc]; 78 va1_v=[Va1 Va1]; 79 vaf_v=[Vaf Vaf]; 80 vdf_v=[Vdf Vdf]; 81 y=[-3 5.4]; 82 83 %% Plotting th fligth envelope 84 figure; 85 %n=4; 86 plot(Fit1,'b',X1,Y1); 87 hold on; 88 plot(Rx1,Ry1,'b'); 89 plot(Vx1,Vy1,'b'); 90 %n=-2; 91 plot(Fit2,'b',X2,Y2); 92 plot(Rx2,Ry2,'b'); 93 plot(Dx2,Dy2,'b'); 94 %n=2 --> flap 95 plot(Fitf,'r',Xf,Yf); 96 plot(Rxf,Ryf,'r'); 97 plot(Rx3f,Ry3f,'r'); 98 plot(Vxf,Vyf,'r'); 99 % gust envelope 100 plot(x,n3,'--g'); 101 plot(x,n4,'--g'); 102 plot(x1,n4_1,'--g'); 103 plot(x1,n3_1,'--g'); 104 plot(x2,n5,'--g'); 105 plot(x2,n6,'--g'); 106 %Reference velocities 107 plot(vs1_v,y,'--k'); 108 plot(vs2_v,y,'--k');

RA 185 G06-AlOn LSA 3 seats | Project report

109 plot(vc_v,y,'--k'); 110 plot(va1_v,y,'--k'); 111 plot(vaf_v,y,'--k'); 112 plot(vdf_v,y,'--k'); 113 114 axis([0 Vd+10 -3.5 5.5]); 115 grid on; 116 title('Airspeed\& Gust envelope','Interpreter','Latex'); 117 ylabel('Load factorn[]','Interpreter','Latex'); 118 xlabel('Airspeed[kts]','Interpreter','Latex'); 119 legend off;

A.2.1 Second grade polynomial ﬁtting

1 function [fitresult, gof] = CurveFitting(X1, Y1) 2 %% Fit:'untitled fit 1'. 3 [xData, yData] = prepareCurveData( X1, Y1 ); 4 5 % Set up fittype and options. 6 ft = fittype('poly2'); 7 8 % Fit model to data. 9 [fitresult, gof] = fit( xData, yData, ft ); 10 11 % Plot fit with data. 12 % figure('Name','untitled fit 1'); 13 %h= plot( fitresult, xData, yData); 14 % legend(h,'Y1 vs. X1','untitled fit 1','Location','NorthEast'); 15 %% Label axes 16 % xlabel X1 17 % ylabel Y1 18 end

A.2.2 First grade polynomial ﬁtting

1 function [fitresult, gof] = LineFitting(Dx2, Dy2) 2 %% Fit:'untitled fit 1'. 3 [xData, yData] = prepareCurveData( Dx2, Dy2 ); 4 5 % Set up fittype and options. 6 ft = fittype('poly1'); 7 8 % Fit model to data. 9 [fitresult, gof] = fit( xData, yData, ft ); 10 11 %% Plot fit with data. 12 % figure('Name','untitled fit 1'); 13 %h= plot( fitresult, xData, yData); 14 % legend(h,'Dy2 vs. Dx2','untitled fit 1','Location','NorthEast'); 15 %% Label axes 16 % xlabel Dx2 17 % ylabel Dy2 18 end

A.3 Propeller Design

The following code has been used to design the propeller. Not all code is presented, since most of it deals with making plots and similar tangential functions. Only the chore of the

RA 186 G06-AlOn LSA 3 seats | Project report computation is shown, as well as the variable declaration so as to show the input variables.

A.3.1 Variable declaration

1 % Solver setup 2 n_el = 100;%number of elements in blade 3 4 % Physical info 5 U = 56.5;% Speed[m/s] 6 height = 2000;% Altitude[m] 7 R = 0.72;% Radius[m] 8 N = linspace(1600,2300,501);% Rotation(rpm)(max 2400 rpm) 9 nb = 2;% Number of blades 10 r_min = 0.1;% Radius of central cone 11 AirfoilFile ='NACA4412.csv';% Airfoil 12 theta0 = 0.3;% Root torsion 13 chord0 = 0.12;% Root Chord

A.3.2 Core function This function is sufﬁcient to obtained the desired results, and its inputs are those declared in the previus snippet of code.

