MEng Report Merged - FINAL

MEng Team Project UAS Challenge - 2015

Written By:

Alfred Dzadey, Jonathan Ebhota, Zuber Khan, Tarek Kherbouche, Amit Ramji, Mozammel, Mohammed Mohinuddin, Micky Ngouani, Malwenna Malwenna , Hassan Turabi, Osman Sibanda, Mohammed Rayad

Ullah

Project Supervisor:

Ray Wilkinson, Joanna Rawska, Kate Williams, Steve lines

UAS CHALLENGE 2015

i ACKNOWLEDGEMENTS MEng Team Project Report (7ENT1024) School of Engineering and Technology

ACKNOWLEDGEMENTS We (UAS Challenge MEng Group) would like to thank the supervisors who gave us their support and unconditional attention throughout the course of the project with weekly group meetings and off the clock advice. Their expertise in the subject helped in the successful delivery of this project. Other notable mentions are to the technicians; Chris Childs and Andrew Curl whose expertise, skill and experience were invaluable in the manufacture of the components of the UAS. We would also like to thank Gordon Banks from Ensinger for supplying the materials use for the project at a discounted prize and very swiftly too; Howard Ash for his assistance in the procurement of materials and components; Yigeng Xu for giving the MEng group permission to use E131B for assembly and testing purposes; Clive Borhem for giving technical assistance to the Propulsion Engineer and also everyone who supported the group directly and indirectly.

UAS CHALLENGE 2015

ii TABLE OF CONTENTS MEng Team Project Report (7ENT1024) School of Engineering and Technology

TABLE OF CONTENTS ACKNOWLEDGEMENTS ............................................................................................... i TABLE OF CONTENTS ................................................................................................ ii LIST OF FIGURES ...................................................................................................... vii GLOSSARY ................................................................................................................ xii 1 Introduction ........................................................................................................... 1

1.1 Competition Overview ..................................................................................... 1 1.2 Project Aims .................................................................................................... 1 1.3 The Project Objectives .................................................................................... 2

2 Design Rationale .................................................................................................. 2 2.1 Design Convergence ....................................................................................... 2

2.1.1 Stage 1 Convergence ............................................................................... 2 2.1.2 Stage 2 Convergence ............................................................................... 3

2.2 Further analysis ............................................................................................... 3 3 Project Management ............................................................................................ 4

3.1 Role of the Project Manager ............................................................................ 4 3.2 The Team Structure ......................................................................................... 4 3.3 Project Planning .............................................................................................. 5

3.3.1 Milestones ................................................................................................. 7 3.4 Leadership ...................................................................................................... 7 3.5 Team Communication ..................................................................................... 9 3.6 Project Budgeting ............................................................................................ 9

3.6.1 Summary of Project Budget .....................................................................10 3.6.2 Source of Funding ....................................................................................10

3.7 Risk Management ..........................................................................................11 3.8 Conflict management ......................................................................................11 3.9 Performance Review ......................................................................................12 3.10 Evaluation ...................................................................................................13

4 Quad-Rotor Design ..............................................................................................14 4.1 Design Rationale - Quad-Rotor ......................................................................15 4.2 Payload Box Design and Mechanism .............................................................16

5 UAV Mass Breakdown .........................................................................................17 6 UAV Cost Breakdown ..........................................................................................17 7 Structural Analysis ...............................................................................................18

7.1 Load Case Definition and Free Body Diagrams ..............................................18 8 UAV Stress Analysis ............................................................................................20

8.1 Stress Reduction Techniques .........................................................................20 8.2 Fatigue Awareness .........................................................................................20 8.3 Fatigue due to induced vibration .....................................................................21 8.4 Pressure Loading on Plates ............................................................................21 8.5 Load Transfer .................................................................................................22 8.6 Fixed and Movable Arm Stress Maximum ......................................................22 8.7 Simplified Plate Deflection ..............................................................................24

8.7.1 Simply Supported Plate Representation ...................................................24 8.7.2 Analytical Method .....................................................................................25 8.7.3 FEA – Simplified Rectangular Approximation ...........................................26

8.8 Plate Deflection - Assembly Contact Model as Built........................................27 8.9 Undercarriage Buckling Calculation ................................................................27 8.10 Undercarriage Bending ...............................................................................27 8.11 Undercarriage Bending - Assembly Contact Model .....................................28 8.12 Undercarriage Torsion .................................................................................29 8.13 Undercarriage Combined Loading - Torsion and Bending ...........................29 8.14 Undercarriage Combined Loading - Assembly Contact Model .....................30

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iii TABLE OF CONTENTS MEng Team Project Report (7ENT1024) School of Engineering and Technology

8.15 FEM Verification – Summary of Undercarriage Results ...............................30 8.16 Modal Analysis of Fixed-arm – Simplified Case ...........................................31

8.16.1 Analytical Modal Analysis – Simplified ..................................................31 8.16.2 Finite Element Modal Analysis – Simplified ...........................................33

8.17 Modal Analysis of Fixed-arm – Actual Parts (As Built) .................................34 8.18 Summary of Modal Frequency Results ........................................................35 8.19 Summarised Margin of Safety Table ...........................................................36

9 Performance, Propulsion & Systems Engineer ....................................................37 9.1 Propeller Diameter Selection ..........................................................................38 9.2 RC Motor Selection Maximum RPM ...............................................................40 9.3 Propeller Pitch Selection ................................................................................41 9.4 Power Supply Voltage Selection .....................................................................43 9.5 Power Supply Capacity Selection ...................................................................44 9.6 RC Motor Selection Power .............................................................................44 9.7 Electronic Speed Controller Selection.............................................................45

10 Unmanned Aircraft System - Subsystems ............................................................46 10.1 Introduction .................................................................................................46 10.2 Navigation Systems.....................................................................................46

10.2.1 Potential Issues with the Navigation systems ........................................47 10.2.2 Solutions ...............................................................................................47

10.3 Mission Control System ...............................................................................47 10.4 Flight Control System ..................................................................................48 10.5 Communication System ..............................................................................50

10.5.1 Serial Connection .................................................................................51 10.5.2 Telemetry Kit Connection ......................................................................51 10.5.3 Radio Connection .................................................................................52

10.6 Systems Integration.....................................................................................52 10.6.1 Communications Systems Test .............................................................52 10.6.2 Interference test ....................................................................................52 10.6.3 Range Test and Altitude Test ................................................................53 10.6.4 Post Manufacture and Assembly Design Checks ..................................53 10.6.5 Post Assembly Control System Calibration ...........................................54

11 Stability and Control I...........................................................................................55 11.1 PID Tuning ..................................................................................................56

11.1.1 Loiter mode ...........................................................................................56 11.1.2 Altitude Hold Mode (AltHold) .................................................................57

11.2 Verifying the performance of PID values .....................................................58 12 Safety Case .........................................................................................................59

12.1 Overview .....................................................................................................59 12.2 Flight Controller Safety Mechanism .............................................................59

12.2.1 Safety Measurements for Flight Testing ................................................59 12.3 Hazardous Components ..............................................................................60 12.4 Battery Fail Safe ..........................................................................................60 12.5 Radio Fail Safe ............................................................................................61

13 Environmental Impact ..........................................................................................62 13.1 Hazardous Material .....................................................................................62 13.2 Air Quality ...................................................................................................62

13.2.1 Emissions .............................................................................................62 13.2.2 Noise ....................................................................................................62

13.3 Infrastructure ...............................................................................................63 13.4 Disposal of Material .....................................................................................63

14 Stability and Control II..........................................................................................65 14.1 Ideal CG location .........................................................................................65

15 Flight modes and tuning .....................................................................................66 15.1 Simulink model ............................................................................................66

UAS CHALLENGE 2015

iv TABLE OF CONTENTS MEng Team Project Report (7ENT1024) School of Engineering and Technology

15.2 Test rig PID Testing.....................................................................................68 15.2.1 Pitch and Roll tuning .............................................................................69 15.2.2 Yaw tuning ............................................................................................71 15.2.3 Waypoint navigation tuning ...................................................................71

15.3 Tuning during flight ......................................................................................72 15.4 Future Work ................................................................................................72

16 Flight Termination Case .......................................................................................73 16.1 GPS Loss ....................................................................................................73 16.2 Communication loss from Ground Station ...................................................73 16.3 Geofence Breach ........................................................................................74 16.4 Maximum Pressure Altitude Breach ............................................................74

17 Systems Layout ...................................................................................................75 17.1 System block diagram .................................................................................75

17.1.1 Hardware Systems................................................................................75 17.1.2 Software Systems .................................................................................76

17.2 Communication ...........................................................................................78 18 Image Processing ................................................................................................79

18.1 Image Recognition ......................................................................................79 18.1.1 The Requirements ................................................................................79 18.1.2 Testing ..................................................................................................79 18.1.3 Results..................................................................................................80 18.1.4 Analysis ................................................................................................80 18.1.5 Shape recognation ................................................................................80

18.2 Video ...........................................................................................................81 18.3 On Screen Display Board (OSD) .................................................................82 18.4 Video transmitter .........................................................................................82 18.5 Video Receiver ............................................................................................82

19 Verification and Validation ...................................................................................83 19.1 Verification Matrix ........................................................................................83 19.2 Validation test..............................................................................................83

20 Future work .........................................................................................................84 20.1 Partial control of Quad-rotor positioning ......................................................84 20.2 Full Autonomy .............................................................................................84

21 Preliminary Payload Box Concept & Servo Integration ........................................85 21.1 Initial designs ..............................................................................................85

21.1.1 The Hinge-clamp Method ......................................................................85 21.1.2 The electro-magnet method ..................................................................85 21.1.3 The Hinge-pin method ..........................................................................86 21.1.4 Others ...................................................................................................86 21.1.5 Payload box mechanism integration .....................................................87

21.2 Servo ..........................................................................................................89 21.2.1 Specifications........................................................................................89 21.2.2 Rational ................................................................................................89

21.3 BEC ............................................................................................................90 21.3.1 Specification .........................................................................................90 21.3.2 Rational ................................................................................................90

21.4 Schematics of connections from battery to servo through pixhawk ..............91 21.5 Controlling the servo as a servo ..................................................................92 21.6 Testing with the Mission Planner .................................................................93

22 Other Involvements .............................................................................................94 22.1 Telemetry Kit ...............................................................................................94 22.2 Design Convergence ...................................................................................94 22.3 Challenges ..................................................................................................94

23 Manufacturing ......................................................................................................95 23.1 Machining Selection ....................................................................................95

UAS CHALLENGE 2015

v TABLE OF CONTENTS MEng Team Project Report (7ENT1024) School of Engineering and Technology

23.1.1 Machines ..............................................................................................95 23.1.2 Tools .....................................................................................................95

23.2 Manufacturing process of Quad-rotors components ....................................96 23.2.1 Fixed Bracket ........................................................................................96 23.2.2 Motor arm end bracket ..........................................................................96 23.2.3 Movable arm vertical fixed bracket /support bracket ..............................96 23.2.4 Landing gear top/bottom support bracket ..............................................97 23.2.5 Top/Bottom half T-joints ........................................................................97 23.2.6 Landing Gear Lug Bracket/ Pivot ..........................................................97 23.2.7 Arm pivot ..............................................................................................97 23.2.8 Main Body Plate ....................................................................................98 23.2.9 PVCs tubes ...........................................................................................98 23.2.10 Motor mount plate ..............................................................................98 23.2.11 Overview of Machining ......................................................................99

23.3 Challenges ................................................................................................ 100 23.4 Manufacturing Plan ................................................................................... 100 23.5 Machining Cost ......................................................................................... 100 23.6 Other involvements in the project .............................................................. 100

24 Test Rig ............................................................................................................. 101 24.1 Initial Conceptual Design of Gimbal Test Rig ............................................ 101 24.2 Octagonal Gimbal Test Rig ....................................................................... 102

24.2.1 Octagonal Model Mount Frame ........................................................... 103 24.2.2 Octagonal Mid Frame ......................................................................... 104 24.2.3 Octagonal Outer Frame ...................................................................... 105

24.3 Weight Estimation for Octagonal Test Rig ................................................. 105 24.4 Cost Breakdown for Octagonal Test Rig .................................................... 105 24.5 Manufacturing Stage of the Octagonal Test Rig ........................................ 106

25 Structural Testing ............................................................................................. 107 25.1 Material Testing ......................................................................................... 107 25.2 Component Testing ................................................................................... 107 25.3 Payload Drop Testing ................................................................................ 108 25.4 Initial Ball socket test rig ............................................................................ 108 25.5 Manufacturing assistance .......................................................................... 109

26 Business Case................................................................................................... 109 26.1 Executive Summary .................................................................................. 109 26.2 Business overview .................................................................................... 110 26.3 Mission statement ..................................................................................... 110 26.4 UAS key design features ........................................................................... 111 26.5 Market Assessment ................................................................................... 111

26.5.1 Potential market – Emergency Service ............................................... 111 26.5.2 Market size and growth ....................................................................... 112 26.5.3 Regulation restriction .......................................................................... 113 26.5.4 Challenges for market entry ................................................................ 113 26.5.5 Competition ........................................................................................ 114

26.6 Financial Forecasts ................................................................................... 115 26.7 Key assumptions ....................................................................................... 115 26.8 Costs ......................................................................................................... 115

26.8.1 Financial statements ........................................................................... 117 26.8.2 Profitability .......................................................................................... 118

Conclusion ................................................................................................................ 119 REFERENCES .......................................................................................................... 121 Appendix. A ............................................................................................................... 125 Appendix. B UAV Design ..................................................................................... 141 Appendix. C UAV Detailed Mass Breakdown ....................................................... 171 Appendix. D UAV Detailed Cost Breakdown ........................................................ 177

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Appendix. E Material Properties .......................................................................... 179 Appendix. F Load Cases and Load Transfer ........................................................... 181 Appendix. G Stress Analysis ................................................................................ 182 Appendix. H Performance & Propulsion ............................................................... 219 Appendix. I UAS System Set Up ............................................................................ 261 Appendix. J Systems .............................................................................................. 269 Appendix. K Altitude control ................................................................................. 281 Appendix. L Verification and validation ................................................................... 289 Appendix. M Telemetry kit Specification ............................................................... 302 M.1. Servo calculation .......................................................................................... 307 Appendix. N Manufacturing .................................................................................. 308

N.1. Machining by milling machine .................................................................... 309 N.2. Machining by XYZ 1330 Lathe .................................................................. 309 N.3. Laser Cutting by Tortec Laser cutter ......................................................... 310 N.4. Cutting blocks by vertical bandsaws machine............................................ 310

Appendix. O Test Rig ........................................................................................... 311 O.1. Initial Gimbal Test Rig Conceptual Design ................................................ 312 O.2. Updated Octagonal Gimbal Test Rig Assembly ......................................... 314 O.3. Octagonal Model Mount Frame Technical Drawing ................................... 316 O.4. Octagonal Mid Frame Technical Drawing .................................................. 318 O.5. Octagonal Outer Frame Technical Drawing ............................................... 320 O.6. Octagonal Gimbal Test Rig Stand Technical Drawing ............................... 322 O.7. Gimbal Test Rig Weight Estimation ........................................................... 324 O.8. Gimbal Test Rig Manufacturing Cost ......................................................... 326 O.9. Qualification test plan ................................................................................ 326 Electrical Performance Tests (Initial, In-Process, Final) ......................................... 326 Storage Temperature Cycling ................................................................................ 326 Thermal Shock ...................................................................................................... 326 Random/Sine Vibration .......................................................................................... 327 Operational Temperature Cycling .......................................................................... 327 O.10. Initial Involvement in the MEng Team Project ............................................ 327 O.11. Tri Angular Bracket Technical Drawing...................................................... 328 O.12. T-Bracket Technical Drawing .................................................................... 330

Appendix. O Design features for business case ................................................... 332

UAS CHALLENGE 2015

vii LIST OF FIGURES MEng Team Project Report (7ENT1024) School of Engineering and Technology

LIST OF FIGURES Figure 1 - Initial Concepts for Stage 1 convergence ..................................................... 2 Figure 2 - Concepts considered in the stage-2 convergence ........................................ 3 Figure 3 - Project Organization Chart ........................................................................... 5 Figure 4 - Progress (to date) of the project ................................................................... 7 Figure 5 - Leadership area of priority – Semester A (Left), Semester B (Right) ............ 8 Figure 6 – Performance Charts for Jonathan (a) and Zuber (b) ...................................12 Figure 7 - Quad-rotor design .......................................................................................14 Figure 8 - Stowage Instructions ...................................................................................15 Figure 9 - Quad-rotor in Stowed Configuration ............................................................15 Figure 10 – Removable Lightweight Payload Box .......................................................16 Figure 11 - Removable Lightweight Payload Box ........................................................16 Figure 12 - Payload Box with simple construction and failsafe mechanism .................16 Figure 13 – Payload Box with Payload Clearance .......................................................16 Figure 14 – Free Body Diagram - Flight and Landing Cases .......................................18 Figure 15 - Free Body Diagram - Landing Cases ........................................................18 Figure 16 - Free Body Diagram - Flight and Gust Load Cases ....................................19 Figure 17 – Fixed Arm Cross Section – See also Appendix G.7 ..................................22 Figure 18 – Mesh for Fixed-arm Assembly – Values as per Appendix G.6 ..................23 Figure 19 - Deflection of Fixed-arm Assembly (Flight Loads) with 7.6mm Deflection ...23 Figure 20 - Stress of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa (Peak) .......................................................................................23 Figure 21 – Stress (Close-up) of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa (Peak) ........................................................................24 Figure 22 - Simplified Plate Representations ...............................................................25 Figure 23 - Simple Plate Deflection Carried out on CATIA showing 4.54mm deflection ....................................................................................................................................26 Figure 24 - Flight and Gust condition of Main Body with 0.13mm Deflection................27 Figure 25 - Lateral Impact Case on Single Leg - 60.6MPa Stress ...............................28 Figure 26 – Stress Element A with Principal Stress for - Analytical – Undercarriage Combined Loading – Bending, Buckling and Torsion (G.13) .......................................29 Figure 27 - Arm and Mass for Rayleigh Method ..........................................................31 Figure 28 – Mass Representation of Motors, Blocks, Plates, Fasteners and ESC .......33 Figure 29 – Simplified FE analysis with 1st Nat freq as 19.64Hz – 69.3mm Deflection (Left) and 164MPa Stress (Right) ................................................................................33 Figure 30 – Simplified FE with 2nd Nat freq as 20.06 Hz (Left) and 3rd Nat freq as 134.6 Hz (Right) ....................................................................................................................33 Figure 31 – Simplified FE with 4th Nat freq as 224.1 Hz (Left) and 5th Nat freq as 411.9 Hz (Right) ....................................................................................................................33 Figure 32 – As Built FE Analysis - Mass Representation of Motors, Fasteners, Cables and ESC ......................................................................................................................34 Figure 33 – As Built FE analysis with 1st Nat freq as 451 Hz – 69.0mm Deflection (Left) and Stress (Right) .......................................................................................................34 Figure 34 - As Built FE analysis with 2nd Nat freq as 736 Hz (Left) and 3rd Nat freq as 1707 Hz (Right) ...........................................................................................................34 Figure 35 - As Built FE analysis with 4th Nat freq as 2 KHz (Left) and 5th Nat freq as 4.1 KHz (Right) .................................................................................................................34 Figure 36 - Prototype Quad Rotor ...............................................................................44 Figure 37: Waypoint Command File ............................................................................47 Figure 38: Telemetry Information transmitted to ground control station .......................50 Figure 39: Transmission Link Statistics (Serial Connection) ........................................51 Figure 40: Transmission Link Statistics (Telemetry Kit) ...............................................51 Figure 41 – PID System (Oscar, 2013) ........................................................................55

UAS CHALLENGE 2015

viii LIST OF FIGURES MEng Team Project Report (7ENT1024) School of Engineering and Technology

Figure 42 Loiter PID values .........................................................................................56 Figure 43 AltHold mode PID values .............................................................................57 Figure 44 Dataflash log in Stabalized mode opened in Mission planner ......................58 Figure 45 Battery fail safe settings chosen in Mission Planner ....................................60 Figure 46 Battery monitor settings chosen in Mission Planner .....................................61 Figure 47 Side view of the Quad-rotor .........................................................................65 Figure 48 Simulink model used ...................................................................................66 Figure 49 Quad-rotor oscillating with only the P gain (left), with P and D gain (right) ...67 Figure 50 PID values on Simulink ................................................................................68 Figure 51 Values that require change (3DR Robotics, 2015) .......................................69 Figure 52 Quad-rotor on the test rig ............................................................................70 Figure 53 Results of what Pixhawk should output (Copter.Ardupilot, 2015) .................70 Figure 54 Geofence configuration on Mission Planner ................................................74 Figure 55 Overall System Hardware Block Diagram ....................................................75 Figure 56 Overall Software Block Diagram ..................................................................77 Figure 57 Matlab alphanumeric code processing letter at 22.98cm .............................80 Figure 58 Shape recognition .......................................................................................81 Figure 59 Circuit Diagram of Ardunio ..........................................................................84 Figure 60: Hinge clamp ...............................................................................................85 Figure 61: electro-magnet ...........................................................................................85 Figure 62: Hinge-pin ....................................................................................................86 Figure 63: Other concept .............................................................................................86 Figure 64: CAD ...........................................................................................................87 Figure 65: Overall payload box ....................................................................................87 Figure 66: Horn and door connection ..........................................................................88 Figure 67: Start up release ..........................................................................................88 Figure 68: Fully Unlocked door ....................................................................................88 Figure 69: Complete release .......................................................................................89 Figure 70 - MG90S servo ............................................................................................89 Figure 71 - SBEC26 Turnigy .......................................................................................90 Figure 72: Schematics of connections .........................................................................91 Figure 73: Configuration of the servo on Pixhawk .......................................................92 Figure 74: Mission with GPS dropping points ..............................................................93 Figure 75: Verification of the performance of the Servo ...............................................93 Figure 76: Machined fixed bracket ...............................................................................96 Figure 77: Machined end bracket ................................................................................96 Figure 78: Machined Fixed bracket .............................................................................96 Figure 79: Machined bottom support bracket...............................................................97 Figure 80:T-joint on foam ............................................................................................97 Figure 81: Lug bracket ................................................................................................97 Figure 82: Landing gear pivot ......................................................................................97 Figure 83: Arm pivot for movable arm..........................................................................97 Figure 84: Cutting nylon plate in Laser machine ..........................................................98 Figure 85: Melted edges ..............................................................................................98 Figure 86: Plate after cutting .......................................................................................98 Figure 87: Assembled motor mount .............................................................................98 Figure 88: Motor mount plate ......................................................................................98 Figure 89.1-3: CNC practice sessions .........................................................................99 Figure 90.1-3: Failed attempts ................................................................................... 100 Figure 91 - Gyroscope Test Rigs ............................................................................... 102 Figure 92 - CAD Drawing of the Quad-rotor .............................................................. 104 Figure 93 - Test Rig Components .............................................................................. 106 Figure 94 - Test Rig Assembly .................................................................................. 106 Figure 95 - Nylon Material and Main Body Plate ........................................................ 107 Figure 96 - Compression Test conducted on Hounsfield Tensometer ....................... 107

UAS CHALLENGE 2015

ix LIST OF FIGURES MEng Team Project Report (7ENT1024) School of Engineering and Technology

Figure 97 - Initial ball socket test rig .......................................................................... 108 Figure 98 - Autoquads Inspection Ltd Logo ............................................................... 109 Figure 99 - Permissions required for different UAS sizes ........................................... 113 Figure 100 - Break Even Graph ................................................................................. 118 Figure 101 - Overall View of Quad-Rotor ................................................................... 142 Figure 102 - Motor Mount Design (Left) & Undercarriage T-Joint (Right) ................... 142 Figure 103 - Undercarriage Pivot Design (Left) & Main Body Sandwich Design (Right) .................................................................................................................................. 142 Figure 104 - Movable Arm Pivot Design .................................................................... 143 Figure 105 - Project Main Body Area ........................................................................ 181 Figure 106 – SOLID187 Element (Ansys, November 2013c) ..................................... 185 Figure 107 – PLANE182 Element (Ansys, November 2013c) .................................... 185 Figure 108 - Arm Cross-section for Stress Calculation .............................................. 190 Figure 109 - Tension & Compression Stress in Arm .................................................. 190 Figure 110 – Mesh for Fixed-arm Assembly – Values as per Appendix G.6 .............. 191 Figure 111 - Deflection of Fixed-arm Assembly (Flight Loads) with 7.6mm Deflection .................................................................................................................................. 191 Figure 112 - Stress of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa (Peak) ..................................................................................... 192 Figure 113 – Stress (Close-up) of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa (Peak) ...................................................................... 192 Figure 114 – Mesh for Arm Assembly (With additional Tab) – Mesh Values as per G.6 .................................................................................................................................. 193 Figure 115 - Modified FB-002 for reduction in point contact stress concentration ...... 194 Figure 116 - Stress Concentration at Arm (without addition) Contact (a) & Close-up (b) .................................................................................................................................. 194 Figure 117 - Deflection of Modified Movable Motor Arm of 7.88mm for flight loads with SF ............................................................................................................................. 195 Figure 118 - Stress of Modified Movable Motor Arm of 20.8MPa for flight loads with SF .................................................................................................................................. 195 Figure 119 - Modified Movable Motor Arm with Stress of 20.8MPa for flight loads with SF (a) & Close-up (b) ................................................................................................ 195 Figure 120 - Load on the Lug (Niu, 1988) .................................................................. 196 Figure 121 - Components of the Load (Niu, 1988) ..................................................... 196 Figure 122 - Areas on the Lug ................................................................................... 196 Figure 123 - Lug Bracket Without Flange (Left) & with additional Flange (Right) ....... 197 Figure 124 – Lateral Unit Load Deflection (Left) & Stress (Right) of Lug Bracket Without Flange ....................................................................................................................... 198 Figure 125 – Lateral Unit Load Deflection (Left) & Stress (Right) of Lug Bracket With Flange ....................................................................................................................... 198 Figure 126 - Mesh for MP-001 (Appendix B.7) with values as per Appendix G.6 ....... 199 Figure 127 – Motor Plate Deflection (0.038 mm) and Stress (41.7 MPa) for flight case with SF at start-up ..................................................................................................... 199 Figure 128 - Error Elements in Model - Due to Separation at FB-001 and EB-001 .... 199 Figure 129 - Simplified Plate Representations ........................................................... 200 Figure 130 - Simple Plate Deflection Carried out on CATIA structural analysis ......... 201 Figure 131 - Mesh of Main Body Plate - Values as per Appendix G.6 ....................... 202 Figure 132 – Single Main Body Plate Analysis – with 17.8MPa Stress at contact holes for flight case with pressure load ............................................................................... 202 Figure 133 – Mass Representation of components and payloads as per Appendix. C .................................................................................................................................. 203 Figure 134 - Mesh of Main body assembly with Values as per Appendix G.6 ............ 203 Figure 135 – Contact model Flight Case for Main body assembly Deflection (left) and Equivalent Stress (right) ............................................................................................ 203

UAS CHALLENGE 2015

x LIST OF FIGURES MEng Team Project Report (7ENT1024) School of Engineering and Technology

Figure 136 - Contact model Flight Case for Main body assembly - Equivalent Stress with predicted locations ............................................................................................. 204 Figure 137 - Resolving Component to Determine Vertical Load ................................ 205 Figure 138 - Undercarriage Leg Under Pure Bending ................................................ 205 Figure 139 - Undercarriage Leg Under Pure Torsion ................................................. 206 Figure 140 - Stress Element A (Warren C. Young) .................................................... 207 Figure 141 - Plan View of Stress Element A .............................................................. 207 Figure 142 - Stress Element A with Principle Stresses .............................................. 208 Figure 143 - Undercarriage Mesh for Contact Model with values as per G.6 ............. 209 Figure 144 – Lateral Landing on Single Undercarriage Leg with 53.6mm Deflection . 209 Figure 145 - Lateral Landing on Single Undercarriage Leg with 60MPa Bending Stress .................................................................................................................................. 210 Figure 146 - Lateral Landing on Single Undercarriage Leg with 60MPa Bending Stress (Close-up) ................................................................................................................. 210 Figure 147 - Tip Landing on Single Undercarriage Leg with 60MPa Bending Stress . 211 Figure 148 -Tip Landing on Single Undercarriage Leg with 66mm Combined bending and torsion deflection ................................................................................................ 211 Figure 149 - Tip Landing on Single Undercarriage Leg with 71MPa Combined bending and torsion stress ...................................................................................................... 212 Figure 150 – Entire Quad-Rotor Flight Deflection of 7.9mm at Motor Arm Tips ......... 213 Figure 151 - Entire Quad-Rotor Flight Deflection of 7.9mm at Motor Arm Tips (Close-up) ............................................................................................................................. 213 Figure 152 - Entire Quad-Rotor Flight Stress of 28.8 MPa at Motor mount plates ..... 213 Figure 153 - Entire Quad-Rotor Flight Stress with Plate Stress peak at 14.42Mpa .... 214 Figure 154 – Downward Load - 1kg Payload and 10N Additional Load onto PB-005 Plate .......................................................................................................................... 215 Figure 155 - Side Load - 1kg Payload and 10N Additional Load onto Hinge Plate at 45deg to horizontal .................................................................................................... 215 Figure 156 - Side Load - 1kg Payload and 10N Additional Load onto short edge 45deg to horizontal............................................................................................................... 216 Figure 157 - Side Load as per Figure 156 - Showing Pre-mature Release due to global deflection ................................................................................................................... 216 Figure 158 – Downward Load as per Figure 154 - new design showing 0.73mm Deflection .................................................................................................................. 217 Figure 159 - Side Load as per Figure 155 –new rigid design and Deflection of 1.56mm .................................................................................................................................. 217 Figure 160 – Side Load as per Figure 156 and Figure 157 – with new design and deflection of 0.41mm* ................................................................................................ 217 Figure 161: Proof of Connection ................................................................................ 261 Figure 162: Mission Planner top menu ...................................................................... 262 Figure 163: Initial Setup for all components ............................................................... 262 Figure 164: Mission Planner Waypoint Entry Point .................................................... 262 Figure 165: Secondary Commands ........................................................................... 263 Figure 166: Area for writing flight plans into Pixhawk's Memory ................................ 263 Figure 167: Stability Tuning for Quad-rotor Control ................................................... 264 Figure 168: Mission Plannner environment for changing parameters ........................ 264 Figure 169: Fail Safe parameters .............................................................................. 265 Figure 170: Typical Set Fail Safe Values ................................................................... 266 Figure 171: Monitoring System Values ...................................................................... 267 Figure 172: Flight ready monitoring system ............................................................... 267 Figure 173: Quad-rotor Acceleration and Velocity parameters .................................. 268 Figure 174 Minim OSD V2.1 (unmannedtechshop, 2015).......................................... 276 Figure 175: 3DR uBlox GPS with Compass Kit (unmannedtechshop, 2015) ............. 278 Figure 176 CG calculations for the x and y-axis ........................................................ 281 Figure 177 CG calculations for z-axis ........................................................................ 281

UAS CHALLENGE 2015

xi LIST OF FIGURES MEng Team Project Report (7ENT1024) School of Engineering and Technology

Figure 178 Overview of the Simulink model .............................................................. 286 Figure 179 Section to change PID values .................................................................. 287 Figure 180 Quad-rotor control mixing ........................................................................ 287 Figure 181 Quad-rotor dynamics ............................................................................... 288 Figure 182 GUI of the Quad-rotor general parameters .............................................. 288 Figure 183: Other CAD views .................................................................................... 306 Figure 184: schematics for the force calculations ...................................................... 307 Figure 185: Machined fixed bracket is CNC Router Pro 2600 .................................... 308 Figure 186: Dry assemble of landing gear lug, pivot and the vertical landing strut ..... 308 Figure 187: Slot bracket Figure 188: Turn button for servo motor ....................... 308 Figure 189: Support corners machined in CNC Figure 190: Triangle payload support glued with hinges 308 Figure 191: Drilling centre hole in fixed bracket Figure 192: Milling arm Pivot ........ 309 Figure 193: Chamfering of movable arm support Figure 194: Smoothing surface by fly cutter 309 Figure 195.1-2: Drilling using slot drills ...................................................................... 309 Figure 196: High speed steel tool .............................................................................. 309 Figure 197.1-2: Machining arm pivot on lathe ............................................................ 310 Figure 198.1-2 Laser Cutting of Nylon 6 sheet for main body plate ........................... 310 Figure 199: Cutting Nylon 6.6 cast block in vertical band saw machine ..................... 310 Figure 200 -OXV in storage configuration .................................................................. 332 Figure 201 - Electro-optic camera on the OXV .......................................................... 332 Figure 202- Main body of the OXV ............................................................................ 332

UAS CHALLENGE 2015

xii GLOSSARY MEng Team Project Report (7ENT1024) School of Engineering and Technology

GLOSSARY

3D Three Dimensional

AFS Advanced Failsafe

Atl Altitude

BEC Battery Eliminator Circuit

BOM Bill of Material

CAA Civil Aviation Authority

CAD Computer Aided Design

CCW Counter-Clockwise

CG Cenre of Gravity

CNC Computerised Numerical Control

COTS Commercial of the Shelf

CPR Cardiopulmonary Resuscitation

CW Clockwise

D Derivative

EMI Electro-Magnetic Inteference

ESC Electronic Speed Controller

EU European Union

FBD Free Body Diagrams

FE Finite Element

FEA Finite Element Analysis

FEM Finite Element Method

FPV First Person View

ft Feet

GPS Global Positioning System

Hz Hertz

I Integral

IAS Indicated Airspeed

IMechE Institution of Mechanical Engineers

km kilometer

knots Nautical Miles

KV kilo-volts

Li-PO Lithium Polymer

m metres

MEng Masters of Engineering

mm Millimeter

MPa Mega Pascals

MTOM Maximum Take-Off Mass

N Newtons

OSD On Screen Display

P Proportional

PA Polyamide (Nylon)

PDR Preliminary Design Review

PVC Polyvinyl Chloride

PWM Pulse Width Modulator

UAS CHALLENGE 2015

xiii GLOSSARY MEng Team Project Report (7ENT1024) School of Engineering and Technology

Quad Quadcopter

RC Radio Controller

RC Receiving (Radio)

RPM Revolutions Per Minute

RTL Return to Launch

SCA Sudden Cardiac Arrest

SUA Small Unmanned Aircraft

TX Transmission (Radio)

UAS Unmanned Aircraft System

UAV Unmanned Aerial Vehicle

VAT Value Added Tax

VLOS Visual line of Sight

W Watts

WBS Work Breakdown Structure

WP Waypoint

UAS CHALLENGE 2015

1 Introduction MEng Team Project Report (7ENT1024) School of Engineering and Technology

1 Introduction

The Unmanned Aircraft system (UAS) challenge is being introduced by the Institution of

Mechanical Engineers (ImechE) for the first time. Teams entered by universities will only constitute

of members from the undergraduates cohort. The competition will provide students from different

universities to develop and demonstrate leadership, teamwork and technical competencies. It’s

being held during the academic year 2014 till 2015. During this period the universities participating

in the competition will follow a structure of designing, developing and demonstrating. It will also

include design reviews, presentations and flight demonstration that will contribute to point scoring.

1.1 Competition Overview

The competition this year is built around a scenario of a natural disaster occurring with a large

areas distressed by an earthquake or tsunami. The scenario could involve many thousands of

people being cut off from supplies and in need for humanitarian aid. The job at hand is to supply

these areas with humanitarian aid food and first aid supplies. Time is critical and the UAS launch

site is some distance away from the affected areas. The UAS operates autonomously via pre-

determined waypoints to areas affected with the capability of image recognition to identify the

supply drop zone. The UAS can be programmed to carry out circuit trips and return to base and

repeating the mission.

For the competition, the UAS can have a maximum take-off mass of 7kg with Commercial Off The

Shelf (COTS) products not exceeding £1000. The UAS will need to perform a series of tasks such

as take-off, climb to an altitude between 100-400ft, cruise, follow a predefined route, drop two

payloads (Bag(s) of flour) weighing 1kg each at any reachable location and land back completely

autonomously.

1.2 Project Aims

The aims intended at the initial stage of this project were to;

Develop and demonstrate leadership, teamwork, technical competence, as well as

commercial skills.

Develop a complex system that will require design, development and demonstration with

regards to a demanding mission requirement.

Apply the knowledge learnt from previous academic years during the course of the

undergraduate engineering degree.

Represent the university in the UAS challenge, successfully compete and win.

Section by Alfred Dzadey

UAS CHALLENGE 2015

2 Design Rationale MEng Team Project Report (7ENT1024) School of Engineering and Technology

1.3 The Project Objectives

The main objectives set, in order to be successful are;

To design and develop a UAS with a MTOM less than 7kg and achieve autonomy in all

phases of flight and tasks.

Expenditure on COTS must not exceed £1000

To be able to switch between manual and autonomous flight.

To develop a UAS that can accommodate and deliver a 1 kg bag of flour.

To create an image recognition system to identify the drop zone and read alphanumerical

characters.

To complete the task in the fastest time possible.

2 Design Rationale

This section begins to discuss the solutions to the mission requirement outlined in the previous

chapter. It assesses various design options and converges to an ultimate solution. Upon evaluation

of various design concepts, it was conclusive that the best approach to tackle the problem was by

going forwards with a Quad-rotor and ground control station.

2.1 Design Convergence

A 2-stage design convergence approach was used to conclude which concept best meets the

requirements. The design concepts selected were compared against set criteria as discussed in

2.1.1 and 2.1.2.

2.1.1 Stage 1 Convergence

Application of the design solutions were analysed using a weighting system under criteria such as

manoeuvrability, structural integrity, stability

during flight and payload accuracy to name a

few. A workable concept was congregated

after a two-stage design convergence; Stage-

1 convergence was to determine the type of

aircraft to be used where the concepts

assessed were:

Fixed Wing

Helicopter

Osprey

Multi-Rotor

Figure 1 - Initial Concepts for Stage 1 convergence(UAVClub, 2015)


UAS CHALLENGE 2015

3 Design Rationale MEng Team Project Report (7ENT1024) School of Engineering and Technology

The results from the Stage-1 convergence demonstrated that a multi-rotor would be the best option

to meet the product design specification.

2.1.2 Stage 2 Convergence

Stage-2 design convergence was to determine which multi-rotor system would best meet the UAS

requirements. Criteria used in stage-2 included redundancy (motor failure), manufacturing

complexity, power consumption, noise, payload capacity, structural integrity and costs to name a

few. The concepts considered during the stage-2 design convergence were:

Quad-Copter

Hex-rotor

Octacopter

3 arm – 6 rotors

Figure 2 - Concepts considered in the stage-2 convergence (UAVClub, 2015)

Upon comparison between the multi rotors under the above criteria, the Hex-rotor was found to be

the best concept that would meet the set criteria. Appendix C details the rationale and justification

for this selection.

2.2 Further analysis

Upon initiation of the design, a mass breakdown of all components for this proposed system was

established, it was found that the Hex-rotor would be overweight. The only solution forward was to

lose two arms to reduce it’s weight hence reducing from a Hex-rotor to a Quad-rotor design. This

allows the design to be approximately 6kg with all possible components and a single payload

accounted. The quad rotor design is 1 kg under the constrained maximum weight of 7 kg, providing

contingency for any miscalculations or any unplanned additional weight.

UAS CHALLENGE 2015

4 Project Management MEng Team Project Report (7ENT1024) School of Engineering and Technology

3 Project Management

To achieve the project objectives, effective organisation, planning, budgeting and management

styles were adopted. This section describes the organisational structure and the key management

tasks undertaken to deliver the project successfully. It describes the leadership, organisation

structure and role selection, project planning, budgeting, people and conflict management, finally a

review on both team leads is discussed.

3.1 Role of the Project Manager

The project management role comes with responsibilities involving the following;

Progress –ensuring the deliverables are being completed within the set timescale.

Budgeting – control the money being spent to ensure the deliverables are being completed

within the baseline cost.

Performance – ensuring the team is performing enough to achieve the goals set out.

Reporting – scheduling regular meetings with team members and supervisors to report

progress and resolve issues.

Planning/Change – handling and resolving any unexpected changes to project without

hugely affecting the outcome of the project or delivery.

Risk – to implement any contingencies into the time plan and budget to manage any

unforeseen risks affecting the project delivery.

Leadership and motivation – Motivating and maintaining morale during the duration of the

project.

Purchasing – dealing with orders being placed, tracking and informing the group of the

delivery progress of the order.

Welfare – taking into account the commitment of individuals while setting actions without

jeopardising the progress of the group.

Conflict – resolving any disrupt between team members and allowing a good working

environment.

Presenting – handling the compilation of all group reports and presentations in terms of

collating, proof reading individual reports and structuring.

3.2 The Team Structure

Project organisation structure needs to be one that facilitates the coordination and implementation

of project activities. The project organisation needs to create an environment in which there are

interactions among team members with minimal conflict, disruption or overlapping. The team

comprises of twelve students; nine studying aerospace, two studying aerospace with space

technology and one studying aerospace systems. With the project being systems related, the team

lacked expertise in that area meaning more work needed to be carried out. Figure 3 shows an

organisational structure to highlight each person’s responsibility and tasks carried out.


UAS CHALLENGE 2015


Figure 3 - Project Organization Chart

As with any large project it is advisable to split project team into sub teams to enable the project to

be manageable. This allows deliverables to be split into smaller tasks with clear objectives within

sub teams. It enables the team members in the sub teams to know exactly what actions are

required for an effective contribution. Another advantage of this set up is that there is a clear line of

authority and also team members will become familiar with each other since they work together in

the same area. Effective communication channels allow for the project manager and team leaders

to effortlessly interact and report back any difficulties or progress updates. Zuber and Jonathan

were appointed sub team leader due to both of them being extraverts and possession leadership

qualities as assessed using MBTI results. The structural team handles tasks relating to the design,

quality control, compliance, manufacture, assembly, test and certification of the UAS. The systems

team handles tasks relating to performance and propulsion, stability, control systems, flight and

navigation, imaging system, mission control, safety and payload deployment system.

3.3 Project Planning

The key to a successful project is in the planning, hence continual involvement and forward

planning must be carried out prior to project initiation. It involves the use of schedules such as

Gantt charts for planning and subsequently to report project progress. Initially, the project scope

was defined and the suitable method of successful delivery of this project was determined. The

following step was working out the durations and having contingency for all the various tasked

Alfred Dzadey

Project Manager

Zuber Khan (Chief Signatory - Quality)

Structural Team Leader

Structural / Stress / Cost / Weights / Assembly Engineer

Osman Sibanda

Marketing/Bussiness Specialist

Mozammel

Manufacturing Engineer

Amit Ramji (Chief Engineer)

Structural / Stress / Design / Hardware & Electrics

Integration and Assembly Engineer

Mohammed Mohinuddin

Structural and Testing Engineer

Jonathan Ebhota

Systems Team Leader

System Engineer

Micky Ngouani

Servo Selection Engineer

Kasun Malwenna

Safety / Stability and Control engineer

Tarek Kherbouche

Camera / Imaging Systems Engineer

Reyad Mohammed Ullah

Stability and Control Engineer

Hassan Turabi

Performance and Propulsion engineer

UAS CHALLENGE 2015


needed to complete the project. Major objectives were subsequently listed and implemented into a

Work Breakdown Structure (WBS) as shown in Table 1 below.

The WBS details the main steps that are required to complete this project. Stages involving design,

manufacture, purchasing and delivery of products may involve several delays that creates

difficulties and hence prevents the scheduled delivery. Strict time management and contingencies

such as overestimating time frames for completion of such tasks have been implemented into the

project plan to account for these delays.

