Design and Development of a Hybrid UAV

Design and Development of a Hybrid UAV
BRUNEL UNIVERSITY
Design and Development of a
Hybrid UAV
Abbinaya T Jagannathan (1037583)
Arturs Dubovojs (1008780)
Bennie Mwiinga (1020511)
Brett McMahon (0813201)
Carlos B. Calles Marin (1018922)
Camilo Vergara (1010295)
Primary Supervisor: Prof. Ibrahim Esat
Secondary Supervisor: Dr Mark Jabbal
Word Count: 45,854

Arthur D.
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ME5308 – Major Group Project
Abstract
The report of the project describes various design stages in detail as it was carried
out from conceptual design stage all the way to the final aircraft testing. It describes
the unique concept of fixed wing aircraft hybridised with tri-copter into a hybrid UAV.
The report describes how the configuration of such aircraft was achieved through
careful design stages, build and implementation, testing and further improvements
and suggestions.

Brett M.
Acknowledgements
The group would like to thank our supervisors and the lab technicians for their
understanding, advice and assistance during the design and build of the aircraft.
We owe a big thank you to lecturers Ibi Esat and Mark Jabbal for taking the time to
meet with the team on a weekly basis to discuss problems and solutions during the
aircraft’s design. We thank the lab technicians, in particular Kevin Robinson in the
aerospace lab for his endless advice and assistance and good humour in building
the aircraft. Additionally we thank the technicians Keith Withers, Steven Riley and
Chris Ellis for their assistance in manufacturing and testing many of the specialised
components of the aircraft, often to short deadlines.
We thank Dr Alvin Gatto for his advice and expertise in preparing for and performing
the flight testing of the aircraft and finally, thank you to the many external part
suppliers for their effort in delivering the parts needed to build the aircraft.

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Statement of Relative Contributions
Every contribution to this report has been clearly marked in the header of each page
with the author’s first name and initials of the surname.
I confirm that all work presented is original and where other sources or reports have
been referred to in the text have been referenced appropriately.
Abbinaya T
Jagannathan
Bennie
Mwiinga
Carlos B
Calles
Marin
Arturs
Dubovojs
Camilo
Vergara
Brett
McMahon
Shown on the table below is a personal statement of each individual’s contribution
and role during the project
Name Responsibilities
Abbinaya T Jagannathan
In the first term I started getting involved with
initial geometric and performance
requirements of the aircraft and later started
to move into aircraft tail sizing, aircraft
stability and control surface sizing. During the
second term, I was involved with design and
testing of components and connections
required for the aircraft. I was also involved
with the build manufacture of the fuselage
and other aircraft components. Overall, I was
working mainly on the theoretical side in the
first term and more towards the practical build
in the second term
Arturs Dubovojs

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At the beginning of the project I have
extensively contributed to different aspects of
the design process by selecting and
analyzing key aircraft parameters and
researching various topics during the
preliminary design phase such as suitable
motors selection and combination of aircraft
stability for tri-copter and fixed wing
configuration. During the detail design phase,
I have designed the tail of the aircraft as well
as contributed to the rod to rod connections. I
further Contributed by finding suitable
manufacturers of the sourced parts and
handled the orders as well as other segments
building and testing of the aircraft.
Bennie Mwiinga
In the first term I was involved in the concept
and preliminary aircraft design of the aircraft. I
also concentrated on the research and
procurement of flight control and avionics
systems that would be on the UAV so that it
would to be fully functional. I was involved in
initial sizing focusing on the VTOL propulsion
system and what would be needed to achieve
a stable VTOL UAV. In the second term I was
mainly focused on the preparation and
installation of avionics and the FCS onto the
UAV. Additionally working on developing
computer testing codes needed for
component tests such as Motor
Characterization.
Brett McMahon

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During the design and build phases of the
project, I assisted in creating virtual models of
the aircraft and its components using 3D CAD
software including Autodesk’s Inventor,
AutoCAD and Dassault Systems’ Solidworks.
During the second term, I assisted with hands
on building.
Carlos B Calles Marin
During the first term I was involved in the
initial sizing and development of the
aerodynamics of the aircraft, mainly focusing
on the wing parameters, drag estimations and
performance calculations. As the project
evolved I helped with the motor selection and
detail design of connectors and wing. As the
group leader it was also my job to manage
the other team members making sure there
was active communication between design
phases.
Camilo Vergara
During early stages of the project I was
involved in the initial geometric sizing of the
aircraft and components through the use of
Computer Aided Design (CAD) software. As
the design progressed I became responsible
for all aspects of the aircraft fuselage design,
in addition to estimation of weight and center
of gravity. Throughout the whole of term two I
was heavily involved with the build of the
UAV, and ultimately also helped out with the
avionics towards the end.

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Contents
Abstract....................................................................................................................... i
Acknowledgements .....................................................................................................ii
Statement of Relative Contributions...........................................................................iii
List of Figure .............................................................................................................. x
List of Tables.............................................................................................................xv
Nomenclature..........................................................................................................xvii
1. Introduction ......................................................................................................... 1
1.1. Motivation...................................................................................................... 1
1.2. Project Description........................................................................................ 3
2. Literature Review ................................................................................................ 4
2.1. State of The Art Technology.......................................................................... 4
2.2. Market Analysis............................................................................................. 9
3. Requirements.................................................................................................... 12
3.1. Regulations ................................................................................................. 12
3.2. Aims and Objectives ................................................................................... 13
Aim.................................................................................................................... 13
Objectives ......................................................................................................... 13
3.3. Mission Profile............................................................................................. 14
4. Design Process................................................................................................. 16
4.1. Concept Design........................................................................................... 16
4.1.1. Individual Proposals ............................................................................. 16
4.1.2. Quality Function Deployment................................................................ 35
4.1.3. Group Concept ..................................................................................... 37
4.2. Preliminary Design ...................................................................................... 41
4.2.1. Weights................................................................................................. 43
4.2.2. Aircraft Sizing: Constraint Analysis....................................................... 44
4.2.3. Aerodynamics....................................................................................... 48
Lift Equation.......................................................................................................... 48
Wing Geometry..................................................................................................... 48
Aerofoil Selection.................................................................................................. 52
Tail Sizing............................................................................................................. 58
The tail configuration......................................................................................... 59

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Volume Coefficients: ......................................................................................... 60
Optimum tail arm and tail plan form area: ......................................................... 61
Tail Aerofoil: ...................................................................................................... 62
Drag...................................................................................................................... 63
Control Surface Sizing: Wing ............................................................................ 70
Lift Curve Slope Numerical Prediction: Wing and Aileron/Flaps........................ 73
4.2.4. Centre of Gravity .................................................................................. 78
4.2.5. Stability and Control: Standard Take-Off and Landing ......................... 81
4.2.6. Stability: Vertical Take-Off and Landing ............................................. 100
4.2.7. Structures ........................................................................................... 113
4.2.8. Computer Aided Design and Technical Drawings .............................. 131
4.2.9. Propulsion........................................................................................... 139
5. Avionics and Flight Control ............................................................................. 150
Components ....................................................................................................... 150
Main Control Scheme and sensor array ............................................................. 151
PID Controller ..................................................................................................... 153
Related software and Full system schematic...................................................... 154
Sonar and Noise Reduction................................................................................ 155
APM Anatomy..................................................................................................... 156
Manual Transition ............................................................................................... 158
6. Component Testing......................................................................................... 160
6.1. Stress Tests .............................................................................................. 160
6.1.1. Rods ................................................................................................... 160
6.1.2. Connections........................................................................................ 166
6.2. Motor Characterisation.............................................................................. 170
6.2.1. Set-Up ................................................................................................ 170
6.2.2. Calibration of Equipment .................................................................... 174
6.2.3. Procedure for Testing ......................................................................... 174
6.2.4. EDF Results ....................................................................................... 177
6.2.5. Motor Results ..................................................................................... 179
7. Build & Manufacturing Methods & Materials in Chronological Order............... 182
7.1. Logistics .................................................................................................... 182
7.1.2................................................................................................................ 183

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Key materials used.......................................................................................... 184
7.2. Fuselage ................................................................................................... 187
7.3. Connections .............................................................................................. 193
7.4. Wings & Tail.............................................................................................. 194
7.5. Propulsion ................................................................................................. 195
7.5.1. VTOL Propeller motors Mount............................................................ 195
7.5.2. Lander 90 mm EDF Mount ................................................................. 197
7.6. Avionics..................................................................................................... 198
Servo Installation............................................................................................. 198
Camera, Live stream & OSD Installation......................................................... 198
GPS & Compass installation ........................................................................... 199
Sonar Installation ............................................................................................ 199
Flight Control System & Air speed sensor Installation..................................... 199
ESC Installation and calibration ...................................................................... 199
Battery and Power Distribution Harness (PDH) Installation ............................ 200
Telemetry (MAVlink) Installation...................................................................... 200
8. After Build Testing........................................................................................... 201
8.1. Flight Tests................................................................................................ 201
8.1.1. Horizontal Flight Test 1....................................................................... 201
Horizontal Flight Test 2....................................................................................... 205
8.1.2. Vertical Flight Tests ............................................................................ 208
9. V-n Diagram.................................................................................................... 210
10. Budget.......................................................................................................... 212
11. Conclusion.................................................................................................... 214
12. Improvements and Further Research........................................................... 215
Landing Gear...................................................................................................... 215
STOL Propulsion ................................................................................................ 215
Power Plant Upgrade.......................................................................................... 215
Transmitter/Camera Range ................................................................................ 216
Aircraft Systems.................................................................................................. 216
Material upgrade................................................................................................. 216
Transition attempt............................................................................................... 216
Autonomous Flight capabilities ........................................................................... 216