1 function [T,P_consumed, P_useful, efficiency] = ... propeller(n_el,U,height,R,omega,nb,r_min,AirfoilFile,theta0,chord0) 2 [Cl_mat,Cd_mat] = importAirfoil(AirfoilFile); 3 [dr,r, Theta, Chord, Sigma] = discretization(nb, n_el, r_min, R,theta0, ... chord0); 4 [lambda_forwards, lambda_sound] = nonDymensionalization(U,omega,R,height); 5 [Lambda_induced] = ... getLambdaInduced(n_el,lambda_forwards,lambda_sound,r,Theta,Sigma,Cl_mat,Cd_mat); 6 [T,M] = getSpecs(n_el,nb,r,dr,R,Theta,Chord, omega, Lambda_induced, ... lambda_forwards, lambda_sound,height, Cl_mat, Cd_mat); 7 [P_consumed, P_useful, efficiency] = Energy(T,M,U,omega); 8 end

In here we see that several functions are sequentially called, this was done to keep the code tidy and errors better localized.

A.3.3 importAirfoil Tihs function imports the angle of attack vs. lift and drag coefﬁcient curves from any csv ﬁle.

1 function [Cl_mat,Cd_mat] = importAirfoil(AirfoilFile) 2 sourceFile = readtable(AirfoilFile); 3 4 %Airfoil data 5 Polar(:,1) = table2array(sourceFile(:,1))*pi/180; 6 Polar(:,2) = table2array(sourceFile(:,2)); 7 Polar(:,3) = table2array(sourceFile(:,3)); 8 9 Polar_extra = [%Approximated values for large AoA 10 -pi/2 0 0.5; 11 -3/4*pi 0.5 0.2; 12 pi 0 0.1; 13 3/4*pi -.5 0.2; 14 pi/2 0 0.5]; 15

RA 187 G06-AlOn LSA 3 seats | Project report

16 Polar = [Polar; Polar_extra]; 17 sort(Polar); 18 19 Cl_mat = get_coefMat(Polar(:,1),Polar(:,2)); 20 Cd_mat = get_coefMat(Polar(:,1),Polar(:,3)); 21 end

The get_coefMat function is used to interpolate the curves more precisely, with cubic polynomial interpolation instead of linear interpolation. This is also necessary for the solver, since it uses a method that requires the function to be differentiable everywhere (which this method guarantees but linear interpolation does not).

A.3.4 discretization This function creates some variables necessary to work with blade elements.

1 function [dr,r, Theta, Chord, Sigma] = discretization(nb, n_el, r_min, R, ... theta0, chord0) 2 dr = (1-r_min/R)/(n_el); 3 r = r_min/R + dr/2 : dr : 1 - dr/2; 4 Theta = thetaDistribution(r,theta0); 5 Chord = chordDistribution(r,chord0); 6 Sigma = Chord*nb / (pi * R); 7 end

We see two functions being called here. These deﬁne the geometry of the blade regarding torsion and chord as a function of the distance from the center.

1 function [ Theta ] = thetaDistribution(x, theta0) 2 Theta = theta0./x; 3 end

1 function [ Chord ] = chordDistribution(x, chord0) 2 Chord = chord0*(1-x/2); 3 end

A.3.5 nonDymensionalization This function creates some nonDymensional variables. Working with nondymensional mag- nitudes simpliﬁes the code.

1 function [lambda_avance, lambda_sonido] = ... nonDymensionalization(U,omega,R,height) 2 lambda_avance = U/(omega*R); 3 [T,¬,¬] = atmosphere(height); 4 lambda_sonido = sqrt(1.4*287*T)/(omega*R); 5 end

A.3.6 getLambdaInduced This function is the heart of the program. It has only one output: Lambda_induced. This is an array which equals the nondymensional downwards airspeed caused by the propeller at each distance form the center. This is essential to obtain the angle of attack of each blade element.