Work Breakdown Structure

1 Scope 4.3 Structural material and sizing ready for purchase

1.1 Determine project scope 4.4 Design purchase readiness

1.2 Define resources 5 Order parts

1.3 Scope complete 5.1 Send out order list for components and delivery

2 Design Specification/System Requirements 6 Manufacturing & Assembly

2.1 Create Design specification for a UAV 6.1 Machine structural frame

2.2 Review system specifications 6.2 Integrate systems components

2.3 Create system requirements 6.3 Integrate structural frame, system and propulsion components

2.4 Obtain approvals to proceed (concept, timeline, budget)

7 Testing and Validation

2.5 Analysis complete 7.1 Develop unit test plans using design specifications

3 Preliminary Design 7.2 Develop integration test plans using design specifications

3.1 Review specifications 8 Integration Testing

3.2 Payload Delivery System 8.1 Test system integration

3.3 Propulsion System design 8.2 Integration testing complete

3.4 Systems design 9 Critical Design Review (CDR) and Flight Readiness Review (FRR)

3.5 Concept Structural design 9.1 Draft CDR report

3.6 Preliminary Safety Case consideration 9.2 Deliver CDR report

3.7 Preliminary Weights estimation 9.3 Draft FRR report

3.8 Obtain approval to proceed 9.4 Deliver FRR report

3.9 Preliminary Design complete 10 Competition

3.10 Deliver PDR to IMeche 10.1 Design Presentation

4 Final Design ready for purchase 10.2 Flight Readiness Review

4.1 System components finalised ready for purchase

10.3 Competition day

4.2 Propulsion components ready for purchase 10.5 UAS CHALLENGE FINISH

Table 1- Work Breakdown Outline Once the work breakdown structure was established, the project schedule was created and is used

as a baseline schedule for the whole duration of the project life. Using the project plan, a graph

representation of the current progress has been created and is shown in Figure 4. This is a

simplified overview of the progress made so far which is detailed in the project plan shown in

Appendix A.2. The progress made so far and completion of tasks can be seen in more detail in the

project plan.

UAS CHALLENGE 2015


Figure 4 - Progress (to date) of the project

3.3.1 Milestones

The major milestones set for this project are as follows:

30 October - Defining scope of project

16 November - Complete Design Analysis

05 December – Deliver PDR to IMechE

16 December – Design ready for purchase

1 April – Deliver CDR report

30 May - Integration testing complete

12 June – Deliver FRR report

1 July Design presentation

July – Competition Day and End of UAS Challenge

3.4 Leadership

Leadership involves creating an inspiring vision and managing the delivery of the vision.

Leadership brings together the skills needed to achieve this vision. Therefore, it is vital that the

style of leadership is rightly chosen for team performance and effective quality. The style of

leadership may vary during the duration of the project. ‘The three circle model’ is a concept that is

used to represent the dynamics of a group displaying the percentage of effort in terms of team,

task and individual. (Adair, 2012) It is critical for the leader to monitor these areas to ensure that

one area doesn’t needlessly become dominant. An example is where the group may take long to

make decisions due to the size of the group and differences of opinions. This is mitigated by

creating a cut of point whereby the group is no longer being effective in the decision making

process.

0 10 20 30 40 50 60 70 80 90 100

Scope

Design Specification/System Requirements

Preliminary Design

Final Design ready for purchase

Critical Design Review (CDR)

Order parts

Manufacturing & Assembly

Testing and Validation

Integration Testing

Flight Readiness Review (FRR)

Competition

Progress (%)

Project Progress to date

UAS CHALLENGE 2015


Consequently, the leader will conclude what has been discussed and make the final decision. In

Figure 5 the three circle diagrams depicts in which area leadership was stressed during the

academic year.

Figure 5 - Leadership area of priority – Semester A (Left), Semester B (Right) During the academic year the leadership style varied between a democratic and an authoritarian

style. Semester A involved initial stages of the project whereby there were a lot of group

discussions. It involved the development of the design concept hence a democratic style was

chosen to allow everyone’s input in decision making. This method allows members to feel free to

express their opinion. For people who were intrinsic it was encouraged for them to voice their

opinion in all decision making by actually asking what their thoughts were. This allowed team

members to grow in confidence and voice their opinion, which was good for the group dynamics. At

times, it made decision making problematic but it’s the responsibility of the leader to step in and

make the final decision based on the majority vote.

In semester B, the approach of leadership changed. It required a leader of a more authoritarian

style. This is due the fact that the project had shifted from a design phase to a development and

manufacturing stage. This stage is on a critical path hence an authoritarian style of leadership was

needed to help mitigate any delays. This involved a lot of communication on a daily basis to

establish what was set out to achieve and what was actually accomplished at the end of the day.

Furthermore, constant checking up on individuals was needed to ensure progress and also to deal

with unforeseen circumstances. For example, a time came when there was an issue with the

machining of the main plates for the airframe. The team required more material but there was none

left. As the leader, it was essential to step in and resolve the solution. This circumstance was

handled immediately by contacting the supplier and explaining the situation at hand and how

urgent the material was needed. As a result, the supplier ‘Ensinger’ was able to send out an order

as a free sample for next day delivery.

TEAM

TASK

INDIVIDUAL

TEAM

TASK

INDIVIDUAL

UAS CHALLENGE 2015


3.5 Team Communication

Throughout the project, weekly meetings with team members were undertaken to discuss any

updates, complications and actions required. Also during semester A, we had weekly meetings on

Tuesday noon with our supervisors to discuss the updates, complications and new actions set for

the week coming and where a register of attendance is taken. Ours meetings are made effective,

by using agendas and minutes. Minutes are used to record the discussions, conclusions and

actions set whereas the agenda was used to structure our meetings by having a schedule stating

exactly what topics are to be discussed and who is presenting the topic of discussion. An example

of the minutes, agenda can be seen in Appendix A.3 and A.4. Communication is essential for the

progression and success of a group. Without effective means of communication the group

production comes to a standstill. Communication methods used in the project are as follows. A

breakdown of the various group communications methods are presented in Table 2

Communication Aids

Types/Techniques Description Email Agendas are always sent out 24 hours before our official meetings

with our supervisors and also minutes are also sent out 24 hours after the meeting as a follow up of what was discussed and agreed in the meeting.

WhatsApp

It is used a form communication where all group members can discuss about findings or issues

Google drive An account was made for sharing files between members in the group. Each individual in the group has a folder with their name and hence can share their work to the group

Text messages and phone calls

For contacting individuals in the group privately for any needs regarding the project

Group meetings It’s used as a way to meet up face to face to discuss and updates or issues and to check progress of work and make decision.

Table 2 - Forms of communication used in project

3.6 Project Budgeting

For this project, there was a need for managing the funds to stay within the financial range of

£1390. A budget was used to project the costs and also to track the funds. A comparison of the

actual funds and the budget estimation has been made to see how much has been spent. Table 3

shows the operational budget. On the left are the projections for the budget as of November 2014.

On the right hand side we have the actual unit prices and quantities purchased. The final column

presents the difference between the two. The budget also includes a contingency factor of 1.2 to

anticipate any failures crashes or even unforeseen costs. A more detailed representation of the

product cost can be found in Appendix. D.

UAS CHALLENGE 2015


Budget Estimation as of

01/11/2014 Actual as of 23/04/2015

Part Unit Price Quantity Unit Price Quantity Difference

Flight controller £150.00 1 £159.98 1 -£9.98

Telemetry kit £40.00 1 £35.80 1 £4.20

GPS Module £50.00 1 £53.94 1 -£3.94

ESC £30.00 5 £27.16 5 -£2.84

Propellers £5.00 6 £3.95 6 £6.30

Brushless Motors £20.00 5 £19.16 5 £0.84

Camera £50.00 1 £56.41 1 -£6.41

OSD £30 1 £29.99 1 £0.01

Batteries £90.00 2 £60.40 3 -£1.20

RC Transmitter £30.00 1 £14.99 1 £15.01

Air frame including landing gear and payload box

£150.00 1 £146.30 1 £3.70

Extra cable and connectors £50.00 1 £20.95 1 £29.05

Test Rig* £150.00 1 £132.08 1 £17.92

Unplanned Quad Parts £0.00 0 £21.02 1 -£21.02

Delivery Costs* £100.00 1 £125.06 1 -£25.06

Total: £1,157.63

C. Factor (x1.2) 1389.16

Current Total: £1,100.94

Remaining: £231.53

*Not Part of COTS Percentage: 79.252438 Table 3 - UAS Challenge 2015 Budget

3.6.1 Summary of Project Budget

The main outcome of the budget that can be identified is that the project is £231.53 (21%) within

budget. This includes the majority of the UAS components, materials and also a test rig with

minimal additional items left to purchase. The flight controller is the team’s most expensive COTS

due to aspiring for a flight controller that was widely used. This will allow us access to open-source

information about autonomous control of the UAS. A complex alternative was to make use of an

Arduino board costing approximately £60 and to program the flight plan manually, hence potentially

saving £100. The team has had to spend some money for items that were not considered initially.

This has accumulated to a total of £210.03 which has been put that as unplanned Quad parts. We

have also gone over budget slightly on delivery cost which was unplanned. A detailed expenditure

of the project to date can be seen in Appendix. D

3.6.2 Source of Funding

There were two main sources funding this project. One was the funding from the university and the

other was from the team members. The university provided the team with £1000 to cover

everything from designing manufacturing, testing and the development of the product. Later, it was

found that this budget was insufficient and that it would only cover the purchase of materials and all

the components required for a Quad rotor design that would achieved good results for the

UAS CHALLENGE 2015


competition. The team decided to invest in the project extending the budget to £1500 with the aims

of winning an award and being reimbursed.

3.7 Risk Management

Due to a team member previously dealing with the safety case, the task of risk management was

delegated to that member. The rationale behind this decision was that the safety case and risk

management interlink so it was logical to delegate these tasks to an individual. The risk

management can be found in Appendix J.3.

3.8 Conflict management

Conflict is a common phenomenon in group projects. It’s inevitable and hence important as a

leader to understand the various conflict resolution techniques. Conflict isn’t always a bad thing

because it can present opportunities for improvement. Teams usually try to avoid or ignore

conflicts rather than addressing it. It can also have an effect on the team performance, as a leader

it is my duty to prepare for conflicts by creating an atmosphere that allows for dealing with conflicts

without relationship and emotional problems, for instance forming an atmosphere that supports

constructive criticism so that discrepancies can be expressed. Conflicts in the group were resolved

mostly using the following steps: Firstly, the people were separated from the problem by

diagnosing what was causing the issue and then various options were developed in order to

resolve the problem at hand. Secondly, the options were evaluated and the unimportant issues

were distinguished from the vital problems. Additionally, a common ground between each side was

found and a solution for both sides was mutually agreed upon. Finally, the agreement was

monitored to ensure that it was kept.

An example whereby this process was exercised was a situation involving to members in the team.

The issue sparked during a previous report submission where Micky had made a copied CAD

version of the payload box design. Amit had previously made a design for the payload housing due

to severe schedule delays in the project and was not happy about Micky’s contribution and

informed me that for the MEng report he doesn’t want Micky to present that design because he did

not design the payload box and was not involved during the decision making. To resolve this, a

conference call was made between the leader and the two team members. Both sides of the

argument was requested, Micky said that the only reason he had made a new CAD file was to

show how the servo mechanism was going to be applied to the box. It was clear that it was a lack

of communication between the group and absentee Micky had led to this confusion. It was

concluded that Amit would discuss the payload box design itself and its structural analysis whereas

Micky would discuss the servo mechanism and how it interacts with Pixhawk. The conversation

was concluded by further asking if both team members were happy with the decision made and

they were both were pleased with the outcome.

UAS CHALLENGE 2015


0

1

2

3

4

5

Jonathon Ebhota

0

1

2

3

4

5

Zuber Khan

3.9 Performance Review Performance management allows a business, or in this case, allows the project manager to

determine the strengths and weaknesses of each member in the team. It provides feedback back

to the person being performance managed in areas where they need to develop their skills and

knowledge which they can apply to improve the project in either project delivery or team dynamics.

A performance review of each team member was made to see their strengths and weaknesses

which are broken down into seven different categories; Enthusiasm, Team value, Planning,

Execution, Delivery, and Contribution. As the team itself is quite large, two members were

appointed as team leaders to overlook specific sections; structural and systems. For these two,

three further criteria’s were included; Coaching, Managerial skills and Motivational skill. The

breakdowns of each criterion are as follows:

Enthusiasm: passion and interest for the role and subject

Coaching: training and guiding other team members through their work performance and

subject knowledge

Managerial skills: ability to plan and delegate workload, communicate between team

members, and solving issues between members fairly and objectively

Motivational skills: being able to understand what motivates each member and keeping

them motivated

Team value: quality and information being passed on, insight in topics, availability for help

when asked and general team sportsmanship

Planning: time management, planning for delays and possibly additional workload

Execution: method of execution, holding up other members

Delivery: quality of final work produced for individual role

Contribution: overall workload taken, experience and insight provided and contribution to

the team

Figure 6 – Performance Charts for Jonathan (a) and Zuber (b)

UAS CHALLENGE 2015


Figure 6 presents the performance chart of the two sub team leaders. It describes on a scale from

0 to 5 how they are rated against each category, where 5 is classed as the best and 0 is classed

as the worst. On average it can be seen that Zuber was excellent in most of these categories

whereas Jonathan mostly was good all round and both lacked in motivating team members. A

more detailed individual performance review of each team members can be seen in Appendix A.5.

3.10 Evaluation

Self-evaluation allows one to reflect on how effective their performance was during the project.

During this process, the performance of the leader can be assessed to see how effective it was,

noting areas that need improvement. It should also list the skills developed and what skills need to

be worked on in order to be a better leader if they were to do the project the second time round.

Using the Project manager evaluation form in the Appendix A.6, it was possible for me to evaluate

my performance. It was found that the team scored me 5 in regards to management of the

team/project, having the ability to work with others, ability to present options and reach decisions

and the ability to locate and utilize resources effectively. As a leader of the team and project these

were my strongest areas. An area in which I was marked to be average was the ability to anticipate

and analyse problems. It seems like this is an area in which I need to focus on if I were to manage

a project again. Moreover, as part of the evaluation there were phrases regarding the likelihood to

work again with the project manager on another project. It was reported back that they would be

willing to work with me again on another project with some changes applied.

Further points were asked for regarding any specific strong points/ weak points about my

performance. The positives were: Firstly, I had good form of communication skills and

persuasiveness. Secondly, I was always going the extra mile and continuously standing up for the

team in front of supervisors and I was very supportive. Evidently, it is clear that I possess essential

skills such as good communication, persuasiveness and I’m very supportive of the team member

because I believe as the leader of the team it is my duty to be the voice of the team and the person

held accountable for the team. The negatives were: primarily, I need to keep within deadlines and

secondly, I am far too lenient. These are areas that need to be worked if I’m to manage another

project. I’m apparently too lenient when it comes to deadlines. During the period of the project, we

have had team members that have had personal issues outside the academic work. It meant that if

I had set deadlines for work to be completed and a team member said they couldn’t complete the

task due to personal problems, I would be reluctant to give them more time. This could potentially

cause delays in the initial time plan and hence maybe keeping to a strict deadline regardless of the

personal background situation might be necessary and something I could consider next time.

Overall, I would say I have done a good job in managing this group so far and I have learnt some

vital lessons regarding what to do and what not to do as a project manager.

UAS CHALLENGE 2015

14 Quad-Rotor Design MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Amit Ramji 4 Quad-Rotor Design A Hex-Rotor had been considered during the early stage of the design convergence process,

however during the detail design stage this had been changed to a Quad-rotor design. The reason

for such a dramatic design change is due to mass and cost constraints and is detailed in Appendix.

C and Appendix. D respectively.

Upon detailed consideration of the mass and materials involved with the Hex-rotor, it had been

decided to significantly modify the design and produce a Quad-rotor. As detailed in Appendix. C,

the reduction in mass by alterations in geometry, reduction of parts and optimising the use of

materials results in a very lightweight structure as shown in Figure 7 below. The use of extruded

Nylon 6 main body plates (Appendix B.7) allows for a lightweight structure that is fastened together

into a sandwich design to provide a significantly rigid structure. The use of Carbon Fibre has been

entirely eliminated due to financial constraints; hence a suitable strengthened alternative is

selected. The use of M3 bolts and Nylon 66 blocks (Appendix B.7) allows for a rigid main structure

with multiple load paths. Using the machined Nylon 66 blocks in compression allows for the

majority of the loads to remain in-plane of the main body plates and allows the fasteners to take up

most of the load.

Details of the design architecture and in-depth features are found in Figure 7 through Figure 9 and

Appendix B.7.

Figure 7 - Quad-rotor design

UAS CHALLENGE 2015


4.1 Design Rationale - Quad-Rotor

Figure 8 - Stowage Instructions

Figure 9 - Quad-rotor in Stowed Configuration

Fixed Nylon bracket in compression

Moving Nylon tube position support bracket

Rotating Nylon

Mount with

Spacers and

Through Bolt

Sandwich Design to minimise bending effect with rigid links (M3 bolts)

In-Plane Shear for plates

Remove Quick Release pins (2-Off) for compact stowage.

UAS CHALLENGE 2015


Section by Amit Ramji 4.2 Payload Box Design and Mechanism Figure 10 through Figure 13 show the design of the payload housing with a simple trap-door type

design activated by gravity with release of a servo. The design can be adapted to use either rotary

servo motors or linear actuators. The structural analysis of the payload compartment and its

development is carried out in Appendix G.15 and G.16. Dimensions of the payload compartment,

component parts and BOM can be found in Appendix B.7.

Figure 10 – Removable Lightweight Payload Box

Figure 11 - Removable Lightweight Payload Box

Figure 12 - Payload Box with simple construction and failsafe mechanism

Figure 13 – Payload Box with Payload Clearance

UAS CHALLENGE 2015

17 UAV Mass Breakdown MEng Team Project Report (7ENT1024) School of Engineering and Technology

5 UAV Mass Breakdown Initially a Hex-rotor was considered with a structure mass of 1777.5g, an all-up mass of 7371.9g

with a single payload. The mass of the UAV must not exceed 7kg, hence a complete redesign as a

Quad-rotor has been fulfilled.

A detailed Quad-rotor mass calculation (Appendix. C) has been carried out to ensure the UAV is

within CAA certifiable weights limits to enable flight and to ensure the requirements are met

(IMechE, Jan 2015). The total mass of the Quad-rotor is 6026.2g with single payload. An itemised

breakdown shown in Appendix. C.

UAV Structural Mass

The total mass of the structure is calculated to be 1012.5g including all the materials and fixings

depicted in Appendix B.7. The structure mass is well below the target mass of 1.5 Kg, due to the

extensive and detailed stress analysis carried during the detailed design stage. The entire itemised

breakdown can be observed in Appendix. C.

UAV Electrical / Miscellaneous Components Mass

The total mass of the Electrical / Misc. components is calculated to be 5013.7g including all the

motors, batteries and additional wiring and soldered joints. The itemised breakdown can once

again be observed in Appendix. C.

6 UAV Cost Breakdown Initially a Hex-rotor was considered, which would inherently have increased cost compared to a

Quad-rotor due to increased structural, electronic and propulsion components. It was therefore

unequivocal that a Quad-rotor was the tactic forward to achieving a solution within budget

requirements.

A detailed cost calculation for the Quad-rotor (Appendix. D) has been carried out to ensure the

UAV is within IMechE budget limits (IMechE, Jan 2015). The total cost of COTS items within the

Quad-rotor is £824.84, structure cost of £81.34, hence a total cost of £906.18 with an itemised

breakdown provided in Appendix. D. The above cost summary is inclusive of VAT, less delivery

and is accurate to retail prices at the time of purchase.

Section by Zuber Khan


UAS CHALLENGE 2015

18 Structural Analysis MEng Team Project Report (7ENT1024) School of Engineering and Technology

7 Structural Analysis

7.1 Load Case Definition and Free Body Diagrams

Figure 14 – Free Body Diagram - Flight and Landing Cases

Figure 15 - Free Body Diagram - Landing Cases


UAS CHALLENGE 2015

19 Structural Analysis MEng Team Project Report (7ENT1024) School of Engineering and Technology

Figure 16 - Free Body Diagram - Flight and Gust Load Cases

UAS CHALLENGE 2015

20 UAV Stress Analysis MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Amit Ramji


8 UAV Stress Analysis

8.1 Stress Reduction Techniques The following design techniques have been adopted to maximise efficiency of the material and

ensure a lightweight and stress reduced structure at local discontinuities and overall load paths.

Further comprehensive methods of stress reduction and material optimisation can be found in

Appendix G.1, below is a summarised list of methods.

Align known material properties with major load direction where possible. Hence the use of

Nylon 66 Blocks being used in compression (FB-001 & 002) and fasteners being used in

shear and tension (M3’s & M5’s).

Stiffen or reinforce unsymmetrical features to minimize flexure. An example of this

consideration is the use of the Nylon 66 Fixed Blocks (FB-001 & 002) used in the main

body alongside the M3 Brass spacers which act as rigid links between the main body plates

(BP-001 and 002) to reduce total body deflection.

Encourage smooth transitions in cross section and stress levels, avoiding hard points in the

primary load path. In some cases this could not be avoided (MA-001 contacting FB-002 –

See Figure 115 through Figure 119), therefore an additional local support (MB-001 –

Appendix B.7) is incorporated.

Where appropriate, distribute the load pathways between multiple components to avoid

bulky structure and concentrated stress distributions on single components. An example of

such situation is the multiple load paths in the main body, where a sandwich type design is

achieved. The stiffness of the main body structure is greatly increased with rigid links (M3

Fasteners, FB-001, FB-002, MB-001 and M3 spacers).

8.2 Fatigue Awareness A gain in fatigue life can in most situations be achieved without an increase in cost, simply by

attention to design detail. Further comprehensive methods of fatigue resistance with material

optimisation can be found in Appendix G.1, below is a summarised list of methods considered. The

following should be taken into account when considering the Quad-Rotor structure:

Avoiding sharp edges, corners and sudden changes in cross-section can reduce stress

concentrations. Fillet and intersection radii should be as large as possible as such used in

the Lug Bracket (LB-003) and Pivots (AP-001 & LP-001).

The majority of fatigue cracks will start at stress concentrations such as holes, notches, etc.

Any design features or processes that can be applied to reduce the severity of such stress

concentrations should be used.

Ensuring design of joints are such as not to give rise to built-in stresses on assembly, or

load some portions of the joint unduly. The use of M3 and M5 from the same supplier to

avoid mixing fasteners of dissimilar material/strength and those that require differing

UAS CHALLENGE 2015


tolerances of fit. Fasteners with tighter tolerances will load the local structure during

repeated flexure more than a loose tolerance fastener due to the miniscule freedom of

movement of the joint.

In fatigue critical areas, interference fit fasteners shall be used whenever possible in

preference to clearance fit. A close tolerance for clearance/transition fit fasteners will

improve the fatigue performance of the joint, as this will minimize the risk of individual holes

being over-loaded. For the current Quad-Rotor design, fasteners are loaded axially hence

introducing a bolt pre-load and reducing the miniscule movement if any existed.

8.3 Fatigue due to induced vibration A gain in fatigue life due to induced vibration can also be achieved simply by attention to design

detail, material selection, edge distances and overall geometry. Further comprehensive methods of

optimisation can be found in Appendix G.1, below is a summarised list of methods considered. The

following should be taken into account when considering the Quad-Rotor structure and rotating

components:

Fatigue damage can often arise from induced vibration from the motors as compared with

fatigue damage arising from directly applied structural stresses. Often this vibration is not

sustained for long periods of time, a modal analysis case has been considered for the

Fixed-arm assembly as shown in section 8.17 and compared to analytical methods as

shown in section 8.16. Such calculated modal frequencies should be avoided or swiftly

passed through the first 3 natural frequencies when powering up the motors to idle and can

be programmed into the ESC’s as “soft, medium, hard” starts.

Avoiding the use of long cantilevered members, as these will experience high inertia forces

in vibration. The modal analysis of the Arm has been the main concentration for the

purpose of frequency response analysis, as the cantilever of the Arms are more susceptible

to vibration than any other components.

8.4 Pressure Loading on Plates A complete structural analysis was carried out on the UAV with the main stresses and loads

summarised below. The first scenario to be analysed was the UAV in flight, flying at maximum

speed allowable with maximum head on gusts off 25knots. The distributed load calculated in

Appendix F.2 approaches to 5.27Kg, which has a 1.5 global load safety applied to it. This was then

used to determine the deflection and stress of a simplified UAV model.



UAS CHALLENGE 2015


8.5 Load Transfer Loads are transferred from the arms to the Nylon clamps using a moment balance shown in Figure

108. Reaction loads passing through the clamps could then be calculated, the Fixed-arm clamp

having 65.18N passing through it and the Movable-arm having 63.29N.

Figure 17 – Fixed Arm Cross Section – See also Appendix G.7

8.6 Fixed and Movable Arm Stress Maximum The maximum bending stress experienced on the Fixed-arm is 14.42MPa as shown in Appendix

G.7 and the maximum bending stress experienced by the Movable-arm is 15.26MPa as shown in

Appendix G.8. Refer to Appendix B.7 for parts list, Appendix. E for material properties, G.4 for

boundary conditions, G.5 for Finite Element solver method, G.6 for mesh types and properties and

G.7 - G.8 for results of the contact model for bending case of the UAV Arms. A Sample calculation

for the Fixed-arm is shown below:

𝜎 =𝑀𝑦

𝐼=

25 × 0.17 ×0.016

2𝜋

64(0.0164 − 0.01154)

= 14.42𝑀𝑝𝑎

D1 F1

F2

D2


FEA by Amit Ramji Analytical by Zuber Khan

UAS CHALLENGE 2015


Stress analysis at Fixed-arm – FEA Method

Mesh: Values as per section G.6

Figure 18 – Mesh for Fixed-arm Assembly – Values as per Appendix G.6 Results:

Figure 19 - Deflection of Fixed-arm Assembly (Flight Loads) with 7.6mm Deflection

Figure 20 - Stress of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa (Peak)


UAS CHALLENGE 2015


Figure 21 – Stress (Close-up) of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa (Peak)

FEM Verification: Tube Stress Comparison

One can observe the results from the above analytical stress calculation being 14.42MPa and the

stress level as seen in the far field stress contour of the tube in Figure 21 (15.8MPa) being very

close. Substantiation of the numerical modelling and contact constraints can be deemed as

accurate as a very small difference is observed between the methods.

8.7 Simplified Plate Deflection Plate deflection has also been calculated analytically to enable comparison to an FEA model,

ensuring the modelling techniques are correct and establishing meshing and connection properties

to be used on the entire UAV FEA model. The analytical method calculated a deflection of

4.555mm, whereas the FEA package calculated 4.54mm (Appendix G.11). These results are in the

same order of magnitude and are marginally different; therefore the modelling technique is deemed

correct and usable throughout.

8.7.1 Simply Supported Plate Representation

A simple plate deflection was determined of a 2mm thick Nylon plate with dimensions of 315mm by

280mm. This was the largest the plate would go to on the UAV if necessary therefore was used for

the purpose of this analysis. The reason for this was to compare the analytical results with the

results produced by the FEA model. If the results were similar or close to the analytical method, the

modelling method could be applied to the whole UAV model where the plates are used.



UAS CHALLENGE 2015


Figure 22 - Simplified Plate Representations

All edges simply supported for this analysis.

8.7.2 Analytical Method

Below are the Navier stokes equations used to work out the plate deflection at the centre, where

the maximum deflection will take place.

𝐷 = 𝐸𝑡3

12(1 − 𝑣2)

Equation 1 - Flexural Rigidity of the Plate (Ventsel and Krauthammer, 2001)

𝑤(𝑥, 𝑦) = ∑ ∑ 𝑤𝑚𝑛 sin𝑚𝜋𝑥

𝑎sin

𝑛𝜋𝑦

𝑏

∞

𝑛=1

∞

𝑚=1

= 𝑤11 sin𝜋𝑥

𝑎sin

𝜋𝑦

𝑏+ 𝑤12 sin

𝜋𝑥

𝑎sin

2𝜋𝑦

𝑏+ 𝑤21 sin

2𝜋𝑥

𝑎sin

𝜋𝑦

𝑏+ …

Equation 2 – Navier solution (Ventsel and Krauthammer, 2001)

𝑎𝑚𝑛 =16𝑞0

𝑚𝑛𝜋2

Equation 3 - Navier stokes coefficient 1 (Ventsel and Krauthammer, 2001)

𝑤𝑚𝑛 =1

𝜋4𝐷

𝑎𝑚𝑛

[(𝑚2

𝑎2 ) + (𝑛2

𝑏2)]2

Equation 4 - Navier Stokes coefficient 2(Ventsel and Krauthammer, 2001)

First the pressure distributed on the whole plate surface was calculated.

𝐹𝑜𝑟𝑐𝑒

𝐴𝑟𝑒𝑎=

33.8445

88200 × 10−6= 383.72

𝑁

𝑚2

Followed by calculating the flexural rigidity

𝐷 = 3300 × 106 × 0.0023

12(1 − 0.42)= 2.61905

The Navier coefficients 1 and 2 could be calculated for when mn = 1 1, 1 3, 3 1, 3 3

𝑎11 =16×383.72

1×1×𝜋2 = 622.063 𝑎13 = 207.35 𝑎31 = 207.35 𝑎33 = 69.12

X = a = 315mm

Y =

b =

280m

m Youngs Modulus, E = 3300MPa

Thickness, t = 0.002m

Poisson’s Ratio, v =0.3

Distributed Force = 33.8445N

Area = 88200 x 10e-6 m2

UAS CHALLENGE 2015


w11 =1

π4 × 2.61905

622.063

[(12

0.3152) + (12

0.282)]2 = 4.67689 × 10−3

w13 =1

π4 × 2.61905

207.35

[(12

0.3152) + (32

0.282)]2 = 5.21215 × 10−5

w31 =1

π4 × 2.61905

207.35

[(32

0.3152) + (12

0.282)]2 = 7.59333 × 10−5

𝑤33 =1

𝜋4 × 2.61905

69.12

[(32

0.3152) + (32

0.282)]2 = 6.41566 × 10−6

The coefficients were then input into the Navier solution equation to calculate the deflection at the

centre.

𝑤(𝑥, 𝑦) = 4.67689 × 10−3 × sin (𝜋 × 0.1575

0.315) × sin (

𝜋 × 0.14

0.28) + 5.21215 × 10−5 × sin (

π × 0.1575

0.315)

× sin (3π × 0.14

0.28) + 7.59333 × 10−5 × sin (

3π × 0.1575

0.315)

× sin (π × 0.14

0.28) + 6.41566 × 10−6 × sin (

3π × 0.1575

0.315) × sin (

3π × 0.14

0.28)

w(x, y) = 4.67689 × 10−3 − 5.21215 × 10−5 − 7.59333 × 10−5 + 6.41566 × 10−6

w(x, y) = 4.555 × 10−3m = 4.555mm

8.7.3 FEA – Simplified Rectangular Approximation

Using Catia the same plate was modelled with the same constraints and loads to see the deflection

it would cause.

Figure 23 - Simple Plate Deflection Carried out on CATIA showing 4.54mm deflection From the FEA model (Figure 23) it was found that the deflection has been calculated to be

4.54mm.

UAS CHALLENGE 2015





The mesh used was set to a size of 2mm with absolute sag of 1.5mm. Therefore any further plate

bending analysis carried out on CATIA, should be set to the same mesh size and constraints as it

has been substantiated to provide accurate answers.

Method Deflection

Analytical (Rectangular Plate) 4.555mm

FEA CATIA (Rectangular Plate) 4.54mm

Table 4 – Comparison of Simplified Plate Deflection for Model Substantiation

8.8 Plate Deflection - Assembly Contact Model as Built

To enable an accurate understanding of plate deflection as an assembly, a non-linear contact

model has been modelled in Ansys and shows a very small deflection of ≈0.13mm. The reason for

such a reduction in deflection compared to the simplified substantiation is due to the presence of

rigid bodies (Fasteners and FB/MB

series blocks). Refer to Appendix B.7

for parts list, Appendix. E for material

properties, G.4 for boundary

conditions, G.5 for Finite Element

solver method, G.6 for mesh types

and properties and G.12 for results of

the contact model for in-flight case of

the Quad-rotor.

Figure 24 - Flight and Gust condition of Main Body with 0.13mm Deflection

8.9 Undercarriage Buckling Calculation The undercarriage is also analysed to check whether it is suitable for heavy landings and repeated

loadings. The critical load was calculated in Appendix G.13 which was 393.7N = 40.13Kg. Meaning

the UAV could land on a single undercarriage and be able to withstand a load of ≈40Kg before

buckling. A sample calculation from G.13 is shown below:

𝑃𝐶𝑅

𝜋2 × 𝐸𝐼

(2𝑛𝐿

𝜌 )2 =

𝜋2 × 3100 × 106 × 9.7193 × 10−5

(2 × 0.18

4.926 × 10−3)2 = 556.78𝑁

8.10 Undercarriage Bending Analysis on pure bending has also been carried out in Appendix G.13, to represent a pivot jam or

lateral sideward landing on a single undercarriage leg. With the applied 1.5 global load safety

factor the stress experienced by the undercarriage leg was in the region of 62.2MPa, being higher

than the yielding properties of the PVC material (Appendix. E). However this analysis has assumed

UAS CHALLENGE 2015



a worst-case scenario with the UAV landing on a single leg, which can now be avoided. The UAV

would also share multiple load paths if a misbalanced landing were experienced therefore reducing

the stress. Additionally, the entire Quad-rotor structure would deflect as a result of such bending

impact, highlighting that a parent non-linearity has not been considered. To further analyse such

parent non-linearity on a single undercarriage leg, spring constraints at the Lug bracket (LB-003)

bolt holes with the stiffness of the main body structure can be modelled

8.11 Undercarriage Bending - Assembly Contact Model In order to obtain an accurate understanding of landing conditions, a 1-second impact case has

been created on Ansys to highlight potential failure points. It is worth noting the analytical

technique described above in section 8.10 with a stress of 62.2 MPa is very close to that shown in

Figure 25 (60.63MPa). From this similarity in analytical and numerical methods, it is conclusive that

the analytical modelling techniques are substantiated and can be relied upon for further analysis if

required. Refer to Appendix B.7 for parts list, Appendix. E for material properties, G.4 for boundary

conditions, G.5 for Finite Element solver method, G.6 for mesh types and properties and G.13 for

results of the contact model for bending case of the undercarriage.

Figure 25 - Lateral Impact Case on Single Leg - 60.6MPa Stress

UAS CHALLENGE 2015




8.12 Undercarriage Torsion Torsional analysis has also been carried out to determine the twist the undercarriage would

experience if the UAV landed on the tip of one horizontal leg (UH-001 - Appendix B.7). Appendix

G.13 calculates a pure torsion case to be used for a combined loading effect in section 8.13 and

8.14. The calculated twist angle is 0.6257rad or 35.85°, the twist angle being of such high

magnitude indicates a high stiffness constraint at the boundary condition or a significantly high load

due to single leg impact assumptions. However the assumption of a single leg impact is a rare

occasion and can now be avoided. The shear experienced by the undercarriage due to the twist is

calculated to be 30.57MPa which is significantly low compared to the PVC yielding properties in

shear being 1099.3MPa (Appendix. E).

8.13 Undercarriage Combined Loading - Torsion and Bending

A combined loading analytical method is also carried out on the undercarriage leg representing 3

loads being applied at the same time including a torsion, buckling and bending loads as shown in

“Analytical – Undercarriage Combined Loading – Bending, Buckling and Torsion” of Appendix

G.13. The principle stress is calculated as 27.1MPa and -34.5MPa, which is acceptable due to the

yielding strength of the PVC being 55MPa (Appendix. E). The loads were calculated with an

applied 1.5 global load safety factor and the over engineered assumption of a single leg impact.

The principle angle of the stresses were -41.55° and 48.45° respectively and a sample calculation

is shown below:

𝜎1 =−1.059798 − 6.34

2+

1

2√(−1.059798 − 6.34)2 + 4 × 30.5722 = 27.1 𝑀𝑃𝑎

The maximum shear caused by the

combined loading is calculated to be

30.795MPa, which is also well within

the capabilities of the material.

Figure 26 – Stress Element A with Principal Stress for - Analytical – Undercarriage Combined Loading – Bending, Buckling and Torsion (G.13)

𝜏𝑥𝑦

𝜎𝑥

𝜎𝑦

𝜏𝑥𝑦

𝜏𝑦𝑥 𝜎𝑦

𝜎𝑥

𝜏𝑦𝑥

A 𝜃1

𝜃2

27.095

-34.495

UAS CHALLENGE 2015




8.14 Undercarriage Combined Loading - Assembly Contact Model

An FEA method with combined torsion, bending and shear loads have been applied to a single

undercarriage leg in Appendix G.13 titled “FEA Results – Combined Torsion and Bending – Tip

Contact”. Refer to Appendix B.7 for parts list, Appendix. E for material properties, G.4 for boundary

conditions, G.5 for Finite Element solver method, G.6 for mesh types and properties and G.13 for

results of the contact model for combined tip loading of a single undercarriage.

8.15 FEM Verification – Summary of Undercarriage Results

Case Description Deflection (mm) or (deg)

Equivalent Load (N) or Stress

(MPa)

Buckling Analytical Axial loading of UV-001 N/A 393.7N

Bending Analytical Bending of UV-001

N/A 62.2MPa

Bending FEA 53.6 mm 60.63MPa

Torsion Analytical Torsion of UV-001 35.85 deg 30.57MPa

Combined Analytical

Combined Bending and Torsion of UV-001

N/A 34.495MPa

Combined FEA 66.76mm 71.76MPa

Table 5 – Summary of Undercarriage Results – See G.13

UAS CHALLENGE 2015



8.16 Modal Analysis of Fixed-arm – Simplified Case Parts: As per Appendix B.7

Materials: As per Appendix. E

8.16.1 Analytical Modal Analysis – Simplified

Modal analysis was carried out to determine the natural frequency of the UAV arm with the full

assembly of parts with their corresponding weights. Once the natural frequency is known, one can

program the autopilot system (Pixhawk) and ESC’s to ramp through the primary natural

frequencies to ensure excessive vibration is not encountered. The ESC’s can control the motors to

have a “soft/ medium/hard” start to idle for this reason and the modal frequencies can be avoided

to protect the structure (loosening fasteners, fatigue and instability during flight).

Figure 27 - Arm and Mass for Rayleigh Method To determine the natural frequency of the arm with the weight of all attached components, the

following equations were used.

∅1(𝑥) = 𝑎1(3𝐿𝑥2 − 𝑥3)

Equation 5 -Static Deflection Curve (MEGSON, 1999)

𝜔2 =∫ 𝐸𝐼 (

𝑑2∅𝑑𝑥2)

2

𝑑𝑥 + ∑ 𝑘𝑗∅2(𝑥𝑗)𝑁𝑗=1

𝐿

0

∫ 𝜌𝐴∅2𝑑𝑥 + ∑ 𝑚𝑗∅2(𝑥𝑗)𝐽𝑗=1

𝐿

0

Equation 6 - Rayleigh's Natural Frequency Equation (MEGSON, 1999)

To be able to calculate the natural frequency using Equation 6, the static deflection equation

requires to be differentiated twice.

𝑑∅1(𝑥)

𝑑𝑥= 𝑎1(6𝐿𝑥 − 3𝑥2) ∴

𝑑2∅1(𝑥)

𝑑𝑥2= 6𝑎1(𝐿 − 𝑥)

The deflection where the concentrated mass is attached:

∅1(𝑥 = 0.231) = 𝑎1(3𝐿𝑥2 − 𝑥3) = 0.025625808

Using that and inputting some of the values the equation becomes:

𝜔2 =12𝐸𝐼𝐿3

𝜌𝐴 [9𝐿2𝑥2

5−

6𝐿𝑥6

6 +𝑥7

7 ]0

𝐿

+25𝑚𝐿6

64

𝜔2

=12 × 3100 × 106 × 2.35845 × 10−9 × 0.2343

1.4 × 9.7 × 10−5 × [9 × 0.2342 × 0.2342

5−

6 × 0.234 × 0.2346

6 +0.2347

7 ] +25 × 0.36846 × 0.2346

64

L = X = 0.234m

m = 0.36846Kg

UAS CHALLENGE 2015


𝜔2 =1.124131676

(1.4 × 9.7 × 10−5 × 5.361854627 × 10−3) + 2.362901004 × 10−5 =

1.124131676

2.43571499 × 10−5

𝜔 = √1.124131676

2.43571499 × 10−5= 214.83𝑟𝑎𝑑/𝑠 = 34.19𝐻𝑧 = 2041.93𝑅𝑃𝑀

From this it can be concluded that the natural frequency of the simplified arm is 34.19Hz.

Rayleigh’s method usually always over predicts, therefore in reality the natural frequency will be

slightly lower.

UAS CHALLENGE 2015


Section by Amit Ramji 8.16.2 Finite Element Modal Analysis – Simplified

Parts: As per Appendix B.7

Materials: As per Appendix. E


Results:

Figure 28 – Mass Representation of Motors, Blocks, Plates, Fasteners and ESC

Figure 29 – Simplified FE analysis with 1st Nat freq as 19.64Hz – 69.3mm Deflection (Left) and 164MPa Stress (Right)

Figure 30 – Simplified FE with 2nd Nat freq as 20.06 Hz (Left) and 3rd Nat freq as 134.6 Hz (Right)

Figure 31 – Simplified FE with 4th Nat freq as 224.1 Hz (Left) and 5th Nat freq as 411.9 Hz (Right)

UAS CHALLENGE 2015



8.17 Modal Analysis of Fixed-arm – Actual Parts (As Built) Parts: As per Appendix B.7 Materials: As per Appendix. E Mesh: Values as per section G.6 Results:

Figure 32 – As Built FE Analysis - Mass Representation of Motors, Fasteners, Cables and ESC

Figure 33 – As Built FE analysis with 1st Nat freq as 451 Hz – 69.0mm Deflection (Left) and Stress (Right)

Figure 34 - As Built FE analysis with 2nd Nat freq as 736 Hz (Left) and 3rd Nat freq as 1707 Hz (Right)

Figure 35 - As Built FE analysis with 4th Nat freq as 2 KHz (Left) and 5th Nat freq as 4.1 KHz (Right)

UAS CHALLENGE 2015


Section by Amit Ramji 8.18 Summary of Modal Frequency Results

1st Nat Freq (Hz)

2nd Nat Freq (Hz)

3rd Nat Freq (Hz)

4th Nat Freq (Hz)

5th Nat Freq (Hz)

Simplified Analytical (8.16)

34.19 N/A N/A N/A N/A

Simplified FEA (8.16)

19.64 20.06 134.6 224.1 411.9

As-built FEA (8.17)

451 736 1707 2000 4100

Table 6 – Summary of Modal Frequencies for Fixed Motor Arm As predicted from the Rayleigh method in Section 8.16, the actual natural frequency will be slightly

lower between the 34.19 Hz Vs the 19.64 Hz. From this simplified analysis, one can substantiate

the modelling techniques used in the FEA for more complex assemblies. The As-built cases have

significantly higher modal frequencies and was also predicted due to the increased stiffness when

considering fastened motor plates and blocks. Additionally it is worth noting that the higher less

important frequencies have modal excitation closer to the motor mount plates, hence the reason

for selecting Aluminium Alloy plate as a mounting material for the motors (Appendix B.7 and

Appendix. E). Aluminium Alloy compared to the cast mild-steel motor brackets which are supplied

with the motors are less susceptible to fatigue damage due to repetitive vibration.

UAS CHALLENGE 2015


Section by Zuber Khan 8.19 Summarised Margin of Safety Table Below is a margin of safety table which has maximum loads and stresses which could be applied

onto the Quad-rotor and also the maximum allowable loads and stresses. Using the maximum and

allowable loads and stresses, safety factors were obtained.