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Tilt rotors............................................................................................................. 217
Ultimate load testing on wings and fuselage....................................................... 217
Stress testing of 3D printed parts ....................................................................... 217
13. Bibliography.................................................................................................. 218
14. Appendices................................................................................................... 228
14.1. Appendix A – Technical Details ............................................................. 228
Roskam Constraint Analysis ........................................................................... 228
Wing Profile Analysis: Additional Aerofoils...................................................... 229
Comparison between the two analysis methods in XFLR5 ............................. 230
Results for Vortex Lattice Method (VLM) and the Panel Method with a
percentage comparison................................................................................... 230
Motor Test Code ............................................................................................. 235
14.2. Appendix B – Project Plan and Management ........................................ 237
14.2.1. Gantt Chart...................................................................................... 237
14.2.2. Logistics .......................................................................................... 239

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List of Figure
Figure 1 The IAI Mini Panther in level cruise flight [1] . .............................................. 4
Figure 2 Sitter type UAV the V-Bat from MLB [4]. ...................................................... 5
Figure 3 The Orbis from Santos Labs in hover [6]...................................................... 5
Figure 4 The Latitude Engineering HQ hybrid Prototype [8]....................................... 6
Figure 5 The Wingcopter V13CH in VTOL mode. ...................................................... 7
Figure 6 Advanced VTOL Technologies' Hammerhead [10] ...................................... 7
Figure 7 Bell Eagle Eye Tiltrotor UAV [11]. ................................................................ 8
Figure 8 QTW-UAV developed by Chiba University, Japan [14]. ............................... 8
Figure 9 Graph showing basic relation between small UAVs..................................... 9
Figure 10 Maximum Endurance vs. Maximum Take-off weight for a range of UAVs
[19] ........................................................................................................................... 11
Figure 11: The STOL mission profile........................................................................ 14
Figure 12: The VTOL mission profile........................................................................ 15
Figure 13 approximate sketch of the concept idea................................................... 18
Figure 14 Individual Proposal Concept by Bennie Mwiinga...................................... 21
Figure 15 The Doak VZ-4 by Doak Aircraft Company [24]. ...................................... 21
Figure 16 Blown Flight Control Concept................................................................... 23
Figure 17Tube and Tray Fuselage Concept............................................................. 25
Figure 18 Tube and Tray Fuselage as Used by Hobby Flyers ................................. 26
Figure 19 Concept Sketch of an initial idea.............................................................. 29
Figure 20 Concept Design Sketch............................................................................ 33
Figure 21 A quality function deployment (QFD) Matrix............................................. 35
Figure 22 Flow Diagram showing Design Stages..................................................... 42
Figure 23 Force Diagrams: a) Forces on a climbing aircraft, b) Forces on aircraft at
constant bank angle [27]. ......................................................................................... 44
Figure 24 Constraint Analysis for the UAV............................................................... 46
Figure 25 Wing Geometry, Note: dimensions in millimetres..................................... 50
Figure 26 Induced Drag Factor Vs. Taper and Aspect Ratio [28]............................. 51
Figure 27 Typical Cambered Aerofoil [31]................................................................ 52
Figure 28 Lift Coefficient Vs. Moment Coefficient Analysis of different profiles........ 53
Figure 29 Initial and Final profile Comparison. ......................................................... 54

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Figure 30 XFLR results for the final wing configuration. (a) Moment Force and chord
wise lift distribution. (b) Spanwise lift distribution. (c) ISO view of lift and lift
distribution................................................................................................................ 55
Figure 31 Polars: (a) Variation of Lift coefficient with AoA. (b) Variation of the drag
coefficient with lift coefficient. (c) Variation of lift to drag ratio with AoA. .................. 56
Figure 32: Tail design procedure as illustrated by Mohammad Sadraey. [40].......... 59
Figure 33 Total Drag Decomposition........................................................................ 63
Figure 34 Drag velocity curve................................................................................... 66
Figure 35 CD Vs. CL Polar for the wing and the aircraft........................................... 68
Figure 36 Comparison of the Lift Curve Slopes using different predicting methods:
Online database, XFLR5 and ESDU sheets............................................................. 74
Figure 37 How to obtain Trailing Edge Angle ...................................................... 76
Figure 38 Wing Curve slopes with control surface deflections. ................................ 77
Figure 39 Force balance kit to acquire aircraft CG location...................................... 79
Figure 40 Front load with dual dead weight batteries and back-up 4000 mah main
battery ...................................................................................................................... 80
Figure 41: Tail incidence angle vs. Moments generated. ......................................... 82
Figure 42: Graphs indicating the derivatives and for stable and instable
aircraft conditions. .................................................................................................... 84
Figure 43 Wing and tail forces.................................................................................. 86
Figure 44 Statically Stable and Unstable pitching moment curves........................... 87
Figure 45 Final aircraft CG Lift configuration............................................................ 88
Figure 46: control surface effectiveness parameter vs. control surface to lifting
surface chord ratio. [40]............................................................................................ 92
Figure 47 Shows the rudder curve slope with deflection angles of ±20 degrees...... 97
Figure 48 Shows the elevator curve slope with deflection angles of ±20 degrees. .. 97
Figure 49 Longitudinal CG Envelope for Project vehicle .......................................... 98
Figure 50 Tri-copter configuration with reference axes. ......................................... 101
Figure 51 Pitch up by using Rotor 1. ...................................................................... 102
Figure 52 Roll in the Clockwise direction................................................................ 102
Figure 53 Roll in the Counter Clockwise direction.................................................. 102
Figure 54 Yaw authority of a tri-copter. .................................................................. 103
Figure 55 Mass Flow of air through rotor in hover.................................................. 104
Figure 56 Altitude Hold (Hover) with all 3 rotors..................................................... 106

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Figure 57 Mass Flow of air through rotor in vertical climb. ..................................... 107
Figure 58 A level vertical climb by the tri-copter..................................................... 108
Figure 59 Flow of air through the rotor in forward flight.......................................... 109
Figure 60 Rotor Disc showing Azimuth angle......................................................... 110
Figure 61 Full model of UAV at a hover. ................................................................ 111
Figure 62 Full model of UAV in transition............................................................... 111
Figure 63 UAV model in full horizontal flight........................................................... 112
Figure 64 Monocoque fuselage design [61] ........................................................... 114
Figure 65 Truss fuselage structure [32].................................................................. 114
Figure 66 Semi-monocoque Fuselage [32] ............................................................ 115
Figure 67 Global Hawk Cutaway [64]..................................................................... 115
Figure 68 Falco Cutaway diagram [64]................................................................... 116
Figure 69 Cutaway of the ScanEagle [64].............................................................. 116
Figure 70 Bonding in progress of the Demon UAV composite structure [65] ......... 117
Figure 71 Loading on a triangular structure [68]..................................................... 118
Figure 72Skeletal frame of the fuselage................................................................. 119
Figure 73 Landing Gear Positioning for Proper Weight Distribution [71] ................ 121
Figure 74 Moveable Landing Gear Concept........................................................... 122
Figure 75 ABS Landing Gear Mount - Broken During Aircraft Assembly................ 123
Figure 76 ANSYS principle stress analysis on bulkhead displaying key on the left 124
Figure 77 demonstration of typical wing structure [75] ........................................... 127
Figure 78 Single Spar Wing Connection ................................................................ 131
Figure 79 Double Spar Wing Connection............................................................... 132
Figure 80 Moveable Landing Gear Mount.............................................................. 133
Figure 81 Computational Stress Test Result for Basic Landing Gear Mount ......... 134
Figure 82 Computational Stress Test Result for Lightweight Landing Gear Mount 134
Figure 83 Nose Vertical Lift Fan Skeletal Structure................................................ 135
Figure 84 Detail View of the Tongue and Groove Assembly Method ..................... 135
Figure 85 Initial fuselage concept........................................................................... 136
Figure 86 First Full Group CAD Aircraft Design ..................................................... 136
Figure 87 Structure and connections of various components within the aircraft..... 137
Figure 88 Top, front and side views of the final CAD model................................... 138
Figure 89 T/W vs Maximum amp draw................................................................... 140
Figure 90 T/W vs EDF price ................................................................................... 141