RA 188 G06-AlOn LSA 3 seats | Project report The objective is to solve for lambda induced in the blade element equation at each blade element: 2 2 8(λ f + λi)λir = (r + (λ f + λi) )(Clcos(φ) − Cdsin(φ)) (A.1) To do so an initial lambda induced is proposed and both sides of the equation calculated. The two are subtracted and, if the proposed value for lambda induced was correct the result will be zero. Otherwise, we’ll obtain some non-zero value used as a measure of the error. This is done in the BEMequation function. The objective is, therefore, to minimize this error.

2 2 Error = |8(λ f + λi)λir − (r + (λ f + λi) )(Clcos(φ) − Cdsin(φ))| (A.2)

To do so a method called gradient descent is used. This method consists in calculating the derivative of some unknown function (in this case error as a function of lambda induced), and to push the independent variable towards the direction where its value decreases (increased when the derivative is negative, decreased otherwise; in proportion to the magnitude of this derivative).

Because calculating a derivative with solely two data points is problematic (heavily distorted by noise), three of them are obtained and the minimum value is used. Among the methods tested, this was the quickest to converge.

Gradient descent is performed on the Error function until it reaches a threshold (in this case a 0.1% error). Because Lambda_induced(r) is a continuous function, we can use the value obtained in a blade element to initialize the value for the following element, thus propagating the results forwards and considerably reducing the time to converge.

1 function [Lambda_induced] = ... getLambdaInduced(n_el,lambda_forwards,lambda_sound,r,Theta,Sigma,Cl_mat,Cd_mat) 2 %Lambda is calculated witha gradient descent algorithm 3 Lambda_induced = zeros(1,n_el);%Starting value 4 dLi = 1E-4;%Size of the step to compute derivative 5 steps = zeros(1,n_el);%Number of steps per blade element(for debugging) 6 for e=1:n_el%Algorithm ran per each blade element 7 for i=1:1E4 8 %A first attempt at guessing value is computed 9 [error] = BEMequation(r(e),Theta(e), ... Sigma(e),Lambda_induced(e), ... lambda_forwards,lambda_sound,Cl_mat,Cd_mat); 10 11 if abs(error)<1E-3 12 %If the error is within bounds, lambda_induced is propagated to ... the next element and the program moves on 13 if(e6=n_el) 14 Lambda_induced(e+1) = Lambda_induced(e); 15 end 16 steps(e) = i; 17 break 18 else%If the error is too high, we slightly change the value of ... lambda_induced to aproximate the derivative 19 [error0] = BEMequation(r(e),Theta(e), ... Sigma(e),Lambda_induced(e)-dLi,lambda_forwards, ... lambda_sound, Cl_mat,Cd_mat); 20 [error1] = BEMequation(r(e),Theta(e), ... Sigma(e),Lambda_induced(e)+dLi,lambda_forwards, ... lambda_sound, Cl_mat,Cd_mat); 21 22 %The derivative is calculated thrice to smooth out the noise 23 dE_dLi(1) = (abs(error) - abs(error0)) / dLi; 24 dE_dLi(2) = (abs(error1) - abs(error)) / dLi; 25 dE_dLi(3) = (abs(error1) - abs(error0)) /(2*dLi);

RA 189 G06-AlOn LSA 3 seats | Project report

26 27 [¬,j] = min(abs(dE_dLi));%The most conservative value is ... chosen 28 k = dE_dLi(j)*dLi; 29 Lambda_induced(e) = Lambda_induced(e) - k;%The program ... descends down the error function 30 end 31 end 32 end 33 end

The function that obtains the Error is the following:

1 function [error] = BEMequation(r,Theta, ... Sigma,Lambda_induced,lambda_forwards,lambda_sound,Cl_mat,Cd_mat) 2 %Obtaining relevant angles 3 Phi = atan((Lambda_induced + lambda_forwards)/r); 4 Alpha = Theta - Phi; 5 6 %Obtaining compressibility correction factor 7 Mach = sqrt((Lambda_induced + lambda_forwards)^2 + r^2)/lambda_sound; 8 9 %Obtaining dC_Fz(e) 10 Cl = eval_coefMat(Alpha, Cl_mat) / sqrt( 1 - Mach^2); 11 Cd = eval_coefMat(Alpha, Cd_mat); 12 dC_Fz = Cl*cos(Phi) - Cd*sin(Phi); 13 14 %BEM theory tells us that F1= F2: 15 F1 = (r^2+ (lambda_forwards + Lambda_induced)^2)*dC_Fz.*Sigma; 16 F2 = 8*(lambda_forwards + Lambda_induced)*Lambda_induced*r; 17 error = F1 - F2; 18 end

A.3.7 getSpecs Once the airﬂow is well deﬁned, this function obtains the the thrust(T) and torque on the shaft (M).

1 function [T,M,Alpha] = getSpecs(n_el,nb,r,dr,R,Theta,Chord, omega, ... Lambda_induced, lambda_forwards, lambda_sound, height, Cl_mat, Cd_mat) 2 %Computation of relevant angles 3 Phi = atan((Lambda_induced + lambda_forwards)./r); 4 Alpha = Theta - Phi; 5 6 %Computation of aerodynamic coefficients 7 Cl = zeros(1,n_el); 8 Cd = zeros(1,n_el); 9 for e=1:n_el 10 Cl(e) = eval_coefMat(Alpha(e),Cl_mat); 11 Cd(e) = eval_coefMat(Alpha(e),Cd_mat); 12 end 13 14 %Glauert compressibility factor 15 Mach = sqrt((Lambda_induced + lambda_forwards).^2 + r.^2)/lambda_sound; 16 if max(Mach) > 1 17 error('Supersonic flow achieved'); 18 end 19 Cl = Cl./sqrt(1 - Mach.^2); 20 21 E = Cl./Cd; 22 23 %Nondymensional forces 24 dC_Fz = Cl.*cos(Phi) - Cd.*sin(Phi);%Vertical force

RA 190 G06-AlOn LSA 3 seats | Project report

25 dC_Fx = Cl.*sin(Phi) + Cd.*cos(Phi);%Resistive forces 26 27 %Computation of speed and density to re-dymesionalize 28 [¬,¬,rho] = atmosphere(height); 29 V = sqrt(((Lambda_induced + lambda_forwards).^2 + r.^2)) * omega.*r; 30 31 %Computation of Thrust(T) and Torque(M) gradients along the radius 32 dFz = 0.5*rho*V.^2.*Chord .*dr .* dC_Fz; 33 dMz = 0.5*rho*V.^2.*Chord .*dr .* dC_Fx .*r; 34 35 %Integration of gradients 36 T = nb*sum(dFz); 37 M = nb*sum(dMz); 38 end

A.3.8 Energy This function calculates power used and consumed, as well as its ratio (the efﬁciency).

1 function [P_consumed, P_useful, efficiency] = Energy(T,M,U,omega) 2 P_consumed = M*omega; 3 P_useful = T*U; 4 if P_consumed<0 || P_useful<0 5 efficiency = 0; 6 else 7 efficiency = P_useful / P_consumed; 8 end 9 end

A.3.9 Other secondary functions The outputs of the last two functions are enough for the core function to return what it was re- quested to: Thrust, Used power and Power consumed. The program can therefore end here. However, there are some functions that are more tangential to the topic (yet still essential to the program). These are exposed here.

The atmosphere function obtains various air properties as a function of height, according to the International Standard Atmosphere.

1 function [T,P,rho] = atmosphere(h)%h in meters 2 T = 288.15 - 0.0065*h;%T inK 3 P = 101325 * (T/288.15)^-5.2586;%P in Pa 4 rho = 1.225*(T/288.15)^-6.2586;% rho in kg/m3 5 end

The following two functions are used to interpolate using third degree polynomials. Matlab has a built-in function already, but one better suited for periodic functions was written. They both are heavily commented so they are not further detailed.