Part No. (Appendix B.7)

Case / Calculation / Section

Loading Description

Maximum Applied Load/Stress

Maximum Allowable Load/Stress Appendix. E

Safety Factor, SF= Allowable /Applied

FA-001 Case 1 (G.7)

Maximum Thrust from Motors

14.42MPa 55MPa 3.81

MA-001 Case 1 (G.8)

Maximum Thrust from Motors

15.26MPa 55MPa 3.60

UV-001 Case 2 (G.13)

Undercarriage Pipe Under Buckling

10.5Kg 56.76Kg 5.41

Case 4 (G.13)

Undercarriage Pipe Under Torsion

30.57MPa 1099.3MPa 35.96

LB-003 (G.9) Undercarriage Lug Under Maximum Loading

72.84N 1765.15N 24.23

UV-001 Analytical – Undercarriage Combined Loading – Bending, Buckling and Torsion (G.13)

Combined Loading 𝜎1 on Undercarriage Vertical Leg

27.09MPa 55MPa 2.03

Combined Loading

𝜎2 on Undercarriage Vertical Leg

34.5MPa 55MPa 1.59

BP-001 & BP-002 Assembly.

Appendix G.12 Main Body Deflection due to Maximum Thrust and Gusts

5.83MPa 55MPa 9.43

Table 7 - Summarised Margin of Safety Table

UAS CHALLENGE 2015

37 Performance, Propulsion & Systems Engineer

MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Hassan Dzadey 9 Performance, Propulsion & Systems Engineer

As a practicing performance and propulsion engineer the key parameters that were vital within this

report was to investigate and identify possible propeller, motor, esc (electronic speed controllers)

and power supply combinations that are efficient and also cost effective with the ability to achieve

the mission objectives set by the IMechE UAS challenge. It is also within the interest of this report

to point out the work that has been conducted as a systems engineer to improve the navigation of

the UAS (Unmanned Aircraft Systems) and target tracking.

At the start of the MEng project IMechE had set specific limitations to which partly involved the

performance of the UAS along with mission details. These specifications are identified below which

were strictly followed:

Maximum Take-Off Mass (MTOM) must be equal to, or less than 7kg

Must have the capability to fly under 20knots wind and 25knots gust conditions

Maximum airspeed of 60knots (IAS) must not be exceeded

Must be capable of operating within altitude range of 100ft-400ft

Must have the ability to complete 2km round mission

On top of the IMechE specification there were specifications set by the MEng group, which are

listed below and also strictly adhered to.

Initial cost limitation of £550 after taking into account structural other electrical components

Initial propulsion and power supply weight limitation of 3.7kg was set after taking into

account structural, payload and electrical components weights

At the start of the MEng project various design concepts such as aeroplane, helicopter, Quad-rotor,

Hex-rotor, octocopter and osprey tilt rotor were put forward and analysed and after careful

alliteration the Hex-rotor was chosen as the design that the group would like to construct and put

thought to the IMechE UAS Challenge. Hence for the PDR the performance and propulsion

calculations were based on Hex-rotor as shown in appendix A. During mid-January it was identified

that while the cost would be under the £1000 limit set by the IMechE, the maximum take-off mass

of 7kg would be exceeded by 500g. From this point it was decided to change the design to Quad-

rotor, hence from Appendix H.2 onwards the calculations will be based on and around the Quad-

rotor design.

Most of components that are investigated in this report involves two unknown variables that are

required e.g. a propellers two variables involves its diameter and pitch, the RC motor has KV and

power, and the power supply requires voltage and capacity calculations. For this reason this report

has been split into sections which address one variable per component at a time by process of

elimination.

UAS CHALLENGE 2015



9.1 Propeller Diameter Selection To perform any type of performance and propulsion calculations the MTOM is vital, for this initially

7kg is used which was the estimated mass and also the maximum permissible mass from IMechE

specifications. Also 7kg is used as it would be easier to down grade the performance at a later

stage in the project if it is required than to do it the other way around. Having initialised the MTOM,

the lift required to hover per motor can be calculated using Equation 7.

(MTOM ∗ 9.81

Number of motors) = 17.18N/motor

Equation 7 - Lift Required Using the same principle as helicopter the Quad-rotor must sustain lift and also move forward by

changing pitch, therefore the propellers must have the capability to sustain lift required and also

thrust for forward movement which results in Equation 7 being insufficient and it has to be

modified. The modification can be seen in Equation 8 to account for lift and thrust. Equation 8 has

been obtained and validated of its use from different Quad-rotor builders and hobbyists alike.

(MTOM ∗ 2 ∗ 9.81

Number of motors) = 34.34N/motor

Equation 8 - Modified Lift Equation Although propellers are the one of the cheapest components that will be integrated onto the Quad-

rotor they are single handily the most vital components to achieving efficient performance. There

are 11 different companies that produce propellers from different material properties for the RC

enthusiasts, but there is only 5 companies (Aeronaut, APC, DJI, EMax and Graupner) that

manufacture multicopter props which are in the orientation of CCW (Counter Clock Wise) and CW

(Clock Wise-normally referred to as ‘pusher’). Using Equation 7, (Staples, 2014) the lift that each

propeller can produce at different RPM (Revolutions Per Minute), in this case between 0 –

20,000RPM will be analysed and a plot of Lift Vs RPM will be produced. A sample of this equation

at work can be seen in Appendix H.2 figure: 1.1. Other investigated propellers are documented in

appendix B

L = 4.392399*10-8*RPM*𝐷𝑖𝑎𝑚𝑒𝑡𝑒𝑟 3.5(𝑖𝑛)

√𝑃𝑖𝑡𝑐ℎ (𝑖𝑛)*(4.23333 ∗ 10−4 ∗ 𝑅𝑃𝑀 ∗ 𝑃𝑖𝑡𝑐ℎ (𝑖𝑛))

Equation 9 - Length of Propeller From figure 1.2, in Appendix H.2 shows that as the propeller diameter increases the lift required

which in this case is 34.34N can be achieved at a lower RPM value of 6,200RPM therefore making

the system more efficient as the current draw will be lower but although having a larger propeller

would be more efficient the velocity will be effected as higher RPM results in faster flying. In this

case the ideal propeller will be in the top right hand corner of the black box below which

corresponds to propeller dimensions of 9*4.7 at 20,000RPM that can also achieve lift required.

UAS CHALLENGE 2015



Now that a RPM range 6,200 to 20,000 is identified in Appendix H.2 1.2, the power versus RPM

graph can be produced to identify the power required from a specific RC brushless motor. Equation

10 (S, 2014) is used to establish the power vs RPM graph:

𝑃𝑜𝑤𝑒𝑟 (𝑊) = 𝐾𝑝 ∗ 𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝐷𝑖𝑎𝑚𝑒𝑡𝑒𝑟4(𝑓𝑡) ∗ 𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑃𝑖𝑡𝑐ℎ(𝑓𝑡) ∗ 𝑅𝑃𝑀3 ∗ (1 ∗ 10−9)

Equation 10 - Power Produced by Propeller

Where Kp = Propeller constant listed in 221 H.2table 1.1 as each propeller manufacturer has its

own propeller constant.

From figure: 1.3 in Appendix H.2 it identifies that a propeller with dimensions 7*6 rotating at

20,000RPM is an appropriate match for having the lowest power consumption of 473W for

efficiency coupled with highest RPM value for high speed. But now looking at figure 1.2 lift vs RPM

it can be seen that a propeller with dimensions of 7*6 will not produce the required lift of 34.34N

but only attain 17.34N. Although this propeller would be great for speed and efficiency it would not

have the required lifting capability to sustain flight. This process is repeated for every propeller and

the results are shown in Appendix H.2 table 1.2 shows the results for different propeller dimensions

together with RPM, power consumed and lift produced which is obtained from figures 1.2 and 1.3.

The maximum RPM used for each propeller to calculate its lift capability is 20,000RPM as

mentioned earlier, but even though the propeller is spinning at 20,000RPM there are still certain

propellers that cannot achieve the minimum lift required per motor of 34.34N and therefore

assigned with the letter N in the acceptability section of the table. This represents that the propeller

performance is not acceptable, this range falls from propellers 7*6 to 9*3.8.in Appendix H.2 Table

1.2 also shows the some propellers at 20,000RPM can produce in excess of 34.34N of lift and

therefore it was required to reduce the RPM to obtain the required lift, one example of this would

be propeller 9*6 at 20,000RPM produced 39.84N of lift and required 1265W to achieve this. As

there this no need to have the excess 4.82N of thrust the RPM can be reduced down to 18,750

which in turn makes the whole system run more efficiently and the power consumption reduces

down to 1042W which in turn results in reduced current draw. The results can be seen on table 1.3

in Appendix H.2

This type of analysis can be seen in larger propeller dimensions such as 17*10, table 1.4 in

Appendix H.2 has the results for this propeller size and it can be seen that a propeller of this size

would achieve well in excess of the 34.34N lift required, 476.45N with 20,000RPM while requiring

2685.2W. This would results in excess thrust of 442.11N which is not required, therefore the RPM

can be reduced down to 5,450 which in turn produces the lift required of just over 34.34N, hence

less current draw making the whole system more efficient.

UAS CHALLENGE 2015



Looking at the two propellers sizes in tables 1.3 and 1.4 in Appendix H.2 it can be clearly seen that

having a larger propeller definitely increases the endurance time of the Quad-rotor because the

power consumption required is reduced by half from 1042W to 543W which in turn means lower

current draw for the same amount of lift produced. Lower current draw results in longer flight time.

One of the disadvantages of increased propeller diameter is the fact that the RPM is reduced

therefore effecting the velocity of the Quad-rotor. Finding the best propeller combination between

these two propellers that would give the lowest current draw with the highest velocity while

maintaining the lift required of 34.34N was the key engineering challenge that was faced though

out this project.

9.2 RC Motor Selection Maximum RPM

Now that a range of different propeller dimensions are identified in 9.1 we will now look into the

motors that are available for use that can be matched to the identified propellers to achieve the

best combination in terms of performance.

Looking back at table 1.2 in Appendix H.2 it can be identified that the investigated motors must

have an RPM range between 5,450RPM to 20,000RPM and also capable of supplying power

between the range of 543W to 1195W so that the Quad-rotors performance abilities can be

achieved. To achieve the best performing RC motor for this project different motors are researched

and investigated from different manufactures, the results are shown in Appendix H.3, table 1.5

Table 1.5 shows the performance details of each motor stated by the manufacturers at the time of

build. One of the details that is not give is the maximum RPM of the motor once a propeller is

attached to it. To calculate this Equation 11 (Bernhard, 2009) is used.

RPM= KV*maximum cell voltage*reducing factor

Equation 11 - Determining RPM

The KV of a motor is specified by the manufacturer and it represents revs per minute per voltage

e.g. taking the details specified in table 1.6 located in Appendix H.3. It show that this specific motor

EMax GT2820/07 has an RPM rating of 850 per voltage supplied. And again from table 1.6 in

Appendix H.3 it can be seen that the manufacturer has stated that a cell range of between 3s and

4s is permissible where 1s is the equivalent of 3.7V therefore 4s (4*3.7V = 14.8V). The reducing

factor in Equation 11 represents the drop in motor maximum RPM capability when a propeller is

attached which is in the region of 0.83. An example of how the maximum RPM is calculation can

be seen below.

850*(4*3.7)*0.83 = 10064RPM

UAS CHALLENGE 2015



By analysing the data obtained from table 1.5 in Appendix H.3 the maximum RPM that these RC

motors can achieve is identified, therefore any propeller that requires higher RPM than what the

motors can achieve is eliminated, this is certainly true for propellers have small diameters and

require high RPM to attain the lift required. From table 1.5 in Appendix H.3 it can also be seen that

the maximum RPM that can be achieved is 14,800, therefore by looking at Appendix H.3 in table

1.2 it can be seen that propellers that are in dimension range of 7*6 to 10*7 and 11*3 can be

eliminated as they require higher than 14800 RPM. Therefore table 1.2 can be reduced down to

1.2.1 in Appendix H.3

9.3 Propeller Pitch Selection To eliminate more propellers from table 1.2.1 in Appendix H.3 we now look at the velocity that the

Quad-rotor will be designed for. At the start of the project a specification was written that stated,

“Maximum airspeed of 60knots (IAS) must not be exceeded” and also “Must have the capability to

fly under 20knots wind and 25knots gust conditions”. In this analysis the decision to assume that

the Quad-rotor would be traveling under the most extreme case scenario was taken throughout the

2km course.

Initial calculation can be performed to evaluate the velocity required to complete the course in the 2

minute time frame using speed equation stated in Equation 12 (Anon, 2014). (Although 2 minutes

is stated here the Quad-rotor will have the capacity to fly for 5 minutes as a contingency)

𝑆𝑝𝑒𝑒𝑑 (𝑚

𝑠) =

𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 (𝑚)

𝑇𝑖𝑚𝑒 (𝑠)=

2000 (𝑚)

120 (𝑠)= 16.6𝑚/𝑠

Equation 12 - Determining Speed From Equation 12 - Determining Speed it can been seen that under ideal conditions (zero wind)

the Quad-rotor is required to fly at a velocity of 16.6m/s to achieve the 2km in 2 minutes, but the

above equation has not considered wind speeds and gust conditions of up to 25 knots and

therefore will be considered below.

20knots wind speed = 10.28m/s

25knots gust speed = 12.86m/s

To complete the course within the time frame stated of 2 minutes the Quad-rotor must be capable

of travelling at velocity of between 26.88m/s (52.25knots) under maximum wind speed and at

29.46m/s (57.27knots) under maximum gust conditions.

To calculate the maximum velocity that can be obtained requires the maximum tilt angle, this can

be achieved by using Equation 13 (Anon, 2014).

F*cos(𝜃) = Weight

Equation 13 - Force Produced at Maximum Tilt Where:

F = Force (N)

W = Quad-rotor weight (N)

UAS CHALLENGE 2015



𝜃𝑚𝑎𝑥 = Maximum tilt angle (degrees)

The total force that the Quad-rotor will produce is obtained by the propellers lifting capabilities in

this case each of the four propeller chosen produces 34.34N, 137N in total. By rearranging

Equation 13, the force required at a particular angle can be calculated, and by trial and error (as

long as it doesn’t exceed 137N *0.7 = 95.9N because we need excess thrust for sudden gusts) we

can find the maximum tilt angle.

𝐹𝑜𝑟𝑐𝑒 (𝑁) =𝑊𝑒𝑖𝑔ℎ𝑡 (𝑁)

cos (𝜃)

Using trial and error When 𝜃 = 10; F= 69.7N 𝜃 = 20; F=73.0N 𝜃 = 30; F=79.29N 𝜃= 40; F=89.6N 𝜃=

44; F=95.4N

From this trial and error section it can be seen that a maximum angle of 44 degrees can be

achieved, but if this angle is exceed than there is the possibility that the Quad-rotor will stall

therefore it is advisable to use an angle setting of less than 44 degrees, in this case 32 degrees is

used.

Now that the maximum flight angle is obtained Quad-rotor maximum speed in straight flight can be

calculated using Equation 14 (Andy, 2014).

𝑉max 𝑆&𝐿𝐹 = 𝑅𝑃𝑀𝑚𝑎𝑥 ∗ 𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑃𝑖𝑡𝑐ℎ ∗ 0.000954 ∗ 0.44704

Equation 14 - Quad Rotor Maximum Speed Equation 14 assumes that the Quad-rotor will be travelling parallel with the x-axis like an aircraft,

but for a Quad-rotor Equation 14 has to be modified to take into account the angle setting that the

Quad-rotor will be travelling at. Equation 15 shows this modification.

𝑉𝑚𝑎𝑥=𝑉max 𝑆&𝐿𝐹 * Cos(𝜃𝑚𝑎𝑥)

𝑉𝑚𝑎𝑥=(𝑅𝑃𝑀𝑚𝑎𝑥 ∗ 𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑃𝑖𝑡𝑐ℎ ∗ 0.000954*0.44704) * Cos(𝜃𝑚𝑎𝑥)

Equation 15 - Quad Rotor Maximum Speed at Angle Setting Where

𝑉𝑚𝑎𝑥 (𝑚

𝑠) = 𝑀𝑎𝑥𝑖𝑚𝑢𝑚 𝑉𝑒𝑙𝑜𝑐𝑖𝑡𝑦, 𝑖𝑛 𝑡ℎ𝑖𝑠 𝑐𝑎𝑠𝑒 29.46𝑚/𝑠

𝑅𝑃𝑀𝑚𝑎𝑥 = 𝑀𝑎𝑥𝑖𝑚𝑢𝑚 𝑅𝑃𝑀 𝑤ℎ𝑖𝑐ℎ 𝑖𝑠 𝑢𝑛𝑘𝑛𝑜𝑤𝑛 𝑎𝑡 𝑡ℎ𝑖𝑠 𝑠𝑡𝑎𝑔𝑒

𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑃𝑖𝑡𝑐ℎ 𝑅𝑎𝑛𝑔𝑒 (𝑖𝑛) = 3.8 − 13 𝑖𝑛𝑐ℎ𝑒𝑠 𝑢𝑠𝑖𝑛𝑔 𝑎𝑝𝑝𝑒𝑛𝑑𝑖𝑥 𝐶

Equation 15 must be rearranged to calculate the maximum RPM required at different pitch to

achieve 20.46m/s

𝑅𝑃𝑀𝑚𝑎𝑥 =(

𝑉𝑚𝑎𝑥𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑃𝑖𝑡𝑐ℎ ∗ 0.000954 ∗ 0.44704

)

Cos(𝜃𝑚𝑎𝑥)

From table 2.0 in Appendix H.4 it can be seen that as propeller pitch increases, RPM required

reduces to obtain 29.46m/s. As calculated earlier the maximum RPM that can be obtained from the

brushless motors is 14,800RPM, this shows that any propeller that has a propeller pitch setting that

is under 6 inches can be eliminated. Table 1.2.1 in Appendix H.4 can now be modified to table

1.2.2 in Appendix H.4

UAS CHALLENGE 2015



9.4 Power Supply Voltage Selection From Appendix H.3 table 1.5 it can be seen that each brushless motor has an operating power

supply cell range generally the higher the cell count the more efficient the system will be and also

the higher the RPM will be. In this section different power supplies will be investigated to determine

the ideal power source that can be used in this project. From table 1.5 it can be seen that the

power supply cells range from 2s to 9s, but 2s cells will not provide the RPM required and it will not

be considered in this analysis, therefore 3s to 9s will be the main point of the research. The

investigation will give importance to cost, weight and coulomb rating, the full range of power

supplies analysis can be seen in table 1.7 in Appendix H.5

From table 1.8 in Appendix H.5it can be concluded that as the number of cells increases so does

the cost and the weight. Although a 3s cell is desirable because of its low cost and weight the

system will be inefficient due to Equation 16 (Anon, 2015).

𝐶𝑢𝑟𝑟𝑒𝑛𝑡 (𝐼) = 𝑃𝑜𝑤𝑒𝑟 (𝑊)

𝑉𝑜𝑙𝑡𝑎𝑔𝑒 (𝑉)

Equation 16 - Current Draw Detailed analysis can be obtained from table 2.1 in Appendix H.5which shows the RPM required to

sustain lift and RPM required to achieve forward velocity of 29.46m/s coupled with current draw

using different lithium ion cells. Using a propeller that requires 1000W a sample calculation can be

conducted, the results are shown in table 1.9, Appendix H.5. This table also represents power

consumption required by propeller dimension of 10*8 to achieve the RPM required for lift of 35N

which is obtained from table 1.2.2 in Appendix H.4 also the RPM to obtain the forward velocity of

29.46m/s that was obtained from table 2.0 in Appendix H.4. There are propellers that cannot

acquire the RPM required for forward velocity but has sufficient RPM to sustain lift a case of this

can be seen in table 2.2, Appendix H.5 this case the RPM to sustain lift of 35N has to be increased

to match the same RPM to achieve forward velocity. The result of increasing RPM means that

power consumption and current draw required is increased as it can be seen in table 2.3, Appendix

H.5. Table 2.1 in Appendix H.5 can now be updated to take into account the increase in RPM on

certain propellers, the new data in presented in table 3.9, Appendix H.5. From table 3.9 certain low

current propellers can be identified, these propellers are also seen in table 2.4, Appendix H.5

The propellers identified in table 2.4 are bought from a local hobby store and tested on the test rig

that was build. From testing it was identified that propeller dimensions of 12*6 is the most efficient

propeller, the justification for this is shown in Appendix H.10

UAS CHALLENGE 2015



9.5 Power Supply Capacity Selection The current draw, flight time and weight will be the deciding factor in choosing the battery. But

firstly the current draw is discussed, because RC wiring that are available for sale are from turnigy

where each cable has its own rating based on maximum current that is permissible the listing is

provided in table 2.5 Appendix H.6. From this it can be identified when all the motors are working

at its full capability that the current draw per motor cannot exceed 50Amps at any point as it could

cause the cables to burn up and this would result in catastrophic failure. Figure 1.4 identifies the

locations were 200Amp cannot be exceeded. Therefore by this method it can be identified from

table 2.4 in appendix H.5 that lithium-ion cells 3s and 4s cannot be used.

Figure 36 - Prototype Quad Rotor

Using the 50A current limit per motor battery capacity required can be calculated using flight time

of 5 minutes. Equation 17 can be used to calculate battery capacity required.

𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 (𝐴ℎ)

𝑇𝑜𝑡𝑎𝑙 𝐶𝑢𝑟𝑟𝑒𝑛𝑡 𝐷𝑟𝑎𝑤 (𝐴)∗ 60 = 𝑇𝑜𝑡𝑎𝑙 𝑓𝑙𝑖𝑔ℎ𝑡 𝑡𝑖𝑚𝑒 (𝑚𝑖𝑛𝑢𝑡𝑒𝑠)

Equation 17 - Battery Capacity Required

𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 (𝐴ℎ) = 𝑇𝑜𝑡𝑎𝑙 𝑓𝑙𝑖𝑔ℎ𝑡 𝑡𝑖𝑚𝑒 (𝑚𝑖𝑛𝑢𝑡𝑒𝑠) ∗ 𝑇𝑜𝑡𝑎𝑙 𝐶𝑢𝑟𝑟𝑒𝑛𝑡 𝐷𝑟𝑎𝑤 (𝐴)

60

𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 (𝐴ℎ) = 5𝑚𝑖𝑛𝑢𝑡𝑒𝑠 ∗ 200𝐴

60

Battery capacity required = 16.6Ah

Table 1.7 in appendix H.5 has been updated and documented as table 1.7.1 in appendix H.6 to

show addition information such as total cost and total weight that the battery capacity required is

known. Also as mention earlier 3s and 4s lithium-ion cells has been disregard due to high current

draw. From table 2.6 in appendix H.6 it can be seen that 5s lithium-ion power supply will be ideal

for this project were its lowest is weight and also the cost is one of the lowest.

9.6 RC Motor Selection Power When selecting the motor all the analysis that has been conducted up to now has to be

considered. Information to consider involves power consumption has to be greater than 811W,

RPM has to be greater than 11,270, must be capable of working with 5s, and also must able to

work with the propellers identified in table 2.4, appendix H.5 Other important data to be considered

will be weight and cost. The required criteria is applied to table 1.5 in Appendix H.3 which reduces

the number of RC motors that can be used for this project, these available motors are identified in

UAS CHALLENGE 2015



table 2.7, appendix H.7. Table 2.7 can be further reduced to table 2.8 in appendix H.7 when only

5s lithium-ion power supplies are considered.

From table 2.8 it can be seen that there are three closely matched motors that can be used for this

project, power 46, Quantum MT 3510 V2 and EMax GT2826-06. Power 46 although has some of

the specs that are required it cannot be considered for this project because the maximum RPM

while using 5s lithium-ion power supply is too lower to consider. Quantum MT 3510 V2 has very

attractive specs such as lowest cost out of the three and also the lowest weight but one of the main

issues any the reason for why it cannot be considered for this project is the fact that the power

consumption value if very low. EMax GT2826-06 is a motor that has most of the specs that are

required for this project, power consumption is perfect, 5s lithium-ion cell and propeller range from

10-14 can be used without a problem, cost and weight are ideal when compared to others that

weigh 200-290g. The only issue with this motor is that the maximum RPM cannot be used when

flying and the thrust setting will be based on 83% therefore the RPM will drop from 12987RPM to

11168RPM. Which in this case a shortage of RPM will occur (12,901RPM-11,168RPM =

1,733RPM) therefore recalculating based on 11,168RPM new maximum thrust and velocity

obtained. Results are presented in table 2.9, appendix H.7by using equation 1.3 maximum thrust is

calculated as 34N per motor As the maximum thrust has changed so does the maximum angle

using equation 1.7, 43°. Using equation 1.9 maximum velocity can be calculated 20.9 m/s. Power

consumption using equation 1.4 696W, with power consumption 15% extra has to be added 800W.

Using equation 1.6 current draw can be calculated 43A. Finally by using equation 1.7 flight time

can also be calculated 5.6 minutes which more than the expected 5 minutes.

9.7 Electronic Speed Controller Selection ESC (Electronic Speed Controller) is the next component to be selected for this project. Esc’s are

used to vary the RPM of the motors, as seen in earlier stages that vary RPM would mean that

thrust, velocity and pitch angle can all change just by varying the RPM. Esc’s are the only way in

which the Quad-rotor can be controlled autonomously as they will be connected up into pixhawk

directly which is the autopilot chosen for this project by other team members.

When it comes to selecting esc’s the general rule used by hobbies is to know the maximum current

draw that the motor can handle, in this case 52A then to add 15% to obtain the esc current

required. In this case its 59.8A, therefore esc’s that are rated at 60A will suffice for this project. As

with everything in this project there are other factors such as weight and cost that needs to be

taken into account, an analysis of this is shown in table 4.0, appendix H.8

From table 4.0 two esc’s are identified and presented in table 3.0, appendix H.8, one which is

lowest in weight and another that is lowest in cost. Both of the esc’s would have been ideal for this

project but as the motors were being bought from the same company that sells the robotbirds pro it

was decided that to save further cost on postage and packaging the 3g difference in weight will not

affect the project, therefore robotbirds pro 60A esc’s were ordered.

UAS CHALLENGE 2015

46 Unmanned Aircraft System - Subsystems


Section by Jonathan 10 Unmanned Aircraft System - Subsystems

10.1 Introduction

This chapter would discuss the different systems on board the UAS such as navigation and

communication systems as well as schematics showing detailed information on how the system

components are integrated with each and with other systems. A detailed specification sheet is also

provided in the chapter below. A list of systems aboard the UAS is shown below:

Navigation control system

Mission control system

Image recognition system

Flight control system

Communication system

Details on how to configure and operate all systems on-board the UAS through the autopilot

system are shown in Appendix. J.

10.2 Navigation Systems

The navigation system comprises of the following components:

Global Positioning System

Telemetry Kit

Radio Controller

Autopilot flight control system

Ground Control Station

Camera

On Screen Display

The function of the navigation system of the UAS is to provide the information need for the flight

controller to control the UAS to its mission destination. In this case, the mission is to deliver a

payload at a particular spot at pre-specified GPS coordinates. The GPS unit on board is used to

get the GPS lock on the co-ordinates, the on board compass gets the direction of the co-ordinates

and the gyro on board the flight controller determines motion on the relevant axis and then this

information is fed to the motors through the ESCs which regulate the voltage supply to the motors

to control the attitude of the UAS by either reducing or increasing the RPM of the motor. The GPS

coordinates are programmed into the navigation system with the use of waypoint files. The

navigation commands can be entered into the notepad and then loaded to the autopilot system as

shown in Figure 37: Waypoint Command File.

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Figure 37: Waypoint Command File The ground control station is used to input the commands GPS coordinates and payload release

mechanism in the form of a mission plan. This mission plan can be saved and edited if there is a

need to change the mission parameters for the UAS. The ground station is also used to monitor

the data generated by the sensors on board the UAS and it is transmitted back via the telemetry

kit. The ground control station consists of a laptop, telemetry transmission antenna and mission

planner software.

10.2.1 Potential Issues with the Navigation systems

An issue with the GPS unit is the HDOP (Horizontal Dilution of Precision) which

reduces the accuracy of the horizontal position of the UAS and this poses a problem for

mission deployment. The HDOP continuously varies depending on a number of factors

such as number of satellite count picked up by the GPS unit and weather conditions.

Another issue that can affect the performance of the navigation system is the

transmission rate and range of the telemetry kit as there may be a lag in the

transmission of data between the UAS and the ground control station.

Electromagnetic interference (EMI) from electrical components affecting the

performance of the compass on board the flight controller.

10.2.2 Solutions

To correct the flight condition for the HDOP accuracy of the UAS, the UAS flight control

can be switched from automatic flight control to manual flight control and the UAS can

be flown to the exact position where the payload is to be deployed.

The compass and the GPS unit that would be affected by EMI would be placed away

from components that generated magnetic fields.

10.3 Mission Control System

The mission control system comprises of the following components:

Autopilot flight controller

Payload Release mechanism

Ground control Station

Camera

The function of the mission control system is to deliver the payload at a particular position. The

payload is delivered when the autopilot control system determines the UAV is at the correct

position (correct altitude, correct GPS coordinates). A command inputted into the mission plan

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would then be sent to the servo and the payload release mechanism would then be activated and

the payload released.

The main issue with the mission control system is the accuracy of the navigation control system

and also the autopilot where the command to deploy is stored. If any problem is encountered, the

UAS can be flown manually and the payload can also be deployed manually with the use of a radio

controller.

10.4 Flight Control System

The flight control system consists of the following components:

Autopilot control systems

Electronic Speed Controller

Batteries (Avionics and propulsion)

Motors and propellers

The flight control system is used to control the UAS attitude and altitude. It comprises of the

propulsion system and the autopilot system working in conjunction from the data received from the

navigation system. To control the altitude or attitude of the UAS, a command is sent from the

ground control station to the autopilot. The autopilot then calculates the voltage output from the

battery that would be required to carry the command. The autopilot then regulates the voltage

supply from the battery to the motors with the use of ESCs. Yaw, pitch and roll are carried out due

to differential RPM of the motors on the Quad-rotor.

The flight control system also carries out the stability and control function for the UAV. The

autopilot system has an in built controller which has been reprogrammed to correct errors and

make adjustments in flight control. The controller is the PID (Proportional Integral and Derivative)

variant and this is done using the auto tune function when flying the UAV with the use of the radio

controller. The PID values have to be calculated before being inputted into the UAV before its initial

flight and the methods used to get the PID are:

Matlab Model to simulate flight conditions of the Quad-rotor

Selecting the right PID values for the different flight conditions.

When the simulation is run, the Matlab model is then put through a series of different

flight conditions and data is collected from these simulations.

The PID gain values are changed constantly in order get the control system to respond

the right way to disturbances in flight conditions, Table 8 shows the guiding principles for

choosing PID Values.

The simulations are monitored in forms of graphs and hence the values can be changed

when they are needed to be.

Testing the UAV system using a test rig

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Set the Quad-rotor inside the test rig and make sure everything is correctly bolted and

connected for safety and all the propellers are fixed and not within the reach of the test rig

arms

Input the PID values gotten from simulations using MATLAB

Test the Quad-rotor under multiple conditions

Use a high airflow fan to replicate strong gusts to see how well the Quad-rotor responds

to extreme flight conditions.

Test fly the Quad-rotor

Take the Quad-rotor to an open area for test flying

Start with simply manoeuvres before moving onto more extreme manoeuvres

The autopilot control system on board the Quad-rotor is capable of learning and during

the first flight test which would also be used for auto-tuning the control board, the Quad-

rotor would learn the appropriate response time and record it.

This method is used to program in the PID values for flight readiness.

Controller

Response

Rise time Overshoot Settling time S-S error

Kp Decreases increases No change decreases

Ki Decreases Increases Increases Eliminates

Kd No change Decreases decreases No change

Table 8 Effects on the close loop response from PID (University of Michigan, 1996)

To create the MATLAB model, the physics behind Quad-rotor behaviour is modelled such as the

torque and forces produced by the motors, the Quad-rotor’s inertial frame in relation to non-linear

dynamics. With the above information equations of motion can be generated by using a rotation

matrix to simulate the motion of the Quad-rotor. An appropriate controller can then be designed to

reduce any error produced by the Quad-rotor system. The model is not a 100 percent accurate

representation of the Quad-rotor due to different assumptions made in the course of modelling the

Quad-rotor. For this reason, a test rig will be used to improve the PID gain values as a simulation

on MATLAB will only take us so far without. The test rig will be used to fine-tune our close-to-final

PID values before we can actually test the Quad-rotor in actual flight.

An integral part of the flight control system is the autopilot system. The autopilot system comprises

of three layers of ware:

Firmware

Software

Hardware

To fully utilise the capability of the autopilot system, the firmware and software aspects are edited

to make the application flexible in terms of navigation and mission control. The autopilot system

used is Pixhawk which is built on the open source px4 platform. The autopilot system is capable of

carrying out functions such as autonomous flight, computer vision operations and robotic functions.

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The autopilot system has enough processing power to carry out the above mentioned functions at

the same time.

The autopilot systems also has on board sensors which generate and provide information about

different systems on board the UAV and also data about flight performance, this information

(Figure 37) is transmitted to the ground station for observation and control with a telemetry kit

operating at 433Hz. To improve flight conditions of future flights, telemetry data is logged by the

autopilot system and the data gathered can be analysed to make adjustments to any system to

raise the performance of the UAS.

Figure 38: Telemetry Information transmitted to ground control station

10.5 Communication System

The communication system for the UAS consists of:

Radio Controller

Telemetry Kit

Minim OSD

Autopilot System

The communication system is used to transmit telemetry data from all components on the UAS to

the ground station for observation and control. There are three methods of connecting the UAS to

the ground control station:

Serial Connection

Telemetry Kit Connection

Radio Connection

The different connection methods have different transmission rate and therefore different functions.

The UAV and the ground control station communicate using a protocol called MAVLINK. This

communication protocol is the main protocol for the Pixhawk unit and this determines the

transmission rate for different types of transmission methods and format of data transmitted.

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10.5.1 Serial Connection

The serial connection is used to connect the Pixhawk autopilot to a ground control station through

a Universal Serial Bus connection. The baud rate for the transmission is 115200 bits per second

and this connection is used to configure the autopilot system for the first time. The extremely fast

connection is used to load the firmware and software needed to run the autopilot system and also

to calibrate all on board sensors for the first time. Other components of the UAV can be connected

and also configured through the serial connection. The serial connection is also useful when

running diagnostics on the autopilot or any connected component as the transmission rate and

quality would prevent loss of data or useful information through data packet loss in transmission.

The transmission rate can be monitored by the link statistics as shown in Figure 39.

Figure 39: Transmission Link Statistics (Serial Connection)

10.5.2 Telemetry Kit Connection

The telemetry kit is used to connect the Pixhawk autopilot to a ground control station through a

radio connection over a frequency of 433Hz. The baud rate for the transmission is 57600 bits per

second. This is the primary method of connecting to the autopilot for flight purposes and any other

secondary purpose of the UAV. The connection can also be used to configure the autopilot system

to calibrate on board sensors but due to the connection speed, it is advisable to use the serial

connection for that. For autonomous flight, the flight plan is uploaded to the autopilot through this

connection and with the use of a ground control station. During flight, any secondary mission plans

for the UAV are also sent through the telemetry kit connection; this can range from servo activation

to camera functions. The strength in telemetry connection would decrease as the UAV moves

further away from the ground control station. During flight, all the telemetry generated from all

components is sent to the ground control station through the telemetry kit. The transmission rate

can be monitored by the link statistics as shown in Figure 40.

Figure 40: Transmission Link Statistics (Telemetry Kit)

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10.5.3 Radio Connection

The radio controller is used to connect to the Pixhawk autopilot and the UAV through a frequency

of 2.4 GHz. The radio controller is used to fly the UAV manually without the need for a ground

control station or GPS based command input to the autopilot system. The radio controller is also

used to configure some stability and control criteria such as PID through a method known as auto-

tune. The radio controller has a number of channels that are used to carry a number of secondary

UAV functions such as servo control, camera control etc. The radio controller also acts as a

backup flight controller when the autonomous flight system fails or acts as a safety flight measure

when the UAV flies out of range of telemetry range of the ground control station.

10.6 Systems Integration

To make sure that all the systems to be used on the UAV can work together and can also

accomplish the primary and secondary objectives of the UAS and that the components to be used

are also compatible, a series of tests are carried out on the each system and its respective

components. Some examples of the tests are:

Communication systems test

Servo test

Propulsion system test

Image recognition system test

Post Assembly Design Checks

Post Assembly Systems calibration

Some of the tests listed above are discussed in different chapters such as the propulsion test in the

chapter dealing with propulsion and performance and the servo test in the chapter dealing with

UAS mission delivery. Every other test is explained below:

10.6.1 Communications Systems Test

The tests carried out on the communication systems are of the following types:

Interference tests

Range tests

Altitude Tests

10.6.2 Interference test

To carry out the interference test, the UAS communication systems are operated near areas or

devices of high magnetic interference, near devices that give off radio waves such as Wi-Fi

devices and TV antennas. The UAS communication systems also tested indoors and outdoors but

in close proximity to a building. The result of these tests is shown below:

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Test System Component Result

Magnetic

Interference

Telemetry kit

Radio Controller

The rate of transmission is reduced and also

the number of bits (information transmitted)

lost is increased.

Near

Buildings

Telemetry kit

Radio Controller

Operation in a building has little or no effect

on the radio controller. The telemetry loses

range and quality of transmission especially

when there is a wall between the transmitter

and receiver.

Radio Waves Telemetry kit

Radio Controller

There are severe consequences due to the

difference in the transmitting frequencies

telemetry kit and the radio controller.

Table 9 - UAS Interference Tests

The UAS is designed for open field flight and as such the tests carried out above do not affect the

objectives of the UAS mission, the reason for the test is for future use of similar UAVs used for

different purposes as stated in the business case. These tests were done to show the durability of

the UAS control systems and its adaptability to different operating environments.

10.6.3 Range Test and Altitude Test

The telemetry kit to be used on the UAS is designed to be used at ranges of about 1.5 kilometres;

the farthest point on the UAS is approximately 500 metres from the ground control station. The

range of the UAS telemetry was tested in an open field as well as during the interference tests. The

largest open field used for the test was 600m at its farthest point and the UAS remained in contact

with the ground control station during the test. The antennas for the telemetry kit are Omni-

directional and thereby transmit data in all directions and also upwards.

The altitude test for the UAS was carried out by taking the UAS receiver to the fourth floor of a

multi-storey building of approximately 60 feet. The communication system worked well even with

interference with the Wi-Fi in the building. The radio controller was also tested for both range and

altitude and the tests results show that the radio controller is capable at operating distances of the

UAS mission.

10.6.4 Post Manufacture and Assembly Design Checks

The post assembly design checks were carried out after the UAV had been built, assembled and

the electronic components are connected and ready for testing. The post assembly design checks

include the following:

Inspect structure of UAV to make sure that there is adequate space and protection for

electronic components.

Inspect assembly to make sure components are assembled neatly and safely.

Inspect assembly to make sure electronic components are connected to their proper ports

or power sources.

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Inspect health of all propulsion system components (motors, propellers, Escs, batteries).

Inspect wiring and make sure that wiring on the assembly match the wiring diagrams.

Inspect Assembly to make sure that the design specifications were met by comparing the

UAV to the design specification sheet.

Inspect assembly to make sure that all safety precautions were taken into consideration

during the assembly and manufacture of the UAV

10.6.5 Post Assembly Control System Calibration

The electronic components were configured when they were bought in order to carry out various

tests but after assembly the memories of the autopilot system and all other components are

deleted. The main reason for reconfiguring the control system equipment is that sensor error as a

result of being calibrated before the component is assembled on the frame. When the assembly is

done and all the components and their sensors recalibrated, such error is reduced. The UAS

recalibration was done with the use of the 3 axis test rig and the following sensors were calibrated.

Accelerometer

Compass

Radio Controller

Joystick

Gyroscope

Fail-safe systems

Arming Checks

After all the calibration was done and all other system integration checks carried out, the UAV was

then set-up to tune its PID values for flight.

UAS CHALLENGE 2015

55 Stability and Control I MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Malwenna 11 Stability and Control I

Quad-rotors can be regarded naturally stable compared to fixed wing aircraft by the nature of their

design. That is mainly due to the thrust being generated by all four corners where the resultant will

act on the meeting point of three main axes. However, natural stability is only achieved if the CG of

the quad is designed to be on above mentioned intersecting point, so that the vertical forces on the

quad will originate from the same point with no moments about the CG when it's stable. Even if the

CG is not at the intersecting point, quad can be stabilized by simply changing the RPM of the

motors so the moments will be balanced. Therefore, the first step of making a Quad-rotor stable is

the placement of CG. Stability about yaw is achieved by having counter rotating propellers to zero

the resultant torque created by rotating propellers.

Controllability on the other hand did require more attention. There are four rotating parts indicating

more control is needed. Only control input will be the thrust change by changing the RPM of the

motors. But the problem lies within the accuracy of the input due to various factors such as human

error, mechanical error and disturbances by outside forces. This is where the control board

(Pixhawk) takes over to minimize the errors and aid the copters controllability in achieving the

desired output. This is done by a system of three independent Proportional Integral Derivatives

also referred to as PID controllers. As shown in Figure 41, it is a closed loop system where the

error is corrected by subtracting the output from input to identify the error and running the error

through three PID gains. This is a very quick process which will be repeated until the error is

corrected.

Figure 41 – PID System (Oscar, 2013)

P (Proportional gain coefficient) –This controls the sensitivity of the quad to the angular

change being input and therefor, most important controller.

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I (Integral gain coefficient) – controls the precision of the angular input, especially when

outside disturbances are present such as gust. This controller will identify the disturbance

and minimize the effect caused by it.

D (Derivative Gain coefficient) – By having an input on the quad, there can be accelerations

towards the desired output and this gain will dampen if they are unwanted or amplified if

they helps in achieving the output. Therefore, it helps in predicting errors and mitigates

them (Hove, 2013).

Although PID controllers seem simple, the mathematics behind these is complex to grasp. PID

gains will depend on the weight, size and purpose of the Quad-rotor. Therefore, the main

responsibility of Stability and Control role is to obtain correct PID values for particular quad using

mathematical models, MATLAB simulation or PID tuning. Later is regarded as the most reliable

method.

As stability was a joint role between Mohammed and Malwenna, the work was split between these

two and so was the report. Please refer to 14 “Stability and Control II” for CG placement and

MATLAB model.

11.1 PID Tuning Refer to section 15 “ Flight modes and tuning” for information on test rig.

11.1.1 Loiter mode

The main purpose of tuning for loiter mode is for Pixhawk to automatically keep the current

heading and altitude, especially at payload deployment until character recognition identifies the

target. During loiter tuning, the pilot would fly the quad manually as in stabilize mode, but releasing

the stick would keep the Quad-rotor in the same position. However, in order to achieve good loiter

characteristics, there are three main requirements to be fulfilled

GPS positioning – GPS is normally positioned elevated from the Pixhawk and other electrical

components. This is to lower the magnetic interference caused by

other components so that GPS positioning hold will be accurate.

Ideal position for the GPS will be decided when the GPS protective

case and the mast have arrived.

Magnetic interference on the compass – Original GPS unit

decided for the quad was ‘3DR uBlox GPS’ which also includes the

‘LEA-6H compass’. Since it will be mounted high, magnetic

interference will be minimized. However, given the availability,

timeliness and budget restraints, GPS Crius CN-06 v2 was

purchased which does not have a compass and Pixhawk inbuilt

compass will be used instead which will have magnetic

interferences. Figure 42 Loiter PID values

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Vibration –Analysis has been done by the structural team to minimize the vibrations and

therefore, lower vibrations will help in loiter.

The Loiter PID P value at the top of the Figure 42 refers to the conversion of difference between

desired and actual position as a speed towards the targeted position. Rate Loiter PID values will

then convert the desired speed to desired acceleration towards the targeted position and desired

acceleration would result in quad obtaining a lean angle to correct the position. These values do

not require changing as advised in the Ardupilot tuning guide (Copter.ardupilot.com, 2015), but will

be changed just to observe in later testing.