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Figure 91 EDF unit weight vs Thrust Capability ..................................................... 141
Figure 92 STOL mission current comparison for the initial and final endurance
calculations ............................................................................................................ 148
Figure 93 VTOL mission current draw comparison for the initial and final endurance
calculations. ........................................................................................................... 148
Figure 94 General control scheme of the UAV [87]................................................ 151
Figure 95 ArduPilot Mega 2.6 from 3D Robotics .................................................... 152
Figure 96 Schematic of MPU-6000. ....................................................................... 152
Figure 97 Block diagram of tri-copter control include 2 gain values [57]. ............... 153
Figure 98 Control allocation by a controller on a tri-copter..................................... 153
Figure 99 Example of Cascade Control.................................................................. 154
Figure 100 Cascaded PID used by APM [88]......................................................... 154
Figure 101 MaxBotix XL MaxSonarEZL0. .............................................................. 155
Figure 102 Sonar EM Noise reduction modification. .............................................. 156
Figure 103 APM 2.6 anatomy................................................................................. 156
Figure 104 Phases of flight during the transition maneuver from hover to horizontal
flight........................................................................................................................ 158
Figure 105 Cantilever Load Testing Arrangement.................................................. 160
Figure 106 Cantilever Physical Stress Test Results Graph.................................... 161
Figure 107 Three Point Physical Stress Test Results Graph ................................. 162
Figure 108 Three Point Physical Stress Test Results Graph ................................. 163
Figure 109 demonstration of carbon fiber rod deflection with cantilever point loading
............................................................................................................................... 165
Figure 110: Experimental setup of the test conducted (left) and a drawing of the
component (right) ................................................................................................... 167
Figure 111: Load vs Tensile extension for the 10mm diameter hole ...................... 168
Figure 112: Load vs Tensile extension for the 20mm diameter hole. ..................... 169
Figure 113 Thrust bench and NI High USB carrier used for motor characterisation.
............................................................................................................................... 171
Figure 114 EDF Mount for thrust bench. ................................................................ 171
Figure 115 Motor Mount for thrust bench. .............................................................. 171
Figure 116 80A ESC Turnigy Superbrain............................................................... 172
Figure 117 Turnigy KV-RPM Meter. ....................................................................... 172
Figure 118 National Instruments Hi-Speed USB Carrier. ....................................... 172

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Figure 119 Turnigy 4000 mAh LiPO Battery (6s). .................................................. 172
Figure 120 PWM changing the angle of a dc motor [95]. ....................................... 173
Figure 121 Sample calibration curve for the test bench. ........................................ 174
Figure 122 Numeric Loading for EDF and trend line. ............................................. 177
Figure 123 Thrust Results for the EDF................................................................... 178
Figure 124 Thrust efficiency of two and three bladed propellers [96]. .................... 179
Figure 125 VTOL Motor test with different propellers............................................. 180
Figure 126 Current Draw of the motor for any given thrust. ................................... 181
Figure 127 Laser cutting the aft EDF bulkhead ...................................................... 187
Figure 128 Fuselage during initial Epoxy resin stage of construction (left), tilting
Propeller mount (Right) .......................................................................................... 188
Figure 129 rear view of the front Bulkhead displaying the nose gear mechanism.. 189
Figure 130 drilling axle holes on the non-vertical mounting plate of the carbon fiber
Landing gear .......................................................................................................... 191
Figure 131 Rear landing gear assembly................................................................. 192
Figure 132 Fuselage structure with back-up rear undercarriage (left), Nose gear
(right)...................................................................................................................... 192
Figure 133: Schematic of assembly of the aluminum VTOL motor mounts............ 193
Figure 134: Load tests conducted on the P400 ABS plastic (left) and the 3mm (right)
plywood motor mounts. .......................................................................................... 196
Figure 135 EDF Mount to the fuselage, Side view (left), top view (right)................ 197
Figure 136 Reinforced rear landing gear mount..................................................... 203
Figure 137 Strengthened Retro-fit Nose Landing Gear.......................................... 204
Figure 138 Second flight test ground roll demonstration ........................................ 205
Figure 139 Second flight test tip stall demonstration.............................................. 206
Figure 140 Second flight test landing stall demonstration ...................................... 207
Figure 141 UAV in Tri-copter mode........................................................................ 209
Figure 142 V-n Diagram and Gust Loading graph.................................................. 211
Figure 144 Roskam Constraint Analysis ................................................................ 228
Figure 145 To obtain for Step 4 in Table 17 [49]....................................... 231
Figure 146 To obtain for Step 5 in Table 17 [49].................................. 231

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List of Tables
Table 1: Sketch of concept design idea.................................................................... 16
Table 2: Key parameters of individual concept design ............................................. 17
Table 3 Key parameters of individual proposal ........................................................ 18
Table 4 Individual proposal by Bennie Mwiinga ....................................................... 20
Table 5 Individual Concept Design........................................................................... 28
Table 6 Individual Concept #4.................................................................................. 31
Table 7 Typical Aircraft Parameters. [26] ................................................................. 33
Table 8 House of Quality table, How’s vs How’s ...................................................... 36
Table 9 Group Concepts .......................................................................................... 38
Table 10 Constraint Analysis Equations, obtained from Mattingly et All [27]............ 45
Table 11 Constraint Analysis Parameters. ............................................................... 46
Table 12 Different Wing Geometry Design Aspects ................................................. 49
Table 13 Wing Geometry Parameters...................................................................... 50
Table 14: Effects of changes in tail volume coefficients ........................................... 61
Table 15 Drag Components of the aircraft for cruise, 22.2 m/s. ............................... 69
Table 16: Time to achieve specific bank angles...................................................... 71
Table 17 Process to attain the lift curve slopes of the wing and the deflected control
surface. ..................................................................................................................... 75
Table 18 Parameters and Results............................................................................ 76
Table 19 Lift variation with control surface deflection............................................... 77
Table 20: Horizontal and vertical tail design details.................................................. 82
Table 21: Static and dynamic stability requirements. [40] ........................................ 83
Table 22: Methods of determining the location of neutral point [38] [53] .................. 85
Table 23: Control Surface Functions........................................................................ 89
Table 24: Rudder deflection required during various landing at various crosswind
velocities. ................................................................................................................. 95
Table 25 Rudder and elevator curve slope results using ESDU method, to be used in
the control surface sizing.......................................................................................... 96
Table 26 showing properties of similar thickness plywood material strength [72] .. 125
Table 27 properties comparison of foam core wing reinforced with carbon fibre spars
to balsawood ribbed structure reinforced with carbon fibre spars [77] ................... 128
Table 28 Battery properties for a suitable range of products [86]........................... 146

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Table 29 VTOL and STOL endurance.................................................................... 149
Table 30 Necessary Avionics Components for the UAV. ....................................... 151
Table 31 APM Anatomy Glossary. ......................................................................... 157
Table 32: Results obtained from the stress test conducted on the 3D printed
component. ............................................................................................................ 168
Table 33 Testing Procedure for Motor Test............................................................ 176
Table 34 List of Suppliers and any comments surrounding orders and components
delivered................................................................................................................. 183
Table 35 Mid-Project Budget.................................................................................. 212
Table 36 Final-Project Budget................................................................................ 213
Table 37 Wing profile: Additional Analysis ............................................................. 229
Table 38 VLM and Panel Method Result comparison from XFLR5 ........................ 230
Table 39 STOL Mission profile and current specifications...................................... 232
Table 40 VTOL Mission profile and current specifications...................................... 232

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Nomenclature
 VTOL - Vertical Take-Off and Landing
 STOL - Standard Take-off and Landing
 EDF - Electronic Ducted Fan
 MAC - Mean Aerodynamic Chord
 CAD Computer Aided Design
 AR - Aspect Ratio
 Reynolds Number
 - Weight Force
 - Mass
 - Air density
 - Aspect Ratio of Horizontal Tail
 - Volume Coefficient of horizontal tail
 - Volume Coefficient of Vertical tail
 - Area of Horizontal Tail
 - Area of Vertical Tail
 - Optimum arm of the Horizontal Tail
 - Optimum arm of the Vertical Tail
 - Area of Wing
 - Centre of Gravity
 - Static longitudinal stability
 - Dynamic longitudinal stability
 - Static directional stability
 - Dynamic directional stability
 - Location of Neutral Point
 - Location of Centre of Gravity
 - Location of Aerodynamic Centre
 - Static Margin
 - Efficiency of stabiliser
 - Wing Curve Slope
 - Horizontal tail Curve Slope
 - Vertical tail Curve Slope
 - Aircraft static longitudinal Stability Derivative
 - elevator effectiveness directive
 – control surface chord effectiveness parameter
 - Wing Root Chord
 - Wing Tip Chord
 - Inboard location of ailerons
 - Outboard location of ailerons
 - Aileron deflection
 - Aircraft rolling moment coefficient
 - Approach Velocity
 - Rolling Moment
 - Steady state roll rate
 - Wing Area