1 function [coefMat] = get_coefMat(x,y) 2 % This funcion has as inputs: 3 %-x is any angle between-infinity and+infinity. 4 %-y the dependent variable ofa periodic functionf(x) with period2pi 5 % 6 % There is just one output 7 %- An array with the following columns: 8 % X0|A|B|C|D 9 % --> such thaty=A+ Bx+ Cx^2+ Dx^3 within the domainx(i):x(i+1) 10 % --> X0 isx(i), to map each polynomial to its lower bound

RA 191 G06-AlOn LSA 3 seats | Project report

11 % --> The last element wraps around the first one, so the 12 % properties of continuity and differntiablity are mantained for the 13 % entirety of the period 14 % 15 % To use this output array one must use the function 16 %- eval_coefMat(X,coefMat) 17 positions = [mod(x,2*pi), y]; 18 19 %This prelimnary array is sorted alongx: 20 positions = sortrows(positions); 21 x = positions(:,1)'; 22 y = positions(:,2)'; 23 24 % Obtaining gradients at lower and upper bounds: 25 dydx = zeros(1,length(x)); 26 for i=2:length(x)-1 27 dydx(i) = (y(i+1) - y(i-1))/(x(i+1) - x(i-1)); 28 end 29 dydx(1) = (y(2) - y(end))/(x(2)+2*pi - x(end)); 30 dydx(end) = (y(1) - y(end-1))/(x(1)+2*pi - x(end-1)); 31 32 % Coefficients obtained fromy and dy/dx at x_lowerBound and x_upperBound 33 coefMat = zeros(length(x),5); 34 for i=1:size(coefMat,1)-1 35 coefMat(i,:) = get_coefRow(x(i),x(i+1),y(i),y(i+1),dydx(i),dydx(i+1)); 36 end 37 coefMat(end,:) = ... get_coefRow(x(end),x(1)+2*pi,y(end),y(1),dydx(end),dydx(1)); 38 end 39 40 41 function [coefRow]=get_coefRow(X1,X2,Y1,Y2,dY1,dY2) 42 %This function adjusts the polynomialy=A+ Bx+ Cx^2+ Dx^3 43 % such that it passes through(X1,Y1) and(X2,Y2) with derivatives 44 % dY1 dY2 at these points 45 if X1 == X2 46 coefRow = [X1, (Y1+Y2)/2, 0, 0, 0]; 47 else 48 Xdif = X2-X1; 49 % The equations to adjust this can become quite cumbersome, but in 50 % matrix form they are somewhat more bearable 51 A = 1/Xdif^2*[ 52 (X2^3 - 3*X2^2*X1)/Xdif -X2^2*X1 (3*X2*X1^2 - X1^3)/Xdif ... -X2*X1^2; 53 6*X1*X2/Xdif X2^2+2*X1*X2 -6*X1*X2/Xdif ... X1^2+2*X1*X2; 54 -3*(X1+X2)/Xdif -2*X2 - X1 3*(X1+X2)/Xdif ... -X2-2*X1; 55 2/Xdif 1 -2/Xdif ... 1; 56 ]; 57 B = [Y1, dY1, Y2, dY2]'; 58 Polynomial = A*B; 59 coefRow = [X1, Polynomial']; 60 end 61 end

Now the one that used the output:

1 function [coef] = eval_coefMat(x, coefMat) 2 % This function recieves two inputs: 3 %-x independent variable 4 %- coefMat is the matrix generated by get_coefMat that contains the 5 % polynomial coefficients 6 %