Loiter speed refers to the maximum horizontal speed achieved by the quad in loiter mode and is in

the units of cm/s. Therefore 500 refers to 5m/s. Max acceleration at loiter mode is limited to half the

loiter speed by Mission Planner.

11.1.2 Altitude Hold Mode (AltHold)

Engaging in this mode will enable Pixhawk to take control of the throttle and automatically maintain

the altitude present at engaging moment. The pilot will still be keying pitch, roll and yaw to stabilize

the quad. This mode will be useful when hovering to deploy the payload. Correct “AltHols” tuning

was not possible to obtain so far in the current test rig since the altitude is fixed. Pixhawk uses the

inbuilt barometer to measure the pressure difference in order to correct the altitude. Therefore, it’s

important to take the Quad out from the test rig and test AltHold in a secured and open area,

according to rules of regulatory bodies and also not on whether sensitive days which can cause

pressure readings to fluctuate. Therefore, this will be conducted in later test stages when Stabalize

mode is properly tuned. When AltHold is engaged, the throttle would be automatically set between

40% -60%. The pilot can take control of the throttle anytime and throttle input over 60% will cause

to ascend and below 40% will cause to descend. However, if the landing is performed in AltHold

mode, it would take a few more seconds than normally to disarm the motors after a touchdown.

Maximum climb and descent rates are set to a lower value of 2.5 m/s during testing since it

requires practice and experience to control the quad manually without causing any damage.

Purpose of Altitude Hold P is to convert altitude error into a climb/descent rate. The higher rate is

suitable to correct altitude aggressively, but too high can cause oscillations. Throttle rate PD

converts earlier rates to accelerations. Throttle acceleration will feed the acceleration error back

Figure 43 AltHold mode PID values

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into system to further reduce the altitude error. It is notable that in this setting, D value is kept at

zero. D gain dampens the unwanted acceleration toward desired variable and in this particular

case, acceleration is required. Therefore, it will be kept at zero. Further P to I will have a 1:2 ratio

(3DRobotics, Altitude Hold Mode, 2015), which will be maintained during testing. Built quad is more

powerful than a normal therefore, reducing PI values by 50% will be a good starting point to initiate

testing. Hence, better performance is expected at P = 0.5000 and I = 1.0000. See sections 15.2.1

Pitch and Roll tuning, 15.2.2 Yaw tuning and 15.2.3 Waypoint navigation tuning for other flight

modes.

11.2 Verifying the performance of PID values

Primarily this would be done by observation and there are two stability engineers to confirm the

result. Since there is still a human element involved, preferable method would be to use flight

record.

In Mission Planner, there are two ways to record the flight data. Through Dataflash logs which

uses on-board flash memory and can be downloaded using MAVLink and through telemetry logs

which is recorded in the mission planners since we are using the 3DR Radio telemetry. Dataflash

logs will be used to verify the performance of the PIDs by opening log in the mission planner which

will open value graph as shown in Figure 44 depending on the flight mode.

Figure 44 Dataflash log in Stabalized mode opened in Mission planner

(3DRobotics, Verifying performance with dataflash logs, 2015)

Figure 44 is the graph from the stabilizer mode where major concern is to achieve good roll and

pitch. So in order to evaluate performance, we need a comparison between desired role and actual

roll, desired pitch and obtained Pitch. It is clear that the units of X and Y axes are the same and the

shape of the two lines are similar and track well. It is obvious that if PIDs are not good, the shapes

of the graphs would not be similar. Therefore, achieved PID values are good. A similar process will

be used in other flight modes as well. For an example, in Alt Hold mode, Barometric altitude (Baro

ALT), Waypoint altitude (desired altitude) and the GPS altitude (inertial nav at estimate) will be

compared see if they track well.

UAS CHALLENGE 2015

59 Safety Case MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Malwenna 12 Safety Case

12.1 Overview The UAS can possibly cause property and individual damage to its Pilots, spectators and parts of

the overall population and surroundings. The harm may be brought on by the UAS's contact with

the ground or due to equipment falling out. Therefore, UAS is only allowed to fly in UK airspace if

they are considered safe in operation. UAS in this particular competition being less than 7Kg

MTOM, they will fall under SUA (Small Unmanned Aircraft) category and should comply with UK

Air Navigation Order 2009 articles 138, 166, 167 and CAA CAP 722, and CAP 393. (UK CAA

Safety and Airspace Regulation Group, 2014) (Civil Aviation Authority, 2012)

The main requirements extracted from those articles are as bellow;

The UAS should not operate above 400 feet (122 m)

The UAS should always be in Visual Line of Sight (VLOS) since collision avoidance is

primarily based on this

Maintain a "pilot in control", which is to take control and fly the UAS in case of failure of

autonomy

Operate 150m away from congested areas

Should not operate within 50m of person, vehicle or structure except 30m at takeoff and

landing

Apart from this, it is made sure that team is referring to the University UAS Challenge 2015

competition rule book while designing, manufacturing, testing and demonstration of UAS.

12.2 Flight Controller Safety Mechanism The Pixhawk’s flight controller we have chosen has a number of safety mechanisms; It includes a

motor arming safety feature when manually controlling the copter. At take-off, throttle stick should

be held up for several seconds to safely arm the motors and vice-versa at landing. It also includes

safety modes such as RTL (Return to Launch), Failsafe and GeoFence. In the event of a signal

lose to the UAS, it can be programmed to return to launch location using RTL while Failsafe will

ensure its safety and GeoFence will transmit its current location. Stabilize or Stabilize plus modes

can be triggered to land the Quad-rotor safely in case of a motor failure.

Please see section 16 “Flight Termination Case” for more information.

12.2.1 Safety Measurements for Flight Testing

Ensuring no personnel are near propellers when they are powered, especially when

performing PID tests.

Terminating the flight before battery’s safety capacity is reached.

After landing, ensure battery power to the components has been stopped either by

removing cables or using a switch before handling the UAV.

Prior testing, ensuring the home location shown in the mission planner software is correct.

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Using a staggered flight test approach, increasing speed and height with each test.

Use of checklists for mechanical and electrical components, systems and assembly before

every flight test to ensure they are connected correctly and working.

12.3 Hazardous Components 1) High speed propellers – detachment of propellers in flight can cause serious injuries to

people and animals. Therefor it is suited to avoid composite made props and use breakable

and flexible props. The downside to this is it will reduce the performance of the propeller.

However, given the reliability and safety, plastic props were ultimately chosen which will

break in an event of a crash without serious damage to personnel or structures

2) Batteries – lithium polymer batteries are often seen exploding due to misuse, which can

cause serious structural damage to the aircraft. Use of high build quality batteries and

monitoring their charging and temperature regularly can avoid such failure

(Rogershobbycenter.com, n.d.). Purchased batteries will be made brighter in colour to

identify them in a crash and they are mounted using Velcro Straps for easy removal.

12.4 Battery Fail Safe Battery Fail Safe mode in

Mission Planner is to land

the Quad-rotor or return to

launch if battery voltages

drop down a certain

percentage. But the

requirement for this fail safe

to activate is that the battery

should be connected to

Pixhawk power module.

Three 5s batteries are being

used to power the motors and ESCs which are not connected to Pixhawk since Pixhawk can only

support up to 4s batteries. Therefore, this fail safe is used only for the system battery which is a

2200MAH 4s Lipo.

Fail safe will trigger at two occasions

1) If the 4s battery voltages goosed below 12volts – Minimum safe voltage for a 4s battery to

operate is assumed to be 3 volts per cell.

2) If the 4s battery remaining capacity goes below 440MAH – This is 20% of the capacity of

the battery which is 2200MAH and 440MAH is being set as the configurable Reserved MAH

(reserved for land or RTL)

Figure 45 Battery fail safe settings chosen in Mission Planner

UAS CHALLENGE 2015


Battery Fail safe can be

disabled anytime,

however, there is a

separate low battery

warning option being set

up. Low battery massege

will appear on the

ground control station

following a loud beep if

the battery percentage

goes below 23%. Having 23% will give sufficient time for ground control to prepare for the fail safe

method selected since failsafe will initiate at 20%

12.5 Radio Fail Safe

This fail safe is used if communication between RC transmitter and receiver is broken. Given the

nature of the mission to be fully autonomous, it is highly unlikely that this will be in use. The RC will

only be used if autonomy fails and if decided to continue with the mission without triggering other

Fail safes. However, this can be very useful in testing stages especially in AltHold and auto tuning

if quad becomes unresponsive or uncontrollable.

There are four occasions where Radio Fail Safe is possible

If the RC transmitter is switched off accidentally

If quad exceeds the maximum RC range

Malfunction in RC receiver wiring or PPM encoder

If the RC transmitter runs out of power (Turnigy 9x RC is powered by eight AA batteries)

When the Fail safe is triggered, motors will automatically disarm if Quad-rotor is on the ground. It

will return to launch (RTL) if it has a GPS fix and more than 2m away from launch location or if has

no GPS fix and within 2m, it will land. Even if the communication link is restored prior to landing, it

will still continue with the fail safe unless the flight mode is changed using Mission Planner or RC

switches. It will continue with the mission if the mode is on Auto.

Figure 46 Battery monitor settings chosen in Mission Planner

UAS CHALLENGE 2015

62 Environmental Impact MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Malwenna 13 Environmental Impact In order to UAS design to be a success, it is important to assess the environmental impact it has

from the initial concept of the design. One of the aims in the design process was to have minimum

impact on the environment without compromising the performance. This report covers the main

Environmental factors affected by the use of UAS and proposes action taken and will be taken to

lessen them.

13.1 Hazardous Material 1) High speed propellers

2) Batteries

3) Plastic material – most of the Quad-rotor structure will be built using plastics such as Nylon

6.6, Nylon 6 and PVC. Reason for choosing these were presented in the section. The

environmental impact of using these materials is high and there are three conceivable

ecological issues to be considered. Plastics are generally produced using natural resources

which must be conserved, such as oil, gas or coal and increasing the use will drain the

natural resources. As a by-product of the manufacture of plastics, various pollutants will be

created which have to be dealt with properly by manufacturing companies.

13.2 Air Quality

13.2.1 Emissions

Air pollution due to UAV usage is primarily from gas emissions during flight. Therefore, Reduction

in emissions was considered in the initial planning of the power plant. The end result was to

discard the use of any fuel and use battery powered motors which will not only minimize air

pollution, but eliminate it. Therefore, this Quad-rotor design will have zero air pollution due to

emissions.

13.2.2 Noise

There are two types of noise originating from a UAS. Aerodynamic noise is the noise due to

vortices at the blade tips. Higher blade loading and speed will result in a higher noise. But the most

significant noise is the noise from the power plant. Specially noise from a fuel engine airplane

where noise arouses due to combustion and exhaust compared with a similar set up electrical

engine, is higher. However, in this case we are using four motors powered by 5s batteries and

provides a lift to carry 2kg payload. Therefore, power requirement and work done is higher, so is

the noise than in a normal Quad-rotor. However the noise is being minimized by proper weight

distribution and propeller balancing to reduce the vibrations causing the noise. In this particular

competition, high noise can be advantageous as well since UAS has to remain at VLOS always

and noise will aid in locating the vehicle. However UAV would be under the permitted noise level of

Elvington Airfield area and would not significantly affect the air quality.

UAS CHALLENGE 2015


13.3 Infrastructure Quad-rotor is a small UAS system and therefore, does not require major infrastructure changes.

But upon impact, it can cause serious damage to infrastructures. Also, uncontrolled radio

frequencies can cause interferences for civil operations. To prevent such incidents, the quad will

be operated 50m away from structure, personnel and 150m away from congested areas as

required by the CAA regulations.

13.4 Disposal of Material Plastic material - The most common method of disposing these materials is by burying in landfill

sites, but since they have a low decay rate, increasing use of plastics will create a build-up in

landfills. The materials that have been used in the design are high quality materials manufactured

using the correct method which means they are not degraded. Therefore, burying them in landfill

sites will not produce harmful gasses such as methane, which are normally produced by low

quality materials. Alternative way of disposal is to incinerate it. Burning plastic can reverse the

process to obtain raw materials such as crude oil, gasses and coal. These gassed can also be

recycled separately after. However, this process will also generate some harmful gasses.

Incinerating Nylon will produce carbon monoxide, ammonia, aliphatic amines, ketones, nitriles and

hydrogen cyanide and later in exceeding room temperature is a highly poisonous gas

(schoolworkhelper.net, 2014)Therefore, this process should be carried out in a controlled

environment. Another way of recycling them is by reprocessing, which will produce materials which

are inferior from previous quality, but can be used for products such as bags and dust bin bags

where quality is not that important. At the end of the lifetime, Quad-rotor can be disassembled and

plastic material can be taken to the numerous plastic recycling companies available in the UK for

them to be properly reprocessed.

Lithium Polymer batteries – Since Lipos are the most hazardous equipment used, they have to be

disposed in a responsible manner. Earlier method of disposal was dumping the Lipos in a salt-

water bucket and letting it to degrade and disposal though drainage system. This started causing

problems to Water Authorities since they use lithium to trace the water leaks and degrading Lipo in

salt-water will add lithium salts to the drainage system (Smith, n.d.)Therefore the best ways to

dispose Lipos are the procedures via local authority Environment Waste Department where they

have a special used battery collection system on going. But prior to handing over the Lipos,

following steps should be carried out.

Lipos should be discharged to a minimum voltage – Suitable resistance should be used to

avoid overheating of both the battery and resistor and battery should be drained as closes

to zero volts.

Discharging the battery using shorting the leads should not be attempted.

UAS CHALLENGE 2015


Once discharging is completed, they should be secured in a stout cardboard box or similar

and clearly labelled with "SPENT LITHIUM BATTERIES FOR RECYCLING" (Smith, n.d.).

Also batteries can be returned to the battery retailers whom are obligated to accept spent batteries

under the National Battery Back Scheme.

Table 10 Impact of Quad-rotor on environment

Environmental

Factor

No Impact No Significant Impact Significant

Impact

Hazardous Materials X

Emissions X

Noise X

Infrastructure X

Waste X

UAS CHALLENGE 2015

65 Stability and Control II MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Reyad 14 Stability and Control II Due to the complexity of the Quad-rotor stability, joint roles were required in order to carry out the

complex tasks. The stability testing sections are split, along with the reporting.

Refer to Section 11 “Stability and Control” and 11.1 “PID Tuning” for more information on the

introduction of Quad-rotor stability.

14.1 Ideal CG location The Centre of Gravity (CG) placement on a Quad-rotor needs to be taken into consideration early

on as it affects the flight performance, speed and stability. Ideally, the CG should be at the centre

point of the multi-copter, at 0 on the x and y axis. As this may not be feasible depending on the

size and weight of the systems and batteries, the CG may be off centre by up to 1-2cm. If the CG

is at the aft of the Quad-rotor then it will naturally try to pitch back, increasing the time it takes for

the quad to pitch forward, therefore sacrificing forward speed and forward acceleration, although it

will be very effective at reducing forward speed. The same principle also applies if the CG

placement is closer to the front of the Quad-rotor. However, unlike an aft CG, a fore CG placement

can be beneficial as unless the Quad-rotor is hovering, it will be always at forward flight during the

main event of the IMechE competition, which will allow the Quad-rotor to accelerate faster, improve

forward speed and reduce the thrust required from the motors to remain at the desired pitch which

will help increase the flight time. If, however, the CG is too far in front the motors will require

additional thrust from stopping the Quad-rotor from pitching too much. These same rules apply to

roll, if the CG is on the left or right side, it would allow the Quad-rotor to roll to that specific side

faster but react slower if it was to roll to the other side. On top of this, the Quad-rotor will need to

provide additional thrust so that it does not roll to one side continuously.

As previously mentioned, the Quad-rotor might be able to get away if the CG placement is no

greater than 1-2cm away from the centre, any more and the additional thrust required to

compensate for the stability will reduce the flight performance by a large amount. Unless major

changes are made to the Quad-rotor, from here on, which is highly unlikely, for each cm the CG is

off by; the motors will need to provide an additional 155g of thrust to compensate.

Figure 47 Side view of the Quad-rotor

UAS CHALLENGE 2015

66 Flight modes and tuning MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Reyad

The z-axis is far less black and white, compared to the x and y-axis. The lower the CG on the z-

axis the more stable the Quad-rotor becomes, on the other hand, the more stable the Quad-rotor

becomes, the more thrust will be required to manoeuvre the Quad-rotor. Assuming that the top

plate is the datum (see Figure 47), the CG is at 0.9 cm with one payload and two batteries and -1.7

cm after payload has been deployed. As these points are below the propeller, they should allow for

some level of stability with very little compromise to the manoeuvrability.

15 Flight modes and tuning

15.1 Simulink model Initially there was a plan to create a mathematical model on Simulink to simulate the behaviour of

the Quad-rotor while it’s flying, or at the least; hovering. The reason for this is that it allows for safe

testing without any crashes since the test can simply be restarted. However, later on it was

decided that this may not be needed as a simply test rig will allow for tuning of the Quad-rotor

without the requirement of a mathematical model and certain changes to the Quad-rotor will allow

the group to test the Quad-rotor in relative safety. Nevertheless, a ready made Simulink model

(Figure 48), (full model in Appendix. J) was found which was previously made by the winning

students at Drexel University for the Mathworks ‘MATLAB and Simulink Student Design Challenge

2014’ (Mathworks, 2014). While the model the group has created has the possibility of not being

perfectly accurate, they have verified it with Mathworks using a real Quad-rotor, making this small

model somewhat credible. Still, the purpose of the model was to check the changes in PID values

and their effect on stability, before the Quad-rotor and test rig was built, and get their ratio of PID

numbers for testing on the test rig once it has been made.

Figure 48 Simulink model used

To get started on MATLAB, an m-file of all size and weight of the main components will be required

as well as the thrust and torque coefficients of the motor. These will be used to create a transfer

UAS CHALLENGE 2015


function for the PID values (full model in Appendix. J). Another m-file will be required for the initial

starting conditions.

Once the Quad-rotor parameter and initial conditions file have been created, the files will be loaded

onto the altitude control file on Simulink. At first the PID values will all be set to zero before being

adjusted one at a time, first for the throttle command and then followed by the Roll and Pitch on the

‘Position Control’ Simulink file. The test was done in the same fashion as one would do for the

physical test, by pushing the P value (in throttle command) until the throttle response is deemed

acceptable. Without an I or D value this would cause the Quad-rotor to oscillate with little to no

damping, as seen in Figure 49. To introduce damping, the D gain must be increased, and as

Figure 49 once again indicates, after multiple iterations, the D gain allows the Quad-rotor to

stabilise at 100ft, the minimum height for the UAS Challenge, at a reasonable amount of time.

Figure 49 Quad-rotor oscillating with only the P gain (left), with P and D gain (right)

Pitch and Roll should be theoretically be the same as a Quad-rotor in a ‘+’ or ‘x’ shape, so should

ideally be symmetrical and therefore the PID values for one mode should be very same for the

other. However, at the time of Simulink testing the Quad-rotor had not been manufactured for

validation but for the test it was assumed that the CG was at the centre for the x, y and z-axis and

all moments were the same for all the arms. For the roll and pitch the P value were increased until

there were oscillations before the D and I values were increased one at a time for the Quad-rotor to

fly with a good response. Full test data can be found in full model in full model in Appendix. J and

also for Throttle/altitude and for pitch and roll .

When it was time to do the yaw tuning, it appeared that regardless of what value the PID were,

even 0, the Quad-rotor would still yaw on the Simulink model and therefore no further attempt was

made on yaw. Fortunately, yaw is less of a concern on a Quad-rotor as it can be manually

controlled without any issues to the Quad-rotor’s flight path.

After extensive testing on Simulink, the final values for Throttle, Pitch and Roll are shown on the

figure below, Figure 50.

UAS CHALLENGE 2015


Figure 50 PID values on Simulink

15.2 Test rig PID Testing To test all the PID values from the test rig, the test rig itself must be able to move freely in all axes.

However, to achieve the values for only the pitch, roll and yaw, each axis must be tied down, so

not to interfere. Having said this, the placement of the test rig may be of an issue, as the Quad-

rotor flies through the air, it should pivot around its CG, but while it is on the test rig, the CG will be

slightly above the pivot point which means that PID values we get from the test rig would be great

while the Quad-rotor is on the test rig, their effectivity will be greatly reduced once the Quad-rotor is

removed from the test rig. What this means is that once the Quad-rotor is removed from the test

rig, the must be further fine-tuned to make it suitable for the UAS challenge.

The test rig was not made in time for comprehensive testing, a plan was created which involved

the PID values on Pixhawk being slowing increased, same as with the MATLAB model, one at a

time until satisfactory results. While Pixhawk comes with their own values for PID, they’re designed

for 3DR’s own Quad-rotor (3DR Robotics, 2015). For this reason, the team’s Quad-rotor requires

its own set, calibrated through Mission Planner. The PID numbers will need to be adjusted for the

typical Roll/Pitch (yellow), as seen below, yaw (orange)(Figure 51), altitude hold (green), loiter

(pink), and waypoint navigation (blue).

UAS CHALLENGE 2015


Figure 51 Values that require change (3DR Robotics, 2015)

Stabilise mode, one of the initial recommended modes for tuning the Quad-rotor, as it handles

some level of control over the Quad-rotor over the pilot such as maximum roll and pitch. In ‘acro’

mode, the pilot has full control and could therefore push the Quad-rotor to overturn itself and if the

pilot was tuning the Quad-rotor without a test rig then the pilot could cause some serious damages

to the Quad-rotor and if the propellers snap off, damage to anyone nearby like the pilot. In stabilise

modem, the Quad-rotor will automatically try to stabilise itself once the pilot releases the stick,

making it the ideal starting point for tuning. As progress was made, more and more control would

be taken back from Pixhawk before testing it under acro mode.

As it’s a Quad-rotor, the pitch and roll values can remain the same since it can also be flown

sideways in the same way for forward flight. However, this will be tested later on if they require a

different set of PID values. As Pixhawk does not allow for a P value below 0.08, that will be what

the test will start with. I and D will be set to zero to minimise their effect and will be incorporated

once the P value is satisfactory.

15.2.1 Pitch and Roll tuning

To tune the pitch and roll PID values, the process follows a similar process as the one for

MATLAB:

1. Set the Quad-rotor up on the test rig, as seen on Figure 52, and make sure all arms are

fastened, all loose cables are tied and the batteries are well charged

2. Make sure that the batteries are well placed and not causing the Quad-rotor to tilt to one

side from the misaligned CG

3. Test the Quad-rotor with the given PID values for 3DR’s own Quad-rotor (3DR Robotics,

2015) to see how it handles

4. Test with the values from MATLAB to see if the Quad-rotor response time are the same

5. Test with minimal PID values to see how stable it may be naturally

6. Start increasing the P value to improve the response time to that close to which ever it was

most stable to (IRIS+ or MATLAB version)

7. Increase the P, I, and D values until the Quad-rotor stabilises whilst in the test rig

UAS CHALLENGE 2015


Figure 52 Quad-rotor on the test rig

For a more detailed version of the test plan, please see full model in Appendix. J.

The test rig was finally completed approximately one week before the submission date which gave

the group two days to do some quick testing. During these two testing days the Quad-rotor

managed to pitch and return to level within a very good time frame, approximately 2.3 seconds,

however, due to the placement of the batteries, the roll took much longer than expected to stabilise

(just over 6 seconds). This is due to the Quad-rotor not being fully assembled as it was designed to

be but put together for testing purposes and the batteries were not closely placed at the centre to

minimise the moments and the placement of the CG. After some realignment, the pitch, roll and

yaw results were much improved and better than expected in some cases. However, Pixhawk has

its own method of providing data on how the Quad-rotor should stabilise, see Figure 53, and

unfortunately there wasn’t enough time to improve the result nor could the data be extracted (at the

time) for further analysis.

Figure 53 Results of what Pixhawk should output (Copter.Ardupilot, 2015)

UAS CHALLENGE 2015


15.2.2 Yaw tuning

The Yaw tuning will involve a similar procedure to the pitch and roll tuning, although unlike the

pitch and roll values, it won’t require fine-tuning as yaw has less of an impact on stability. Having

said that, there are still good reasons for improving the yaw PID values to improve the response

(start and stop), reduce the overshoot and add a damping to the yaw acceleration.

Please see section 11.1.1 Loiter mode and 11.1.2 Altitude Hold Mode (AltHold) for other flight

modes.

15.2.3 Waypoint navigation tuning

In auto mode, the Quad-rotor will follow a pre-set path, from Mission Planner, and is capable of

doing certain tasks, such as deploy payload, taking video of flight path and pictures of current

locations. Tuning auto mode includes altitude and position from AltHold and loiter modes and as a

result should only be tuned after those two have been tuned. In the configuration menu the

maximum horizontal and vertical up/down speed can be changed in 10mm/s, so 25m/s will be

written as 2500. There is an issue where Pixhawk cannot maintain control of both altitude and

horizontal speed simultaneously whilst going over certain speeds, which can vary from Quad-rotor

to Quad-rotor. For this to be checked, the group’s Quad-rotor must be flown in auto mode to see

how much of a compromise this may be before steps are taken to overcome this issue.

Auto mode can be setup so that the Quad-rotor starts the mission from the ground or whilst flying.

If the Quad-rotor is set on the ground, then the throttle must be set to zero as the moment the

throttle is increased, the Quad-rotor will be set to auto mode and make its way to the first waypoint.

If the Quad-rotor is starting whilst in the air, it will start moving towards its first waypoint once the

controller has been set to auto. If, after it has been to auto, the first command is ‘take-off’, it will

recognise that command as completed and move to the next one. While the Quad-rotor is set to

auto mode, Pixhawk will overlook all inputs from the pilot as long as not disable auto mode and

yaw. As some pilots may decide to take pictures of the location, the pilot still has some control over

the yaw control, although Pixhawk will try to regain control once it has reached its next waypoint.

Waypoints can be set up as ‘fast waypoints’ (Ardupilot, 2015), which operate in the same way as

regular waypoints but without any delay or loitering at that waypoint. For both cases, a radius must

be inputted into Pixhawk, so that Pixhawk can recognise that way point as complete once it is

within a certain range.

UAS CHALLENGE 2015


15.3 Tuning during flight Once all PID values have been gathered for pitch, roll, yaw, stabilize mode, altitude hold, loiter,

and waypoint navigation, the Quad-rotor will be removed from the Quad-rotor and taken to a

controlled environment for flight testing. If required, the all the PID values for all modes will be

further fine-tuned during flight to check the response, stability and how well it handles disturbances

such as mild wind. Once it can fly well with a pilot, the Quad-rotor will be set in auto mode and

made to fly through waypoints, up to 2km in length, slow at first before increasing speed to

maximise the thrust from the motors and minimise the time taken. Once the above has been

completed, the same process will be continued but with payload deployment until satisfied.

15.4 Future Work As the test rig has been made in the final week leading up to the submission, no successful testing

was done. Now that the test rig and Quad-rotor has been built, from here on out, it will solely be

testing and fine-tuning the PID values for maximum performance and response time and ideally

without losing too much stability during the process. The Quad-rotor will fly up to 2km during the

UAS Challenge and after the changes in the latest briefing from IMechE, the Quad-rotor has a

large advantage due the quick acceleration and stoppage of a Quad-rotor over a fixed wing

aircraft.

UAS CHALLENGE 2015

73 Flight Termination Case MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Reyad 16 Flight Termination Case The core functionality of the Pixhawk software Mission Planner is to return to launch (RTL) if it

loses contact with the ground station or manual control. If more advanced options are required

then Pixhawk has an on-board Advanced Failsafe (AFS) system. The pilot can setup for failsafe

conditions so that the multi-copter can loiter for a short period of time before RTL, automatic

landing or termination (Plane.Ardupilot, 2014). If termination is chosen, then this will apply to all

modes of flight termination cases, whether that is GPS loss, communication loss, Geofence breach

or altitude breach. Once the aircraft has entered termination mode, it is no longer recoverable so

for this purpose the Quad-rotor will not be set on termination but land as a last resort.

16.1 GPS Loss The AFS system monitors the strength of the GPS receivers throughout the flight. If both GPS, on-

board and external, lose position lock for over 3 seconds, then the Pixhawk AFS initiates

(Plane.Ardupilot, 2014). This involves the system looking into one of the parameters called

‘AFS_WP_GPS_LOSS’ which instructs the multi-copter on its next action, ranging from loiter for a

period of time, disarming the motors and landing, or in a sequence of two or more of these actions.

It is also possible to specify a mission waypoint number which Pixhawk will use as a reference

point for where it should head to next if it loses GPS signal, similar to RTL. If the GPS regains

positioning then the multi-copter will continue its mission from where it left off.

16.2 Communication loss from Ground Station The AFS system constantly monitors the strength of the data-link between the Quad-rotor and the

ground station using the ‘HEARTBEAT MAVlink’ (Plane.Ardupilot, 2014) messages being

transmitted by the ground station. If for a period of 10 seconds or greater the multi-copter does not

receive a HEARTBEAT message then it enters AFS state. During AFS state, it looks for the

‘AFS_WP_COMMS’ parameter, which will contain a waypoint number to navigate to on

communication loss (APM Plane, 2014). The MAVlink messages are purely for Pixhawk, which

gets informed that it is receiving communication from the ground station. The ground station itself

will not see these messages from Pixhawk

If the multi-copter loses GPS positioning and connection with the ground station then this is

considered ‘dual loss’ and the multi-copter will immediately terminate. The user can, however,

override Pixhawk and enable manual mode and take control regardless of GPS loss, ground

station loss or both as long as Pixhawk has not already started flight termination. If all connections

are lost, including manual control then the multi-copter will terminate flight after a specified time in

milliseconds, in our case for 30,000 milliseconds or 30 seconds.

UAS CHALLENGE 2015

74 Flight Termination Case MEng Team Project Report (7ENT1024) School of Engineering and Technology

16.3 Geofence Breach Geofence allows the user to set boundaries of where the multi-copter can operate in terms of

distance and height. If the multi-copter goes outside the set boundaries it will switch to guided

mode and fly back to a pre-defined location (Plane.Ardupilot, 2014) or a failsafe condition such as

report back as seen in Figure 54.

Figure 54 Geofence configuration on Mission Planner

16.4 Maximum Pressure Altitude Breach When the airspace is being shared by multiple UASs, the flight altitude will be measured by a

common reference pressure, typically the QNH, defined as barometric pressure adjusted to sea

level. The AFS system can force a pressure altitude limit, as a value in millibars in the

AFS_AMSL_PRESSURE parameter, while the pilot can set the pressure altitude limit in the

AFS_AMSL_LIMIT (Plane.Ardupilot, 2014) in metres. If both parameters are set and are exceeded

then the AFS will initiate a termination process.

The AFS system will also monitor the barometer, and if it shows to be unhealthy for 5 seconds then

the AFS system will look at the AFS_AMSL_ERR_GPS (Plane.Ardupilot, 2014) parameter. The

multi-copter will enter flight termination immediately if it is set at the default value of -1 otherwise it

will continue flight and use the value as a margin to add to the GPS height and allow the flight to

continue if the GPS altitude plus the AFS_AMSL_ERR_GPS value, in meters, is below the

AFS_AMSL_LIMIT value. This margin value is to account for the inaccuracies of GPS altitudes and

according to APM, a value of 200 is reasonable for safety to ensure AFS_AMSL_LIMIT pressure

altitude is not breached (Plane.Ardupilot, 2014).

See full model in Appendix. J for more information on the safety case.

UAS CHALLENGE 2015

75 Systems Layout MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Tarek 17 Systems Layout

17.1 System block diagram

The over view of the operation and systems of the Quad-rotor are shown in the following two

diagrams. The first diagram shows the hardware and how the subsystems interact with each other,

while the second focuses on the operation of the software subsystems operation with the Quad-

rotor. The schematics of the systems are in Appendix K.3(ardupilot) (S@M, 2014).

17.1.1 Hardware Systems

The following block diagram is off the hardware of the Quad-rotor, the diagram has two sections

which operate in their own unique way. The first section is the ground control station which has

two subsystems, the controller and the ground control station computer with the communication

system. The base station computer is what stores the Mission planner, this software allows for the

control of the flight path and the operation of the Quad-rotor. The communication system linked to

the computer communicates with the Quad-rotor which allows for transmitting and receiving data.

The radio control allow for manual control of the Quad-rotor by the pilot.

The second section is of the Quad-rotor consists of motors, flight controller, power distribution,

camera, GPS Module, and video graphics processing unit (VGPU) or the minimOSD along with

pixhawk. The motors flight controller and power distribution resample the systems that of

propulsion systems. The GPS module is for pinpointing the location of the Quad-rotor and flight to

the desired location. Finally the camera and the minimOSD with pixhawk are part of the

transmission of the video and data feed to the ground control station.

Figure 55 Overall System Hardware Block Diagram

UAS CHALLENGE 2015


Signal name Description

User input The user input is to turn the Quad-rotor on/off, toggle on/off video stream, activate/

deactivate the autopilot and arm/disarm the Quad-rotor. This is achieved through

ground control station Mission Planner and using the radio controller allows for

manual flight of the Quad-rotor.

Power The power supply of the Ground Control Station is from the laptop, where it must

have the battery fully charged before the mission. At the ground station a portable

power supply will be available for the computer to be connected to for recharging.

GPS Satellite

Signal

The GPS system on the Quad-rotor receives a GPS signal from global orbiting

satellites and on the ground station it is connected to WIFI where it updates its

mapping and positioning.

RC Control

Signal

The manual control is through the transmitter and receiver of the radio controller.

Telemetry

Data

The communication between the Quad-rotor and Ground Control Station is through

the 3DR telemetry kit operating at 433MHz.

Video Data The video data is sent through the video transmitter and receiver kit which is

connected to the minimOSD which includes extra video data such as altitude,

attitude and direction.

Motor Thrust The motors function is provide thrust in order to lift the Quad-rotor and travel

around the course.

Sensor Data The flight sensors record various data such as accelerometer, magnetometer and

gyroscope which there information is sent to pixhawk which are then processed to

meet the flight conditions.

Table 11 Overall System Hardware Block Diagram Description

17.1.2 Software Systems

The following diagram and table are of the software block diagram of the Quad-rotor systems. This

section also has two subsystems, as there is software running on the ground control station and on

the Quad-rotor. The ground control has two main subsystems and some have further subsystems.

The first subsystem is the Radio Controller (RC) transmitter, which transmits manual pilot control

commands to the receiver on the Quad-rotor to control the flight conditions on the Quad-rotor. The

second subsystem contains the ground control station computer, running operating windows 7

using the Mission Planner for the mission planning. The Mission planner receives video and data

information from the Quad-rotor, the Mission planner then displays that information and stream for

the user. On the Quad-rotor there are three main subsystems. The first is the RC receiver which

receives signal from the radio controller on ground which operates on a tuned frequency for the

receiver and transmitter to operate coherently. The second subsystem is Pixhawk, it receives

signal from RC receiver with command to control the Quad-rotor. The second is the video graphics

processing unit (VGPU), the minimOSD receives data from Pixhawk such as the altitude, attitude,

and heading etc which are processed with the video feed and transmitted to the ground control

UAS CHALLENGE 2015


station. Pixhawk is the main computer or brain on board the Quad-rotor, Pixhawk receives data

through the telemetry kit which contains flight commands such as GPS coordinates or signal when

to release the payload. The table below contains more descriptions on individual systems.

Figure 56 Overall Software Block Diagram

Signal name Description

Video

Commands

Video command will be sent through the video link which makes the camera take a

photo of the target for it to then be processed. This will read the alphanumeric

information at the target and display it at the ground control station.

Video Data The video data is transmitted from the camera on bored the Quad-rotor, with

information from the minimOSD. The video will be displayed on the ground station

Mission planner. The video is transmitted through video transmitter which will be

operating in the same frequency as the video receiver on the ground station.

Telemetry

Command

The telemetry command is sent from the Ground Control Station Mission planner

through the telemetry transmitter to the receiver which then sends the information

to pixhawk to be processed.

Telemetry

Data

The telemetry data sends data from the Quad-rotor with pixhawk data to ground

station. The data from pixhawk includes information such altitude, attitude, location

and speed which are displayed on the Mission planner page.

User Data The user data is the collection of the flight information which is displayed on the

Mission planner with information regarding current flight conditions.

RC

Commands

The RC commands are the commands transmitted by the transmitter to the

receiver with flight control commands to control the flight conditions of the Quad-

rotor

Table 12- Overall Software Block Diagram Description

UAS CHALLENGE 2015


17.2 Communication

The range and performance of the radio frequency (RF) link are critically dependent on the

antenna used. The radiation pattern of a quarter wave monopole antenna is heavily dependent on

the design and layout. Therefore selecting the correct antenna and placing in the most efficient

location on the Quad-rotor is crucial. The mounting of the telemetry and the video transmitter, must

take into account of the possibilities of shadowing, as this can be a factor when mounted in an

obstructed area for example between the two structural plates of the Quad-rotor. The effects of

shadowing will hinder the range and coverage of the transmission range. For this reason the most

common set up on a UAV or aircraft is the vertical polarization. As the advantage of a vertical

polarization, waves propagate much more effectively in this orientation near the earth, whereas

horizontal polarized the waves will be cancelled out by the reflection from the earth.

Electromagnetic Interference (EMI) is common issue that occurs with electronic devices as they

might interfere or interrupt the performance of a device, due to radiation and the source could be

from nature or manmade devices. If the EMI intervenes with the aircraft systems it could turn out to

be a very serious issue during flight, especially if a system such as navigation are disrupted this

would lead to a loss of signal and would lead to missing the flight path hence increasing flight time.

Therefore, on the Quad-rotor the mounting of the GPS will be placed on most elevated location on

the Quad-rotor, and the telemetry and video transmitter will be placed some distance away from

each other to avoid interference (Wyatt & Tooley, 2008).

When testing the GPS and telemetry kit loss of performance was identified, and the reason for this

was that the operating frequency of the laptop is up to 400MHz and the telemetry operates at

frequency of 433MHz. During testing the laptop was used as the power source for the GPS and

telemetry kit. This could also be due to path loss as the test was carried out on a long narrow field

with trees obstructing the line of sight signal. Nonetheless a range of more than 400metters was

achieved as this was down to the maximum length of the field. To protect signal strength aluminum

shielded wires are used to protect against EMF to help reduce cable loss. Cable loss is the amount

of signal lost due to the cable, another measure taken to reduce this effect it to have long enough

cables to reach each connection point because the longer the cable the higher signal loss (Bailey,

2003).

UAS CHALLENGE 2015

79 Image Processing MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Tarek 18 Image Processing

18.1 Image Recognition

18.1.1 The Requirements

Image recognition code will be used to read the letter at the target and displayed letter on the

ground station screen. Earlier competition requirements stated that there would be a mixture of

alphanumeric characters at the target which should be recognized and displayed at the ground

station (Barragan, 2014). However the march 2015 rules state that there will only be one letter at

the target in a target area of 2m by 2m.

18.1.2 Testing

For testing purpose the target has been scaled down to resemble real life operation. The

parameter of the square target is 2m by 2m and the Quad-rotor cruise altitude is at 100ft. To verify

the code ability to recognize the target letter, tests were carried out at different altitudes to compare

the results. The reason for testing at different altitudes is because for payload deployment the

Quad-rotor would need to descend to an altitude to safely deploy the payload at the target. The

altitudes that have been selected for testing are at 100ft, 50ft and 20ft. The target will be elevated

at 1.5 m above the ground.

Scaling

As the delivery box is elevated above ground at 1.5m = 150cm, taking scale at 1/20 therefore

testing is as follows:

For 100ft:

100ft = 3048cm → 3048 − 150 = 2898 cm

∴2898

20= 144.9 𝑐𝑚

For 50ft:

50ft = 1524cm → 1524 − 150 = 1374 cm

∴1374

20= 68.7 𝑐𝑚

For 20ft:

20ft = 609.6cm → 609.6 − 150 = 459.6 cm

∴459.6

20= 22.98 𝑐𝑚

As the altitude was scaled down the target character should also be scaled down. The target is 2m

by 2m.

2𝑚 = 200𝑐𝑚 →200

20= 10𝑐𝑚

Therefore the target will be 10cm by 10cm with the letter in the middle of the square target.

UAS CHALLENGE 2015


18.1.3 Results

To represent the altitude of 100ft, 50ft and 20ft, the test will be carried out at 144.9cm, 68.7cm and 22.98cm.

For testing only two letters where tested, H and Z.

Altitude (cm) Alphanumeric Letter Result

144.9 H 1317 characters where displayed

144.9 Z 1400 characters where displayed

68.7 H 2489 characters where displayed

68.7 Z 1568 characters where displayed

22.98 H H

22.98 Z Z

Table 13 - Alphanumeric processing at different height

Figure 57 Matlab alphanumeric code processing letter at 22.98cm

18.1.4 Analysis

From the result obtained in Table 13, it shows that when taking a picture above 50ft the results are

not consistent and hence the Quad-rotor is required to descend to an altitude of lower then 20ft to

achieve a more accurate result. Descending to altitude lower than 20ft is beneficiary for the

payload deployment because it allows for greater possibility of a safer deployment.

18.1.5 Shape recognation

The shape recognation code was planned to be used to identify the target during flight, when the

Quad-rotor reaches the coordinate set at the ground station using the Mission planner the program

would be able to identify the target and center itself ontop of it. The code measures the properties

of image regions and will scalar the actual number of pixels in the region of the image, which then

can be identified as a shape, in the figure bellow it demontraights its operation. The setback of

running the code during the mission is its time to process the image, in this case it took more then

30mins as the code removes any connected components pixels that have fewer then 60 pixels.

UAS CHALLENGE 2015


Normally it is running at a minimum of 20-30 pixels which takes 2-5 minutes, however as the test

will be in an open field the grass causes a lot of interferance during processing hence a higher

pixel setting is required (Samieh, 2007).

The figure below is taken from a height of 22.95cm to resemble 20ft and the size of the box is

10cm by 10 cm to resemble the target of 2m by 2m:

Figure 58 Shape recognition

18.2 Video

The camera model selected for the live feed video to the ground station is through Mobius

ActionCam. The Mobius camera is commonly used on such UAVs, the camera provides a high

quality video feed and the quality can be altered from three possible choices. This can be useful if

needed to transmit over a long range but operating at its highest resolution of 1080p-30fps will be

needed as the ground station is within a reasonable distance and it would need to process the feed

to determine the alphanumeric at the target. The camera is needed to provide a still image of the

target from the Quad-rotor and transmitted to ground station (Mobius, 2015).

The camera has the ability to record while streaming, this would allow for playback of the flight at

another time. However this feature is not crucial but it may be used to analyze the flight condition

of the Quad-rotor. The camera has five video recording cycles’ time settings they are 3, 5, 10,

15mins or “Max”. The max will record until the 4GB memory has reached its limit and if recording at

1080p-30fps that would equate to 30mins of footage this can be increased by the use of an

additional memory card.

Initially Boscam TR1 was going to be used as the FPV camera but this was dismissed as the

camera is not compatible with the minimOSD. The camera would require an additional transmitter

for it to transmit video data from the minimOSD. Mobius ActionCam was then selected for its video

quality its size and weight.

UAS CHALLENGE 2015


18.3 On Screen Display Board (OSD) The Quad-rotor will be fitted with On Screen Display bored to view the flight data at the ground

station. The model chosen is the MinimOSD due its compatibility with Pixhawk, configuration ease

and the error indication and warning system (lost GPS fix, over speed, battery voltage and

percentage and the received signal strength indication). The MinimOSD also will display the

direction, altitude, attitude, current waypoint and heading. The displays can be changed by using

MinimOSD-extra Firmware to reprogram the OSD to display additional features such as vertical

speed and way point distance. The OSD must be connected through a FTDI Breakout board which

can then be connected to a computer for programming. The video feed must be connected to the

MinimOSD for it to contain the additional video data provided from pixhawk, the OSD will then

need to be connected to the video transmitter to transmit the video full diagram in (S@M, 2014).