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 - Horizontal Tail Area
 - Vertical Tail Area
 Se - Elevator Area
 ce – Elevator Chord
 be - Elevator Span
 – Induced Drag
 - Bank Angle
 - Second moment of area
 ̇ - Steady State Roll Rate
 - Lift at Take-off
 - Rotational Velocity
 - Moments about the aerodynamic centre
 - Horizontal tail curve slope
 - Aircraft Lift coefficient at take-off
 - Maximum profile lift coefficient.
 – Wing angle of attack
 - Angle of attack
 - downwash angle
 - Horizontal tail incidence angle
 - Angle of attack horizontal tail
 - Elevator chord effectiveness parameter
 - Elevator deflection
 - Elevator effectiveness derivative
 - Static longitudinal stability derivative
 - Distance between aerodynamic centre and main landing gear
 - Distance between centre of gravity and main landing gear
 -Curve slope of wing-fuselage combination
 - Vertical distance between thrust provider and centre of gravity
 - Lift coefficient at cruise incidence angle
 - Lift coefficient at zero wing incidence angle
 - Crosswind velocity
 - Aircraft side force due to crosswind
 - Sideslip angle
 - aircraft sideslip derivative
 - aircraft sideslip derivative
 - Vertical tail lift curve slope
 - Aircraft control derivative
 - Efficiency of vertical tail
 - Vertical tail side wash gradient
 - Rudder deflection
 - Aircraft crab angle during crosswind landing
 - Centre of aircraft side projected area
 - Aircraft centre of gravity
 - Aerofoil training edge angle

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 - Slope of lift-coefficient curve with incidence for two-dimensional
aerofoil in incompressible flow
 - Theoretical slope of lift-coefficient curve with incidence for two-
dimensional aerofoil in inviscid, incompressible flow
 - Slope of lift-coefficient curve with control deflection for two-
dimensional aerofoil in incompressible flow
 - Theoretical slope of lift-coefficient curve with control deflection for
two-dimensional aerofoil in inviscid, incompressible flow

Arturs D.
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1. Introduction
1.1. Motivation
The UAV industry is developing rapidly and currently is a very popular topic due to
the broad variety of applications of this technology. This increase in popularity
creates higher demands in the field and calls for constant technological advance.
There have been a number of projects which have involved designing, building and
programming of unmanned aerial, some of these types of projects concentrated on
developing autonomous flight and obstacle avoidance techniques. Usually such
projects concentrate on one aspect since it is very time consuming especially as a
university project where time is very limited and not all of it be dedicated to a project.
Combining few of such aspect together is a lot harder and challenging due to time
limitations and limited resources.
Airports and aircraft carriers take up a lot of space due to lengthy runways which,
creates some problems finding the airfields launching aircrafts even for home built
RC planes. On the other hand, fixed wing aircrafts are very efficient for distance
travelling and staying in the air longer comparing to rotor crafts. After individual
concept designs have been proposed, the group selected a collaborated idea and it
was decided to design, build and program a hybridised UAV of VTOL aircraft and
fixed wing aircraft. This idea combines both concepts and enables using the benefits
of both. After further research into hybrid UAVs, the decision was to develop a
combination of tri-copter with fixed wing aircraft. At the time there was a quad-copter
hybridised with a fixed wing aircraft however, a tri-copter has not been done before.
Quad-copter combined with a fixed wing aircraft has at least 5 thrust generators
unless it is a tilt rotor where, tri-copter has one less motor which can decrease
overall weight of the vehicle. Such design could be developed further to improve
specifications and achieve better performance and parameters than the existing
UAVs on the market.
A project like this have not been done at Brunel University previously therefore,
success of this project would be a great achievement for the university and could
even attract publicity and improve university rating.

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The project’s success could give a big contribution to university’s teaching curriculum
regarding UAVs and improve it for future students. Due to tight time constraints of
this project there will be a lot of room for improvement of this project, further
development and expansion therefore, this project could be used as a dissertation
topic for future years for individuals as well as groups. Once the UAV is fully ready it
could be used as a learning platform for students about UAVs.
At last, a project like this is a great way to apply the knowledge gained through 4
years of university where theory is applied to a real life problem to which the solution
is yet to be found. Not only it is a way to apply the knowledge but also, there is a lot
to be learned during the course of the project, aspects which have not been covered
during the course of education. Besides the application of theoretical knowledge it
allows to compare the theoretical input to outcome of the result and feasibility of
theory in practice. Most important such project would allow each group member to
carry out self-assessment and evaluate what they have achieved over the 4 year
period of the course.

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1.2. Project Description
Initially the project idea was to design and build a UAV however, each group
member had a different idea and view of the project. After proposing individual
concepts, a combined idea based on individual inputs of the group members was
carried forward, to design a fixed wing aircraft with short take-off/landing (STOL)
ability as well as a vertical take-off/landing (VTOL) capability. Further decisions
were made to design an electrical vehicle rather than using gas/fuel due to health
and safety regulations and limiting time constraints. The fixed wing part of the aircraft
is straight forward, conventional concept which have been used for almost a century
now however, for VTOL is there were few considerations such as the number of
rotors and their configuration. The optimum tri-copter configuration was selected to
decrease the stability complexity. Also, tilt rotor configuration was excluded due to its
increased complexity with additional servos and mechanics for tilt mechanisms. So
the final decided concept design was of a fixed wing aircraft with one thrust
generator for horizontal flight combined with 3 vertical motors (tri-copter
configuration) for VTOL.
The final, fully developed aircraft was planned to have the option of programmable,
fully autonomous flight which does not need external, manual inputs to operate as
well as a remote control capabilities. The aircraft required to be equipped with a
camera allowing live streamed video to the user for surveillance purposes. However,
due to very limited time constraints of this project, the realistic objectives had to be
decided which involved designing and building the aircraft with functioning tri-copter
configuration as well as the fixed wing, horizontal flight configuration. The aircraft
had to be remote controllable for both configurations. If the main objectives are
achieve, the secondary, optional objectives such as transition segment between
VTOL and horizontal flight can be worked on. If the main objectives are achieved the
project would be considered successful.

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2. Literature Review
2.1. State of The Art Technology
The idea of an aerial vehicle that can perform both VTOL like a helicopter as well as
STOL like a fixed wing aircraft is not a new one. However limitations in technology as
well as the complexity of the associated control system has prevented widespread
development of these types of vehicles especially those with an autonomous nature.
Recently a handful of fully autonomous hybrids have been unveiled some are still
prototypes while others are fully functioning production models. Below are some
examples of these state of the art V/STOL aircraft. Being able to hybridize rotorcraft
and fixed wing aircraft provides the opportunity for further applications of UAV/S in
roles that would normally be exclusive to either one or the other.
The mini panther is a smaller version of its larger relative ‘The Panther’ and weighs
12 kg (shown in figure 1). It is however the newest Iteration in the family to date. It
has 3 propeller motors, the front two motors tilt upwards. In conjunction with the aft
propeller; that is permanently in the normal position relative to the aircraft, the mini-
panther is able to perform vertical take-off and transition into cruise, as well as
transition from cruise to stable hover flight. The third motor on the aft section of the
fuselage acts as the third arm of a tri-copter when the front two are tilted, this allows
for yaw control in hover and vertical flight.
Figure 1 The IAI Mini Panther in level cruise flight [1] .
Another approach to producing a VTOL aircraft is the V-Bat from MLB (shown in
figure 2). It is an alternate solution to a VTOL UAV as it is designed as a sitter type
aircraft. As a sitter type aircraft the V-bat begins its flight in a vertical position on the
ground and then hovers to a predetermined altitude [2]. At said altitude it is able to

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transition autonomously from hover to cruise and vice versa. Developed with funding
from DARPA, the military version also includes a 6 foot extending arm to pick up
objects whilst at hover close to ground level [3]. This UAV is capable of flight up to
15,000 feet altitude, with a maximum endurance of 10 hours.
Figure 2 Sitter type UAV the V-Bat from MLB [4].
Similar to the MLB V-Bat are systems that only use a ducted body design (Figure 3).
These systems are sitter-type VTOL that also have a number of rotors (normally 4)
placed in an X or + formation similar to a quad rotor. The rotors and control surface
are all enclosed within a duct body. These duct type UAV are able to transition from
hover to horizontal flight by tilting themselves forward and increasing and vectoring
the thrust generated by one motor. An example of this design is the Santos Lab
Orbis which currently uses a hydrogen fuel cell and has a span of 3.8m [5].
Figure 3 The Orbis from Santos Labs in hover [6]
It is well known that rotor craft are able to perform VTOL in the most efficient
manner. By hybridizing a rotor craft with a fixed wing craft the benefits of both types
of vehicles can be maintained. Latitude Engineering, a small drone company from

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Tucson, Arizona USA has developed such a hybrid [7]. Calling it the Hybrid Quad
rotor (HQ) (shown in Figure 4 below) it is simply a hybrid between a quad rotor and a
fixed wing aircraft. It weighs 27 Kg and has 4 electric motors that it uses to hover and
1 gas powered motor mounted on the aft of the aircraft to provide thrust for forward
flight. The HQ is still in development but Latitude Engineering has been able to
maintain a hover and transition to forward flight.
Figure 4 The Latitude Engineering HQ hybrid Prototype [8].
Another project that has taken the same approach as Latitude Engineering is the
Wing copter V13CH project by Jonathan Hesselbarth. The Wing copter (shown in
Figure 5) also hybridizes a fixed wing and a quad rotor, however, the Wing copter
utilizes its 4 electric brushless motors to produce thrust in forward flight and in VTOL.
Utilizing a set of swivel arms; on which the four motors are mounted, the Wing copter
is able to perform transition from hover to horizontal forward flight and vice versa.
This a novel solution that also requires a control system that can keep the aircraft
stable while transition is being performed. The transition performed by the
Wingcopter however not an automated one is and is instead manually controlled
(with some assistance from on board flight control). By utilizing the rotary heads on a
radio control transmitter, the controller is able to vary the tilt of the four motors and
transition into horizontal flight.