RA 192 G06-AlOn LSA 3 seats | Project report

7 % There's only one output: 8 %-y is the dependent variable 9 % 10 11 x = mod(x,2*pi); 12 13 if xcoefMat(end,1) 14 coef = eval_Poly3(x,coefMat(end,2:5)); 15 else% Biunary search forx 16 notFound = true; 17 min_i = 1; 18 max_i = size(coefMat,1); 19 while notFound 20 i = min_i + floor((max_i-min_i)/2); 21 bigger = x > coefMat(i+1,1); 22 smaller = x < coefMat(i,1); 23 if bigger 24 min_i = i; 25 elseif smaller 26 max_i = i; 27 else 28 notFound = false; 29 end 30 end 31 coef = eval_Poly3(x,coefMat(i,2:5)); 32 end 33 end 34 35 function [y] = eval_Poly3(x,Poly) 36 y = Poly(1) + Poly(2)*x + Poly(3)*x*x + Poly(4)*x*x*x; 37 end

A.4 Gantt diagram

RA 193 Oct '18 Nov '18 Dec '18 2 3 4 5 8 9 10 11 12 15 16 17 18 19 22 23 24 25 26 29 30 31 1 2 5 6 7 8 9 12 13 14 15 16 19 20 21 22 23 26 27 28 29 30 3 4 5 6 7 10 11 12 13 14 17 18 19