The minimOSD will display the battery life remaining of the Pixhawk, the battery connected to the

Pixhawk will also be supplying power to the video transmitter and servomotor. The setback of this

setup is that at the ground station the user will not be able to observe the battery life of the motors

as they are not connected to pixhwak. The reason for the motors having a separate power source

is due to the fact that the motors are powered by a 5s battery where as Pixhawk can only operate

with a maximum sized battery of 4s.

18.4 Video transmitter The video transmitter Boscam TS351 operates at a frequency of 5.8GHz and transmission range

of 500m with a standard antenna. The transmitter will be mounted on the Quad-rotor legs to reduce

interference with other signal transmitters and receivers. The video transmitter will also be

operating in a vertical polarization position. Transmitter will be powered through the pixhawk

battery however the transmitter will require a voltage regulator to reduce the voltage provided from

the 4s battery to 12v from 14.8v.

18.5 Video Receiver The ground station will receive its video feed through the Boscam RC305 which contains 8

channels that similar to the transmitter, also the frequency band is 5.8GHz. The video receiver will

be connected to an external USB video capture card which will then display the video on the

laptop. The capture card contains its own software where it’s able to play video or toggle the video

stream to capture the image. The benefits this setup is that it allows for the image processing to be

done on one laptop and the mission management on a separate laptop. This will allow for a

delegation of reasonability distributing the work load between two users and will reduce the

possibilities of reduction in laptop performance. The video receiver contains two output pins, this

also allows for an optional additional FPV screen for the pilot, the output of the receiver is an AV

cable for that reason it can easily be connected to an FPV screen but it requires an adapter to

stream video on a laptop.

UAS CHALLENGE 2015

83 Verification and Validation MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Tarek 19 Verification and Validation The validation and verification stage is from the V model, which is used in all forms of engineering

projects. The V model operates from a hierarchical perspective starting from requirements,

standards to testing. The benefit of the V model allows for easy tracking of the phase where the

product is currently held, for example when the product reaches the verification stage it measures

how the system was built to the system requirements (Monhem, 2010). The flow direction of the V

model is all interchangeable as after one stage is complete one can check if the outcome suite the

previous stage requirements as it’s a good method of defect tracking. Also it’s a cost effective

method of making sure the right product is built as once the product reaches the validation or

operation stage and spot that it does not match the requirements or regulations this will hinder the

progress of the project.

19.1 Verification Matrix

The verification stages starts off with verifying the systems requirements document and analyzing

the requirements and verify if they satisfied every “shall”, “may”, or “should” statements. The

statements are collected and in a document called Verification Matrix. The document will define

each requirement and the verification method it will show the type of test methods to be carried to

verify the product matches the requirements. The following key terms are used in the Verification

Matrix in Appendix K.1.

Inspection: Visually verify form and configuration of the hardware or software. Inspection involves

the use of measuring tools to retrieve values such as mass, dimensions and other physical

characteristics.

Analysis: includes computation or comparison to previous or experimental data. It verifies the

conformance by the use of analytical tools, modelling or simulations which will allow for a

predication of performance with use of calculation or subsystem testing.

Demonstration: Is to verify the required operability of a software or hardware that does not require

qualitative measurement or the aid of a test device. Test device can be used to contribute to

demonstration of the function.

Test: to verify the conformance of the performance, physical characteristics and characteristics to

requirements with use of technical operates to retrieve detailed quantification of the performance.

19.2 Validation test

The validation tests require relating the results of the verification back to the requirements in

evidence to show the compliance of the product to the requirements and to meet the rules

specified by IMechE. Some of the validation tests are yet to be completed but all the specified

validations in the table will be tested before the competition. The validation table contains the dates

when each test should be carried out are in Appendix K.2. Some of the tests that were carried out

did not meet the requirements and they are rescheduled for testing at a different date after further

development.

UAS CHALLENGE 2015

84 Future work MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Tarek 20 Future work The following concepts where considered but where not investigated or developed further because

of complexity, time and mainly due to the limitation of remaining funding. The main reason for not

proceeding with these concepts is due to cost, as the spending on the Quad-rotor is very close to

the maximum value of COTS which is set at £1,000. If the component list goes above the

maximum COTS the Quad-rotor will not be allowed to enter the competition, hence these where

not established or expanded on further due to the cost issue.

20.1 Partial control of Quad-rotor positioning For partial control of the Quad-rotor during the deployment of the payload, it would require an

additional two small motors that allow for the maneuverability of the Quad-rotor in the +x,-x,+y and

–y axis. It was planned for the two motors to be controlled through a Bluetooth connection as the

deployment area to the ground control station is within a 60meters distance. The Bluetooth module

allows for transmission range of up to 60 meters. The motors can be controlled using a mobile

phone or tablets. The Bluetooth module would be connected to an ardunio bored which will allow

for the control of the motors. Pixhawk will still be controlling the altitude and attitude of the Quad-

rotor, this is only possible when the GPS is deactivated allowing for the manual control of the

motors to adjust the positioning of the Quad-rotor (Santos, 2013).

20.2 Full Autonomy To achieve full autonomy the shape recognition code must be used to detect the square target.

The camera can be mounted on a gimbal, which allows for the camera to adjust the target on the

screen by centralizing the target. Then as the gimbal aims towards the target and laser will be

mounted on to the gimbal with another sensor which will pick the laser and that would then operate

the two motors similar to section 20.1but without the need of the Bluetooth receiver. Therefore as

the camera tilt to centralize the target the laser will be pointing at an angle, the laser sensor would

then detect the laser and operate the motors until the Quad-rotor is directly above the target and

the laser is pointing vertically down in the z axis. This would also require deactivating the GPS for a

set time so pixhawk will not correct its positional hold (techbitar, 2013).

Figure 59 Circuit Diagram of Ardunio

UAS CHALLENGE 2015

85 Preliminary Payload Box Concept & Servo Integration


Section by Micky 21 Preliminary Payload Box Concept & Servo Integration

As part of the UAS design challenge, it has been given out to design a payload delivery system

and mechanism. Initially, 3 methods were considered to in the delivery of the payload which is the

hinge-clamps system, the electro-magnet method and the hinge-pin method. However, following a

radical change in the design of the Hex-rotor with a box capable to accommodate 2 bags, it was

concluded to fit the new Quad-rotor with a box able to accommodate only 1 bag to flour.

21.1 Initial designs

21.1.1 The Hinge-clamp Method

In this concept the bags of flour are to be put in pre-designed cases which have small holes in

which the clamps tooth will integrate. In addition, the cases of different sizes would be superposed,

with the bottom case slightly wider than the top one.

Then, the system would be coupled to a set of 4

clamps actuated by 2 servo motors.

The advantage of the system would be no variation

of the C.G in the XOY plane. In addition, the

structure is robust in case of turbulences, vibrations

or sudden movements

The disadvantage of this method is that the hinge-

clamp and servo motor can be tedious to implement

21.1.2 The electro-magnet method

In this second concept, the pre-designed cases

come with metallic bars placed at precise

places in the 2 cases that is to say; the top case

would have a metallic rod at its bottom top face.

The second case would have 2 metal rods

located at the corners of its top face. The metal

rods would interact with 3 electro-magnet

placed at specific locations on the airframe.

The advantage is that the C.G will not vary in the XOY plane

The main disadvantages of this method are first of all, once the electro-magnet are on, the

magnetic field might interfere with the overall electronics on board. Secondly, the electro-magnet

will drain the power greatly. Finally, the additional rods will add more weight on the UAV and the

structure might not be appropriately robust in turbulence and vibration scenarios.

Figure 60: Hinge clamp

Figure 61: electro-magnet

UAS CHALLENGE 2015



21.1.3 The Hinge-pin method

This last concept relies upon a set of pins

coupled to hinges to release the payload. In

this scenario, the loads are placed beside

each other and one after the other are

released once the UAV arrive at their

respective drop-off point.

The advantage of this system is that it is a

relatively simple system to operate

The disadvantage of this method is that, once a load has been released, it would cause an offset of

the initial centre of gravity point which the UAV would have to adjust and thus, will use the power

greatly.

21.1.4 Others

This design came as the quad was still meant to carry 2 payloads and a set of linear servos.

However, the linear servos was found to be too big, heavy and very expensive. Moreover due to to

weigh restricted it was decided to use only 1 payload at the time.

Figure 63: Other concept

Figure 62: Hinge-pin

UAS CHALLENGE 2015



Section by Micky Conceptual design by Zuber Khan

Actual design by Amit Ramji

Figure 64: CAD

21.1.5 Payload box mechanism integration

A representation of the payload box final designs is as show in Appendix B.7. Once the servo is

powered, it actuates the horn that rotates of an angle of 90 degrees. Subsequently, the movement

release the movable door which lets the bag of flour fall. The detailed design can be found in

section 4.2 and structural analysis in Appendix G.15 to G.16.

Figure 65 to Figure 69 is an illustrative reproduction of the final design carried out in section 4.2,

these aim to show the servo integration.

Figure 65: Overall payload box

UAS CHALLENGE 2015



Figure 66: Horn and door connection

Figure 67: Start up release

Figure 68: Fully Unlocked door

UAS CHALLENGE 2015



Figure 69: Complete release

21.2 Servo

It was decided to choose the MG90S servo, Metal gear with one bearing for relevant reason

towards the UAV specifications. Tiny and lightweight with high output power, this tiny servo is

perfect for RC airplane, helicopter, Quad-rotor or robot. This servo has metal gears for added

strength and durability. The servo can rotate approximately 180 degree (90 in each direction), and

works just like the standard kinds but smaller. This servo is controllable with any code, hardware or

library to control; these servos. This servo is appropriate to make part moves without building a

motor controller with feedback and gear box, especially since it will fit in small places

21.2.1 Specifications

Weight: 13.4 g

Dimension: 22.5 x 12 x 35.5 mm approx.

Stall torque: 1.8 kgf.cm (4.8V), 2.2 kgf.cm (6V)

Operating voltage: 4.8 V – 6.0 V

21.2.2 Rational

This digital servo uses switched mode power which is considerably more efficient than the

analogue power alternative. A small microprocessor inside the servo analyses the receiver signals

and processes these into very high frequency voltage pulses to the servo motor. Instead of 50

pulses per second, the motor will now receive upwards of 300 pulses per second. The pulses will

be shorter in length of course, but with so many voltage pulses occurring, the motor will speed up

much quicker and provide constant torque.

The result is a servo that has a much smaller dead band, faster response, quicker and smoother

acceleration, and better holding power. In order for the servo to operate smoothly, the force it

generates should have be greater that the friction force Ffrict-sliding as its horn slide along the

Figure 70 - MG90S servo

UAS CHALLENGE 2015



payload box door. Thus it should be greater that Ffs = 0.013 Kgf. The calculation can be found in

Appendix M.1

21.3 BEC They require +5V to power the opto-isolator and while the Pixhawk can be powered from the servo

rail, it does not provide +5V to the servo rail. The ESCs must be powered by a BEC or with a

jumper from an unused connector on the board. In this case, it was decided to use an SBEC to

power the electronic rather than a jumper.

Turnigy 5A SBEC is an advanced switching DC-DC regulator which will supply a constant 5A. It

works with 2 - 7 Cell Lipoly pack and supplies a constant 5 or 6v to your receiver and is

interference-free, perfect for confined spaces

21.3.1 Specification

Type: Switching

Input protection: Reverse polarity protection

Output (Constant): 5v/5A or 6v/5A

Input: 8v-26v (2-7cell lipo)

Weight: 18g

21.3.2 Rational

There are 3 main types of power regulator or battery elimition circuit which are BEC, UBEC and

sBEC.

The BEC and the UBEC are good power regulator for small specifications such as those involving

curreent bellow 10A and voltage difference across the BEC less than 5V otherwise there is a risk

of short circuit or melting circuit thus damaging the flight control.

In this case the Swicthing sBEC26 has been chosen because it design make it prone for hight

voltage discharge with considerably less heat emanation thus less waste of power. Moreover, it

can a providd more power throught the intensity of curent it supports. The crucial importance of a

voltage regulator for the system is that the high voltage supplied by the batteries (26 - 18V) would

damage the flight control components which opperates below 6V.

Figure 71 - SBEC26 Turnigy

UAS CHALLENGE 2015



21.4 Schematics of connections from battery to servo through pixhawk

Figure 72: Schematics of connections

The Quad-rotor will run with 2 batteries. The battery pack 1 (18.5V, 16Ah, 3s LiPo) will only run the

motors whereas the RC receiver and the payload servo will be run by the battery pack 2 (11.1V,

2.2Ah, 2s cell LiPo). The reason for this arrangement is that once the motors are switched off, the

flight control system Pixhawk is still reading its mission.

UAS CHALLENGE 2015



The battery pack 2 will power the servo and other receiver through the SBEC which will drop its

voltage to 5V-6V. The SBEC is connected on the AUX OUT pin 6 and the servo will be connected

on the AUX OUT pins from 1~4 since the platform is Arducopter. The RC receiver is connected at

the RC pin.

21.5 Controlling the servo as a servo

Firstly, the Quad-rotor will perform a loiter in a figure of 8 before engaging into releasing the loads.

As the servo will be used to operate the payload box door during the delivery phase, it will be set

as servo in the mission planner of Pixhawk. The way to control a servo under this type only works

as part of the mission that is to say autonomously. To do so, the Pixhawk should be connected to

the mission planner as follow:

On the Config/Tuning > Full Parameter List page, ensure that the RCXX_FUNCTION is set

to zero for the servo that’s to say RC9_FUNCTION as the servo is connected to the

Pixhawk’s AUX OUT 1).

Then Press the Write Params button

Figure 73: Configuration of the servo on Pixhawk

Following Create the mission to be fly and add a DO_SET_SERVO command and include the

servo number ( “10”) in the “Ser No” field and with the PWM value (usually between 1000 ~

2000) in the “PWM” field.

09

http://dev.ardupilot.com/wp-content/uploads/sites/6/2013/08/MissionPlanner_ServoSetup.jpg

UAS CHALLENGE 2015



Figure 74: Mission with GPS dropping points

The DO_SET_SERVO command is a “do command” which means that it can only be run between

waypoints so it must not be the first or last command in the mission. It will be executed

immediately after the waypoint that precedes it. After the first payload is dropped, the Quad-rotor

will return to the ground station location to be fitted with the 2nd payload and perhaps a new battery.

21.6 Testing with the Mission Planner This verification phase involve testing whether the servo are moving as expected. The mission

planner’s Flight Data screen includes a “Servo” tab on the bottom right that can be used to test that

the servos are moving correctly.

Figure 75: Verification of the performance of the Servo

http://dev.ardupilot.com/wp-content/uploads/sites/6/2013/08/MissionPlanner_MoveServo.png

http://planner.ardupilot.com/wp-content/uploads/sites/5/2013/08/Servo_TestingWithMP.png

UAS CHALLENGE 2015

94 Other Involvements MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Micky 22 Other Involvements

22.1 Telemetry Kit The telemetry kit provides and ground-to-air data link between the auto pilot and your station

laptop or tablet. There are mainly 2 wireless telemetry kits which are a radio set kit and a Bluetooth

data link set. The latter is certainly cheap, however, it is only intended for pre-flight ground use

only, and it is not a replacement for a RC transmitter and receiver. The main disadvantages of the

Bluetooth set are its limited range of around 50m and its overall weight of 9.5 g. Therefore, it would

be appropriate to use a radio set telemetry kit. The 3DR radio set has been chosen for the purpose

of this project and details of the prices of the parts are included below in the Appendix. M The

range of the radio set could be increased by replacing the ducted original antenna with a high gain.

22.2 Design Convergence At the start of the project, we had to go through a design convergence method in order to choose

the right vehicle. I was responsible to look into an osprey design with titling rotors. It was found that

the mechanical design has a high complexity level. Furthermore, my involvement was noted on the

choice of the autopilots which were Ardupilot-Mega and Pixhawk.

22.3 Challenges

It was particularly challenging to find information related to the system integration for many

reasons. First of all, the use of Pixhawk has made it particularly difficult to find and integrate other

system component because it is relatively new. The majority of information available online are

related to APM (ardupilot Mega) which is the earlier version.

UAS CHALLENGE 2015

95 Manufacturing MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section written by Mozammel Material, design and manufacturing decisions by Amit Ramji

Section by Mozammel

23 Manufacturing

In order to achieve an efficient structure, manufacturing methods were identified at a primary stage

of the project. The manufacturing plan included materials to be used, joining methods, machines to

be used and the best possible way to carry out the tasks on time. Initially composite laminates and

tubes were intended to be used however the complexity in manufacturing and cost restrictions did

not allow this.

23.1 Machining Selection

Acknowledging the weight and budget limit for the project, the manufacturing process includes

milling, lathe, laser cutting and CNC machining which are available within the lab facility of the

university.

23.1.1 Machines

The following machines are used to manufacture the parts depending on their operating functions.

Machine Type Functions

Milling machine (Bridgeport Series 2)

Use end mills to obtain precise dimensions

Use centre/slot drill to do holes

Use fly cutter to obtain smooth surface

XYZ 1330 Lathe Use high speed steel tooling to obtain smooth surface

on the nylon 6.6 rod

Use high speed steel tooling to machine centre holes

on the nylon 6.6 rod

Trotec Laser Cutter Use laser to cut the Nylon 2mm thick plate for main

body plate

Vertical Bandsaws machine Use to cut raw materials into required dimensions

Denford Router 2600 Pro Milling Machine

Use to obtain components directly from CAD model

Denford VMC 1300 Milling Machine Use to obtain components directly from CAD model

Table 14: List of Machines

23.1.2 Tools

Tools Functions

1 High Speed Steel Tooling For precise cutting in XYZ 1330 Lathe

2 End mills For precise cutting in milling machine

3 Centre drill For accuracy in drilling holes

4 Slot drills For drilling holes in milling machine

5 Fly cutter(single point) For precise cutting in Bridgeport series 2 milling machine

UAS CHALLENGE 2015


6 Centre Finder Complete To setup the datum (X,Y,Z directions) in XYZ 1330 lathe, Bridgeport series 2 milling machine

7 Metric slip gauges To obtain accurate measurements

8 Precision Parallel Set For accurate setup

9 Micrometre To measure dimensions

Table 15: List of tools and their functions

23.2 Manufacturing process of Quad-rotors components

The machining of the components includes different machines but identifying the most simple yet

better finishing quality was preferred. Due to the limitation of technical facilities and knowledge

components are marginally modified. All the sharp edges are smoothened to obtain edge fillets and

radii features by using sandpaper machine; by doing so the possibility of cracks and fatigues in the

structures is reduced.

23.2.1 Fixed Bracket

The fixed brackets are made of nylon 6.6, which has the favourable

characteristics to hold the arms in place. Also bearing in mind, the

finishing quality is more emphasised and that is why using the milling

machine, the brackets are manufactured. Figure 76: Machined fixed

bracket shows the fixed bracket machined in milling machine.

Due to some limitations in CNC machining, smooth edges were not

obtained as shown in Appendix. N

23.2.2 Motor arm end bracket

The compressive characteristics of nylon 6.6, makes it an ideal

material to be used to securely hold the motors into the motor mount

plates. The brackets are drilled by using end mills of 13 mm diameter

followed by 16 mm diameter. The 3mm diameter holes are drilled

through to be fastened with motor mount plates.

23.2.3 Movable arm vertical fixed bracket /support bracket

From nylon 6.6, the brackets are machined in milling machine as per

the technical drawings. All dimensions are carefully machined

according to the technical drawings but the edge fillets of radii of 12.5

mm were comprised due to complexity and availability of appropriate tools.

Figure 76: Machined

fixed bracket

Figure 77: Machined

end bracket

Figure 78: Machined Fixed bracket

UAS CHALLENGE 2015


23.2.4 Landing gear top/bottom support bracket

Considering the impact of the landing gear, nylon 6.6 (16mm cast sheet)

is used to manufacture the brackets which makes it more reliable to

support the landing gear. The bottom support bracket has a 2 mm

counter bored of 8 mm diameter to attach a spring which has

compression stiffness of 300 N/mm and the damping of 0 N.s/mm. The

spring is considered to withstand the impact from the landing gear which

is bridged between landing gear bottom support bracket and the landing

gear pivot.

23.2.5 Top/Bottom half T-joints

Evidently these parts were the most challenging to manufacture

considering the design detailing. Since the joints are designed to hold

the landing gear strut and stabilizer, the dimensions are critically

important. The parts could have been 3D printed but the materials

properties would have been different since the 3D printer at the

University of Hertfordshire only uses Acrylic.

23.2.6 Landing Gear Lug Bracket/ Pivot

From a 30 mm cast sheet of nylon 6.6, the part is machined to the

designed dimensions by the milling machine. The edge fillets and the

radii were achieved by using the sand paper through visual

inspection.

The landing gear pivot was machined by the lathe machine followed

by the milling machine to get the diameter of 25 mm from 30 mm cast sheet

of nylon 6.6 and the pivotal section respectively. The groove for the spring

was modified by making a counter bored of 2 mm depth.

Refer to 308Appendix. N for the landing gear lug bracket with the landing

gear pivot.

23.2.7 Arm pivot

The arm pivot required two types of machining; centre lathe and milling. The

nylon 6.6 rod was clamped into the four-jaw chuck and the desired length

and diameter was cut by high speed steel tooling; then the rod was

bored 22mm deep and milled 5.5 mm from both sides on the other end.

The remaining flat part was then drilled to make it suitable for pivotal

function for the movable arm.

Figure 79: Machined bottom support bracket

Figure 80:T-joint on foam

Figure 81: Lug bracket

Figure 82: Landing

gear pivot

Figure 83: Arm pivot

for movable arm

UAS CHALLENGE 2015


Section by Mozammel Motor Mounts Manufactured and Assembled by Zuber Khan

23.2.8 Main Body Plate

Due to ease of use in Trotec Laser machine setup, laser cutting was

attempted. Even though accurate dimensions were obtained but during

machining the heat of the laser melted the edges and clear smoke was

observed.

To overcome the challenge of heat damaged edges, the main body

plates were machined in Denford Router 2600 Pro. The advantage of

such machining has high accuracy and the material properties are not

affected as much by any thermal energy; deburr and polishing of sharp

edges was carried out using hand file.

23.2.9 PVCs tubes

All the tubes were roughly cut down by hand and then by lathe machine precise dimensions were

achieved.

23.2.10 Motor mount plate

The positions of the holes are really important as to align with the motors accordingly. The

machining of motor mounts plates were attempted on the CNC machine but since the thickness of

the aluminium plate is 1mm so clamping was not achieved properly with current tool constraints.

Therefore the holes positions were carefully marked by hand and piloted by a 1.5 mm drill and

finished off by a 3 mm drill. Figure 16 shows the brackets are screwed in with the motor mount

plates.

Figure 84: Cutting nylon plate

in Laser machine

Figure 85: Melted edges

Figure 86: Plate after cutting

Figure 88: Motor mount plate Figure 87: Assembled motor mount

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23.2.11 Overview of Machining

1. Milling Machines (Bridgestone Series 2)

Mostly milling machine is used to obtain precise cutting (by end mills), smooth surface (by fly

cutter), and holes (by centre drill and slot drills). Centre finder complete, precision parallel sets,

micro-meter and metric slip gauges are the usual tools that are used while milling the components.

Refer to Appendix. N for figures of some machined components by milling.

2. XYZ 1330 Lathe

Arms (fixed and movable), arm pivots, landing gear pivot, landing gear strut and stabilizers are

machined in lathe machine. The arms were drilled in 20 mm by a 13mm slot drill on one side to

install the LED lights. The nylon 6.6 rod (diameter of 25mm) was machined to diameter 22 mm by

high speed steel tooling and then bored into 22 mm at the centre of the rod with a diameter of

16mm.Refer Appendix. N for figures regarding lathe machine.

3. Tortec Laser cutter

The laser cutter is used to machine the main body plates but it has been identified that the heat

has melted the edges of the plates so subsequently it is decided to machine on CNC machine

Refer Appendix. N for figures on laser machining set up.

4. Vertical Bandsaws Machine

Vertical bandsaws machine is used to cut the purchased block or sheet (aluminium alloy) into

required dimensions for the components Refer to Appendix. N

5. CNC Machines (Router 2600 Pro and VMC 1300)

CNC machines are used for machining the main body plate and the turn button for payload box

after several practise sessions. Although simulations were carried out to obtain motor mount

plates, brackets and T-joints but unfortunately due to lack of knowledge at that point of time,

machining was countered with many known and unknown errors. Since the building of the Quad-

rotor solely depends on the manufacturing timeline so with the supervision of the technicians

majority of the parts are machined in the milling and lathe machines.

Figure 89.1-3: CNC practice sessions

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23.3 Challenges

Compared to acrylic and wood foam which are mainly used in practise session, nylon and

aluminium are proven to be hard to machine. Alongside desired tools for clamping, machine

planning and drill sizes were not available during manufacturing. Hence manufacturing of the parts

were mostly dependent on milling and lathe machines and as a result constrained some desired

features of the parts. Pace of manufacturing was also affected as supervision was required while

using the milling and lathe machine in the machine lab.

Below show failed attempts on motor mount when tried on VMC 1300 and router 2600 Pro.

Figure 90.1-3: Failed attempts

23.4 Manufacturing Plan

After the hand calculations and numerical analysis (Finite Element Analysis) carried out by, the

materials were purchased. As shown on the Gantt chart below, practise session is the longest due

to the limited availability of machining lab and lack of hands on experience with the machines. As

per the manufacturing plan the assembly is scheduled to be done on the first week of April.

23.5 Machining Cost

All the components were machined at the university labs thus eliminating any labour and

operational cost for the machines.

23.6 Other involvements in the project

In course of this project and due to the necessity of the project progression, the following

contributions we made.

Design Convergence – Manufacturing techniques for multi-rotors (3-8)

Initial estimation of Cost – Pre-PDR

Conceptual Payload CAD model – Using dimensions of 1 kg flour bag to propose potential

designs

Assisting on Test rig building

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101 Test Rig MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Mohammed Mohinuddin 24 Test Rig

This section comprises intensive testing and manufacture of a 3-axis gimbal test rig. As with any

project of this type, practical testing is vital to collaborate with the theoretical data determined for

the designed product. In general, testing during design of a new product is mandatory to reflect its

key performance factors and capabilities. It is most commonly known that manoeuvrability of a

multi-copter predominantly affects the flight performance. Therefore controlled flight during

manoeuvrability is a major factor to be achieved. In real world applications, achieving a stable

multi-copter during its flight regime is an issue of concern. Design of a test rig would aid in the

demonstration of multi-copter flight within a controlled environment and significantly add advantage

during the testing phase of this project. The test rig will be adopted to demonstrate safe control of

the UAV functions. Hence the design stage of the test rig began with rigorous brainstorming

activities bearing its effectiveness in mind.

To solely rely on systems to operate as efficiently as possible is not good practice, hence testing

the operation of individual system components and post integration would validate the testing

processes. The gimbal test rig would be a beneficial tool for the verification of sub-system tests in

controlled conditions. As part of the competition requirement, the chosen Quad-rotor design is

required to be able to carry two payloads (1kg each bags of flour) and deploy each payload

independently. This independent deployment of payloads at any given time could cause instability

post deployment and hence would affect the weight distribution on the UAS. The stability of the

multi-rotor after imbalance can be verified during testing within the gimbal design discussed later

on. It is anticipated that the test rig will aid to define PID control numbers which will hugely benefit

in the monitoring system and stability side of the project. Although the UAS design specification

does not require building a test rig, it was noted that fabricating a gimbal test rig would be

worthwhile as manoeuvrability and stability of a Quad-rotor is tremendously complex. Therefore a

safe testing method would have to be implemented to avoid damage on such a costly design. The

initial phase of testing using the gimbal test rig aided in the calibration of various sub systems such

as compass, magnetometer, Pixhawk, RC controller, GPS, etc.

24.1 Initial Conceptual Design of Gimbal Test Rig The initial brainstorming for designing a gimbal test rig began with a simplified design that would

establish the gyroscopic motion of the Quad-rotor. The general arrangement with bill of materials

(BOM) for initial conceptual design of the gimbal test rig assembly and further technical drawings

can be found in appendix O.1. However it was brought under notice that this design would be very

heavy and would occupy huge amount of space, therefore to overcome the consequences a more

robust and design specific based on actual Quad-rotor has been designed and is discussed

hereafter.

UAS CHALLENGE 2015


24.2 Octagonal Gimbal Test Rig

In order to overcome a few challenges encountered through initial test rig design mentioned above,

a more compact and robust design has been established as shown below in Figure 91. The

principle aim during redesign of the test rig was to reduce the overall space it would require for

storage and also the cost of manufacture. However it was figured out that the test rig would allow

the model to perform movement about all six degrees of freedom i.e. a similar approach like a

gyroscope. (Experimental Aircraft Info, 2006)

Figure 91 - Gyroscope Test Rigs

Majority of aircraft instruments use the basic principle of gyroscopes to control attitude, compass

and turn coordinates. In course of this project construction wise the gyro is fixed in the instrument

by octagonal rings or gimbals as shown in Figure 91and these rings give the gyro certain motions

of freedom. It is these motions or movement in a plane which allow for the features used in these

UAS CHALLENGE 2015


instruments. (Project, 2015) The gimbal test rig would utilise gyroscopic motion and is a device that

would be used to measure, maintain orientation and most importantly stabilise the Quad-rotor

under investigation.

Moreover the gimbal test rig’s operation is mainly based on the principle of preserving angular

momentum. The outer gimbal or ring (green coloured frame in Figure 91 which is the gyroscope

frame, is mounted so as to pivot about an axis in its own plane determined by the support from the

stand. This outer gimbal possesses one degree of rotational freedom and its axis possesses none.

The middle gimbal or ring (red coloured frame in Figure 91) is mounted to the outer gimbal so as to

pivot about an axis in its own plane that is always perpendicular to the pivotal axis of the gyroscope

frame (outer gimbal). This middle gimbal has two degrees of rotational freedom. The axle of the

spinning inner most gimbal (blue coloured frame in Figure 91) defines the spin axis. The motors

mounted on the Quad-rotor are coupled to spin about an axis, which is perpendicular to the axis of

the middle gimbal. Overall the entire gimbal test rig is meant to allow freedom of movement in all

yaw, roll and pitch axis. (Turner, 2015) See appendix O.2 for the general arrangement for the

updated octagonal gimbal test rig assembly with its bill of materials.

24.2.1 Octagonal Model Mount Frame

The figure alongside represents the general

arrangement for the model mount frame with

an inner spacing of 1000mm and shows the

necessary components used to construct the

frame. Refer appendix O.3 for detailed

technical drawing of the model mount frame.

The figure shown above validates the required inner length of each of the boxes used to construct

the model mount frame. An internet based octagon edge length calculator was used to obtain the

dimensions during the entire design.

UAS CHALLENGE 2015


Figure 92 - CAD Drawing of the Quad-rotor

The figure alongside shows the final

dimension of the Quad-rotor between

propeller tip to tip on either side of the arm.

The specific dimension worked out to be

995mm, which clearly justifies a minimal

clearance of 2.5mm between the inner side of

the model mount frame and the propeller tip.

Regardless of the 2.5mm clearance, the

propellers would be located slightly above the

frame level which is due to the mounting of

the Quad-rotor. Therefore it can be concluded

that a choice of 1000mm inner distance would

be sufficient enough to accommodate the

entire Quad-rotor.

24.2.2 Octagonal Mid Frame

The mid frame figure shown on the left had to

be constructed such that it would freely

accommodate the model mount frame, hence

it was observed that an inner distance of

1139mm would allow the model mount frame

to rotate and spin easily about its designated

axis. A clearance of 44.1mm on either side is

estimated between mid and outer frame.

Refer appendix O.4 for detailed technical

drawing of the mid frame.


the mid frame.

UAS CHALLENGE 2015


24.2.3 Octagonal Outer Frame

The outer frame figure shown alongside had to be

constructed such that it would freely accommodate

the mid frame, hence it was observed that an inner

distance of 1249mm would allow the mid frame to

rotate and spin easily about its designated axis.

Refer appendix O.5 for detailed technical drawing of

the outer frame.


the outer frame.

24.3 Weight Estimation for Octagonal Test Rig

The spreadsheet in appendix O.7 represents the weight estimation for the entire gimbal test rig

conducted analytically, through CATIA estimation and from supplier data sheet. The entire test rig

frames and stand will be fabricated using 1”x1” aluminium box sections with the brackets

manufactured from 1.2mm aluminium sheet.

24.4 Cost Breakdown for Octagonal Test Rig

The overall cost for manufacturing the gimbal test rig was estimated to be £132.08 inc. VAT and

based on sourcing materials from one supplier named Metals4U. However the cost incurred for

materials purchase was raised through collective funding. The spreadsheet detailing the costing

can be found in appendix O.8.

UAS CHALLENGE 2015


24.5 Manufacturing Stage of the Octagonal Test Rig

The entire test rig was fabricated using facilities provided at the university and tools brought from

colleagues. Conversely the initial design for the joint bracket was rectified to produce a much

simpler design to manufacture reducing the costs of water jet cutting. The figures shown below

illustrate the fabricated parts and the entire gimbal test rig assembly. The final test rig was

fabricated under the assistance of almost all the members in the group.

Figure 93 - Test Rig Components

Figure 94 - Test Rig Assembly

UAS CHALLENGE 2015

107 Structural Testing MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Mohammed Mohinuddin 25 Structural Testing

In addition to stability checks on the test rig, other tests such as static material tests, impact/ crash

tests are proposed to be conducted. The selected material used as the base plates were tested to

validate its bending capabilities withholding its structural integrity. However further tests were also

proposed to be conducted to improve the performance of the Quad-rotor and to meet the

conformance.

25.1 Material Testing

It was collectively decided in the group that the material to be used for manufacturing the base

plates would be nylon 6. The figures in this section

represent the exact material used Figure 95 and the

manufactured base plate after cut-outs Figure 95

Figure 95 - Nylon Material and Main Body Plate

A compression test, using the Hounsfield 1kN Tensometer, was carried out on the base plate

material as seen in Figure 96 This test simulates the dominant load type experienced by the plates.

It was observed through this analysis that the plate would survive tremendous load and would not

deform permanently which is a justification to the stress analysis carried out in the structural

loading and analysis chapter. The nylon plate sample was tested in two different orientations and

provided reasonable understanding in the plate bending behaviour.

25.2 Component Testing

On receipt of various components such as motors, pixhawk, telemetry kit, GPS and servos, each

component was individually quality

checked and tested for conformance by

the relevant personnel. The motors were

tested to check the amount of current

drawn and to reflect their performance.

The other components were also tested

to check whether they would perform the

tasks they were purchased for.

Figure 96 - Compression Test conducted on Hounsfield Tensometer

UAS CHALLENGE 2015

108 Structural Testing MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Osman Sibanda

25.3 Payload Drop Testing

It is anticipated that when an object hits the ground with a speed of 2-5 m/s, it would not cause any

substantial damage to the object withholding its structural integrity. Conversely a simple drop test

was used to replicate that the payload would remain intact after impact. To create the same

amount of energy dissipation in the test as there will be at full load the following calculations were

used. A trial drop test with 1Kg bag of flour was conducted and the following schematic represents

this.

Considering the conservation of energy, the potential energy possessed by the bag of flour will be

converted to kinetic energy on impact neglecting air resistance and heat. The following calculations

denote the possible results to be anticipated.

mgh = 0.5mv2

1x9.81x0.98 = 0.5x1xv2, therefore v= 4.4m/s

Hence it can be concluded through above calculation that the payload remains intact and free from

any substantial damage.

25.4 Initial Ball socket test rig Initially the group had intended to use two test rigs for testing and calibrating the systems in the

multi-copter prior to flying it. The idea of using test rigs was dismissed as the gimbal test rig was

enough to carry out all the testing and calibration

needed. Figure 97 below shows the proposed ball socket

test rig at that stage.

Figure 97 - Initial ball socket test rig

UAS CHALLENGE 2015

109 Business Case MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Osman Sibanda

25.5 Manufacturing assistance Due to the amount of work and time needed for manufacturing the Quad-rotor components and

test rig, I was assigned to assist the manufacturing engineer and the testing engineer with the

manufacturing of the UAV and test rig. Refer to manufacturing report for details.

26 Business Case The main part of this report is the business case of our UAS with accordance to the iMechE

requirements. The rest of this report will focus on the business case of our Quad-rotor the

Odonata-XV. Our company name is Autoquads Inspection Ltd.

Figure 98 - Autoquads Inspection Ltd Logo

26.1 Executive Summary Inspections of structures are paramount for the safety of the users and the public on all

infrastructures. Some may have long intervals of inspections but some critical infrastructures

require regular inspections depending on the criticality. The current methods of inspections have

proved expensive and very risky for the inspectors that carry them out.

Autoquads Inspections Ltd proposes using our UAS, the Odonata XV to carry out inspections

autonomously for wind turbines, bridges, rail lines and overhead power lines in the UK with a future

plan of expanding to Europe. The ability for small UAS’s to manoeuvre in confined spaces, hover,

fly at low speeds and altitudes and to perform various manoeuvres at any given time makes them

ideal for inspections tasks. They can provide images in real time and also obtain high resolution

images which can be recorded at the same time for reviewing later on.

The business case will cover the key design features of the UAS and how these features can be

used to our advantage for different inspection purposes. The market research is covered in section

6 of this report and highlights the potential markets that could have been chosen. Also the market

size and predicted market growth by industry experts, regulations involved with this UAS category

and it also covers some of the competition that exists in the market already. The financial forecast

of this project can be found on chapter 26.6. It covers the key assumptions made in calculating the

related costs of this project and how long it would take for the company to break even.

UAS CHALLENGE 2015


26.2 Business overview Inspections of hard to reach areas (etc. power lines, bridges, wind turbines, roofs) using the

traditional methods has proved to be time consuming, expensive and dangerous in most situations

for the inspectors. Structures exposed to the environmental elements will suffer deterioration over

time and therefore require regular inspection to avoid any serious complications occurring and

interrupting the services being provided.

For power lines the current methods being used are by using helicopters, manually controlled

UAV’s or by line inspectors. The line inspectors usually work from the ground using binoculars and

are sometimes required to climb up the power lines to manually inspect them. This causes major

risks on the line inspectors if the power lines are live and this method is not very reliable. For

example the Killmore bushfire that occurred in Victoria, Australia as a result of poor inspection by

one of the inspectors caused casualties of around 119 people in 2009 (ABC NEWS, 2014). For the

business case of our UAS, Autoquads Inspection Ltd is hoping to enter into the inspection market

which will consist of different infrastructural inspections, which will include overhead power line

inspection, train paths, bridges and wind turbines. The UAS was designed to be able to carry a 1kg

payload and be able to drop it at designated targets autonomously. For the UAS to be used for

inspections, some minor changes will have to made in order to carry our inspections autonomously

this will include adding electro-optic and thermal imaging camera in order to inspect insulations on

power lines.

26.3 Mission statement Currently the inspection market uses manually controlled crafts (helicopters) or linemen on foot.

The existing usage of helicopters with crewmen using binoculars for inspections in some cases is

much more expensive and it’s unlikely to get any cheaper unless new technologies are introduced.

The use of helicopters has also caused a lot of damage to landowners around the power lines

especially to the farmers and their domestic animals. Some farmers have been compensated large

sums of money because of disturbances to the livestock by premature births or because of

stampedes leading to property damage (European Commission, 2014) .The use of linemen on the

ground is time consuming, very risky and very expensive since well qualified personnel have to do

the inspections and are usually paid by the hour. An introduction of autonomous UAS should

significantly cut down costs and at the same time cut down on the time taken to inspect structures.

Autonomous inspections will;

Ensure efficient use of experts by using them only when they are needed

Allow for targeted maintenance by prioritizing maintenance where it’s needed the most

therefore cutting down shutdown times. Longer shutdown of services leads to dissatisfied

customers/clients therefore a possibility for the companies to lose out on competitive

advantage

Reduce labour costs by having less staff to monitor the UAS during inspections

UAS CHALLENGE 2015


There is less risk involved with the crew and the surrounding environment including animals

Lesser environmental disturbances etc. noise pollution

A costs analysis carried out by Europa that using UAS’s for inspections compared to the other

methods can significantly reduce the costs involved by about a third. (European Commission,

2014)

26.4 UAS key design features The OXV has been designed to be collapsible by allowing 4 of the arms and landing gear to

rotate about pivots. This allows the OXV to fold into linear and compact configuration allowing

for easy transportation and storage as shown in Appendix B.7. The payload housing is

designed to be removable for ease of transport, increased functionality and the ability to attach

various devices as a payload or for this business case an electro optic and thermal imaging

camera would added on for high resolution images and for insulation inspection. The design of

the OXV was a well thought out process in consideration of the maintenance of the UAS. The

UAS is designed with quick release pins to allow the arms of the quad to fold up and also

allows the UAV to be stored in tighter spaces.

The UAS can also be set up for a perch and stare function. This function will allow the UAS to

land somewhere for an extended time for observation and re-launch itself after observing.

26.5 Market Assessment The growing interest/use of autonomous flying has allowed for opportunities in the inspection

sector. Inspections of structures is a growing market with the introduction of UAS’s as it is vital

for inspections of damaged components/parts, insulations, deteriorating parts and overhanging

trees. Rather than using manually controlled drones or using helicopters, the use of UAS’s has

a promise to be more efficient, less costly and less dangerous in this sector. Autoquads

Inspection Ltd aims to provide the best quality of service to its customers. This UAS will greatly

benefit entities that provide Enterprise Asset Management (EAM) for bridges, power lines, train

paths and wind turbines.

26.5.1 Potential market – Emergency Service

Sudden Cardiac Arrest (SCA) - is a condition where the heart suddenly stops beating and

blood to the brain and other organs stops flowing. Abnormal heart beat rhythms are called

arrhythmias; this is when a heart beats too fast, too slow or at irregular rhythms. 95% of people

who suffer Sudden Cardiac Arrest (SCA) die within minutes and for every 1 minute a persona

suffers cardiac arrest their chance of survival is decreased by 10%. (NIH, 2011)

Cardiopulmonary resuscitation (CPR) can be performed on individuals experiencing SCA but

they are not as effective as AED’s. Automatic external defibrillators (AED) are lightweight,

battery operated; portable devices used to measure heart rhythm and can send electric shocks

to restore the heart to normal rhythm. AED’s are very easy to use and also come with

UAS CHALLENGE 2015


instructions as well as voice commands to alert the user when to send electric shocks. Using

our UAS AED’s can be transported to the patients quickly to deliver the services before the

medics arrive to the scene.

Product Size (H x W x D) (cm) Weight (kg) Average Cost (£)

Zoll AED plus 13.3 x 24.1 x 29.2 3.1 1194

Phillips Heartstart Onsite 19 x 21 x 7 1.5 1199

HeartSine Samaritan PAD 350

AED

20 x 18.4 x 4.8 1.1 kg 1175

Defibtech Lifeline AED 22 x 30 x 7 2 1245

Table 16 Table showing potential AED's

Table 16 shows the potential AED products with the dimensions, weights and the average

costs. This option was not chosen for various reasons including the prices of the AED’s

themselves which are very expensive. The other reason was the time it would take to reach the

patients, the Odonata XV would have been able to reach some patients a little time before the

emergency staff get there but in most cases the emergency staff are improving their response

time therefore this market would not have proved to be profitable over time.