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Figure 5 The Wingcopter V13CH in VTOL mode.
Similarly employing a tilting mechanism to achieve VTOL is the Hammerhead
developed by David Howe & Lyndon Caine of Advanced VTOL Technologies [9]
(shown in Figure 6). It has a canard that helps improve the UAVs stall characteristics
as well as the ability to thrust vector in order to limit pitch divergence [9].The
hammerhead employs twin counter rotating electric rotors stationed on a tilting stub
wing assembly which AVT claims “minimises pitch, roll and yaw coupling” [9]. The
hammerhead is capable of performing either STOL or VTOL. In STOL mode the tilt
wing stub is positioned such that the rotors produce a thrust propelling the UAV
forward. In VTOL mode the tilt wing stub is positioned vertically allowing the
hammerhead to take off and hover like a helicopter.
Figure 6 Advanced VTOL Technologies' Hammerhead [10]
Another example of hybrid VTOL design is the Bell Eagle Eye UAV (see Figure 7).
Developed by Bell Helicopter – Textron, Texas USA. It is a Tilt rotor UAV capable of
VTOL which it achieves by tilting its nacelles in the appropriate direction to either
perform VTOL or STOL. The Bell Eagle Eye has a payload weight of 90kg, a
maximum speed of 200kt and an endurance of 8 hours [11] [12]. It also utilizes an
automated flight control system to assist in performing transition.

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Figure 7 Bell Eagle Eye Tiltrotor UAV [11].
Use of tilting mechanisms can also be found in other hybrid VTOL aircraft such as
the QTW-UAV by Chiba University, Japan (see Figure 8). The QTW-UAV uses 4
rotors placed on tilting wings that swivel to the vertical position to gain altitude and
then swivel forward to transition to horizontal flight. It utilizes 4 electric rotors giving it
a payload weight of 5kg, endurance of 15 min and a max speed of 81kt [13].
Figure 8 QTW-UAV developed by Chiba University, Japan [14].

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2.2. Market Analysis
In order to design a vehicle fit for purpose, market research was undertaken to find a
suitable starting point to work to. Typical dimensions and functionalities of existing
real world UAVs were used to compare the sizes the aircraft under development
should be close to. Below is a diagram showing a triangular relationship between
wing span, total length, and Max Take-off weight with a logarithmic scale.
Figure 9 Graph showing basic relation between small UAVs
AV RQ 11 B Raven [15] Bayraktar Mini UAS [16]

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AV Wasp III [17] Innocon Micro Falcon [18]
The graph on Figure 9 Graph showing basic relation between small UAVs represents
information gathered on similar sized small UAV platforms. This was used during the
early stages of initial geometric sizing to ensure the values that were being
calculated for the aircraft were within a set industry trend. In essence it was an early
rudimentary matching plot to define design tendencies for wingspan, vehicle length,
and Maximum Take-Off weight. Both the first converged group concept as well as
the current project aircraft design are listed on the plot. An important performance
parameter to note is that the cruise velocities were not included in analysis as values
could not be found for all vehicles looked at. However when averaged out along
other smaller UAVs, typical cruise velocities were around 20 .
Once the vehicle size had been constricted to a hypothetical box, performance
characteristics were then researched for a broad range of UAVs. Figure 10below
was sourced out from a thesis from MIT, displaying the relationship between
Maximum take-off weight and endurance in hours for a wide variety of UAVs. The
theoretical project aircraft would lie between 1 kg to 10 Kg along the bottom axis.
This range on the plot has a trend of a maximum endurance of 1-2 hours.
Considering the vehicle in question would be a hybrid, the added dead weights, the
use of all-electronic propulsion as well as a vertical flight mission segment which
would drain a higher amount than the usual battery draw during straight level un-
accelerated flight would all impact this maximum endurance value. As a result it
would be expected to be significantly lower in reality.

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Figure 10 Maximum Endurance vs. Maximum Take-off weight for a range of UAVs [19]
This market research of current UAVs in service was invaluable in producing
information to use as a guideline to the aircraft design process. After every main
phase was completed values of the aircrafts performance, and sizing was checked
with current vehicles such as these. From this initial selection of fixed wing UAVs,
the variety was narrowed down to include specialized aircraft which were applicable
to similar mission profiles, which is to say a fixed wing UAV that has VTOL
functionality.

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3. Requirements
3.1. Regulations
Every country has its own aviation regulations. The Civil Aviation (CAA) Authority
states that any UAV exceeding a weight of 7 kg need to be certified or approved.
Due to this, the main priority of the aircraft was its weight. It was decided that the
aircraft’s maximum take-off weight was kept under 6.5 kg and this was ensured
throughout the design process.
However, the CAP 722 and CAP 393 Air Navigation Order states that aircraft that
weigh less than 7 kg should also follow some regulations depending on whether they
are being used for commercial purposes or not. The regulations for aircraft with a
mass of less than 7 kg states that the aircraft should abide by appropriate
operational constraints in order to ensure public safety. The regulations are based on
the flying operation being conducted and the potential risks to any third party.
General principles for UAV operations outside segregated airspace should follow an
approved “detect and avoid” system and avoid crowded areas. It is also important for
the aircraft to not fly beyond the visual line of sight. The CAP 722 also states that the
aircraft should be flown such that the pilot controlling it can take manual control at
any point of time and fly the aircraft out of danger. It also states that the aircraft
should not be flown above 400 feet at any point of time. Further details of regulations
applicable to this UAV can be found in the CAP 722 and CAP 393 of the Civil
Aviation Authority Air Navigation Order. These regulations were kept in mind
throughout the build of this UAV. [20] [21]

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3.2. Aims and Objectives
Aim
To design, manufacture and test an aircraft of fixed wing configuration hybridised
with a tri-copter. Horizontal flight capabilities of the aircraft have to be demonstrated
as well as the VTOL capability using a remote control transmitter.
.
Objectives
1. To use aircraft design techniques and approaches to design a fixed wing tri-
copter hybrid aircraft
2. To test and prove the suitability of load critical components of the aircraft
3. To build and test the final selected UAV design
4. To program the aircraft and have the avionics ready for both horizontal and
vertical flight
5. Carry out a demonstrative flight for each of the configurations of the UAV

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3.3. Mission Profile
The purpose of the UAV is to be able to perform a surveillance and reconnaissance
role. This requires the aircraft to be equipped with a camera and live stream
capabilities. The main concept of this project is to design an aircraft that is capable of
a Vertical Take-off & Landing as well as a Short take-off and landing (V/STOL). This
would allow for the UAV to be launched from different environments where a runway
is not available such as urban areas with limited airspace or launches from the sea.
Due to the tight time schedule and the complexity of the project the transition
between vertical and horizontal flight is not a priority for the project. Initially the UAV
has to be able to perform a VTOL mission: take off vertically, climb, hover, climb
further, hover at new altitude and descend to land. For the STOL mission it has to
perform a separate mission profile where the aircraft has to: take-off, climb, cruise,
loiter, descend and land. It is expected that parts of the mission profile segments are
performed autonomously by an on-board autopilot.
Figure 11 and 12 below illustrate the mission profiles allocated for the STOL and
VTOL flight. The calculations of the mission profiles were done based on the current
drawn from the batteries to be used. The total time of UAV operation for the STOL
mission should be at least 9 minutes. The UAV should be able to operate for around
3 minutes when performing the VTOL mission.
Figure 11: The STOL mission profile.
0
5
10
15
20
25
30
35
0 100 200 300 400 500 600
Altitude(m)
Time (s)
STOL mission profile