Gantt Diagram G06 0h 0%

Project Deparments. 0h 0% 1. Aerodynamics 0h 0% 111.1 Airfoil study and research 0 0% E.Gago , P.Bernad 111.2 Airfoil definition 0 0% E.Gago , P.Bernad 112.1 Plant form study 0 0% E.Gago , P.Bernad 112.2 Plant form definition 0 0% E.Gago , P.Bernad 112.3 First wing analysis 0 0% E.Gago , P.Bernad 113.1 Twist defintion 0 0% E.Gago , P.Bernad 113.2 Wing -Twist Analysis 0 0% E.Gago , P.Bernad 117.1 Twist vs No Twist & wing ref... 0 0% E.Gago , P.Bernad 114.1 Dhiedral deifintion 0 0% E.Gago , P.Bernad 114.2 Dhiedral - Wing Study 0 0% E.Gago , P.Bernad 117.2 Dhiedral vs No dhiedral & w... 0 0% E.Gago , P.Bernad 115.1 Ailerons study and research 0 0% E.Gago , P.Bernad 115.2 Ailerons definition 0 0% E.Gago , P.Bernad 115.3 Ailerons configurations anal... 0 0% E.Gago , P.Bernad 116.1 Winglets study & definitions 0 0% E.Gago , P.Bernad 116.2 Winglets-Wing analysis 0 0% E.Gago , P.Bernad 117.3 Winglets vs no Winglets & w... 0 0% E.Gago , P.Bernad 117.4 Final wing configurations def... 0 0% E.Gago , P.Bernad 117.5 Final wing configuration ana... 0 0% E.Gago , P.Bernad 121.1 First tail sizing and research 0 0% I.Ugartemendia , P.Nadal 124.1 Configurations study 0 0% I.Ugartemendia , P.Nadal 124.2 Configurations definition 0 0% I.Ugartemendia , P.Nadal 121.2 Tail definition 0 0% I.Ugartemendia , P.Nadal 122.1 Airfoil definition 0 0% I.Ugartemendia , P.Nadal 122.2 Rudder analysis 0 0% I.Ugartemendia , P.Nadal 123.1 Airfoil definition 0 0% I.Ugartemendia , P.Nadal 123.2 Elevator analysis 0 0% I.Ugartemendia , X.Carrillo 131.1 Initial study and research 0 0% 131.2 Configurations definitions 0 0% I.Ugartemendia , X.Carrillo 131.3 Configurations analysis 0 0% I.Ugartemendia , X.Carrillo 141.1 Parasit drag study 0 0% C.Medina , P.Bernad , X.Carrillo 141.2 Drag optimization 0 0% E.Gago , I.Ugartemendia , P.Bernad , X.Carrillo 143.11 Static-stability analysis 0 0% E.Gago , I.Ugartemendia , P.Bernad , X.Carrillo 143.12 Static-stability conclusions 0 0% I.Ugartemendia , X.Carrillo 143.21 Dynamic-stability analysis 0 0% E.Gago , I.Ugartemendia , P.Bernad , P.Nadal 143.22 Dynamic-stability conclusi... 0 0% E.Gago , I.Ugartemendia , P.Bernad , P.Nadal 142.1 Efficencey analysis study 0 0% E.Gago , I.Ugartemendia , P.Bernad , X.Carrillo 142.2 Efficency analysis conclusio... 0 0% E.Gago , I.Ugartemendia , P.Bernad , X.Carrillo 2. Structures Dpt. 0h 0% 210.1 Study the sizing limitations f... 0 0% A.Sans , P.Nadal 210.2 Initial sizing of the aircraft 0 0% A.Sans , X.Carrillo 220.1 Study of the possible materia... 0 0% A.Fernández , C.Méndez , P.Nadal 220.2 Choice of materials 0 0% A.Fernández , C.Méndez 231.1 Sizing of the wings 0 0% A.Sans , P.Nadal 231.2 Building of the wings structu... 0 0% A.Sans , P.Nadal , X.Carrillo 232.1 Analysis of the beams 0 0% A.Fernández , P.Nadal 233.1 Analysis of the ribs 0 0% A.Sans , X.Carrillo 233.2 Joint analysis (wings) 0 0% A.Fernández , A.Sans 241.1 Sizing of the tail 0 0% A.Fernández , X.Carrillo 241.2 Building of the tail structure 0 0% A.Fernández , A.Sans , X.Carrillo 242.1 Analysis of the beams (tail) 0 0% A.Sans , P.Nadal 243.1 Analysis of the ribs (tail) 0 0% A.Fernández , C.Méndez 243.2 Joint analysis (tail) 0 0% A.Sans , C.Méndez , P.Nadal 251.Validation of the chosen engin... 0 0% A.Fernández 251.2 Analysis of the engine bench 0 0% A.Fernández , C.Méndez 261.1 Sizing of the fuselage 0 0% C.Méndez , P.Nadal 261.2 Building of the fuselage struc... 0 0% C.Méndez , X.Carrillo 262.1 Analysis of the fuselage 0 0% A.Sans , C.Méndez 262.2 Joint analysis of the whole st... 0 0% A.Sans , C.Méndez 262.3 Critical points analysis 0 0% A.Sans , X.Carrillo 262.4 Fatigue analysis 0 0% A.Fernández , C.Méndez , X.Carrillo 271.1 Analysis of the landing gear ... 