26.5.2 Market size and growth

The emerging UAV technology is to become key in the future competitiveness of the European

aerospace industry compared to other parts of the world. According to (European Commission,

2014) the common European market will offer a solid base to compete globally with other

leading competitors in the world e.g. USA, Israel, Brazil, China and Russia. It is predicted that if

an enabling legal framework is adopted it will furthermore allow the operations and the

manufacturing of the UAV’s to grow from simple operations to more complicated operations

thus allowing the current businesses to gain valuable practical expertise while developing their

businesses. For example in France the number of approved operators rose from 86 to 400

after the introduction of an initial regulation.UK and Sweden has also seen similar growths in

different markets because of an enabling regulation (QinetiQ, 2013).

Industry experts believe it is really difficult to predict the potential UAS’s have globally but it is

currently predicted to be worth about $5.2 billion and it is expected to grow to about $11.6

billion per year in 2023 (QinetiQ, 2013). Further to boosting businesses across Europe the UAV

market is set to increase jobs globally as well. In Europe around 150,000 jobs are forecasted

by 2050 (European Commission, 2014). According to the Scottish Enterprise, the UK has the

most ambitious project for wind farm plans. This therefore means more market for inspection

since operations and maintenance accounts for over a quarter of the lifetime cost of a wind

farm. Analysis by UK Government predicted that the number of wind turbines will increase to

over 5500 by year 2025 with the operations and maintenance claiming £2bn per year from this

business boost.

UAS CHALLENGE 2015


26.5.3 Regulation restriction

Since this UAS is less than 7Kg MTOM, they will be categorised as a SUA (Small Unmanned

Aircraft) category and should comply with UK Air Navigation Order 2009 articles 138, 166, 167

and CAA CAP 722, and CAP 393. (CAA, 1995)

The following chapter summarizes the UK legal requirements for flying UAV’s in the UK

(Austin, 2010):

The craft should not endanger anyone or anything, including the pilot of the UAV. The pilot

holds the responsibility for the operations to be conducted safely.

The UAV must be in VLOS (visual line of sight taken to be at 122m vertically and 500m

horizontally) of the pilot at all times. For any uses beyond these distances, the pilot must

seek CAA permission and prove the craft can be flown safely at that distance.

CAA permission is required for any aerial work

Should not be flown within restricted airspace

The craft should not be flown;

Above or around 150m of any congested area

Above or around 150m of an assembly of more 1000 people

Around 50m of any vessel, vehicle or structure which is not under the pilot’s control

Within 50m of any person during take-off or landing and within 30m of any persons

during flight except for the pilot

Figure 99 - Permissions required for different UAS sizes

Figure 99 (Austin, 2010) shows the permissions required from the CAA for the different sized

aircrafts. In our case the aircraft is under the 20kg limit therefore the registration and the

airworthiness approval is not required but an operating permission and pilot qualification would

be required.

26.5.4 Challenges for market entry

A lot of factors will heavily affect entering the inspection sector. The biggest challenge would be

the initial capital that would have to be invested in mass production of the Odonata XV. The

subsequent chapters give some details into technological challenges that the organisation

could face;

Safe operation

EU aviation policy defines safety as the paramount objective. The current regulatory system for

UAS based on fragmented rules for ad hoc operational authorizations is an administrative

UAS CHALLENGE 2015


bottleneck and hampers the development of the European UAS market (Austin, 2010). Either

to produce or to operate UAS national authorizations do not benefit from mutual recognition

and do not allow for European wide activities. The principle that safety will not be compromised

is hugely followed by the European aviation system for the integration of UAS. UAS operations

must demonstrate an equivalent level of safety in contrast to manned aviation.

Security

UAS is not resistant to probable unlawful actions. The potential uses of the UAS could be for

military purposes, the navigation or communication system signals of other UAS could be

jammed or ground control stations hijacked. Any identified security requirements needs to be

translated into legal obligations for all relevant players, such as the air navigation service

provider, UAS operator or telecom service provider, under the oversight of the competent

authorities. (Austin, 2010)

Data protection

Fundamental rights must not be trespassed by the UAS operations, including the respect for

the right to family and private life, and the production of personal data. Amongst the wide range

of potential civil UAS applications a number may involve collection of personal data and raise

ethical, privacy or data protection concerns, in particular in the area of the surveillance,

monitoring mapping or video recording.

UAS operators would need to comply with the applicable data protection provisions, notably

those sets out in the national measures established pursuant to the sat protection Directive

95/46/EC and the Framework Decision 2008/977. (Austin, 2010)

26.5.5 Competition

Knowing your competition is a crucial process of a successful product, therefore this chapter

will look into present competitors’ strengths and weaknesses and how the Odanata XV can

gain a competitive advantage.

Existing competitors

Product: Aibot X6 UAS

Aibot uses the X6 UAS to carry out inspections on power lines, wind turbines, bridges, train

paths and oil and gas pipelines. Our company will respond to Aibot by having a much smaller

UAV and providing the service much cheaper than them. (Aibotix, 2015)

Product: Md4-200

Micro drones uses the Md4-200to carry out inspections on oil pipelines, power cables, cooling

towers, forestry, radiation and wind turbines . The Md-200UAS has a very good flight time of

about 30 minutes but our company will respond to MicroDrones by having a heavier payload

capability since the Falcon can only carry 200g. (Micro Drones, 2015)

Product: Asctec Falcon 8

Asctec uses the Asctec Falcon 8 to carry out inspections solar parks, offshore/onshore

turbines, structural integrity and wind parks. Autoquads inspection Ltd will respond to this

UAS CHALLENGE 2015


competitor by having thermal imaging cameras and by providing our service cheaper because

of the cost of the Odonata XV. (AscTec, 2015)

26.6 Financial Forecasts The financial forecast section will aim to estimate the predictable cost of this project for the next

5 years. The section will cover the assumptions to be made in order to predict some costs,

thee pro-forma income statement, cash flow statement, mass manufacturing costs for the

whole UAS and break-even analysis to predict how long it will take until the company makes

enough money to break even.

26.7 Key assumptions In order to make financial forecasts possible to calculate, a number of assumptions had to be

made. For the company to stay privatised the capital will come from a combination of

sponsorship and bank loans. Assuming 20 Quad-rotors and 1 controller in charge of 4

computers therefore only 5 controllers are needed. These controllers will also be in charge of

system maintenance. Inflation will be assumed to be 3% in the second year and will rise by 1%

yearly. Currently 500MW wind farms cost £50,000 to £100,000 averagely a year, we will

assume the cost to be £50,000 as we’re using a cheaper method (UAS) and to attract

customers (National Grid, 2013).The price of power line inspections is £15.46 per mile

(Network Rail, 2015).

Table 17 below shows the amount of wind farms, overhead power lines, bridges and rail lines

in the UK and how much of the market share Autoquads hopes to take over yearly. Autoquads

Inspection Ltd will aim to acquire 2% of the market share the first year and rising by 2% yearly

onwards except for the windfarms. The windfarms will take time to inspect because each

windfarm will contain several wind turbines that need inspecting, therefore for windfarms an

assumption of 20 (0.5% of market) windfarms for the first year and increasing by 30 yearly.

Amount 2% share

1st

year

4% share

2nd

year

6% share

3rd

year

8% share

4th

year

10% share

5th

year

Cost

(£)

Wind farms 4338 (+3000) 20 50 80 110 140 50000

Overhead power lines

4470 miles (National Grid) 89 178 268 357 447

50

Rail 9788 miles 195 391 587 783 978

200

Bridges 1000 20 40 60 80 100 1000

Cost 1063450 2627100 4190800 5754450 7317950

Table 17 UK market estimation and AIL market share

26.8 Costs Autoquads Inspection Ltd hopes to start with 5 crafts for each infrastructure inspection (20 in

total) to start with for the first year and use them to find efficient ways to undertake tasks. This

plan of introducing a few Quad-rotors for the first year will ensure the company establishes

itself and gets used to the way of working in this sector. This period will also ensure that

UAS CHALLENGE 2015


whatever challenges that are encountered can be solved and recorded down for future

reference. After the first year the company hopes to introduce 20 more Quad-rotors yearly for

the next four years, this will largely depend on the business growth so when the time comes it

could be more or less than planned. The table below shows the expected costs for the first 5

years;

Fixed Costs Cost (£) Quantity Total (£)

Manufacturing (materials) 1500 20 30000

Machines and tools 10000 1 10000

Infrastructure Equipment (Ground station.…) 50,000 1 50000

Marketing campaign 10000 1 10000

Administrative costs 30000 1 30000

Infrastructure 3000000 1 3000000

Total 3130000

Table 18 Fixed Cost

Running Costs Monthly Cost (£) Yearly Cost (£) 5 year cost

Building Repairs/maintenance 1000 12000 60000

Utility Bills 3000 36000 180000

Controller/Inspectors/Monitors 10000 120000 600000

Maintenance costs 1000 12000 60000

Labour 7000 84000 420000

Training 2000 24000 120000

Miscellaneous (transportation, 1000 12000 60000

Total 25000 300000 1500000

Table 19 Running Costs

Year Cost (£) Addition of new

crafts (£)

Extra Labour

(£)

Inflation Total (£) Information

1 3430000 0

0 3430000 Fixed, infrastructure

2 300000 30000 120000 3% 463500 Pay, maintenance

3 300000 30000 120000 4% 468000 Pay, maintenance

4 300000 30000 120000 5% 472500 Pay, maintenance

5 300000 30000 120000 6% 477000 Pay, maintenance

Total

5311000

Table 20 Yearly Costs

UAS CHALLENGE 2015


26.8.1 Financial statements

Pro-forma statement

The pro-forma statement table below shows the costs and revenues anticipated for the next 5

years.

Cost/Revenue Year 1 (£) Year 2 (£) Year 3 (£) Year 4 (£) Year 5 (£)

Services revenue 1063450 2627100 4190800 5754450 7317950

Variable cost of services provided 120000 240000 480000 960000 1920000

Fixed cost of services 12000 12000 12000 12000 12000

Gross margin 931450 2375100 3698800 4782450 5385950

Variable operating costs 3130000 270000 510000 990000 1950000

Fixed operating costs 300000 300000 300000 300000 300000

Untaxed income -2498550 1805100 2888800 3492450 3135950

Income tax (40%) - 722040 1155520 1396980 1254380

Net Income -2498550 1083060 1733280 2095470 1881570

Table 21 Pro-forma statement

Cash Flow

The cash flow statement table shows the key costs in and out. The key values in the cash flow

statement are the beginning cash balance and the ending cash balance.

Item Year 1 (£) Year 2 (£) Year 3 (£) Year 4 (£) Year 5 (£)

Beginning Cash balance

0 -2504796.375 -1424444.025 304502.775 2394734.1

Net income after tax -2498550 1083060 1733280 2095470 1881570

Depreciation expense 6246.375 2707.65 4333.2 5238.675 4703.925

Ending Cash balance -2504796.375 -1424444.025 304502.775 2394734.1 4271600.175

Table 22 Cash flow statement

Break even

As shown from the cash flow statement Auto quads Inspection Ltd starts off with nothing (not

including sponsorship and loans) and the ending cash balance is a negative £2.5m, for the

second year the ending cash balance is still negative meaning the company hasn’t started

making profit yet but the value is decreasing. By the end of the third year the ending cash

balance is £304000, this is the break-even point. It will be at this point that the company will

start making profit in this sector, so it will take 3 years for the company to break even.

UAS CHALLENGE 2015


26.8.2 Profitability

Figure 100 - Break Even Graph

As the graph above shows, the profitability of this business looks promising if everything goes

according to plan. Although the capital is a huge amount it should pay off at the end of the third

year with a profit of £300000 after the capital is paid. Beyond the third year the profit will be

expected to rise steadily for a couple of years then the rise will depend on the market growth at

that time. For the time being though the market looks promising so therefore it can be concluded

that this project will be very profitable.

Business risk assessment

The business risk assessment can be found in Appendix N.1.

-£2.50 -£1.42

£0.30

£2.39

£4.27

-£3

-£2

-£1

£0

£1

£2

£3

£4

£5

1 2 3 4 5

Pro

fit

Mill

ion

s

Years

UAS CHALLENGE 2015

119 Conclusion MEng Team Project Report (7ENT1024) School of Engineering and Technology

Conclusion As part of the University’s academic curriculum, engineering projects at Masters Level have both a

documentation aspect (Dissertation or Report) and a physical aspect (product development and

testing); both aspects of the UAS Challenge project were carried out successfully under the

guidance of supervisors and technicians. The UAS challenge is a 1st generation IMechE

competition for Unmanned Aircraft Systems making this project the first of its kind in the University

of Hertfordshire. The UAS was designed and built to have real-world applications and this was

quantified through numerous validation and verification tests as well as quality control processes.

The UAS challenge project was carried out by a team of 12 aerospace engineers who worked

industriously in order to meet the project deadlines and objectives over a course of seven months

to deliver a top-tier product. In order to successfully deliver this project, product development

processes were integrated into the project phases through the creation of the design specification

which was used to keep engineering design process in line with rules from CAA and IMechE.

The management of the project was very professional as the meetings were held with supervisors

to discuss potential project pitfalls and solutions. The project manager also created a project plan

and budget plan to keep the project on schedule and on budget. Project management processes

such as QFD and WBS was used in the requirement analysis carried out on the design

specification in order to determine the right aerial vehicle to use for meeting the requirements.

Frequent requirement analysis were carried out in order to make sure that the product being

development meets the requirements and the management model used is a form of the “V” model.

An example is the change from a quadcopter to a hexacopter after the results of a structural

analysis showed that the weight limit set by IMechE would be exceeded. Another example is the

switch from Arducopter Autopilot System to Pixhawk Autopilot System after critical analysis

showed that processing capacity, safeguard measures and competency of the Pixhawk Autopilot

System was significantly higher than the Pixhawk Autopilot System.

The technical approach to the project was very professional and conservative as every process is

documented properly for analysis and this is shown in the level of testing carried out on electronic

components and the structural analysis of the UAS materials and components (Finite Element

Analysis and Bending Tests). The selection of materials used for manufacture was done after

meticulously analysing different materials and comparing them in areas such as strength and price,

this resulted in the manufacture of a structure below the weight limit. Catia and Ansys were used to

design the UAV structural components and Mission Planner software was used to program the

Autopilot System and every other programmable component. A control system operating manual

was created to enable non-system group members would be able to use the UAS control systems.

UAS CHALLENGE 2015

120 Conclusion MEng Team Project Report (7ENT1024) School of Engineering and Technology

Safety regulations were followed in the course of the project such as following all laboratory

regulations when using the facilities for manufacturing, assembly and testing and also designing

the UAS according to regulations set by IMechE and CAA. Failsafe have been programmed into

the UAS for safe operation and recovery when failsafe conditions such as loss of GPS, loss of

communication are activated.

The project’s objective has been met as the UAV is below the weight limit set by the IMechE and

can also be certified by the CAA. The UAS is also capable of a number of flight modes such as

autonomous flight, semi-manual and manual flight. The project budget was also not exceeded and

the UAS has been built and tested a couple of times. This project would serve as a foundation and

legacy to future generations of aerospace engineers that would partake in the UAS challenge from

University of Hertfordshire in the hope of reaching and surpassing the levels reached in the course

of this project.

UAS CHALLENGE 2015

121 REFERENCES MEng Team Project Report (7ENT1024) School of Engineering and Technology

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125 MEng Team Project Report (7ENT1024) School of Engineering and Technology

Appendix. A

A.1. Initial Project Plan

[PAGE INTENTIONALLY LEFT BLANK]

ID Task Name Duration Start Finish

0 UAS Challenge Project Plan 191 days Fri 10/10/14 Sun 05/07/151 Scope 15 days Fri 10/10/14 Thu 30/10/142 Determine project scope 3 days Fri 10/10/14 Tue 14/10/143 Define resources 12 days Wed 15/10/14 Thu 30/10/144 Scope complete 0 days Thu 30/10/14 Thu 30/10/145 Sponsorship 97 days Fri 10/10/14 Sun 22/02/156 Secure project sponsorship 97 days Fri 10/10/14 Sun 22/02/157 Design Specification/System

Requirements12 days Fri 31/10/14 Sun 16/11/14

8 Create Design specification for a UAV 7 days Fri 31/10/14 Sun 09/11/14

9 Review system specifications 7 days Sun 09/11/14 Sun 16/11/1410 Create system requirements 12 days Fri 31/10/14 Sun 16/11/1411 Obtain approvals to proceed (concept,

timeline, budget)12 days Fri 31/10/14 Sun 16/11/14

12 Analysis complete 0 days Sun 16/11/14 Sun 16/11/1413 Preliminary Design 16 days Sun 16/11/14 Fri 05/12/1414 Review specifications 15 days Sun 16/11/14 Thu 04/12/1415 Payload Delivery System 15 days Sun 16/11/14 Thu 04/12/1416 Propulsion System design 15 days Sun 16/11/14 Thu 04/12/1417 Systems design 15 days Sun 16/11/14 Thu 04/12/1418 Concept Structural design 15 days Sun 16/11/14 Thu 04/12/1419 Preliminary Safety Case consideration 15 days Sun 16/11/14 Thu 04/12/1420 Preliminary Weights estimation 15 days Sun 16/11/14 Thu 04/12/1421 Obtain approval to proceed 15 days Sun 16/11/14 Thu 04/12/1422 Preliminary Design complete 0 days Thu 04/12/14 Thu 04/12/1423 Deliver PDR to IMeche 0 days Fri 05/12/14 Fri 05/12/1424 Final Design ready for purchase 8 days Fri 05/12/14 Tue 16/12/1425 System compents finalised ready for

purchase 7 days Fri 05/12/14 Mon 15/12/14

26 Propulsion components ready for purchase

7 days Fri 05/12/14 Mon 15/12/14

27 Structrual material and sizing ready forpurchase

7 days Fri 05/12/14 Mon 15/12/14

28 Design purchase readyness 0 days Tue 16/12/14 Tue 16/12/1429 Order parts 30 days Tue 16/12/14 Mon 26/01/1530 Send out order list for components and

delivery30 days Tue 16/12/14 Mon 26/01/15

31 Manufacturing & Assembly 47 days Mon 26/01/15 Tue 31/03/1532 Machine structural frame 26 days Mon 26/01/15 Sat 28/02/1533 Integrate systems components 26 days Mon 26/01/15 Sat 28/02/1534 Integrate structural frame, system and

propulsion components 23 days Sun 01/03/15 Tue 31/03/15

35 Testing and Validation 47 days Mon 26/01/15 Tue 31/03/1536 Develop unit test plans using design

specifications47 days Mon 26/01/15 Tue 31/03/15

37 Develop integration test plans using design specifications

47 days Mon 26/01/15 Tue 31/03/15

38 Integration Testing 67 days Sun 01/03/15 Sun 31/05/1539 Test system integration 23 days Sun 01/03/15 Tue 31/03/1540 Integration testing complete 0 days Sun 31/05/15 Sun 31/05/1541 Critical Design Review (CDR) and

Flight Readiness Review (FRR)40 days Mon 09/03/15 Fri 01/05/15

42 Draft CDR report 0 days Mon 09/03/15 Mon 09/03/1543 Draft FRR report 0 days Mon 09/03/15 Mon 09/03/1544 Deliver CDR report 0 days Mon 06/04/15 Mon 06/04/1545 Deliver FRR report 0 days Fri 01/05/15 Fri 01/05/1546 Pre-Competition 26 days Mon 01/06/15 Sun 05/07/1547 Design Presentation 6 days Mon 01/06/15 Sun 07/06/1548 Flight Readiness Review 6 days Mon 08/06/15 Sun 14/06/1549 Certification Test Flight 11 days Mon 15/06/15 Mon 29/06/1550 Competition day 3 days Wed 01/07/15 Fri 03/07/15

Scope complete 30/10Scope complete

Analysis complete 16/11

Preliminary Design complete 04/12

Deliver PDR to IMeche 05/12

Design purchase readyness 16/12

Draft CDR report 09/03

Draft FRR report 09/03

Deliver CDR report 06/04

Deliver FRR report 01/05

T W T F S S M T W T F S S M T W T F S S M T W T F S S M T W T F S S M T W T F25 Aug '14 15 Sep '14 06 Oct '14 27 Oct '14 17 Nov '14 08 Dec '14 29 Dec '14 19 Jan '15 09 Feb '15 02 Mar '15 23 Mar '15 13 Apr '15 04 May '15 25 May '15 15 Jun '15 06 Jul '15

Task

Split

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Summary

Project Summary

External Tasks

External Milestone

Inactive Task

Inactive Milestone

Inactive Summary

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Duration-only

Manual Summary Rollup

Manual Summary

Start-only

Finish-only

Deadline

Progress

Page 1

Project: UAS Challenge Project PlaDate: Wed 03/12/14

UAS CHALLENGE 2015


A.2. Up to date Project plan


ID Task Mode

Task Name Duration Start Finish

1 Scope 15 days Fri 10/10/14 Thu 30/10/142 Determine project scope 3 days Fri 10/10/14 Tue 14/10/143 Define resources 12 days Wed 15/10/14 Thu 30/10/144 Scope complete 0 days Thu 30/10/14 Thu 30/10/145 Sponsorship 97 days Fri 10/10/14 Sun 22/02/156 Secure project sponsorship 97 days Fri 10/10/14 Sun 22/02/157 Design Specification/System

Requirements12 days Fri 31/10/14 Sun 16/11/14

8 Create Design specification for a UAV7 days Fri 31/10/14 Sun 09/11/149 Review system specifications 7 days Sun 09/11/14 Sun 16/11/14

10 Create system requirements 12 days Fri 31/10/14 Sun 16/11/1411 Obtain approvals to proceed

(concept, timeline, budget)12 days Fri 31/10/14 Sun 16/11/14

12 Analysis complete 0 days Sun 16/11/14 Sun 16/11/1413 Preliminary Design 16 days Sun 16/11/14 Fri 05/12/1414 Review specifications 15 days Sun 16/11/14 Thu 04/12/1415 Payload Delivery System 15 days Sun 16/11/14 Thu 04/12/1416 Propulsion System design 15 days Sun 16/11/14 Thu 04/12/1417 Systems design 15 days Sun 16/11/14 Thu 04/12/1418 Concept Structural design 15 days Sun 16/11/14 Thu 04/12/1419 Preliminary Safety Case consideration15 days Sun 16/11/14 Thu 04/12/1420 Preliminary Weights estimation15 days Sun 16/11/14 Thu 04/12/1421 Obtain approval to proceed 15 days Sun 16/11/14 Thu 04/12/1422 Preliminary Design complete 0 days Thu 04/12/14 Thu 04/12/1423 Deliver PDR to IMeche 0 days Fri 05/12/14 Fri 05/12/1424 Final Design ready for purchase8 days Fri 05/12/14 Tue 16/12/1425 System compents finalised

ready for purchase 7 days Fri 05/12/14 Mon 15/12/14

26 Propulsion components readyfor purchase

7 days Fri 05/12/14 Mon 15/12/14

27 Structrual material and sizing ready for purchase

7 days Fri 05/12/14 Mon 15/12/14

28 Design purchase readyness 0 days Tue 16/12/14 Tue 16/12/1429 Order parts 30 days Tue 16/12/14 Mon 26/01/1530 Send out order list for

components and delivery55 days Tue 16/12/14 Mon 02/03/15

31 Manufacturing & Assembly 47 days Mon 26/01/15 Tue 31/03/1532 Machine structural frame 26 days Mon 09/03/15 Mon 13/04/1533 Integrate systems

components 26 days Mon

09/03/15Mon 13/04/15

34 Integrate structural frame, system and propulsion components

23 days Mon09/03/15

Wed 08/04/15

35 Testing and Validation 47 days Mon 26/01/15 Tue 31/03/1536 Develop unit test plans using

design specifications37 days Mon

26/01/15Tue 17/03/15

37 Develop integration test plans using design specifications

37 days Mon26/01/15

Tue 17/03/15

38 Integration Testing 37 days Sun 01/03/15 Mon 20/04/1539 Test system integration 23 days Mon 09/03/15 Wed 08/04/1540 Integration testing complete 0 days Mon 20/04/15 Mon 20/04/1541 configure PID for Quad 67 days Sun 01/03/15 Mon 01/06/1542 Critical Design Review (CDR)

and Flight Readiness Review (FRR)

71 days Mon09/03/15

Mon 15/06/15

43 Draft CDR report 0 days Mon 09/03/15 Mon 09/03/1544 Deliver CDR report 0 days Wed 01/04/15 Wed 01/04/1545 Draft FRR report 11 days Mon 18/05/15 Sun 31/05/1546 Deliver FRR report 0 days Mon 15/06/15 Mon 15/06/1547 Pre-Competition 15 days Mon 15/06/15 Fri 03/07/1548 Design Presentation 0 days Wed 01/07/15 Wed 01/07/1549 Flight Readiness Review 0 days Wed 01/07/15 Wed 01/07/1550 Certification Test Flight 11 days Mon 15/06/15 Mon 29/06/1551 Competition day 3 days Tue 30/06/15 Thu 02/07/1552 UAS CHALLENGE FINISH 0 days Fri 03/07/15 Fri 03/07/15

0%

100%100%

100%

Scope complete 0%Scope complete

0%

0%

100%

100%100%

100%100%

100%

100%

Analysis complete 100%

100%

100%

100%

100%

100%

100%

100%

100%

100%

Preliminary Design complete 100%

Deliver PDR to IMeche 100%

100%

100%

100%

100%

Design purchase readyness 100%

100%

100%100%

0%

0%

0%

0%

0%

50%

50%

0%

0%

Integration testing complete 75%

0%

Draft CDR report 0%

Deliver CDR report 0%

15/06

0%

Design Presentation 01/07

Flight Readiness Review 01/07

0%

UAS CHALLENGE FINISH 03/07

15 22 29 06 13 20 27 03 10 17 24 01 08 15 22 29 05 12 19 26 02 09 16 23 02 09 16 23 30 06 13 20 27 04 11 18 25 01 08 15 22 29 06 13Sep '14 Oct '14 Nov '14 Dec '14 Jan '15 Feb '15 Mar '15 Apr '15 May '15 Jun '15 Jul '15

Task

Split

Milestone

Summary

Project Summary

External Tasks

External Milestone

Inactive Task

Inactive Milestone

Inactive Summary

Manual Task

Duration-only

Manual Summary Rollup

Manual Summary

Start-only

Finish-only

Deadline

Progress

Page 1

Project: Updated UAS Challenge PDate: Thu 23/04/15

UAS CHALLENGE 2015


A.3. Example of Minutes


Supervisors meeting |MINUTES

Meeting date | time 10/28/2014 12:00 AM | Meeting location Sim Laboratory

Meeting called by Alfred

Type of meeting Progress check

Note taker Johnathan

Timekeeper Zuber

Attendees

Alfred, Mohin, Zuber, Tarek, Johnathan, Micky, Reyad,

Kasun, Osman, Zee, Hassan , Amit

AGENDA TOPICS

Time allotted | 50 mins | Agenda topic PDS and design convergence | Presenter Alfred

Discussion Presenting the product design specification and the design convergence to Johanna to update the

supervisors on decision and conclusion has been made by the group.

Conclusion: we still need to validate some criteria’s with numbers and not just use assumptions

Time allotted | 10 mins | Agenda topic |Ordering products | Presenter Alfred

Discussion We asked if it was possible to order products from eBay seen as it would be a lot cheaper ordering

product of their manufacture website itself. A list of product was also shown to Johanna specifying what products

we want

Conclusion Johanna proposed that she would as Howard ash if he could purchase some of the products we want

seen as the aerospace department don’t allow purchases from eBay

Time allotted | 10 mins | Agenda topic The need for sponsors | Presenter Alfred

Discussion We was considering if there was a need for sponsors because seen as we are getting a budget of £1000

from the university, there wouldn’t really be a need because we believe the can easily be made with a budget of

£1000

Conclusion we probably won’t need a sponsorship but the option is still open if need but we need to act soon if we

want a sponsor rather than later

Time allotted | 30 mins | Agenda topic multi rotor concept | Presenter Alfred

Discussion For our final concept of a multi rotor, we need to decide if we are going for a 3, 4, 6 or 8 rotor system as

our finalized concept

Conclusion to come up with another design convergence which has a list of criteria for multi rotor which will

compare different types of multi rotors and hence the win concept will be our final design.

Action items Person responsible Deadline

To improve the numbering system on the Product design spec Alfred 10/11/2014 12:00 PM

Research on manufacturing techniques for 3 to 8 rotor system Zee 10/11/2014 12:00 PM

For one motor failing research the stability for 3 to 8 rotor system

and maneuverability of multi rotor systems

Kasun 10/11/2014 12:00 PM

Page 2

The power requirements ie the thrust produced, time and the

speed

Hassan 10/11/2014 12:00 PM

Research into the costs and strength of material for multi rotors Ozy 10/11/2014 12:00 PM

Research Potential Payload capacity for a series of multi rotor

system

Mohin 10/11/2014 12:00 PM

Look into the Noise levels at which 3 to 8 rotor systems of the

same size produce noise

Amit 10/11/2014 12:00 PM

Look into root sizing and complexity and spacing for a series of

multi rotor system

Zuber 10/11/2014 12:00 PM

Research into Criticality of payload, CofG, stability during flight

and how they differ for 3 to 8 rotor systems

Mo 10/11/2014 12:00 PM

Research optical recognition system to see if an extra board is

required and the potential of using matlab

Tarek 10/11/2014 12:00 PM

Look into systems required for a multi rotor system and present a

list to the group

Jonathan and micky 10/11/2014 12:00 PM

Send an email with updated PDS and Design convergence to

supervisors

Alfred 10/11/2014 12:00 PM

MEng meeting times invitations Johanna 10/11/2014 12:00 PM

UAS CHALLENGE 2015


A.4. Example of Agenda


AGENDA

Finalizing Design Concept

October 28, 2014

12:00PM – 2:30 PM

Meeting called by Alfred Dzadey

Attendees:

Alfred, Mohin, Zuber, Tarek, Jonathan, Micky, Reyad, Kasun, Osman, Mozammel, Hassan,

Amit

Note taker: Jonathan

Please bring: List of ideas/ sketches/ brainstorm for multi-rotor to the table

Location: The Simulation Laboratory

Objective: Discussion of multi rotor concept, finalizing roles of groups and ideas of having sponsors

Introduction

Taking register of attendees and general updates

Schedule

Present design Specification and Design

convergence

Discuss ideas and brainstorm multi rotor idea and

structure

Appoint areas to research for each individual with

regards to multi-rotor discussed

To get a sponsor or not to get a sponsor

Presenter

Alfred

Alfred

Alfred

Alfred

Additional Instructions:

DON’T BE LATE PLEASE!!!

UAS CHALLENGE 2015


A.5. Team members Performance review

Jonathan has from the very beginning, been very enthusiastic about the project, even though he had been

moved around quite a few times from one role to another. Though his enthusiasm had led to great results,

it took some work getting there as the systems group had not properly managed their time. Having said

this, he has guided the group through many hurdles, some of which may have been daunting, and helped

speed up certain processes. Jonathon has on multiple occasions tried to stick to the plan to allow for ample

amount of time for testing, and while there have been delays, the planning for of all the systems

integration has allowed for little room for error. Even though there is room for improvement in managing

the team, Jonathon took great pride in both, his team and work, and this has led to a great contribution on

his part.

Tarek’s strong point, coming into this project, was his experience in electronics and has provided his

knowledge in getting some of the systems up and running quickly. His importance to the team was noted

fairly quickly as he could simply put his head down and make good headway. Nevertheless, there were

times where he was slowed down by delays in purchasing which meant he had to wait. While he had

planned for contingency, he never expected certain things to take as long as it did delaying thing further.

0

1

2

3

4

5

Jonathon Ebhota

0

1

2

3

4

5

Enthusiasm Team value Planning Delivery Contribution

Tarek Kherbouche

UAS CHALLENGE 2015


His knowledge was still valued to the team and his never wavering enthusiasm and support allowed him to

work well with all members of the group very well and help make a large contribution overall.

While Malwenna has not shown as much enthusiasm in the beginning, although this could be attributed to

him being an introvert, it did grow as the group started to become much more comfortable with each

other. Malwenna had constantly researched for a large part of semester one, making sure the rest of the

systems team was going the simplest and most effective route to a successful project. While Malwenna was

a great team player and a great asset to the team, by taking up more workload to help out, he ended up

pausing his own role for a short while, but on the other hand, had he not, the rest of the systems would

have been delayed anyhow making this a no win situation. Nevertheless, he has delivered his role

effectively and made a great contribution to the team.

While Mohammed showed some good levels of enthusiasm, his lack of planning led to rushing and some

late nights to get the task done on time. He initially started to help the propulsion engineer to make sure

that the motors that were chosen were right for the task. While this led to a more objective choice for the

motors, his actual role as a stability engineer had to take a back seat. However, he has since focused on his

role as the stability engineer alongside Malwenna and has made great progress in simulation and testing in

the short time he had. Overall, he started on the right path before getting side-tracked but has come to and

has made some great contribution to the team.

0

1

2

3

4

5


Malwenna Malwenna

0

1

2

3

4

5


Mohammed Rayad Ullah

UAS CHALLENGE 2015


Hassan has continuously worked very hard on his role as the propulsion engineer and has continuously

delivered, driven by his enthusiasm. Hassan has spent a large amount of time making sure that the parts he

chose were the right ones. From an extensive number of calculations through to testing; he has worked

very hard throughout the entire project. However, all this came at the cost of time management, although

he has pulled off all the stops and has completed his work on time to great results.

Amit has consistently pushed for more throughout the entire project, not only from himself but from

others as well. He brought in great level of experience and insight to almost all issues for the Quad-rotor.

His enthusiasm has never wavered and motivated some of the other team members to compete on who

can produce the most quality work. He regularly put his team in front and tried to help out where he could,

and though it may occasionally come as unneeded, and sometimes overbearing, he never let it bother him

and continued forward. His delivery to the project was invaluable and his overall contribution was

exceptional.

0

1

2

3

4

5


Hassan Turabi

0

1

2

3

4

5


Amit Ramji

UAS CHALLENGE 2015


Zuber has been consistently enthusiastic about the project, and while this enthusiasm has led to good

results, the ability to get the structural side of the group to work within deadlines has been good. Having

said this, he has guided the group through certain objectives which had helped speed things up the process,

else would have been hampered by unnecessary delays. Zuber has on multiple occasions stuck to plan and

delivered products on time as promised. While there is still room for improvements in managing the team,

he has executed and delivered his role as a structural team leader very well and has made large

contributions as a whole.

Micky had a lot of personally issues during the duration of the project by has tried to manage this with his academic work. He has some been unable to attend meeting for various reason but when work is asked of

him he delivers good quality work. Weakness I would say is need from contribution from him but his strength are he works had when he commits to doing so.

0

1

2

3

4

5

Zuber Khan

0

1

2

3

4

5


Micky Ngouani

UAS CHALLENGE 2015


While Mohin showed some good levels of enthusiasm, his lack of planning led to rushing and some late

nights to get the task done on time. He generally isn’t a morning person and means that he usually is only

available during the day and late evening. During semester B he has become more involved with the group

and has made some great contribution to the team. His work was mostly of good standard and but

struggled to meet deadlines at times due to various reasons. Mohin is team player and is by far his best

attribute.

Osman has contributed greatly with his all-round support to everyone. He is always willing to help out in

any task and go the extra mile. He completes task in reasonable and expected time and has a good amount

of contribution to the team. He is definitely a team player and has shown some levels of enthusiasm during

the period of the project.

0

1

2

3

4

5


Mohammed Mohinuddin

0

1

2

3

4

5


Osman Sibanda

UAS CHALLENGE 2015


Mozammel has not shown as much enthusiasm in the beginning, although this could be attributed to him

being an introvert. He grew in confidence and began to take responsibility for his role and contribution to

the team. His time management wasn’t great due to personal problems outside academic work but

nevertheless he has produced reasonably standard work. He had put in the extra hours to complete tasks

when asked to and hence has had great contribution to the team progression so far.

0

1

2

3

4

5


Mozammel

UAS CHALLENGE 2015


A.6. Project manager performance review

PROJECT MANAGER EVALUATION FORM

City Project Name: UAS Challenge Dates Evaluated Beginning: September 2014 To: May 2015

Name of Project Manager: Alfred Dzadey Type of Evaluation: MEng Project Final

Code Scale -- Please use the rating code below to appraise the project manager’s work

5 The Absolute Best Project Manager that we have knowledge of and experience with anywhere. Very

Exceptional and Far Superior to others. Value added to the project / Achieved on almost always /

4 Above Average -- Noticeably competent / capable / proficient / adept / knowledgeable / skilled / High

Quality / Achieved on a consistent basis /

3 Average: (satisfactory / acceptable / suitable / reasonable / no major problems / potential is there /

dependable / meets the stands of the job

2 Marginal (Fair: improvement is necessary / deficient in certain area, but potential may be there)

1 Below Average (Needs significant or substantial improvement / really lacking / unsatisfactory)

0 Not observed or applicable

Project Manager’s Rating

5 Management of Team / Project 4 Dependability (can be counted on, return calls/email,

4 Understanding of other PM’s needs 3 Ability to anticipate and analyze problems

4 Professionalism 4 Timeliness (attendance, punctuality, fulfillment of

obligations)

4 Achieved project goals 5 Ability to locate & utilize resources effectively

4 Written communications ability 5 Ability to work with others

4 Oral communications ability 5 Ability to present options and/or reach decisions

OVERALL PROJECT PERFORMANCE OR END PRODUCT (use 1 to 5 scale again on project basis)

4 Adherence to Budget 3 Adherence to Schedule

5 Good Public / Private Team Relationship 4 High Quality Results

Overall Rating 5 in terms of public / private time, resources, and money required to work with them

Circle One:

Definitely looking forward to working with this Project Manager again on another project.

Willing to work with this Project Manager again on another project without any changes

Willing to work with this Project Manager again on another project with some changes

Prefer not to work with this Project Manager again or Project Manager needs significant

improvements

Please note any specific comments here or on a separate sheet (weak points needing improvement, strong

points, instances of going the extra mile.)

- Needs to be kept within the deadlines, you are far too lenient.

+ Good communication skills, and persuasiveness

+ Always going the extra mile

+ Always stands up for teams in front of supervisors and very supportive

UAS CHALLENGE 2015

141 UAV Design MEng Team Project Report (7ENT1024) School of Engineering and Technology



Appendix. B UAV Design

B.1. Weight Reduction - Quad-Rotor

In order to eliminate excess mass, design considerations such as those discussed in Appendix G.1 have been used. The main focus was to achieve a high strength to weight ratio with a fairly high stiffness; hence the use of thin plates in a sandwich design justifies the decision rationale. Using an initial arm bending calculation and iterative process, the best tube diameter was converged to be 16mm x 11.5mm with a wall thickness of 2.25. The Outside diameter of the tube now needs a support to sandwich the plates, a high strength Nylon 66 material is selected for the compression blocks (FB-001, FB-002, EB-001 – Appendix B.7). Decreasing the plate spacing to 25mm proves a challenge for incorporating systems and mechanical pivots, however this reduced the overall weight significantly. Furthermore the Nylon 6 plates (BP-001 & 002 – Appendix B.7) incorporates cut-outs and holes to reduce weight further and allow for a reduced cross section during flight. The gust pressure loading of such cross section has been calculated in 8.4 and added to the maximum flight forces however assuming an opposed direction in order to satisfy worst-case flight conditions. The isolated plate deflection is modelled in Appendix G.11 as an infinite plate assembly. Compared with the analytical technique, the error between results is minimal as is discussed in section 8.7. Main Body Plate sizes (BP-001 & 002) have been sized to be the minimal thickness to allow for stress distribution and maintain a stiffened root structure. Reducing the thickness of these plate further without changing materials would mean the plates would be subject to localised bending and shear deflections (similar to ladder/truss design with weak rail supports). Additionally the planar dimensions consider the contact positions of the Arm support brackets and every attempt has been made to reduce the overall root size of the main body plates. Further cut-outs and weight reduction on most components is still possible however due to time and resource constraints, further material optimisation is not considered. Further mass can be removed from the Undercarriage components (UV-001 & UH-001), along with increased cut-outs on the Main Body plates (BP-001 & 002) and tapering of out-board structures. A further study into the use of Short Fibre Reinforces Composite (SFRC) blocks can also be carried out, however this would be mass produced injection moulded components as detail and finer machining is time consuming and costly.

B.2. Detailed Design and CAD Modelling

The design of the Quad-rotor has been carried out while considering manufacturability and precision of machinability. The overall geometry of the Quad-Rotor is controlled by positions of the Main Body plates (BP-001 & 002), where the CNC process is accurate of 0.2mm. If the Fixed or Movable blocks (FB and MB series) are not accurate to nominal values, the through bolts being used in compression will take up the tolerance as Nyloc Nuts are also being used to ensure no assembly is loosened during flight. The 16mm diameter hole in the blocks for the Arms is also considered at the manufacturing stage during component design; if the manufactured component is loose fit, the gap can be closed by the O-rings and hand finishing of mating half-block surfaces (sandwich of FB-001 x 2 to FA-001 - Appendix B.7). Compression and bolt preload of the fasteners holding this local sandwich together will allow the Arm to be secure during assembly and in flight.

UAS CHALLENGE 2015


Figure 101 - Overall View of Quad-Rotor

Figure 102 - Motor Mount Design (Left) & Undercarriage T-Joint (Right)

Figure 103 - Undercarriage Pivot Design (Left) & Main Body Sandwich Design (Right)

UAS CHALLENGE 2015





Figure 104 - Movable Arm Pivot Design

B.3. Payload Housing Design

The payload housing is designed to be removable for ease of transport, increased functionality and the ability to attach various devices as a payload (eg, Camera and gimbal on a quick release turn button). Parts PB-009 and PB-010 (Appendix B.7) allow the Quad-rotor to be multi-functional and allow for a sleek appearance for mounting accessories. The payload housing is a key component in the design, a hollow truss type design has been converged upon to enable the structure to be lightweight and have high stiffness. Multiple design iterations had been considered during the design stage where Appendix G.15 and G.16 show the changes made to PB-006 and PB-008 (Appendix B.7) to increase the stiffness of the housing during flight conditions to avoid pre-mature deployment of payloads. Rendered views of the payload housing can be observed in Figure 10 though Figure 13.

B.4. Supplier Discount and Advertising Opportunities

The value of structural components such as raw Nylon (PA6 & PA66) blocks / sheets have been demonstrated in Appendix. D where the usage costs have been calculated. The usage cost of materials is equivalent to a buy-back scheme used in industry where off-cuts and machining swarf is sold back to the supplier for recycling. Ensinger Ltd (Watford Plastics division) is one of the largest suppliers globally and has agreed to provide the raw materials at a cost equivalent to supply costs in exchange for advertisement. Buy-back schemes are usually used for long term and large volume purchases, however advertisement has been offered in place of a large contractual order. Costing of plastics is non-standard and a retail price is differing between suppliers, many suppliers can afford to offer the same materials at a fraction of the cost depending on their commercial footprint.

B.5. BOM Assembly Techniques

To save time on assembly level modelling in CATIA, the use of repeated parts is key to a quick design and manufacture. Complexity is also reduced as modifications to single parts can be projected to its upstream parent products. Kits have been arranged in the CATIA model comprising of various repeat components. Such kits include; fixed brackets kits, motor mount kits, fastener kits and overall allows for reduction in possibility of clashes and configuration errors.

UAS CHALLENGE 2015


Section by Amit Ramji B.6. Configuration Control

In order to avoid having multiple versions of the same components with little changes in geometry, a single group member had carried out all modelling. This ensures there is one main CAD model with no chance of duplication of parts and introducing variants. An industry equivalent to this restriction would be a check-in/check-out database such as Siemens Teamcentre or CATIA Enovia, however this could not be possible during the timescales involved in the project for integration. The entire model has undergone a 4-step manufacturing readiness level; where level 1 is conceptual design, level 2 being detail design of components, level 3 being further product level design and manufacturing readiness and level 4 being systems 3D modelling and cable routing.