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Figure 12: The VTOL mission profile.
Even though the transition is not a set priority for the project at the moment, it is most
likely to be attempted once the VTOL and STOL missions have been successfully
completed. In the case of the transition being attempted, the aircraft should be able
to operate for around 4 minutes.
In order to design and develop the UAV an appropriate design procedure must be
performed. Since the project is aerospace orientated, the avionics and electronics to
be used by the UAV are to be off-the-shelf components that are marginally modified
to assist in the completion of the project objectives. The aircraft should be safe to
operate, undergo a safety assessment and meet minimum requirements for this type
of aircraft. The UAV is to be equipped with a flight control system capable of
autonomous flight and manoeuvres. However, for the purposes of build and testing
the UAV will be remotely controlled by a human pilot via a radio receiver and
transmitter controller.
0
2
4
6
8
10
12
14
16
0 10 20 30 40 50 60 70 80 90
Altitude(m)
Time (s)
VTOL mission profile

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4. Design Process
4.1. Concept Design
4.1.1. Individual Proposals
Abbinaya T Jagannathan
Table 1: Sketch of concept design idea
The aim of this design concept is to have an autonomous UAV that is suitable for
reconnaissance and surveillance purposes. The objective of this UAV design was to
have an aircraft that is aerodynamically sound and also attempt to achieve a
vertical/short take-off or landing. A push propeller was to be used at the fuselage to
provide the main forward thrust. The integrated motors at the wing are meant to
provide vertical thrust. Since the design consists of only 2 vertical thrust providers,
the feasibility of VTOL is uncertain and if this is the case then, the goal is to achieve
a short take-off by using the vertical thrust providers.
The fuselage in this design is shaped as an aerofoil to have a body that is able to
contribute to the lift force produced by the aircraft. An aerofoil with a high thickness
to chord ratio is to be used for the fuselage. The figure above illustrates an idea of
the UAV proposal. From the figure it can be seen that the other key aspects of this
concept are the V-tail and the use of winglets. The V-tail was selected mainly
because of the push propeller to be used at the rear end of the fuselage. Other
reasons for V-tail selection are that it has a smaller size therefore; it will be lighter

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and have a smaller wetted area which would result in drag reduction. The V-tail
configuration also uses fewer control surfaces compared to a conventional tail.
These control surfaces are called ruddervators and are a combination of rudder and
elevators. [22]
The use of winglets was also considered in the concept for a number of reasons.
Winglets are small wing-like lifting surfaces that are fitted at the tip of the wings for
the purpose of reducing the trailing-vortex drag. As a result, this would increase the
lift generated on the aircraft. [23]
The materials considered for this design were a combination of foam and carbon
composites (main airframe) which are both lightweight materials that can take high
loads.
Category Abbinaya T Jagannathan
Mission Type Reconnaissance/ Surveillance
Environment Outdoor
Design Type Modular
Modular Options Camera
TO (Take-Off Type) V/STOL
L (Landing) V/STOL
Powerplant 1× Push Propeller & 2 × Rotor Integrated in Wings
Wing Medium/High Fixed Wing
Tail V-Tail
Airframe Foam & Carbon Composite
Landing Gear Fixed
Endurance(Prospective) 60 min
Altitude 50-100 m
Glide Capability (Inc. Design) Yes
Radio Controlled Back up Yes
Autonomous Yes
Table 2: Key parameters of individual concept design

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Arturs Dubovojs
Figure 13 approximate sketch of the concept idea
Category Arturs
Dubovojs
Mission Type Reconnaissance/
Payload delivery
Environment Outdoor
Design Type Modular
Modular Options Cargo/Camera
Take-Off Type Catapult
Landing Parachute/STOL
Power plant 1x Push Motor
Wing High Fixed
Wing/Joint Wing
Tail Conventional/
Tailless (Joint
Wing)
Airframe Foam
Landing Gear Fixed
Endurance (Prospective) 2 hours
Altitude 100m
Glide Capability (Inc. Design) Yes
Autonomous Yes
Table 3 Key parameters of individual proposal
Initial idea was to Design and build a high endurance UAV with recon and payload
delivery capabilities for outdoor use. The aircraft had to be able to carry a live

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streaming camera and a small payload. It has to be designed to be able to take off
and land in a standard manner as well as been optimised for catapult mechanism
and a parachute for emergency, vertical landing. Take-off and landing would require
a landing gear, for weight reduction and less complexity a fixed landing gear would
have been used. Aircraft required being equipped with only one push/pull propeller
either on the nose or top of the fuselage with an electric motor capable of providing
enough thrust. The aircraft would have to be able to carry out autonomous flight as
well as a radio controlled option.
After the basic mission and guidelines been set an investigation into wing types was
carried out. A joint wing configuration was selected for the aircraft because of its
ease of implementing it to a small scale UAV in comparison to a full scale aircraft.
Since the aircraft should have a catapult mechanism it should be relatively small and
compact for ease of deploying in unequipped circumstances, joint wing configuration
reduces the required wing span. Since the joint wing configuration is relatively new
invention it has not been implemented on many full scale aircrafts and there is
mainly research being carried out on using it on smaller scale, high altitude UAVs.
The joints between the wings would act as winglets, reducing induced drag. The
option of having a tail additionally to the joint wing was still available. A tail would
improve the aircrafts manoeuvrability by adding yaw capability which joint wing
aircrafts lack. Also by making a separate control surface – tail, the second wing
would become a lifting surface therefore, would generate more lift.

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Bennie Mwiinga
Bennie Mwiinga
Mission Type Recon/Surveillance
Environment Outdoor & Urban
Design Type Modular
Modular Options Camera
L (Landing) V/STOL
Power-plant 3x Ducted Fans
Wing Mid Fixed Wing
Tail H-Tail
Airframe Carbon Composite/Foam
Landing Gear Fixed
Endurance(Prospective) 180 Mins
Altitude 60+ m
Glide Capability (Inc.
Design)
No
Autonomous Yes
Table 4 Individual proposal by Bennie Mwiinga
The current military and commercial applications of UAV/S has increased in the past
14 years at an exponential rate. Also the environments in which these systems are
expected to operate have changed and outdoor operations require novel design
solutions in order to accomplish requirements such as quick deployment, long range
and endurance. Military operations are more frequently being carried out in urban
environments where close quarter combat is conduct. The ability to have a UAV that
can be used in such an environment would assist troops in such operations by being
able to be deployed on the spot to conduct reconnaissance and surveillance. This
proposal is shown in Figure 14.

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Figure 14 Individual Proposal Concept by Bennie Mwiinga.
In order to achieve these requirements a V/STOL type system is proposed. V/STOL
would allow for the UAV to be deployed in any terrain or environment without the
requirement of having a prebuilt runway. The V/STOL system in this proposal is
achieved by having two ducted fans mounted on the wing tips that are able to tilt for
vertical and horizontal flight. This approach was taken by the Doak Aircraft Company
in developing their Doak VZ-4 (Figure 15) in 1958.
Figure 15 The Doak VZ-4 by Doak Aircraft Company [24].
An additional ducted fan is then placed on the aft of the aircraft to provide additional
forward thrust for STOL and transition to horizontal flight. The use of ducted fans
may seem to be the wrong choice due to the lower efficiency when compared to a
purely fan based propulsion. As can be seen in equation 4 Area (A) if increased will
increase the amount of force created. A prop or fan has the advantage that it can
have a wide area and move more mass of air and thus create more force. A ducted
fan however has a smaller area and to move the same amount of air as a prop or fan
it has to increase the rate at which it accelerates the air while moving a smaller
amount.

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(4.1.1.1b)
Where (4.1.1.2b)
∴ (4.1.1.3b)
Where (4.1.1.4b)
(4.1.1.5b)
At the time the Doak was engineered this statement would be true, however,
modern day ducted fans (electric) are designed and engineered to operate at higher
efficiencies as loses are reduced by designing more aerodynamic shrouds and
utilizing more efficient electric motors. These new generation of ducted fans have
been used by many entities ranging from RC Hobbyists to UAVs developed by
companies like Honeywell and Boeing. These ducted fans utilise higher efficiency
motors capable of high rpm and specially designed props to produce high level
performance.
A modular type design would be used in this proposal. It would also allow for the
UAV to service other sectors such as agricultural monitoring and scientific research,
as the UAV would be capable of being outfitted with varying types of payloads such
as FLIR or other EVS.