0 0% A.Fernández , A.Sans 271.2 Sizing of the landing gear 0 0% A.Fernández , C.Méndez 272.1 Landing gear calculations 0 0% A.Fernández , A.Sans , C.Méndez 3. Power Plant Dpt. 0h 0% 311.1 Engine market study 0 0% E.Gómez , M.Marín , P.Nadal 311.2 Engine choice 0 0% E.Gómez , M.Marín , P.Nadal 312.1 Propeller market study 0 0% E.Gómez , M.Marín , P.Nadal 312.2 Propeller design 0 0% E.Gómez , M.Marín , P.Nadal 313.1 Gearbox market study 0 0% E.Gómez , M.Marín , P.Nadal 313.2 Gearbox choice 0 0% E.Gómez , M.Marín , P.Nadal 314.1 Engine bench market study 0 0% E.Gómez , M.Marín , P.Nadal 314.2 Engine bench choice 0 0% E.Gómez , M.Marín , P.Nadal 315.1 Overall performance evaluat... 0 0% E.Gómez , M.Marín , P.Nadal 4. Systems Dpt. 0h 0% 410.1 Research about compulsory ... 0 0% A.Fernández , C.Medina , C.Pérez 410.2 Research about complement... 0 0% C.Medina , C.Pérez 410.3 Global weight assessment 0 0% C.Medina , C.Pérez 410.4 Assessment of the possibility... 0 0% C.Medina , C.Pérez 420.1 Research on possible control... 0 0% C.Medina , C.Pérez 420.2 Weight assessment 0 0% C.Medina , C.Pérez 420.3 Choice of the control system 0 0% C.Medina , C.Pérez 430.1 Sizing of the electric system 0 0% C.Medina , C.Pérez 431.1 Batteries 0 0% C.Pérez 432.1 Power Units 0 0% C.Medina 433.1 Cabling 0 0% C.Medina , C.Pérez 441.1 Communications systems 0 0% C.Medina , C.Pérez 442.1 Flight instruments 0 0% C.Medina , C.Pérez 443.1 Complementary electronics 0 0% A.Fernández , C.Medina , C.Pérez 451.1 Fuel 0 0% A.Fernández , C.Medina , M.Marín 451.2 Fuel tanks 0 0% C.Medina , C.Pérez , M.Marín 452.1 Oil 0 0% C.Medina , C.Pérez , E.Gómez 452.2 Oil tanks 0 0% C.Medina , C.Pérez , E.Gómez 453.1 Complusory illumination 0 0% C.Medina , C.Pérez 453.2 Night flight illumination 0 0% C.Medina , C.Pérez 5. Design and CAD Dpt. 0h 0% 511.1 Search and study of the indo... 0 0% A.Sans 511.2 - Seats design 0 0% A.Sans 511.3 Dashboard 0 0% A.Sans 512.1 Wings surface design 0 0% X.Carrillo 512.2 Fuselage surface design 0 0% X.Carrillo 512.3 Rudder surface design 0 0% X.Carrillo 512.4 Elevator surface design 0 0% X.Carrillo 512.5 Landing gear design 0 0% X.Carrillo 513.1 Final 3D assembly 0 0% X.Carrillo 513.2 Assembly refinement 0 0% X.Carrillo 521.1 Technical sheets 0 0% A.Sans 522.1 Aircraft renders 0 0% A.Sans 522.2 Poster 0 0% E.Gago 531.1 Gcode files for the mock-up 0 0% E.Gago 531.2 Cofigurations for the printing 0 0% E.Gago 532.1 Printing itself 0 0% E.Gago 6. Economics Dpt. 0h 0% 611.1 - Materials' cost 0 0% A.Kalina , D.Rodriguez 611.2 - Facilities cost 0 0% A.Kalina , E.Gómez 611.3 - Human resources' costs. 0 0% D.Rodriguez , E.Gómez 611.4 - Additional costs. 0 0% D.Rodriguez 612.1 Initial marketing Campaign ... 0 0% A.Kalina , D.Rodriguez 612.2 - Study of marketing costs o... 0 0% D.Rodriguez , E.Gómez 612.3 - Study of future Marketing c... 0 0% A.Kalina , E.Gómez 621.1 - Research on Market Oport... 0 0% A.Kalina , D.Rodriguez , E.Gómez 621.2 - Research on Market Threat... 0 0% A.Kalina , D.Rodriguez , E.Gómez 621.3 - Research on Aircraft Stregt... 0 0% A.Kalina , D.Rodriguez , E.Gómez 621.4 - Research on Aircraft Weak... 0 0% A.Kalina , D.Rodriguez , E.Gómez 622.1 - Study of potential custome... 0 0% A.Kalina , E.Gómez 622.2 - Set Marketing goal 0 0% D.Rodriguez , E.Gómez 622.3 - Study of advertisement 0 0% A.Kalina , D.Rodriguez 622.4 - Design of marketing campa... 0 0% D.Rodriguez , E.Gómez 631.1 - Research on possible inves... 0 0% A.Kalina 632.1 - Income vs. time analysis 0 0% A.Kalina , D.Rodriguez 632.2 - Study of profitability margi... 0 0% D.Rodriguez , E.Gómez 7. Team management Dpt. 0h 0% 711.1 Company's hierarchy 0 0% 720.1 Deadlines checking 0 0% A.Kalina , D.Rodriguez , E.Gómez 731.1 Verification of the documents 0 0% A.Kalina , D.Rodriguez , E.Gómez 731.2 Delivery of the current docu... 0 0% A.Kalina , D.Rodriguez , E.Gómez