Part Numbering Scheme

Location Identifiers:

FB = Fixed Bracket EB = End Bracket MB = Movable Bracket AP = Arm Pivot LP = Landing Pivot LB = Landing Bracket BP = Body Plates MP = Motor Plates TJ = T-Joint MA = Movable Arm FA = Fixed-arm UV = Undercarriage Vertical UH = Undercarriage Horizontal

Revision Control:

Revisions of parts are a possibility to introduce under configuration control when the Fit, Form or Function of the part does not change. Due to the constant update of design parts and releasing in a 4 level time-line, revision numbers are not required. Additionally the fact that a single entity is in control of the entire CAD model and configuration control, the potential to introduce part and assembly revisions is unnecessary.

Part and Drawing Release for Manufacture:

Real engineering projects involving a multitude of parts would require a release process, however as the same team member models the Design and carried out the Stress analysis of the structural components, the need for internal release is non-essential. Only one working copy of the entire Quad-Rotor design exists, hence part release and freezing of the design is carried out at internal stage reviews (Levels 1 - 4). Release for manufacture and configuration control again is simplified as a single member is in control of the design and drawing release that also inputs into selecting materials and purchasing. For this reason, drawing release uses the same part-numbering scheme as above and all drawings are deemed as Work-In-Progress until the drawing is assigned a number. An industry equivalent would involve a workflow process where each part and assembly along with material cards and instructions are released at separate departments, however due to project integration constraints, tools such as Teamcentre have not been used.

XY - 00Z

Location / Description Letters Part Number Identifier

UAS CHALLENGE 2015



B.7. Production Support and Drawings


UAS CHALLENGE 2015

171 UAV Detailed Mass Breakdown MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Zuber Khan Appendix. C UAV Detailed Mass Breakdown A detailed mass breakdown was carried out of the whole UAV to ensure that it is within specifications. The itemised breakdown of components and their quantities are shown below.

Structural Part Name

Part No. (Appendix

B.7)

Material

(Appendix.

E)

Density (g/cm3)

Area cm2

Length/ Thickness

(cm)

Volume (cm3)

Mass (g)

Qty Total

Mass (g) Picture

(Appendix B.7)

Tubular Arms MA-001, FA-001

PVC 1.4 0.97 29 28.18 39.46 4 157.84

Fixed-arm Nylon clamps

LB-001 Nylon 1.14 5.2 1 3.56 4.06 16 65.00

Moveable arm half block

clamp FB-002 Nylon 1.14 4.52 1 3.14 3.58 2 7.16

Moveable arm full block

clamp MB-001 Nylon 1.14 8.92 1 5.52 6.29 2 12.59

Moveable arm pivot

AP-001 Nylon 1.14 6.25 5 13.11 14.95 2 29.90

UAS CHALLENGE 2015


Motor clamp full block (end)

EB-001 Nylon 1.14 10.2 1 7.77 8.86 4 35.45

Motor block plate

MP-001 Aluminium 2.7 20 0.1 1.81 4.90 8 39.24

Plates BP-001, BP-

002 Nylon 1.14

351.58

0.2 70.25 80.09 2 160.19

Undercarriage pivot

assembly

LP-001, LB-003

Nylon 1.14 7.37 4 21.67 24.70 2 49.41

Undercarriage tube

UV-001 PVC 1.4 0.97 20 19.43 27.21 2 54.42

Horizontal undercarriage

tube UH-001 PVC 1.4 0.97 35 34.01 47.62 2 95.24

UAS CHALLENGE 2015


Undercarriage T Joint

TJ-001, TJ-002

Nylon 1.14 13.1

2 8 16.06 18.31 2 36.63

Payload box PB-000 Nylon 1.14 101.924 116.19

M3 x 35 Button Head

M3 x 35 x 0.5 Stainless

Steel 7.2 35 0.28 1.7 34 57.8

M3 Nyloc Nut Nyloc M3 x 0.5 Stainless

Steel 7.2 0.09 0.4 34 13.6

M5 x 30 Button Head

M5 x 30 x 0.8 Stainless

Steel 7.2 30 0.4123 4.4 2 8.8

M5 Nyloc Nut Nyloc M5 x 0.8 Stainless

Steel 7.2 0.176 1.4 2 2.8

UAS CHALLENGE 2015


25mm M3 Hex standoff (F/F)

Hexagonal Standoff

Brass 8.45 25 0.55 3.7 4 14.8

M3 Nylon Spacer – 3.2mm

Internal, Outer 6mm, length

25mm

M3 Nylon Spacer

Nylon 1.14 25 0.7 2 1.4


Internal, Outer 10mm, length

10mm

M5 Nylon Spacer

Nylon 1.14 10 0.7 2 1.4


Internal, Outer 12.5mm,

length 10mm

M6 Nylon Spacer

Nylon 1.14 10 1 4 4

M3 x 10 Button Head, pitch 0.5mm

M3 x 10 Button Stainless

Steel 7.2 10 0.6 24 14.4

O-Rings 16mm

internal, O-Rings Rubber 30 2

UAS CHALLENGE 2015


18mm External, c/s

1mm

Quick Release Pin

Quick Release Aluminium 2.7 0.50

3 4.5 2.26 6.11 2 12.21

Springs Springs Steel 7.2 10 2 20

Total (Structural)

1012.5

(g)

Table 23 – Itemised Mass Breakdown of all Structural UAV Components

UAS CHALLENGE 2015


Table 24 - Electronics and Misc Component Masses

Mass for Electronic / Misc Components

Height (mm) Width (mm)

Thickness (mm)

Mass (g)

Qty Total

Mass (g)

Pixhawk 80 49 15 40 1 40

GPS 40 40 9 14.4 1 17.1

OSD 50 18 10 4 1 4

Telemetry kit 50 30 10 50 1 50

Batteries 194 45 47 1848 2 1848

Motors 52.5 35 0 187.4 4 749.6

Propeller Blades

40 4 160

Esc's 80 30 17 75.2 4 300.8

Camera 61 35 18 50 1 50

Lights

150 1 150

Lights control board 46 28 13 73.2 1 73.2

Servos 35.5 22.5 12 13.4 1 13.4

Payload* (Single per trip) 140 105 70 1000 1 1000

Cable ties/ additional cables

100

Power regulator 10 10 7 21.5 1 21.5

Buzzer 30d

5 4.8 1 4.8

Power cable for power module df13

20 15 10 1.5 1 1.5

Power switch 25 7

1.8 1 1.8

Competition GPS Tracker (IMechE, Jan 2015)

59 38 18 50 1 50

XT60 Connectors and Velcro

85 1 85

Motor Extension cable 120 1 120

Additional Systems battery 102 15 35 118 1 118

BEC for servo 50 30 15 55 1 55

Total (Electronics / Misc) 5013.7(g

)

From the summation of all the masses for the electronic and miscellaneous components in Table 24, a total mass of 5013.7 grams was calculated. Total Mass of the UAV 6026.2 grams.

UAS CHALLENGE 2015

177 UAV Detailed Cost Breakdown MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Zuber Khan Appendix. D UAV Detailed Cost Breakdown

Material / Component

Used for (Appendix B.7 for Structural Parts

and Section 10 for Systems)

Raw Cost for Total

Material (£ inc Vat

and Delivery)

Usage cost for Parts

(£ inc Vat and Delivery)

Usage cost for Parts – Excluding Delivery

(£ inc Vat)

PVC Tube MA-001, FA-001, UV-001, UH-001.

£50.34 £37.92 £31.67

10mm Nylon 6.6 Block

FB-001, FB-002, MB-001, EB-001 LB-001, LB-002 PB-009, PB-010

£4.40 £2.25 £1.25


TJ-001, TJ-002 £4.40 £1.01 £0.51


LB-003, LP-001 £6.60 £1.33 £0.67

25mm Solid Circular Bar

AP-001 £4.40 £2.29 £1.145

2mm Nylon 6 Black Plate

BP-001, BP-002, PB-005

£8.80 £7.44 £3.72

1mm Nylon 6 Black Plate

PB-004 £4.40 £3.04 £1.52

Rigid Angle Sections

PB-001, PB-002, PB-003, PB-006, PB-007, PB-008

£39.54 £18.56 £12.31

Aluminium 1mm Plate

MP-001 £9.08 £7.26 £7.26

Pixhawk Pixhawk £159.98 £159.98 £159.98

GPS & Telemetry Kit

GPS & Telemetry Kit

£89.77 £89.77 £89.77

OSD OSD £44.95 £44.95 £43.45

Batteries Batteries £188.76 £188.76 £182.74*

Motors Motors £91.80 £91.80 £91.80

Propeller Blades Blades £12.00 £12.00 £12.00

ESC’s ESC’s £141.75 £141.75 £135.80

Lights & LED panel Board

Lights & LED panel Board

£14.13 £14.13 £14.13

Servo Servo £13.69 £3.42 £3.42

Camera Camera £52.43 £52.43 £47.01

M3 x 35mm x 0.5mm Pitch Bolt

Fasteners £3.95 £2.37 £2.37

M3 Nyloc Nuts Fasteners £1.78 £1.21 £1.21

M5 x 30mm x 0.8mm Pitch

Bolts Fasteners £2.79 £0.56 £0.56

M5 Nyloc Nuts Fasteners £1.19 £0.24 £0.24

UAS CHALLENGE 2015

178 UAV Detailed Cost Breakdown MEng Team Project Report (7ENT1024) School of Engineering and Technology

Table 25 – UAV Itemised Cost Breakdown *Conversion rate accurate as of 26/03/15 - $1 = £0.6678 Total Cost of COTS £824.84 Total Cost of Structure £81.34

M3 Nylon Spacer 3.2mm internal, outside 6mm, length 25mm

Spacer

£4.09

£0.68

£0.68

M5 Nylon Spacer 5.3mm internal, outside 10mm, length 10mm

Spacer £3.39 £0.57 £0.57

M6 Nylon Spacer 6.4mm internal, outside 12.5mm,

length 10mm

Spacer £3.59 £1.20 £1.20

M3 Brass Hexagonal – F/F -

Standoff Spacer £3.09 £2.06 £1.07

M3 x 10mm x 0.5mm Pitch

Fasteners £1.39 £1.11 £1.11

Cable Ties 2.5x100mm

Cable Ties £0.99 £0.99 £0.99

O-Rings – 16mm Internal, 18mm

External, c/s 1mm O-Rings £4.24 £3.60 £3.60

Nylon Hinges 20 x 20mm for

Payload Box Hinge £2.90 £0.97 £0.97

Heat Shrink Tubing Set

Tubing £5.28 £2.53 £0.65

Braided Sleeve Cable Protection

Cable Protection £19.35 £1.55 £1.55

Strobe controller Strobe controller £4.49 £4.49 £4.49

Black Rubber Washers

Black Rubber Washers

£4.39 £4.39 £4.39

M3 x 40mm x 0.5mm Pitch Bolt

Fasteners £1.79 £0.72 £0.72

XT60 Connectors and Velcro

Connectors and Fasteners

£14.29 £14.29 £8.87

Motor Extension Cable

Wires £15.80 £9.88 £9.88

ESC for servo £8.40 £8.40 £8.40

Additional Systems Battery

Batteries £8.50 £8.50 £8.50

Springs £2 £2 £1.60

LED’s x20, Require 4

Lights £11.98 £2.40 £2.40

Total £1062.48 £954.80 £906.18

UAS CHALLENGE 2015

179 Material Properties MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Amit Ramji Appendix. E Material Properties

Material ID Components (Appendix B.7)

Property Value

1 -

Al

All

oy

(MIL

-HD

BK

-5H

)

MP-001 Density ρ= 2770 kg/m3

Young’s Modulus 𝐸= 7.1E10 Pa = 71 GPa

Poisson’s Ratio 𝜈= 0.33

Bulk Modulus K= 6.9608E10 Pa = 69.6 GPa

Shear Modulus G= 2.6692E10 Pa = 26.6 GPa

Tensile Yield Strength σTYS= 280 MPa

Compressive Yield Strength σCTS= 280 MPa

Ultimate Tensile Strength σUTS= 310 MPa

2 -

Bra

ss

(Die

hl, 2

01

5)

Brass M3x25 F/F Spacers.

Density ρ= 8450 kg/m3

Young’s Modulus 𝐸= 1.15E11 Pa = 115 GPa



Shear Modulus G= 4.3201E10 Pa = 432 GPa



3 -

Ny

lon

66

[TE

CA

MID

-66-M

O-

Bla

ck

]

(En

sin

ge

r, 2

01

5b

)

FB, MB, AP, LP, LB, EB & TJ Series. PB-009 & PB-010.


Modulus of Elasticity (Flexural) 𝐸= 3100 MPa






4 -

Ny

lon

6

[TE

CA

MID

-6-M

O-

Bla

ck

]

(En

sin

ge

r, 2

01

5a)

BP-001, BP-002, PB-004 & PB-005








5 -

PV

C H

707

Eq

uiv

(Dir

ect_

Pla

sti

cs,

201

5)

FA-001, MA-001, UV-001 & UH-001








UAS CHALLENGE 2015

180 Material Properties MEng Team Project Report (7ENT1024) School of Engineering and Technology

6 -

Au

sti

nit

ic S

tain

less

Ste

el -

(C

lass

70,

304

gra

de

- c

old

dra

wn

)

(BS

SA

, 2

01

5)

M3 & M5 Fasteners and Nyloc Nuts.


Modulus of Elasticity 𝐸= 193 GPa

Poisson’s Ratio 𝜈 = 0.29

Bulk Modulus K= 134 GPa

Shear Modulus G= 86 GPa

Tensile Proof Strength (0.2% - R1, P0.2)

σ= 450 Mpa

Ultimate Tensile Strength σUTS= 700 Mpa

7 –

PV

C R

igid

An

gle

(Dir

ect_

Pla

sti

cs,

2015)

PB-001, PB-002, PB-003, PB-006, PB-007 & PB-008








UAS CHALLENGE 2015

181 Load Cases and Load Transfer MEng Team Project Report (7ENT1024) School of Engineering and Technology



Appendix. F Load Cases and Load Transfer (Supplement to section 7.1)

F.1. Steady Flight Case A steady flight case scenario during which the UAV would be under maximum flight loads. This would include: the motors producing the maximum amount of thrust, the UAV flying at maximum velocity and maximum gust being applied in the opposite direction of flight. The steady flight case analysis covers various conditions which the UAV will be put under such as, take-off, manoeuvres during flight and hover.

F.2. Drag on the Main Plates Maximum flight speed would be achieved when the UAV is at a maximum tilt angle of 54 degrees (Section 7.1) to the vertical. Using this along with the total surface area of the main body plates, the Drag force could be calculated.

Figure 105 - Project Main Body Area

𝑠𝑖𝑛(54) =𝑇𝑜𝑡𝑎𝑙 𝑃𝑟𝑜𝑗𝑒𝑐𝑡𝑒𝑑 𝐴𝑟𝑒𝑎 (𝑆)

35129.71= 28420.53𝑚𝑚2

Equation 18 - Projected Area For steady flight the motors produce enough thrust to balance the weight. Therefore the mass was

is by 4. 7

4= 1.75𝐾𝑔. However this would not be the thrust when in flight due to the UAV being at

an angle of 54 degrees. Therefore a component was taken as shown below.

𝑐𝑜𝑠(54) =𝑇ℎ𝑟𝑢𝑠𝑡 (𝑇)

1.75= 1.0286 𝐾𝑔

Equation 19 - Thrust at 54 Degrees

To calculate the drag force, the following equation is used:

𝐷 =1

2𝜌𝑉2𝑆𝐶𝐷

Equation 20 - Drag Equation (R. H. Barnard, 2010) Where: D = Drag Force, p = Density, V = Velocity, S = Area, Cd = Coefficient of Drag The maximum gust the UAV has to fly in is 25knots and the maximum allowable flight speed of the UAV is restricted to 60knots. Therefore the maximum wind on the UAV would be 85Knots. A Cd value for the plate was worked out using, ‘(1.28 x sin(angle))’ (NASA, 2014).

1.28 × sin(54) = 1.0355 𝐷 = 1

2 × 1.226 × (85 × 0.5144)2 × 28420.53 × 10−6 × 1.0355

= 34.491𝑁 ∴ 𝐷 = 3.516𝐾𝑔 is the drag force equivalent distributed on the main plate. To this a ‘global load safety factor of 1.5’ was added for the purpose of working out the Maximum stresses and deflections. 1.5 × 3.516 = 5.274𝐾𝑔

Total Surface Area = 35129.71mm2

54° Total Projected

Area

54°

T

1.75Kg

UAS CHALLENGE 2015

182 Stress Analysis MEng Team Project Report (7ENT1024) School of Engineering and Technology



Appendix. G Stress Analysis

G.1. Stress Reduction Techniques The following design techniques have been adopted to maximise efficiency of the material and ensure a lightweight and stress reduced structure at local discontinuities and overall load paths.

Align known material properties with major load direction where possible. Hence the use of Nylon 66 Blocks being used in compression (FB-001 & 002) and fasteners being used in shear and tension (M3’s & M5’s).

Use of flexible joints to avoid excessive stress load transfer (70 Shore Rubber O-Rings and Landing Springs – See Appendix B.7).

Stiffen or reinforce unsymmetrical features to minimize flexure. An example of this consideration is the use of the Nylon 66 Fixed Blocks (FB-001 & 002) used in the main body alongside the M3 Brass spacers which act as rigid links between the main body plates (BP-001 and 002) to reduce total body deflection.

Encourage smooth transitions in cross section and stress levels, avoiding hard points in the primary load path. In some cases this could not be avoided (MA-001 contacting FB-002 – See Figure 115 through Figure 119), therefore an additional local support (MB-001 – Appendix B.7) is incorporated.

Accounting for structural deflections and considering specific threats (Heavy Landing) where compromised integrity of the structure and/or the integrity of the systems installed in the structure could be a cause for concern.

Where appropriate, distribute the load pathways between multiple components to avoid bulky structure and concentrated stress distributions on single components. An example of such situation is the multiple load paths in the main body, where a sandwich type design is achieved. The stiffness of the main body structure is greatly increased with rigid links (M3 Fasteners, FB-001, FB-002, MB-001 and M3 spacers).

G.2. Fatigue Awareness A gain in fatigue life can in most situations be achieved without an increase in cost, simply by attention to design detail. The following should be taken into account when considering the Quad-Rotor structure:

Avoiding sharp edges, corners and sudden changes in cross-section can reduce stress concentrations. Fillet and intersection radii should be as large as possible as such used in the Lug Bracket (LB-003) and Pivots (AP-001 & LP-001).

Avoiding joggles in the load line or catering for joggles by additional stiffening to bridge the joggles. Considering the combined loading effect of cut outs and holes in close proximity as those used in the Main Body Plates and Motor Mount Plates (BP-001, BP-002 and MP-001). A detail hand calculation using Petersons Stress Concentration Factors (Pilkey and Pilkey, 2008) has not been carried out as this complex geometry and cut-outs are previously considered in the Finite Element Model with mesh refinement, inflation and pinch controls.

The majority of fatigue cracks will start at stress concentrations such as holes, notches, etc. Any design features or processes that can be applied to reduce the severity of such stress concentrations should be used.

Ensuring design of joints are such as not to give rise to built-in stresses on assembly, or load some portions of the joint unduly. The use of M3 and M5 from the same supplier to avoid mixing fasteners of dissimilar material/strength and those that require differing tolerances of fit. Fasteners with tighter tolerances will load the local structure during repeated flexure more than a loose tolerance fastener due to the miniscule freedom of movement of the joint.

In fatigue critical areas, interference fit fasteners shall be used whenever possible in preference to clearance fit. A close tolerance for clearance/transition fit fasteners will improve the fatigue performance of the joint, as this will minimize the risk of individual holes being over-loaded. For the current Quad-Rotor design, fasteners are loaded axially hence

UAS CHALLENGE 2015




introducing a bolt pre-load and reducing the miniscule movement if any existed. Pre-tensioning of the bolt can reduce alternating stresses in the bolt and improve its fatigue

performance. The correct seating of the fastener head and nut along with use of the correct installation torque is therefore essential to avoid local bending.

G.3. Fatigue due to induced vibration

Fatigue damage can often arise from induced vibration from the motors as compared with fatigue damage arising from directly applied structural stresses. Often this vibration is not sustained for long periods of time, a modal analysis case has been considered for the Fixed-arm assembly as shown in Appendix 8.17 and compared to analytical methods as shown in Appendix 8.16. Such calculated modal frequencies should be avoided or swiftly passed through the first 3 natural frequencies when powering up the motors to idle and can be programmed into the ESC’s as “soft, medium, hard” starts.

Avoiding the use of long cantilevered members, as these will experience high inertia forces in vibration. The modal analysis of the Arm has been the main concentration for the purpose of frequency response analysis, as the cantilever of the Arms are more susceptible to vibration than any other components.

Rigidly mounted equipment may be vibrated by the structure to which it is attached, hence the use of O-Rings at the motor mounts and dampening foam being used on all sensitive components such as Pixhawk due to its susceptibility to compass excitation during vibration.

G.4. Boundary Conditions - Connection Type and Contact Element Type

Bonded

The bonded connection applies to all contact regions (surfaces, solids, lines, faces, edges). With this connection type there is no sliding or separation between faces or edges (Ansys, November 2013a). This type of contact was used for a quick initial analysis of all assemblies as the solution time and model could be checked. Bonded contact allows for a linear solution since the contact length/area will not change during the application of the loads. Using the bonded contact elements, the contact is determined on the mathematical model where any gaps will be closed and any initial penetration will be ignored (Ansys, November 2013b, Ansys, November 2013e). Correct refinement was carried out once the models were deemed correct and the calculated displacements or stress match analytical methods in sampled areas. The contact types in most regions had been refined to rough or frictionless where appropriate, bonded contact was maintained between LP-001 and UV-001, alongside MA-001 and AP-001 (Appendix B.7).

Frictionless

The Frictionless contact connection is a standard unilateral contact where normal pressure is zero is separation occurs (Ansys, November 2013a). With frictionless connections, gaps can form in the model between bodies depending on the loading criteria and directions. Hence this solution is nonlinear due to the area of contact prone to changing as the load is applied. A zero coefficient of friction is assumed, thus allowing free sliding and is used in the model where pivot regions and open surfaces exist. Such frictionless areas modelled in specific load cases (e.g. landing and entire quad flight cases (Appendix G.14) is associated with parts AP-001, LP-001 and FB-002 (Appendix B.7). For the analysis to converge, all surrounding geometry is well constrained by using bolted connections at bolt surfaces and rough connections at relevant arm brackets. Weak springs are added to the assembly by default during the iterative process to help stabilize the DOF’s in order to achieve a reasonable solution. In many cases the solution processing time was reduced by enabling parallel processing and post compilation of solutions, up to 8-cores had been programmed in some solution cases (Ansys, November 2013f).

UAS CHALLENGE 2015


Rough

The rough connection is similar to the frictionless type however models perfectly rough frictional contact where there is no sliding. Alternative connection types are also possible where friction factors can be modelled, however increases solution time significantly and for the purpose of this analysis is deemed unnecessary. Rough connections apply to regions of faces or edges of plates, brackets and O-ring locations (Appendix B.7). By default, no automatic closing of gaps is performed and corresponds to an infinite friction coefficient between the contacting bodies (Ansys, November 2013a). The rough connection had been replaced by No Separation connections in motor plate regions (MP-001 to EB-001 & FB-001) for the entire quad analysis and landing cases (Appendix G.14). Using this method, the solution time is reduced and allows for accurate stress solutions at the root of the quad-rotor as non-linearity is previously demonstrated in the arm stress analysis (Appendix G.7 and G.8)

No Separation – Rigid Body

The No Separation contact setting is similar to the bonded case and only applies to regions of faces or edges. Separation of the geometries in this contact connection is not permitted (Ansys, November 2013a). The No Separation connection is used in the motor plate regions (MP-001 to EB-001 & FB-001) for the entire quad analysis and landing cases (Appendix G.14). Once again, solution time is reduced and allows for accurate stress solutions at the root of the quad-rotor as non-linearity and friction contact has previously been demonstrated in the arm stress analysis (Appendix G.7 and G.8).

Bolted – Rigid Body

For modelling bolted connections in Ansys Workbench an MPC184 Revolute Joint Element is used instead of Rigid Body Elements (RBE2 or RBE3) used in Ansys Mechanical APDL or NASTRAN. The MPC184 revolute joint is a two-node element that has only one primary degree of freedom, the relative rotation about the revolute (hinge) axis. The Revolute joint is similar to modelling a Beam Line Element at the bolt location alongside using RBE’s to average the bearing pressure loading at hole contact surfaces. This element imposes kinematic constraints such that the nodes forming the element have the same displacements. Additionally, “only a relative rotation is allowed about the revolute axis, while the rotations about the other two directions are fixed” (Ansys, November 2013a, Ansys, November 2013c).

Spring

For the landing consideration a compression spring has been modelled between components (LB-002 and LP-001 (Appendix B.7)). The compression stiffness was set to 300 N/mm and the damping was set to 0 N.s/mm for an initial deflection analysis. The solution is yet to converge due the increased DOF solution from the Ansys modeller.

UAS CHALLENGE 2015



Section by Amit Ramji G.5. Solver Formulation Augmented Lagrange solver method has been used for the majority of contact models involving Bonded and No separation contact, as it is a penalty-based method. In comparison to the Pure Penalty method, this method usually leads to “The Augmented Lagrange method requiring additional iterations, especially if the deformed mesh becomes too distorted” (Ansys, November 2013g). In some analysis cases, Program Controlled or the Pure Penalty method is used for decreasing the solution time and iterations. Such cases include the landing case where solution time is significant due to the Degrees of Freedom of the Undercarriage components.

G.6. Mesh

Element Types Used

SOLID187

The SOLID187 element used as per Table 26 is a high order 3 dimensional, 10-node element. The SOLID187 has a quadratic displacement behaviour and is well suited to modelling irregular meshes (Ansys, November 2013c).

The element allows for having 3 DOF at each node: translations x, y, and z directions. The element has large deflection and strain capabilities; alongside plasticity, hyperelasticity, stress stiffening and creep capabilities. Figure 106 – SOLID187 Element

(Ansys, November 2013c)

PLANE182

The PLANE182 element used as per Table 26 is also known as a QUAD182 [PATRAN Conversion: WEDGE15, HEX20] depending on its use in 2D or 3D configuration. The PLANE182 element can be used for 2D representation of a solid 3D structure.

The element is a 4 node type which has 2 DOF at each node. The element has large deflection and strain capabilities; alongside plasticity, hyperelasticity, stress stiffening and limited 2D creep capabilities (Ansys, November 2013c). Figure 107 – PLANE182 Element (Ansys, November

2013c)

UAS CHALLENGE 2015


Mesh Refinement

The process of mesh refinement is a post mesh generation step in which elements on the selected parts are split and refined. The process of mesh refinement has only been used for small sub-assemblies where the local features are to be studied, for example in the arm stress cases and motor mount plates. Local mesh refinement has been used on main body plates for entire quad-rotor flight analysis cases at the hole and cut-out locations and has been removed at the arms and motor plate regions to decrease computing time. Mesh refinement being removed from such regions is no longer important as the parts have been justified in another upstream analysis case.

Contact Pinch Controls and Inflation

Pinch controls have been used at contact positions where removal of small features (such as short edges and narrow regions) at the mesh level. Pinch control helps to generate better quality elements around such contact positions as the nodes are aligned and shared between mating components. The Pinch control provides an alternative to Virtual Topology modeling used at geometry level. Both Virtual Topology and Pinch Controls work together to simplify meshing constraints due to small features such as edge chamfers and corner radii and grooves. To further ensure the mesh and analysis was efficient, such small features had been removed in a separate simplified CAD model, which also removed fasteners and small non-structural components. Inflation is used in certain locations where high stress concentrations exist and involves additional layers or elements surrounding the feature under question. An example of where inflation has been used is at the Motor plate fastener positions (Table 26).

UAS CHALLENGE 2015


Mesh attributes used in Quad-Rotor Analysis

The following mesh properties have been used to identify the localised stress on individual components, for larger assemblies these values have been changed suited to their location within the load path. e.g. Motor Mounting Plates (MP-001) in the entire quad assembly or Arm assembly has had refinement, inflation and contact pinch controls removed to save on computing time and simultaneously provide accurate results of the global assembly (Ansys, November 2013h, Ansys, November 2013d). The cases where detail analysis of failure points is to be considered, pinch controls, inflation and refinement mesh elements have been used in each analysis case where appropriate. The table below is for reference values of mesh values that should be used for such detail analysis.

Part No. (Appendix B.7)

Material ID (Appendix. E)

Property Value Image

MP-001 1

Element Type

Solid – Tet 10 node

Type of Mesh

TET10 – SOLID187

Size (Aspect Ratio)

1.15 (Min)

9.5 (Max)

2.5 (Ave)

Refinement Level 2 @ 40 Hole and Slot Faces

Inflation 3 Layers at Bolt Interface Positions - Growth Rate 1.2

Pinch Controls

Default at Bolt Locations

FB-001 FB-002 MB-001 EB-001 LB-001 LB-002 PB-009 PB-010

3 Element Type


Type of Mesh

TET10 – SOLID187

Size (Aspect Ratio)

1.3 (Min)

26 (Max)

2.8 (Ave)

Refinement None

Inflation None

Pinch Controls

None

FA-001 MA-001 UV-001 UH-001

5 Element Type

Plane – Quad 4 Node

Types of Mesh

PLANE182 / QUAD182, [PATRAN Conversion: WEDGE15, HEX20]

Size (Aspect Ratio)

1.28 (Min)

26.4 (Max)

2.82 (Ave)

Refinement None

Inflation None

Pinch Controls

None

UAS CHALLENGE 2015


LB-003 3 Element Type


Types of Mesh

TET10 – SOLID187

Size (Aspect Ratio)

1.21 (Min)

52 (Max)

2.84 (Ave)

Refinement None

Inflation None

Pinch Controls

None

LP-001 AP-001

3 Element Type


Types of Mesh

TET10 – SOLID187

Size (Aspect Ratio)

1.28 (Min)

26.4 (Max)

2.82 (Ave)

Refinement None

Inflation None

Pinch Controls

None

TJ-001 TJ-002

3 Element Type


Types of Mesh

TET10 – SOLID187

Size (Aspect Ratio)

1.236 (Min)

44.1 (Max)

3.213 (Ave)

Refinement None

Inflation None

Pinch Controls

None

BP-001, BP-002, PB-004 & PB-005

4 Element Type


Types of Mesh

TET10 – SOLID187

Size (Aspect Ratio)

5mm (Body Size)

1.2 (Min)

15.2 (Max)

3.0 (Ave)

Refinement Level 1 @ 44 Hole and Slot Faces

Inflation 3 Layers at Bolt Interface Positions - Growth Rate 1.2

Pinch Controls

Default at Bolt Locations

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PB-001, PB-002, PB-003, PB-006, PB-007 & PB-008

7 Element Type

Plane – Quad 4 Node

Types of Mesh

PLANE182 / QUAD182, [PATRAN Conversion: WEDGE15, HEX20]

Size (Aspect Ratio) Refinement

1.19 (Min)

63.3 (Max)

5.20 (Ave)

None

Inflation None

Pinch Controls

None

Table 26 – Mesh Attributes for Components

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Section by Zuber Khan G.7. Stationary Motor Arm Stress Analysis Materials: As per Appendix. E

D1 = 0.06m & 0.067m D2 = 0.23m & 0.247m

Transferring loads from F1 to F2 required a moment transfer using: 𝑀 = 𝐹 × 𝑑

Equation 21 – Moment Calculation

Fixed-arm

Full Arm length of 0.23m 𝑀1 = (1.75 × 9.81) × 0.23 = 3.9485𝑁𝑚 Equation 22 - Moment for Fixed Arm The moment can then be transferred to the first Nylon clamp where D1 =0.06m. Which can then be used to find out the force that will be applied on the Nylon clamps. 3.9485 = 𝐹2 × 0.06 𝐹2 = 65.81𝑁

Maximum Fixed Arm Stress

Maximum force was applied to represent maximum thrust produced by the motor. The thrust was then multiplied by the ‘global load safety factor of 1.5’.

Figure 108 - Arm Cross-section for Stress Calculation To work out the stress in the arm the following equation was used.

𝜎 =𝑀𝑦

𝐼=

𝐹 × 𝑑 ×𝐷12

𝜋64 (𝐷1

4 − 𝐷24)

Equation 23 - Stress in a Cylindrical Pipe (Warren C. Young) When the motors are on full thrust the arm will be under maximum compression on the top surface and under maximum tension on the bottom surface as shown in Figure 109.

Figure 109 - Tension & Compression Stress in Arm

Reaction

Force

0.17m

0.23m

Moment

D1=0.016m

D2=0.0115m

Compression

Tension

F1

D1

F2

D2

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Section by Zuber Khan Stress analysis at Fixed-arm – Analytical Solution

𝐹 =𝑀𝑇𝑂𝑊

4 𝐴𝑟𝑚𝑠=

7 × 9.81

4≈ 25𝑁

𝜎 =25 × 0.17 ×

0.0162

𝜋64

(0.0164 − 0.01154)= 14416251.37

𝑁

𝑚2

𝜎 = 14.42𝑀𝑃𝑎 (𝑇𝑒𝑛𝑠𝑖𝑜𝑛) & − 14.42𝑀𝑃𝑎 (𝐶𝑜𝑚𝑝𝑟𝑒𝑠𝑠𝑖𝑜𝑛)

Stress analysis at Fixed-arm – FEA Method


Figure 110 – Mesh for Fixed-arm Assembly – Values as per Appendix G.6 Results:

Figure 111 - Deflection of Fixed-arm Assembly (Flight Loads) with 7.6mm Deflection

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Figure 112 - Stress of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa (Peak)

Figure 113 – Stress (Close-up) of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa (Peak)

FEM Verification: Tube Stress Comparison

One can observe the results from the above analytical stress calculation being 14.42MPa and the

stress level as seen in the far field stress contour of the tube in Figure 109 (15.8MPa) being very

close. Substantiation of the numerical modelling and contact constraints can be deemed as

accurate as a very small difference is observed between the methods.

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G.8. Movable Arm Stress Concentration at Contact Points Materials: As per Appendix. E

Movable-arm

Full Arm length of 0.247m 𝑀1 = (1.75 × 9.81) × 0.247 = 4.2404𝑁𝑚 Equation 24 - Moment for Movable-arm The moment can then be transferred to the first Nylon clamp where D1 =0.067m. This can then be used to find out the force which will be applied on the Nylon clamps. 4.2404 = 𝐹2 × 0.067 𝐹2 = 63.29𝑁

Analytical Stress at Movable-arm

𝜎 =25 × 0.18 ×

0.0162

𝜋64

(0.0164 − 0.01154)= 15264266.16

𝑁

𝑚2

𝜎 = 15.26𝑀𝑃𝑎 (𝑇𝑒𝑛𝑠𝑖𝑜𝑛) & − 15.26𝑀𝑃𝑎 (𝐶𝑜𝑚𝑝𝑟𝑒𝑠𝑠𝑖𝑜𝑛)

The yield strength of the material used for the arms is 55MPa (Appendix. E). One can observe that the arms have a minimum of 3.6 reserve factor remaining, in addition to the added factor of 1.5 for the global safety. From this analysis, it can be justified that the arms at this size and with the properties defined in Appendix. E are suitable for the UAV.

FEA Method for Movable-arm


Figure 114 – Mesh for Arm Assembly (With additional Tab) – Mesh Values as per G.6

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Results:

An additional 0.4129g results in slightly lower stress levels (see Figure 116 & Figure 118). However the main reason for introducing this modification is to eliminate the possibility of piercing MA-001 during repeated loading. By increasing the contact surface area allows for a more distributed loading edge during deflection. Figure 115 - Modified FB-002 for reduction in point contact stress concentration

Figure 116 - Stress Concentration at Arm (without addition) Contact (a) & Close-up (b)

Modified FB-002 Bracket with Tab Addition

Benefits – Increased fatigue resistance and larger Non-Linear contact area, which is important for repetitive loading and general contact stress reduction. Point contact is now a Line contact (for Non-Linear Flexure) and Line contact is now a Surface Contact (for Linear Flexure).

mOrange = rV =1.15g / cm3 ´3.142cm3 = 3.6133g

mBlue = rV =1.15g / cm3 ´3.501cm3 = 4.0262g

\Dm = +0.4129g

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Figure 117 - Deflection of Modified Movable Motor Arm of 7.88mm for flight loads with SF

Figure 118 - Stress of Modified Movable Motor Arm of 20.8MPa for flight loads with SF

Figure 119 - Modified Movable Motor Arm with Stress of 20.8MPa for flight loads with SF (a) & Close-up (b)

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G.9. Undercarriage Lug Bracket Flange Addition Materials: As per Appendix. E

Lug Analysis Using Analytical Methods

Lug analysis was carried out to calculate if the lug design would be able to cope with the loads put upon it.

Figure 120 - Load on the Lug (Niu, 1988) To carry out the analysis the load was split into components as shown in Figure 121.

Thickness of Lug (t) = 4.5mm Width (W) = 25mm Diameter (D) = 5mm 𝑇𝑟𝑎𝑛𝑣𝑒𝑟𝑠𝑒 𝐿𝑜𝑎𝑑 (𝑃𝑇) = 7 × 1.5 × 9.81 × cos 45 = 72.84𝑁

𝐴𝑥𝑖𝑎𝑙 𝐿𝑜𝑎𝑑 (𝑃𝐴) = 7 × 1.5 × 9.81 × cos 45 = 72.84𝑁 Ultimate strength of the material (Ftu) = 85MPa Figure 121 - Components of the Load (Niu, 1988)

The Areas on the lug were determined to be able to calculate the maximum allowable load.

𝐴𝑟𝑒𝑎 𝐴1 = ((𝑊𝑖𝑑𝑡ℎ

2) − (

0.707 × 𝐷𝑖𝑎𝑚𝑒𝑡𝑒𝑟

2)) × 𝑇ℎ𝑖𝑐𝑘𝑛𝑒𝑠𝑠

𝐴𝑟𝑒𝑎 𝐴1 = ((25

2) − (

0.707 × 5

2)) × 4.5 = 48.29𝑚𝑚2

Equation 25 - Area A1 on Lug (Niu, 1988)

𝑨𝒓𝒆𝒂 𝑨𝟐 = (𝑾𝒊𝒅𝒕𝒉−𝑫𝒊𝒂𝒎𝒆𝒕𝒆𝒓

𝟐) × 𝑻𝒉𝒊𝒄𝒌𝒏𝒆𝒔𝒔AreaA2 =

(Width-Diameter

2) × Thickness = (

𝟐𝟓−𝟓

𝟐) × 𝟒. 𝟓 = 𝟒𝟓𝒎𝒎𝟐

Figure 122 - Areas on the Lug Equation 26 – Area A2 on Lug (Niu, 1988)

𝐴𝑟𝑒𝑎 𝐴3 = 𝐴𝑟𝑒𝑎 𝐴2 Equation 27 - Area A3 on Lug (Niu, 1988)

𝐴𝑟𝑒𝑎 𝐴4 = 𝐴𝑟𝑒𝑎 𝐴1 Equation 28 - Area A4 on Lug (Niu, 1988)

𝐴𝑣𝑒𝑟𝑎𝑔𝑒 𝐴𝑟𝑒𝑎 (𝐴𝑎𝑣) = 6

3𝐴1 +

1𝐴2 +

1𝐴3 +

1𝐴4

= 6

348.296 +

145

+1

45+

148.296

= 47.145𝑚𝑚2

Equation 29 - Average Area of Lug (Niu, 1988)

PA

PT

D W

PT

A2

A1 A3

A4

45'

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𝐵𝑒𝑎𝑟𝑖𝑛𝑔 𝐴𝑟𝑒𝑎 (𝐴𝑏𝑟) = 𝑇ℎ𝑖𝑐𝑘𝑛𝑒𝑠𝑠 × 𝐷𝑖𝑎𝑚𝑒𝑡𝑒𝑟 = 4.5 × 5 = 22.5𝑚𝑚2

Equation 30 - Bearing Area on Lug (Niu, 1988)

𝐴𝑎𝑣

𝐴𝑏𝑟= 2.095 Using 2.095 and the tension efficiency graph a Ktu & Kty value of 0.923 was

determined. Therefore the allowable traverse load = Ktu x Abr x Ftu = 1765.15

Therefore the reserve factor for the lug is: 𝑅. 𝐹 =1765.15

72.84= 24.23

From this one can conclude that the lug is more than sufficient for the purpose of this UAV.

Lug Analysis Using FEA Methods with Flange addition

Mesh: Values as per section G.6 Results: Previous Lug Bracket without Flange To improve the stress distribution within the Lug bracket (LB-003) for the Undercarriage, additional flanges have been incorporated to distribute the load evenly to the fastened plate face. It proves beneficial to repeated heavy landings and side impact cases.

Figure 123 - Lug Bracket Without Flange (Left) & with additional Flange (Right)

An additional 0.67g results in lower stress levels (see Figure 124) and is also beneficial for repeated loading and impact consideration during a side impact landing as demonstrated in Appendix G.13. Although the Lug Bracket (LB-003) in Figure 123 (Right) is more complex to machine, the design is a one-off and if a series production part was to be introduced, an injection moulded equivalent would take its place and be simpler and quicker to manufacture. The additional flange demonstrates that the small addition of material can improve the structural performance and repeated loading capability of parts significantly.

mLeft = rV =1.15g / cm3 ´8.833cm3 =10.1579g

mRight = rV =1.15g / cm3 ´9.415cm3 =10.8272g

\Dm = +0.67g

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Figure 124 – Lateral Unit Load Deflection (Left) & Stress (Right) of Lug Bracket Without Flange Modified Lug Bracket with Flange Addition Benefits – Increased fatigue resistance and multiple load paths which is important with repetitive heavy landing and sideward crash cases.

Figure 125 – Lateral Unit Load Deflection (Left) & Stress (Right) of Lug Bracket With Flange

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Section by Amit Ramji G.10. Motor Plate Stress Analysis Materials: As per Appendix. E Mesh: Values as per section G.6

Figure 126 - Mesh for MP-001 (Appendix B.7) with values as per Appendix G.6 Results:

Figure 127 – Motor Plate Deflection (0.038 mm) and Stress (41.7 MPa) for flight case with SF at start-up

Figure 128 - Error Elements in Model - Due to Separation at FB-001 and EB-001

25N Flight Case at 4 RBE3’s 10 Nm of Torque* *(Translated to 4 RBE3’s as in-plane Loads)

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G.11. Main Body Plate Stress Analysis Materials: As per Appendix. E

Simply Supported Plate Deflection

A simple plate deflection was determined of a 2mm thick Nylon plate with dimensions of 315mm by 280mm. This was the largest the plate would go to on the UAV if necessary therefore was used for the purpose of this analysis. The reason for doing this was to compare the analytical results with the results produced by the FEA model. If the results were similar or close to the analytical method, the method could be applied to the whole UAV model.

Figure 129 - Simplified Plate Representations All the edges are simply supported for this analysis.

Analytical Method

Below are the Navier stokes equations used to work out the plate deflection at the centre, where the maximum deflection will take place from engineering judgement.