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Brett McMahon
Figure 16 Blown Flight Control Concept
This design is inspired to some extent by the DEMON concept demonstrator [25]
developed by BAe Systems, Cranfield and other universities which used jets of air to
control the aircraft in flight. Using low profile wing tip fans, the aircraft will be able to
perform roll manoeuvres using pulses of air from the appropriate fan. Short take off
performance would be possible using both fans together to provide lift in addition to
that provided by the main wing. The aircraft would be propelled through the air using
a large pushing motor of sufficient power (with reserve) to achieve a desired flight
speed. The aircraft would be all electrically powered for relative simplicity and safety
compared to petrol powered motors and their required ancillaries. Two batteries are
In-wing motors,
provide short
take off and
airbourne roll
control
Camera
provides visual
target ‘hit’ data
Rod and slot
construction
method, removable
wings if needed
Internal components
laid out with respect
to a selected Centre
of Gravity position
Internal avionics components
mounted on aluminium cruciform,
outer surfaces of foam or vacuum
formed plastic shell pieces
Infra-red range
finder provides
altitude data to
flight controller

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envisaged for the aircraft, one large capacity battery used for the forward thrust
provider and wing tip motors while a second smaller battery would independently
supply the flight control systems. The primary potential benefits of the wing tip fan
arrangement is the level of simplicity that can be achieved over the otherwise
complex arrangements associated with moving flight control surfaces, whilst also
allowing for some moderate weight savings.
The wings will be designed so they can provide the lift required with minimal drag at
moderate flight speeds. Sizing of the aircraft must be sufficient such that the body
can adequately house all the avionic components as well as provide a small degree
of flexibility for adjustments (moving or addition of components) which might be
needed during refinement. The wings must be able to generate more than enough lift
required to support the aircraft’s mass, with a margin of safety for gusting conditions
or lack of performance from the propelling motor.
The flight control system will most likely be based on the Ardupilot series of control
boards available at many remote control hobby shops. The benefit of these control
systems is their general availability, relatively cheap price and extensive
programming support from the remote control operator community. Power
distribution, relays, wiring and programming must be defined separately and
accounted for in the budget to be defined later.
Materials will be determined that possess sufficient strength for the specified
application whilst having minimal weight. All materials must be readily available from
local suppliers and be sourced at the lowest price possible. Construction techniques
used must be such that the overall component weights are kept as low as possible.
Processes and tools required for fabrication and assembly must be simple and
readily available through the university workshops or specially bought in by
ourselves. Total aircraft cost must not exceed departmental budget constraints to be
discussed with the project supervisors.
After discussions within the group, the main problem with this aircraft concept was
the general lack of elevator and tail sections required for pitch and yaw control. Roll
control is managed primarily by the wing motors but additional aileron control
surfaces may need to be added for higher speed flight, to be determined as the
design progressed.

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Conceptual Fuselage Design
The fuselage is the main body of the aircraft. Depending on the aircraft layout, the
fuselage is responsible for housing fuel, weapon stores, cargo, avionics equipment
and passengers. For the UAV aircraft, the fuselage will house all of the electronic
parts such as the flight control boards, navigation systems, cameras and radio
receivers as well as the batteries. In addition to housing of the internal parts, the
fuselage must also provide the fixing structure for the aircraft’s other component
parts such as the wings, motors and the landing gear.
The fuselage concept shown below in Figure 17 uses a thin walled outer skin and a
removable avionics tray, which slides along runners extending the full length of the
fuselage. This approach has previously been used by remote control aircraft builders
as shown by Figure 18. The avionics tray allows for very tight packaging of the
internal electronics and batteries of the aircraft and would be fully removable for
servicing and adjustment, save for a few connections to the aircraft control surfaces
such as ailerons, elevators, rudder and the motors.
Figure 17Tube and Tray Fuselage Concept

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Figure 18 Tube and Tray Fuselage as Used by Hobby Flyers
Dependant on material availability and price, the thin walled (approximately 0.5 to
1mm thick) fuselage material of plastic or aluminium would allow for the required
strength and low weight but could be further reinforced with the addition of stringers
running top and bottom of the fuselage (along with the metal runners for the avionics
tray located along the mid-line).
This simple construction with easy access via the tray arrangement could be made
almost entirely from off the shelf parts and materials. The wing box junction would
have to be specially made to encompass the mid wing section. This junction would
then be mated to the fuselage tubular section using fore and aft fastening wedges.
Screws would be inserted through the fuselage skin and into the wedges, holding it
(and therefore the wing box section) firmly in place.
Toward the nose of the aircraft, a swivelling fan arrangement would be fixed between
two bulkheads allowing the fan to be adjusted using a dedicated servo mechanism to
counteract unwanted yaw from the other lift fans when the aircraft is in vertical flight.
Further ahead of the swivelling fan arrangement would be a clear plastic nose cone,
inside which the video camera and GPS receiver could be mounted, providing a
clear view ahead and upward. The nose section would likely be a two part assembly
made from either specially ordered injection moulded acrylic or made on campus
using a vacuum forming method.
This concept however was not used for the final aircraft as the required plastic or
metal thin walled, large diameter fuselage tubing could not be sourced at reasonable
cost. A case was therefore made for a bespoke fuselage design which could be

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tailored to the changing requirements as the design and the parts specification
evolved.

Camilo V.
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Camilo Vergara
Category Camilo
Mission Type Recon/Multi-role
Environment Outdoor
Design Type Modular
L (Landing) V/STOL
Powerplant 2x Tilt EDF & 1x Fixed
VTOL EDF
Wing Mid/High Fixed Wing
Tail Boom tail
Airframe Metal/Foam
Landing Gear Fixed
Endurance(Prospective) 90-120 Mins
Altitude 100m
Design)
Powered
Radio Controlled Back
up
Yes
Autonomous Yes
Table 5 Individual Concept Design
Every UAV design incorporates various design solutions and ideas as well as a set
level of autonomy that is dictated by its mission parameters and operating
environment. For the purpose of exploring different approaches to the design
problem presented, various types of configurations were looked at.
The potential for a VTOL system on a fixed wing reconnaissance drone is significant.
Not only would it eliminate the need for a runway and be easy to retrieve which
essentially makes it deployable from any location, but from an intelligence aspect, it
would be able to do what no other normal type of fixed wing UAV could do, which is
stop and hover mid-air over a point of interest allowing for a detailed inspection of

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the area ahead instead of having to perform circuits or flyby’s around a target. In
order to achieve this, a particular solution was researched. This initial VTOL system
came in the form of a tilt rotor design, which has its origins from the V-22 Osprey.
Upon First glance there seems to be a very select few UAV’s currently on the market
with this type of technology, the biggest being the Bell 'Eagle eye' which was a true
representative of the twin tilt rotor Osprey. There are a couple more variations to
mention, the first being Israeli Aerospace Industries ‘Panther’ VTOL UAV utilizing a 3
motor design with 2 being mounted on the wings, and a third around the rear section
of the fuselage. Another worth mentioning is a prototype VTOL aircraft called the
‘phantom swift’ by Boeing which incorporated 4 ducted fans, two being located at
opposite wing tips, and two being incorporated into the fuselage itself. Below is a
diagram of the initial concept inspired from existing real world solutions.
Figure 19 Concept Sketch of an initial idea
The challenge of such a design would include aspects like the complexity of the
control system integrated into the autonomous nature of the UAV. From a design
perspective, several considerations must be taken into account for an aircraft of this
nature. with regard to the power-plants themselves, this would include the position of
the rotors from the longitudinal center of gravity, the connections between the servos
and motors themselves, the physical connections to the wing or fuselage depending
on the motors location, and the individual propeller rotation direction. The power
output of the motors utilized would be a design aspect, due to the extra weight of the

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control system for a tilt rotor design, the motors selected must have enough power to
lift the aircraft vertically, as well as perform well at horizontal flight. Electronic Ducted
Fan (EDF) systems have rarely been used for VTOL hybrid applications; so the
aircraft would be experimental by nature.

Carlos C.M.
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Carlos Calles Marin
Category Carlos
Mission Type Recon/Surveillance
Environment Outdoor
Design Type Modular
TO (Take-Off Type) STOL
L (Landing) STOL
Power plant 1x Push Motor
Wing High Fixed Wing
Tail Boom Tail
Airframe Balsawood
Landing Gear Fixed
Endurance(Prospective) 60 min
Altitude 100 m
Design)
Yes
Autonomous Yes
Table 6 Individual Concept #4
The initial idea for the concept was very conservative. The initial requirements for the
design were “very short landing and take-off or hand launched” and some aspect of
autonomous behaviour.
From the design point of view these are the different configurations considered:
1. Type of wing – High, Medium or Low
2. Power plant – Tractor, Pusher or both
3. Tail Type – V-Tail, Standard or Boom Tail
4. Wing – Sweep, Taper, Dihedral, Wash-in/out
A high wing was chosen because the aircraft had to be possibly hand launched,
which means that it needs good stability at low speeds until it reaches cruise. High
mounted wings have better lateral stability than medium or low mounted.
Considering the power plant the pusher configuration was chosen to improve the
aerodynamic performance. The flow behind the propeller no longer has to flow over
the wings, which would be the case with a normal tractor power plant. There are
some situations were both types are used in to increase the thrust provided, acting in

Carlos C.M.
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line with the centre of gravity. In this case it would not be necessary to have so much
thrust.
Regarding the type of tail needed V-Tail was quite interesting, reducing the amount
of drag produced by the tail. For the pusher propeller configuration chosen a Boom-
Tail is required to correctly place the motor. This doesn’t discard the V tail, but it
changes it into inverted V. the problem with V tail is that it requires more expert
knowledge of coding to have autonomous behaviour. The standard Boom Tail was
chosen to avoid any control problems.
In terms of wing design, sweep would not be an option because it is intended for
high speed flight and this aircraft would fly at relatively small speeds. Taper would
increment our performance, by reducing wing tip vortex downwash effect. The
optimal wing shape would be elliptical to have a uniform span wise distribution of lift
with the lowest induced drag possible. Due to its hard manufacturing process the
elliptical wing shape can be approximated by a straight tapered wing, with a taper
ratio of 0.3-0.4, hence it would be desirable to have a taper ratio around those
values. The drawback of uniform lift distribution is that stall is reached evenly
throughout the wing plan form; therefore washout would have to be considered to
have more margin for error. Dihedral would increase the stability, but decrease the
effective span of the aircraft, therefore it is not going not be part of the concept.
To control the aircraft autonomously some readily available micro processing
computers were thought of. There are two options which are Arduino and Raspberry
Pi, these are open source platforms with widely available codes that can perform as
an autopilot for the aircraft.
Figure 20 is a representation of the initial concept where the taper, boom tail and
pusher propeller can be observed.

Carlos C.M.
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Figure 20 Concept Design Sketch.
To achieve good aerodynamic performance and the main mission aim, short take
off/landing, the concept to should have a similar look to that of a glider with high
aspect ratio, minimal weight and streamlined. RC trainer aircraft were also taken into
consideration since they are supposed to be easy to handle, which would benefit the
autonomous nature of the aircraft.
Table 7 shows typical design aspects for a trainer aircraft:
Trainer Aircraft Glider
Wingspan (b) 152 cm 152 cm
AR 6-7 8-10
Overall Length 127 cm 102 cm
Wing Area (S) 0.4216 m2
0.323 m2
Flying Weight 1.81 Kg 0.454 Kg
Wing Loading (W/S) 59 N/m2
20 N/m2
Table 7 Typical Aircraft Parameters. [26]

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To have an idea of what speeds the aircraft would be flying at an initial estimate of
the weights was made as a group effort, and came to the conclusion that the aircraft
would weight about 3.7 Kg.
√
It can be shown that the aircraft would need a velocity ( ) of with a wing area
( ) of to fly. Using the Aspect Ratio formula the span can be determined.
√
The wingspan comes to be around .

Arturs D.
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4.1.2. Quality Function Deployment
Figure 21 A quality function deployment (QFD) Matrix
Figure 21 demonstrates House of Quality, How’s vs How’s which demonstrates the
importance of different parameters in terms of percentage as well as the importance
in relation to other parameters.
Main two parameters were determined to be the Electrical efficiency and Hover
Capability. Hover capability if one of the main objectives of the project therefore it is
one of the main parameter on the other hand if the system is not efficient enough the
current will be drawn very rapidly during hover mode since there are 3 motors that
would be operating at the same time therefore, it is essential for the electrical system
to be efficient otherwise there would not be enough electrical power to fulfil the
mission profile. The third most important parameter for Table 8 is the overall weight
of the aircraft for the same reason the previous one. The lower the weight of the
aircraft the less current it draws, the less power required to operate it which leads to
improve in efficiency. Those are three main aspects of the aircraft which were
concentrated the most on during the design and the built phase of the project.

Arturs D.
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Nevertheless, the other parameters of the aircraft are very important and failure to
reCGnise that could lead to unsuccessful project.
Aircraft Attribute Score Importance (%) Relative Importance (%)
Electrical Efficiency 84.33 100.00 15.99
Hover Capability 71.00 84.19 13.46
Weight 65.67 77.87 12.45
Cruise Speed 41.67 49.41 7.90
Reliability 40.78 48.35 7.73
Range 38.78 45.98 7.35
Drag 37.44 44.40 7.10
Manufacturing Costs 29.44 34.91 5.58
Easy to operate 23.44 27.80 4.45
STOL Distance 23.00 27.27 4.36
Noise 22.11 26.22 4.19
Rate of Climb 17.67 20.95 3.35
Max g-loading 12.11 14.36 2.30
Maneuverability 8.56 10.14 1.62
High quality image 6.11 7.25 1.16
Operation beyond line of sight 5.22 6.19 0.99
Table 8 House of Quality table, How’s vs How’s

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4.1.3. Group Concept
Category Initial Group Concept Design Final Group Concept Design
Mission Type Reconnaissance/ Surveillance Reconnaissance/ Surveillance
Environment Outdoor Outdoor
Design Type Modular Modular
Modular Options Camera Camera
Take-Off Type V/STOL V/STOL
Landing V/STOL V/STOL
Power plant 1x STOL EDF & 3x VTOL EDFs 1x STOL EDF & 3x VTOL Propeller Motors
Wing High Fixed Wing High Fixed Wing
Tail H-Tail Boom Tail
Airframe Balsa Ply wood, Foam, Carbon Composites and Balsa
Landing Gear Fixed Fixed
Endurance (Prospective) 30 min 30 min (depending on motor thrust test)
Altitude 40 - 60m 40 - 60m
Glide Capability (Inc. Design) Yes Yes
Radio Controlled Back up Yes Yes
Autonomous Yes Yes

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Pictures
Initial Group Concept Design Final Group Concept Design
Table 9 Group Concepts

Abbinaya T.J.
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Table 9 outlines the basic ideas of the group concept designs. After discussion and
consideration of the various ideas proposed an initial group concept design was
confirmed. The final concept design was evolved with changes being made to design
in order to make initial design more feasible. When comparing the sketches of the
initial group concept and the final group concept a lot of differences can be noticed.
One such change made is the change in placement of the VTOL motors from the
wing tips to the behind the wings. The VTOL motors were initially placed at the wing
tips and then moved to be integrated into the wing. This change was made in order
to reduce the loads on the wing tips as well as to reduce the moment arm that could
possibly flutter and generate some moments on the wing. This change was also
opted because it could possibly reduce the drag generated on the aircraft.
However, this configuration of the VTOL motors being integrated in to the wing was
again changed to be placed behind the wings just as in the current concept. This
was done due to some stability issues that came up with the VTOL tricopter system.
Another major change in concept design was made with material selection. Balsa
was selected for the initial concept design as it is a traditional material used in small
UAV and remote controlled planes. This changed when in depth research was
conducted on various other materials. With the knowledge gained from research, the
traditionally used balsa was replaced with other modern materials and composites.
A lot of other changes with design had also been made to the aircraft before deciding
on the final concept because of conflicting design issues between the systems
required for the VTOL mission and the systems required for the STOL mission.
Successful completion of the given design concept should provide an UAV that is
capable of doing a vertical take-off and landing as well as a standard take-off and
landing. The propulsion system available for VTOL is a tri-copter made of propeller
motors. The tri-copter is designed such that there are two counter rotating motors
behind each wing and one motor at the nose of the aircraft. The motor at the aircrafts
nose would be equipped with a tilt mechanism in order to counteract the yaw force.
The thrust for the normal flight would be provided by an Electric Ducted Fan (EDF).
The EDF would be placed at the aft of the fuselage in order to provide push
propulsion.

Abbinaya T.J.
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Other aspects of the design such as the wing and tail are kept fairly simple with no
dihedral, twist or sweep. A boom tail was selected mainly to keep it from interfering
with the EDF placed at the aft of the fuselage. Even though the concept design had
been decided an open minded was kept during the phase of the design and
manufacture to cope with unanticipated problems that could arise.

Carlos C.M.
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4.2. Preliminary Design
Once there is a concept idea, the project may move forward into the preliminary
design phase. It consists of making the concept idea reality, all the requirements and
limits are applied in the aircraft/copter design calculations to obtain a unique
outcome to complete the objectives stated.
The preliminary design was decomposed into smaller subsections, allowing different
group member to focus on individual tasks and work more efficiently. These
subsections, along with the rest of the project, may be seen in the workflow diagram
below, Figure 22

Carlos C.M.
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Figure 22 Flow Diagram showing Design Stages
Requirement
s
Market Analysis
Technology
Concept
Design
Initial weight
estimation
Flight control
and avionics
Aerodynamics
Wing Geometry
Initial Tail Sizing
Propulsion
Initial Layout
Initial Costing
Preliminary Design
Initail CG
estimation
Performance
Check
Wing Aerofoil
Selection
Final Wing
Design
Final Tail Sizing
Control Surface
Design
Lift Curve for
Control Surfaces
Fuselage Design
Landing Gear
Optimal CG and
NP
Motor and
Propeller
Selection
Stability and
Control
Frozen Design
Preliminary
Budget and
Costs
Detail
Design and
Build
Structure
Technology
Implementation
and Sourcing
CAD Model
Material and
Equipment
Logistics
Experiments and
Testing
Final Design
Configuration
Fabrication and
Assembly
Proof of
Concept
Ground Testing
Flight Test
Performance
and Results
Design
Optimisation
Final Optimised
Design

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