𝐷 = 𝐸𝑡3

12(1 − 𝑣2)

Equation 31 - Flexural Rigidity of the Plate (Ventsel and Krauthammer, 2001)

𝑤(𝑥, 𝑦) = ∑ ∑ 𝑤𝑚𝑛 sin𝑚𝜋𝑥

𝑎sin

𝑛𝜋𝑦

𝑏

∞

𝑛=1

∞

𝑚=1

= 𝑤11 sin𝜋𝑥

𝑎sin

𝜋𝑦

𝑏+ 𝑤12 sin

𝜋𝑥

𝑎sin

2𝜋𝑦

𝑏+ 𝑤21 sin

2𝜋𝑥

𝑎sin

𝜋𝑦

𝑏+ …

Equation 32 – Navier solution (Ventsel and Krauthammer, 2001)

𝑎𝑚𝑛 =16𝑞0

𝑚𝑛𝜋2

Equation 33 - Navier stokes coefficient 1 (Ventsel and Krauthammer, 2001)

𝑤𝑚𝑛 =1

𝜋4𝐷

𝑎𝑚𝑛

[(𝑚2

𝑎2 ) + (𝑛2

𝑏2)]2

Equation 34 - Navier Stokes coefficient 2(Ventsel and Krauthammer, 2001) First the pressure distributed on the whole plate surface was calculated. 𝐹𝑜𝑟𝑐𝑒

𝐴𝑟𝑒𝑎=

33.8445

88200 × 10−6= 383.72

𝑁

𝑚2

Followed by calculating the flexural rigidity

𝐷 = 3300 × 106 × 0.0023

12(1 − 0.42)= 2.61905

The Navier coefficients 1 and 2 could be calculated for when mn = 1 1, 1 3, 3 1, 3 3

𝑎11 =16×383.72

1×1×𝜋2 = 622.063 𝑎13 = 207.35 𝑎31 = 207.35 𝑎33 = 69.12

X = a = 315mm

Y

=

b

=

280

mm

Youngs Modulus, E = 3300MPa Thickness, t = 0.002m Poisson’s Ratio, v =0.3 Distributed Force = 33.8445N Area = 88200 x 10e-6 m2

UAS CHALLENGE 2015



w11 =1

π4 × 2.61905

622.063

[(12

0.3152) + (12

0.282)]2 = 4.67689 × 10−3

w13 =1

π4 × 2.61905

207.35

[(12

0.3152) + (32

0.282)]2 = 5.21215 × 10−5

w31 =1

π4 × 2.61905

207.35

[(32

0.3152) + (12

0.282)]2 = 7.59333 × 10−5

𝑤33 =1

𝜋4 × 2.61905

69.12

[(32

0.3152) + (32

0.282)]2 = 6.41566 × 10−6

The coefficients were then input into the Navier solution equation to calculate the deflection at the centre.

𝑤(𝑥, 𝑦) = 4.67689 × 10−3 × sin (𝜋 × 0.1575

0.315) × sin (

𝜋 × 0.14

0.28) + 5.21215 × 10−5 × sin (

π × 0.1575

0.315)

× sin (3π × 0.14

0.28) + 7.59333 × 10−5 × sin (

3π × 0.1575

0.315)

× sin (π × 0.14

0.28) + 6.41566 × 10−6 × sin (

3π × 0.1575

0.315) × sin (

3π × 0.14

0.28)

w(x, y) = 4.67689 × 10−3 − 5.21215 × 10−5 − 7.59333 × 10−5 + 6.41566 × 10−6

w(x, y) = 4.555 × 10−3m = 4.555mm

FEA – Simplified Rectangular Approximation

Using Catia the same plate was modelled with the same constraints and loads to see the deflection it would cause.

Figure 130 - Simple Plate Deflection Carried out on CATIA structural analysis

UAS CHALLENGE 2015



From the FEA model it can be observed that the deflection has been calculated to be 4.54mm. The mesh used was set to a size of 2mm with absolute sag of 1.5mm. Therefore any further plate bending analysis carried out on CATIA, should be set to the same mesh size and constraints as it has been substantiated to provide accurate answers.

Method Deflection

Analytical (Rectangular Plate) 4.555mm

FEA CATIA (Rectangular Plate) 4.54mm

Table 27 – Comparison of Simplified Plate Deflection for Model Substantiation

FEA – As Built (Single Plate) Ansys Results


Figure 131 - Mesh of Main Body Plate - Values as per Appendix G.6 Results:

Figure 132 – Single Main Body Plate Analysis – with 17.8MPa Stress at contact holes for flight case with pressure load

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G.12. Main Body Plate Stress Analysis as Built Ansys Results Materials: As per Appendix. E Representation: In order to carry out a quick analysis of the main body assembly, point masses for the payload, systems and batteries had been added to the structure with the masses defined in Appendix. C.

Figure 133 – Mass Representation of components and payloads as per Appendix. C Mesh: Values as per section G.6

Figure 134 - Mesh of Main body assembly with Values as per Appendix G.6 Results:

Figure 135 – Contact model Flight Case for Main body assembly Deflection (left) and Equivalent Stress (right)

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Figure 136 - Contact model Flight Case for Main body assembly - Equivalent Stress with predicted locations

G.13. Undercarriage Buckling and Torsion Cases Materials: As per Appendix. E

Undercarriage Stress Analysis

For carrying out the undercarriage stress analysis the leg was treated as a single entity. The loads were first applied individually to see how the material would react and if it would be able to cope for the initial sizing stage. For all the cases the worst-case scenario would be the full weight of the UAV landing on one leg.

Analytical – Undercarriage Leg Buckling – Without Spring

To calculate the leg buckling stress and critical load, the following equations were used.

𝑛𝐿

𝜌≥ √

𝜋2𝐸

𝜎𝑦𝑠

Equation 35 – Slenderness Ratio (Warren C. Young)

𝜌 = √𝐼

𝐴

Equation 36 - Radius of Gyration (Warren C. Young)

𝑃𝐶𝑅 =𝜋2𝐸𝐼

(𝑛𝐿)2

Equation 37 - Critical Load to Cause Buckling (Warren C. Young)

𝜎𝐶𝑅 =𝜋2𝐸

(𝑛𝐿𝜌 )

2

Equation 38 - Critical Stress to Cause Buckling (Warren C. Young) The assumption was made whilst calculating the buckling load and stress that 1 end was fixed due to a jam and one end free resulting in the equivalent length ‘n value’ to be 2. Working out the second moment of area of the tube.

𝐼 =𝜋

64(𝐷4 − 𝑑4) =

𝜋

64(0.0164 − 0.01154) = 2.35845 × 10−9𝑚4

UAS CHALLENGE 2015



Cross-sectional area of the tube, 𝐴 = 𝜋𝑟2 = (𝜋 × 0.0082) − (𝜋 × 0.005752) = 9.7193 × 10−5𝑚2

Radius of Gyration, 𝜌 = √2.35845×10−9

9.7193×10−5 = 4.926 × 10−3

Slenderness Ratio, 2×0.18

4.926×10−3 ≥ √𝜋2×3100×106

55×106

73.082 ≥ 23.586 Therefore the buckling formula can be used for this scenario. The critical load which would cause the leg to buckle is shown below.

𝑃𝐶𝑅 =𝜋2 × 3100 × 106 × 9.7193 × 10−5

(2 × 0.18

4.926 × 10−3)2 = 556.78𝑁

This demonstrates that approximately 56.76Kg landing on one leg would cause the leg to buckle, if the leg was pointing vertically down. To get a more accurate buckling load, the component of that was taken.

cos 45 =𝐴

556.78= 393.7𝑁 = 40.13𝐾𝑔

Figure 137 - Resolving Component to Determine Vertical Load

Analytical – Undercarriage Leg Bending

Figure 138 - Undercarriage Leg Under Pure Bending

Stress caused on the undercarriage leg due to pure bending has been calculated below. The

assumptions made for the calculation was that the pivot was treated as fixed which considered a

jam or lateral crashing load. Another assumption made was that the t-joint at the bottom of the leg

was also treated as rigid. The force applied on the leg was the full weight of the craft, which was

multiplied by 1.5 (Global load safety factor).

D=178mm

F F

45°

556.78N

A

UAS CHALLENGE 2015



Equation 23 is used to determine the stress in the tube.

𝜎 =𝑀𝑦

𝐼=

𝐹 × 𝑑 ×𝐷12

𝜋64 (𝐷1

4 − 𝐷24)

𝜎 =7 × 1.5 × 9.81 × 0.178 ×

0.0162

𝜋64

(0.0164 − 0.01154)=

0.14668

2.35845 × 10−9= 62193016.6𝑃𝑎 = 62.2𝑀𝑃𝑎

62.2MPa is the maximum stress the leg would undergo under pure bending if the UAV were to land on one leg. This would cause the leg to yield however the load applied is excessive and it is applied only to one leg, which would not occur repeatedly. Additionally this analysis does not consider the entire body deflection that would dramatically reduce the stress levels. In this calculation, the pivot is assumed to be fixed with infinite stiffness, however in reality this cannot be true, as the main assembly would also deflect. Working backwards using Equation 23 the max force could be found out which would cause the undercarriage leg to yield.

55 × 106 =𝐹 × 0.178 × 0.008

2.35845 × 10−9= 91.0918𝑁 = 9.29𝐾𝑔

9.29Kg is the force required in pure bending to cause the leg to yield. This is a significantly low load, however in reality the UAV would land on both legs repeatedly therefore this force could be doubled. The undercarriage design proposes to incorporate springs to help reduce the impact force on the structure and provide some give by allowing for a designed deflection.

Analytical – Undercarriage leg Torsion

Stress caused on the undercarriage leg due to pure torsion has been calculated below. The assumptions made for the calculation was that the pivot was treated as fixed which considered a jam in the pivot mechanism. Another assumption made was that the t-joint at the bottom of the leg was also treated as rigid. The force applied on the leg was the full weight of the craft which was multiplied by 1.5 (Global load safety factor). Figure 139 - Undercarriage Leg Under Pure Torsion

𝜃 =𝑇𝐿

𝐺𝐽

Equation 39 - Angle of Twist (Warren C. Young)

G = Shear Modulus = 1.0993 x 109Pa (Appendix. E)

𝐽 =𝜋𝑑4

32=

𝜋 × (164 − 11.54)

32= 4716.899𝑚𝑚4

Equation 40 - Polar Moment (Warren C. Young)

F

UAS CHALLENGE 2015



T=Torque Applied = 7 × 1.5 × 175 × 9.81 = 18025.9𝑁𝑚𝑚 L= Length = 0.175m Therefore the twist angle was calculated to be:

𝜃 =18025.9 × 180

1.0993 × 103 × 4716.899= 0.6257 𝑟𝑎𝑑 = 35.85°

𝜏 =𝑇𝑟

𝐽

Equation 41 - Shear Stress (Warren C. Young) For shear stress to be maximum, r (radius) needs to be maximum.

𝜏 =18025.9 × 8

4716.899= 30.57𝑀𝑃𝑎

Yield in shear = 1099.3MPa (Appendix. E) Therefore the material is suitable to withstand maximum torque which could be applied on it with an RF=35 in this loading condition.

Analytical – Undercarriage Combined Loading – Bending, Buckling and Torsion

A combined loading analysis was carried out in which 3 different forces are applied to the undercarriage at the same time to see if the material can withstand the loads. The loads which were applied were a bending load, buckling force and a torque at the bottom of the leg with an applied 1.5 global load safety factor. If the material can withstand the loads without yielding it can be assumed that the material is suitable, and can withstand the worst loads the UAV shall face.

Figure 140 - Stress Element A (Warren C. Young) Using the stress element A as shown in Figure 140, the equations of combined load can be used. A plan view of element A has been shown below.

Figure 141 - Plan View of Stress Element A

T = Torque

Buckling Load

Bending Load

𝜏𝑥𝑦

𝜎𝑥

𝜎𝑦

𝜏𝑥𝑦

𝜏𝑦𝑥

𝜎𝑦

𝜎𝑥

𝜏𝑦𝑥

A

UAS CHALLENGE 2015


The formulas used to work out the combined stress and their principle angles have been shown below.

𝜎𝑥 = −𝐹

𝐴

Equation 42 - Compression Stress on Pipe (Warren C. Young)

𝜎𝑦 =𝑀𝑦

𝐼=

𝐹 × 𝑑 ×𝐷12

𝜋64 (𝐷1

4 − 𝐷24)

Equation 23 - Stress in a Cylindrical Pipe (Warren C. Young)

𝜏 =𝑇𝑟

𝐽

Equation 41 - Shear Stress (Warren C. Young)

𝜎1 =𝜎𝑥 + 𝜎𝑦

2±

1

2√(𝜎𝑥 − 𝜎𝑦)

2+ 4𝜏𝑥𝑦

2

Equation 43 - Principle Stress 1 and 2 (Warren C. Young)

The compression stress on the leg using Equation 42: 𝜎𝑥 = −7×1.5×9.81

𝜋

4(162−11.52)

= −1.059798𝑁

𝑚𝑚2

Bending stress on the leg using Equation 23: 𝜎𝑦 =𝑀𝑦

𝐼=

7×1.5×0.178×0.016

2𝜋

64(0.0164−0.01154)

= −6.34𝑀𝑃𝑎

Shear stress on the leg using Equation 41: 𝜏 =7×1.5×0.175×9.81×8

4716.878884= 30.572𝑀𝑃𝑎

Using the above stresses the principle stress could be worked out using Equation 43:

𝜎1 =−1.059798 − 6.34

2+

1

2√(−1.059798 − 6.34)2 + 4 × 30.5722 = 27.095 𝑀𝑃𝑎

𝜎2 =−1.059798 − 6.34

2−

1

2√(−1.059798 − 6.34)2 + 4 × 30.5722 = −34.495 𝑀𝑃𝑎

Using the stresses above the principle angles were determined to show the direction they were in.

tan 2𝜃 =2𝜏𝑥𝑦

𝜎𝑥−𝜎𝑦=

2×30.572

−1.059798−6.34= −8.2629 ∴ 𝜃1 = −41.55° & 𝜃2 = 48.45°

Equation 44 - Principle Stress Angles (Warren C. Young) The principle stresses and their angles could then be applied to Figure 141. Figure 142 - Stress Element A with Principle Stresses

𝜏𝑥𝑦

𝜎𝑥

𝜎𝑦

𝜏𝑥𝑦

𝜏𝑦𝑥 𝜎𝑦

𝜎𝑥

𝜏𝑦𝑥

A 𝜃1

𝜃2

27.095

-34.495

UAS CHALLENGE 2015



The maximum shear caused by the combined loadings has been calculated below using Equation 45.

𝜏 =1

2√(𝜎𝑥 − 𝜎𝑦)

2+ 4𝜏𝑥𝑦

2

Equation 45 - Shear Due to Combined Loadings

𝜏 =1

2√(−1.059798 − 6.34)2 + 4 × 30.5722 = 30.795𝑀𝑃𝑎

It can be concluded that the material would be able to withstand the maximum shear cause by the combined loadings.

FEA Solutions

Mesh: Values as per Appendix G.6

Figure 143 - Undercarriage Mesh for Contact Model with values as per G.6

FEA Results – Bending – Lateral Crash Landing

Figure 144 – Lateral Landing on Single Undercarriage Leg with 53.6mm Deflection

UAS CHALLENGE 2015


Figure 145 - Lateral Landing on Single Undercarriage Leg with 60MPa Bending Stress Figure 145 and Figure 146 show the results of the bending analysis to be 60MPa. The analytical bending calculation from above also results in a similar bending stress of 62.2MPa. The justifications on yielding in the above section still hold true for this analysis.

As above, the entire structure will deform and reduce stress hence 60MPa is not a realistic situation. This infinite stiffness constraint can be solved by finer mesh and adding a spring (Deformable support) with the stiffness of Nylon at the Lug bracket holt-holes. Figure 146 - Lateral Landing on Single

Undercarriage Leg with 60MPa Bending Stress (Close-up)

UAS CHALLENGE 2015


FEA Results – Tip Landing

As with the Lateral bending case above, the entire structure will deform and reduce stress, hence 60MPa is not a realistic situation. This infinite stiffness constraint can be solved by finer mesh and adding a spring (Deformable support) with the stiffness of Nylon at the Lug bracket holt-holes. Additionally the T-Joint in this analysis is considered as a rigid body, however there will be some deflection at the T-Joint, which will reduce the stress upstream. The reason for regarding the T-Joint as rigid in the analysis is to reduce computing time as such a non-linear solution is very lengthy to set-up and run.

Figure 147 - Tip Landing on Single Undercarriage Leg with 60MPa Bending Stress

FEA Results – Combined Torsion and Bending – Tip Contact

Figure 148 -Tip Landing on Single Undercarriage Leg with 66mm Combined bending and torsion deflection As with the Lateral bending and Tip landing cases above, the entire structure will deform and reduce stress, hence 71MPa is not a realistic situation. This infinite stiffness constraint can be solved by finer mesh and adding a spring (Deformable support) with the stiffness of Nylon at the Lug bracket holt-holes. Additionally the T-Joint in this analysis is considered as a rigid body, however there will be some deflection at the T-Joint, which will reduce the stress upstream. The reason for regarding the T-Joint as rigid in the analysis is to reduce computing time as such a non-linear solution is very lengthy to set-up and run

UAS CHALLENGE 2015



Figure 149 - Tip Landing on Single Undercarriage Leg with 71MPa Combined bending and torsion stress

FEM Verification – Summary or Undercarriage Results

Case Description Deflection (mm) or (deg)

Equivalent Load (N) or Stress (MPa)

Buckling Analytical Axial loading of UV-001 N/A 393.7N

Bending Analytical Bending of UV-001

N/A 62.2MPa

Bending FEA 53.6 mm 60.63MPa

Torsion Analytical Torsion of UV-001 35.85 deg 30.57MPa

Combined Analytical

Combined Bending and Torsion of UV-001

N/A 34.495MPa

Combined FEA 66.76mm 71.76MPa

Table 28 – Summary of Undercarriage Results

UAS CHALLENGE 2015



G.14. Entire Quad Non-Liner Contact Model – Flight Case Parts: As per Appendix B.7 Materials: As per Appendix. E Mesh: Values as per section G.6 Results:

Figure 150 – Entire Quad-Rotor Flight Deflection of 7.9mm at Motor Arm Tips

Figure 151 - Entire Quad-Rotor Flight Deflection of 7.9mm at Motor Arm Tips (Close-up)

Figure 152 - Entire Quad-Rotor Flight Stress of 28.8 MPa at Motor mount plates

UAS CHALLENGE 2015


Figure 153 - Entire Quad-Rotor Flight Stress with Plate Stress peak at 14.42Mpa

UAS CHALLENGE 2015



G.15. Payload Housing Non-Liner Contact Model – Old Design Parts: As per Appendix B.7 Materials: As per Appendix. E Mesh: Values as per section G.6 Results:

Figure 154 – Downward Load - 1kg Payload and 10N Additional Load onto PB-005 Plate

Figure 155 - Side Load - 1kg Payload and 10N Additional Load onto Hinge Plate at 45deg to horizontal

UAS CHALLENGE 2015


Figure 156 - Side Load - 1kg Payload and 10N Additional Load onto short edge 45deg to horizontal

Figure 157 - Side Load as per Figure 156 - Showing Pre-mature Release due to global deflection

UAS CHALLENGE 2015



G.16. Payload Housing Non-Liner Contact Model – New Design

Modification of PB-006 and PB-008 to result in a more rigid design to avoid pre-mature deployment of payload and incorporation of two smaller hinge positions. Parts: As per Appendix B.7 Materials: As per Appendix. E Mesh: Values as per section G.6 Results:

Figure 158 – Downward Load as per Figure 154 - new design showing 0.73mm Deflection

Figure 159 - Side Load as per Figure 155 –new rigid design and Deflection of 1.56mm

*No pre-mature deployment during manoeuvres as seen in the previous design from Figure 156 and Figure 157 Figure 160 – Side Load as per Figure 156 and Figure 157 – with new design and deflection of 0.41mm*

UAS CHALLENGE 2015


Section by Amit Ramji G.17. Finite Element Model Checking For the majority of Finite Element Analysis (FEA) modelling used in industry, a sample analytical

calculation should be carried out on a simplified load case or geometry or by correlation with a

physical test. However for the majority of cases, usage of material is required and modal response

is not possible in most laboratories due to costly test equipment and resources. The simplified

geometry cases show substantiation is possible by using the same modelling techniques and

contact types.

As a result, FEA techniques with guidance from NAFEMS and by reference to Ansys

guides*shows good correlation for the analytical solutions and complex non-linear contact models.

Model checks have been carried out at various stages which include material properties, geometry

checks, mesh sizes, boundary conditions and preliminary validation checks such as free modal

analysis in the static workbench, resulting in a zero displacements in all DOF at 0 Hz. Additional

quick model checks as those described in Appendix G.4 make use of initial bonded contacts to

check if all parts of assemblies have been well constrained and later refined before investing

further computing time.

UAS CHALLENGE 2015

219 Performance & Propulsion MEng Team Project Report (7ENT1024) School of Engineering and Technology

Appendix. H Performance & Propulsion

H.1. Propulsion

This section investigates the possible inrunner and outrunner electric motors, propellers, and power sources

that are capable of producing the thrust required to firstly lift the Hex-rotor, and secondly attain the velocity

required to complete the challenge on time before the batteries are exhausted.

1. To calculate Hex-rotor’s performance the MTOW weight is vital, and for this initially 7kg was used

based on 2kg payload, 1.5kg power source, 1.2kg propeller, motors and attachment, the frame and

all other electronics components adding to 2kg plus another 5% for possible unexpected weight

addition.

2. Identified Hover thrust – Using MTOW of 7kg it was identified that for the Hex-rotor to hover it would

require each of the six motors to produced 1.167kg of thrust to hover in 1g

3. Identified thrust for manoeuvrability – Using an equation provided by leading multicopter builders just

as DJI, thrust required for improved manoeuvrability was calculated

(MTOW ∗ 2 ∗ 1.2

Number of motors) = 2.8kg of thrust/motor

4. Identified performance criteria to complete the mission in 2minutes – The mission consists of a

range of 2km, taking into account 2 minutes the velocity that the Hex-rotor requires to travel at is

16.67m/s (32knots) taking into account 12.8m/s (25knot) gust and 10.2m/s (20knot) wind the Hex-

rotor would require to travel at relative speed of 29.7m/s (57knots)

5. Propeller, Motor, ESC and Battery selection- It was very quickly identified that a low rpm/V brushless

electric motor was required so that it can use a large propeller with high pitch so that it can produce

the lift and thrust required but also not too larger of a propeller so the velocity isn’t sacrificed. The

selection of these important components required the use of a sophisticated website called

ecalc.com and theoretical calculations, which identified 20 different combination of producing the

thrust required, velocity required and battery life that can last greater than 2 minutes. By performing

these calculations it helped to narrow down from countless number of propeller, motor, ECS and

battery combinations that would achieve the specification required. Specification of these

components can be found in Apendix E

6. Calculations using xcalc.com

As it can be seen from Appendix C the Turnigy G32-770kv motor has a maximum of 1000Watts. The

1000W is based on no load condition were a propeller is not attached to the motor, but these

conditions change when a propeller is added to the motors, the effect is that the added load reduces

the motors capability to 791.9W. The motor itself also has an efficiency factor of 0.882 which is

specified by the manufacturer that further reduces the motors capability to 698.2W

Performance Calculations

The flight profile for the Hex-rotor has been calculated using the flight path specified in Appendix C and the

performance calculations has been based on using the Turnigy G32-770 motor, APC-E 11x7 composite

propellers, 60A ESC and 16000mAh lipo batteries options include using 5S but cost constraints may result in

using multiple 3S or 4S. Results of the flight profile, velocity, time, distance covered and battery status are

presented below table 1.1.1

UAS CHALLENGE 2015


Table 1.1.1 shows Realistic Calculation Under Windy Condition

Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)

[Runway] 0 0 100 0

[Runway]-[30.5m Altitude] 6.77 3.85 98.7 26.2

[30.5m Altitude]-[1] 15.932 16 92.5 (282-26)=255.8

At [1] half loiter performed 8 0.875 92.2 7

[1]-[2] 15.932 50.78 66.5 (842-7-26)=809

[30.5m Altitude]-[descend to

1m]

6.77 3.85 65.3 26.2

Hover N/A 5 64.5 N/A

From 1m to 30.5m 6.77 3.85 63.3 26.2

[30.5m]-[3] 15.932 24.6 51.8 (418-26.2)=391.8

At [3] half loiter performed 8 0.875 51.5 7

[3]-[Target] 15.932 18.9 42.7 (334-7-26)=301

[30m Altitude]-[descend to

1m]

6.77 3.85 41.3 26.2

Hover N/A 5 40.5 N/A

From 1m to 30.5m 6.77 3.85 39.1 26.2

[30.5m]-[Runway] 15.932 6.13 36.2 (124-26.2)=97.8

Hover N/A 5 35 N/A

Total N/A 152.41

(2.54

minutes)

N/A 2000

Sample Calculation

Max velocity calculation

Total Thrust = Max thrust per motor*number of motors*9.81

Total thrust = 2.8kg*6*9.81m/s2 = 164.8N

Total thrust produced at 49degree tilt angle = Total thrust*Cos(ϴ)

Total thrust produced at 49degree tilt angle = 164.8Cos(49) = 108N

Max velocity (m/s) = √𝑇𝑜𝑡𝑎𝑙 𝑡ℎ𝑟𝑢𝑠𝑡 𝑝𝑟𝑜𝑑𝑢𝑐𝑒𝑑 𝑎𝑡 49𝑑𝑒𝑔𝑟𝑒𝑒 𝑡𝑖𝑙𝑡 𝑎𝑛𝑔𝑙𝑒

2∗𝜌∗𝜋∗𝑟2

Max velocity (m/s) = √108𝑁

2∗1.226∗𝜋∗𝑟2 = 26.22m/s

Taking into account 10.288m/s wind condition Velocity = 15.932m/s

Google maps was used to measure the distance from runway to point [1] = 255.8m

From this time(s) taken is calculated = T=distance/velocity = 16s

UAS CHALLENGE 2015


RPM Range

Velocity Increases with RPM

Lift (N

) Calculating battery percentage remaining

Battery charge state at runway = 16000mAh (16Ah)

Current (I) = 44.77A*number of motors = 44.77A*6 = 268.62A

Time (s) = 16s = 0.267minutes


268.62 (𝐴)∗ 60 = 0.267(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)

Battery Capacity (Ah) = 0.267∗268.62

60 = 1.193Ah

Battery capacity remaining = 16Ah-1.193Ah = 14.8Ah

Battery percentage % = 14.8𝐴ℎ

16𝐴ℎ∗ 100 = 92.5%

H.2. DATA Figure: 1.1 shows lift (N) of 11*8 propeller at different RPM

Figure 1.2 shows Lift Vs RPM graph

RPM

Lift Vs RPM Graph

7*6

UAS CHALLENGE 2015


Po

we

r (W

)

Figure: 1.3 shows power vs rpm graph

Table 1.1 showing propeller constant values for different manufacturers.

Propeller Manufacturer 𝐾𝑝

APC 1.11

Graunper 1.18

Aeronaut 1.31

Table 1.2 showing propeller dimensions and acceptability

Propeller Dimensions (in) Rpm Power (W) Lift (N) Acceptable propeller Y/N 7*6 20,000 463 16.53 N 7*9 20,000 694 20.24 N

7*8.25 20,000 637 19.38 N 7.8*6 20,000 714 24.15 N 7.8*7 20,000 833 27.88 N 8*3.8 20,000 500 21.00 N 8*4 20,000 526 21.50 N 8*5 20,000 658 24.08 N 8*6 20,000 790 26.38 N 8*7 20,000 617 19.09 N 8*8 20,000 619 30.46 N 8*9 20,000 1185 32.31 N 8*10 20,000 1316 34.06 N 9*3.8 20,000 801 31.71 N 9*4.7 20,000 991 35.26 Y 9*6 20,000 1265 39.84 Y

RPM

Power Vs RPM Graph

Power required line for

propeller 7*6

UAS CHALLENGE 2015


18,750 1042 35.02 Y 9*7 20,000 1476 43.04 Y

18,100 1094 35.24 Y 9*7.5 20,000 1582 44.55 Y

17,800 1115 35.28 Y 9*8 20,000 1130 46.01 Y

17,500 930 35.22 Y 9*9 20,000 1898 48.80 Y

17,000 1165 35.26 Y 9*10 20,000 2109 51.43 Y

16,550 1195 35.22 Y 10*3 20,000 965 40.73 Y

18,550 770 35.05 Y 10*4 20,000 1286 47.04 Y

17,000 832 35.20 Y 10*4.7 20,000 1511 50.99 Y

16,570 859 35.00 Y 10*5 20,000 1608 52.59 Y

16,400 886 35.36 Y 10*6 20,000 1929 57.62 Y

15,600 915 35.05 Y 10*7 20,000 2250 62.22 Y

15,000 949 35.00 Y 10*8 20,000 2572 66.52 Y

14,600 1000 35.45 Y 10*9 20,000 2893 70.56 Y

14,100 1013 35.07 Y 10*10 20,000 3215 74.37 Y

13,750 1044 35.15 Y 11*3 20,000 1412 56.87 Y

15,700 683 35.04 Y 11*3.8 20,000 1788 64.00 Y

14,800 724 35.05 Y 11*5 20,000 2353 73.41 Y

13,900 790 35.46 Y 11*6 20,000 2824 80.42 Y

13,200 811 35.03 Y 11*7 20,000 3295 86.87 Y

12,700 843 35.02 Y 11*8 20,000 3765 92.86 Y

12,300 875 35.12 Y 11*9 20,000 4236 98.50 Y

12,000 915 35.46 Y

UAS CHALLENGE 2015


11*10 20,000 4707 103.82 Y 11,700 942 35.53 Y

12*6 20,000 4000 109.06 Y 11,350 731 35.12 Y

12*8 20,000 5333 125.93 Y 10,600 764 35.37 Y

13*4 20,000 3672 117.83 Y 10,900 594 35.00 Y

13*6 20,000 5509 144 Y 9,900 668 35.36 Y

13*8 20,000 7345 166.70 Y 9,200 715 35.26 Y

13*10 20,000 9182 186.31 Y 8,700 755 35.25 Y

14*13 20,000 16056 275.33 Y 7,150 733 35.19 Y

15*6 20,000 9765 238.14 Y 7,700 557 35.29 Y

17*10 20,000 26852 476.45 Y 5,450 543 35.37 Y

Table: 1.3 shows a section of table 1.2 with propeller dimensions 9*6

Propeller Dimensions (in) Rpm Power (W) Lift (N) Acceptable propeller Y/N

9*6 20,000 1265 39.84 Y

18,750 1042 35.02 Y

Table: 1.4 shows a section of table 1.2 with propeller dimensions 17*10


17*10 20,000 26852 476.45 Y

5,450 543 35.37 Y

UAS CHALLENGE 2015


H.3. Motor Specfication and maxium RPM values

Table: 1.5 showing rc motor specifications, and maximum rpm values

Manufacturer

/Model

Power

(W)

KV

(rpm/V)

Max

Current

(I)

Working

Current

(I)

Power

Supply

Cell

Range

(s)

Propeller

Dimension

Range (in)

Weight

(g)

Cost

(£)

Max

RPM

BRC HOBBIES

PRODUCTS

EMax

GT2820/07

600 850 54 48 3-4 9*4.7 – 12*6 140 20.95 10064

A2826-5T 645 840 45 37.6 4 12*6 –13*6.7 175 27.70 9945

Emax

BL2820-07

740 919 59 33.5 3-4 10*5 – 13*6 145 17.95 10880

Emax

GT2826-06

962 710 52 42 4-5 10*5-14-7 175 23.95 12987

Emax

GT3526-04

875 870 69 55 3-4 12*6-13*6.5 265 32.95 10300

Boost 0.50 800 600 55 45 3-5 12-13 295 62.95 8880

Boost 0.60 900 490 60 50 4-6 13-14 345 69.95 8702

Boost 0.80 950 340 60 52 5-7 14-15 395 79.95 7044

Boost 0.90 1000 300 65 55 6-9 16-17 455 79.95 7992

HOBBY KING

PRODUCTS

Turnigy G46 925 550 55 46 4-5 12-15 303 34.43 8140

Turnigy

D3548/4

910 1100 50 45 3-4 159 13.30 13024

Turnigy

D3542/4

690 1450 48 42 2-3 10-12 130 13.60 12876

Turnigy

2834-800

660 800 45 40 2-4 10-12 195 22.38 9472

UAS CHALLENGE 2015


Turnigy

3508-640

550 640 30 25 2-5 10-13 98 20.05 9472

Turnigy

3639

800 1100 45 38 2-3 11-13 136 17.39 9768

Turnigy SK3

3542

670 1000 45 40 3-4 11-12 141 18.63 11840

Turnigy

L3020B

800

600 54 48 3-4 10-12 146 14.59 5328

Turnigy

4250

900 540 60 55 3-5 9-12 236 13.80 7992

NTM 35-30 560 1400 37 32 3 9-13 88.3 12.23 12432

NTM 35-36 722 800 43 34 3-4 9-15 130 16.84 9472

NTM 35-42 600 1250 56 42 3-4 10-11 142 18.64 14800

NTM 35-48 640 1100 70 62 3-4 11-13 173 14.61 13024

Scorpion SII-

3026

1000 710 60 55 4-5 12-15 205 83.74 10508

Scorpion SII-

3014 V2

600 1040 40 35 3-4 11-14 129 67 12313

Scorpion SII-

3014

550 830 30 25 4-5 10-15 129 61.63 12284

Scorpion SII-

3020

780 890 45 40 4-5 10-14 166 73.69 13172

Quanum MT

4010

548.3 580 24.7 20 4-6 9-12 127 23.818 10300

Quanum MT

3510

568.3 630 25.6 18 3-6 9-11 100 16.52 11188

Quanum MT

3510 V2

672.7 700 30.3 24 3-6 10-12 100 16.52 12432

4-MAX

Professional

Series

3542-1000

605 1000 60 55 2-4 12-14 142 33.49 11840

Professional

Series

590 1250 60 55 2-4 11-14 142 33.49 14800

UAS CHALLENGE 2015


3542-1250

Professional

Series

3548-790

850 790 50 45 3-4 11-12 171 35.99 9353

Professional

Series

3548-900

840 900 50 45 3-4 11-13 171 35.99 10656

Professional

Series

3548-1100

850 1100 50 45 3-4 11-12 171 35.99 13024

Professional

Series

4250-650

1150 650 60 55 3-4 12-14 230 48.95 13616

E-FLITE

Power 32 800 770 60 45 4-5 11-14 200 50.27 11396

Power 15 575 950 42 34 3-4 10-13 152 43.57 11248

Power 60 1000 470 80 65 5-6 15-17 230 73.73 8347

Power 46 925 670 55 40 5-6 12-14 290 60.32 11899

Power 25 600 870 44 32 3-4 11-14 190 46.92 10300

Power 25BL 850 1250 58 50 3-4 8-10 183 46.92 14800

Table: 1.6 shows a section of table 1.5 for motor model EMax GT2820/07

Manufacturer

/Model

Power

(W)

KV

(rpm/V)

Max

Current

(I)

Working

Current

(I)

Power

Supply

Cell

Range

(s)

Propeller

Dimension

Range (in)

Weight

(g)

Cost

(£)

Max

RPM

EMax

GT2820/07

600 850 54 48 3-4 9*4.7 – 12*6 140 20.95 10064

UAS CHALLENGE 2015


Table: 1.6 shows a section of table 1.5 for motor model EMax GT2820/07

Manufacturer

/Model

Power

(W)

KV

(rpm/V)

Max

Current

(I)

Working

Current

(I)

Power

Supply

Cell

Range

(s)

Propeller

Dimension

Range (in)

Weight

(g)

Cost

(£)

Max

RPM

EMax

GT2820/07

600 850 54 48 3-4 9*4.7 – 12*6 140 20.95 10064

Table 1.2.1 shows updated version of table 1.2 taking into account motor capabilities


10*8 20,000 2572 66.52 Y

14,600 1000 35.45 Y

10*9 20,000 2893 70.56 Y

14,100 1013 35.07 Y

10*10 20,000 3215 74.37 Y

13,750 1044 35.15 Y

11*3.8 20,000 1788 64.00 Y

14,800 724 35.05 Y

11*5 20,000 2353 73.41 Y

13,900 790 35.46 Y

11*6 20,000 2824 80.42 Y

13,200 811 35.03 Y

11*7 20,000 3295 86.87 Y

12,700 843 35.02 Y

11*8 20,000 3765 92.86 Y

12,300 875 35.12 Y

11*9 20,000 4236 98.50 Y

12,000 915 35.46 Y

11*10 20,000 4707 103.82 Y

11,700 942 35.53 Y

12*6 20,000 4000 109.06 Y

UAS CHALLENGE 2015


11,350 731 35.12 Y

12*8 20,000 5333 125.93 Y

10,600 764 35.37 Y

13*4 20,000 3672 117.83 Y

10,900 594 35.00 Y

13*6 20,000 5509 144 Y

9,900 668 35.36 Y

13*8 20,000 7345 166.70 Y

9,200 715 35.26 Y

13*10 20,000 9182 186.31 Y

8,700 755 35.25 Y

14*13 20,000 16056 275.33 Y

7,150 733 35.19 Y

15*6 20,000 9765 238.14 Y

7,700 557 35.29 Y

17*10 20,000 26852 476.45 Y

5,450 543 35.37 Y

H.4. Propeller data Table: 2.0 shows maximum rpm required at different propeller pitch setting to achieve 29.46m/s

Propeller Pitch

(in)

𝑉𝑚𝑎𝑥

(m/s) 𝜃𝑚𝑎𝑥 (degrees) 𝑅𝑃𝑀𝑚𝑎𝑥

3.8 29.46 32 20,371

4 29.46 32 19,352

5 29.46 32 15,481

6 29.46 32 12,901

7 29.46 32 11,058

8 29.46 32 9,676

9 29.46 32 8601

10 29.46 32 7740

13 29.46 32 5954

UAS CHALLENGE 2015


Table 1.2.2 shows updated version of table 1.2.1 taking into account motor capabilities

Propeller Dimensions (in) Rpm Power (W) Lift (N) Acceptable propeller Y/N 10*8 20,000 2572 66.52 Y

14,600 1000 35.45 Y 10*9 20,000 2893 70.56 Y

14,100 1013 35.07 Y 10*10 20,000 3215 74.37 Y 11*6 20,000 2824 80.42 Y

13,200 811 35.03 Y 11*7 20,000 3295 86.87 Y

12,700 843 35.02 Y 13,750 1044 35.15 Y

11*8 20,000 3765 92.86 Y 12,300 875 35.12 Y

11*9 20,000 4236 98.50 Y 12,000 915 35.46 Y

11*10 20,000 4707 103.82 Y 11,700 942 35.53 Y

12*6 20,000 4000 109.06 Y 11,350 731 35.12 Y

12*8 20,000 5333 125.93 Y 10,600 764 35.37 Y

13*6 20,000 5509 144 Y 9,900 668 35.36 Y

13*8 20,000 7345 166.70 Y 9,200 715 35.26 Y

13*10 20,000 9182 186.31 Y 8,700 755 35.25 Y

14*13 20,000 16056 275.33 Y 7,150 733 35.19 Y

15*6 20,000 9765 238.14 Y 7,700 557 35.29 Y

17*10 20,000 26852 476.45 Y 5,450 543 35.37 Y

H.5. Power supply data Table: 1.7 showing the different power supply analysis

Manufacturer No.Cells (s)

Capacity (mAH) Coulomb (C)

Weight (g)

Cost (£)

Turnigy nano-tech 3 4000 25/50 333 19.26 Turnigy nano-tech 3 5000 35/70 409 30.88

Multistar 3 5200 10/20 325 20.39 Turnigy nano-tech 3 6400 40/80 506 41.51

Zippy 3 8400 30/40 772 40.19 Turnigy nano-tech 3 8400 40/80 641 55.72

Turnigy power 4 4000 40/50 476 32.68 Turnigy nano-tech 4 4000 25/50 433 26.00

UAS CHALLENGE 2015


Zippy 4 4000 20/30 399 19.98 Turnigy nano-tech 4 4500 25/50 467 31.00 Turnigy nano-tech 4 5000 65/130 576 53.71 Turnigy nano-the 4 6000 25/50 623 46.08 Turnigy power 4 7200 40/80 840 48.36

Zippy 4 8000 30/40 845 42.83 Turnigy nano-tech 5 4000 25/50 525 31.99

Zippy 5 5000 45/55 732 46.57 Turnigy power 5 5000 25/30 677 33.12 Turnigy power 5 5000 30/40 695 39.53

Turnigy nano-tech 5 5000 35/70 659 47.73 Zippy 5 5000 20/30 640 29.19 Zippy 5 8000 30/40 1054 53.65

Turnigy nano-tech 5 8000 25/50 924 63.87 Turnigy nano-tech 6 4000 25/50 623 38.36

Turnigy power 6 4500 30/40 745 51.84 Turnigy power 6 5000 35/45 812 44.52

Zippy 6 5000 20/30 754 35.60 Zippy 6 5000 30/40 784 41.30

Turnigy nano-tech 6 5000 25/50 769 51.66 Turnigy power 6 5800 25/35 914 55.58

Turnigy nano-tech 6 6000 25/50 908 57.48 Turnigy nano-tech 6 8000 25/50 1105 78.56 Turnigy nano-tech 7 4500 65/130 895 66.99 Turnigy nano-tech 7 5000 65/130 978 66.99

Zippy 7 5000 25/35 818 46.79 Turnigy power 7 5000 60/120 1025 50.46

Turnigy nano-tech 8 4400 65/130 1012 90.03 Zippy 8 4500 35/45 911 67.19 Zippy 8 5000 25/35 937 53.59

Turnigy nano-tech 8 5000 65/130 1106 110.03 Zippy 8 5800 25/35 1025 60.68

Turnigy power 8 5800 25/35 1216 66.99 Zippy 9 5000 25/35 1021 69.12

Table: 1.8 shows a section of table 1.7 which investigates different power supplies

Manufacturer Number of Lithium polymer cells (s)

Capacity (mAH) Coulomb (C)

Weight (g)

Cost (£)

Turnigy nano-tech 3 4000 25/50 333 19.26 Turnigy nano-tech 4 4000 25/50 433 26.00 Turnigy nano-tech 5 4000 25/50 525 31.99 Turnigy nano-tech 6 4000 25/50 623 38.36

Zippy 7 5000 25/35 818 46.79 Zippy 8 4500 35/45 911 67.19 Zippy 9 5000 25/35 1021 69.12

UAS CHALLENGE 2015


Table 1.9 showing increasing number of lithium cells for a calculated power consumption, current draw will

decrease

Lithium-Ion Cells (s) & Current Draw

(I)

Propeller

Dimensions (in)

Rpm to sustain

lift of 35N

Power

(W)

Rpm to sustain forward

velocity (29.46m/s)

3s 4s 5s 6s 7s 8s 9s

10*8 14,600 1000 9,676 90 67 54 45 38 33 30

Table 2.1 shows the rpm required to sustain lift and rpm required to achieve forward velocity of 29.46m/s

coupled with current draw using different lithium ion cells

Lithium-Ion Cells (s) & Current Draw (I)

Propeller

Dimensions

(in)

Rpm

to

sustain

lift of

35N

Power

(W)

Rpm to

sustain

forward

velocity

(29.46m/s)

3 4 5 6 7 8 9

10*8 14,600 1000 9,676 90 67 54 45 38 33 30

10*9 14,100 1013 8,601 91 68 55 46 39 34 30

10*10 13,750 1044 7,740 94 70 56 47 40 35 31

11*6 13,200 811 12,901 73 54 43 36 31 27 24

11*8 14,088 1316 9,676 118 88 71 59 50 44 39

11*9 12,522 1039 8,601 93 70 56 46 40 35 31

11*10 11,700 942 7,740 84 63 50 42 36 31 28

12*6 11,350 731 12,901 65 49 39 32 28 24 21

12*8 14,088 1864 9,676 167 125 100 83 71 62 55

13*6 9,900 668 12,901 60 45 36 30 25 22 20

13*8 14,088 2567 14088 231 173 138 115 99 86 77

13*10 11,270 1643 11270 148 111 89 74 63 55 49

14*13 8,670 1307 5,954 117 88 70 58 50 44 39

15*6 7,700 557 12,901 50 37 30 25 21 18 16

17*10 11270 4804 11270 432 324 259 216 185 162 144

UAS CHALLENGE 2015


Table 2.2 shows the rpm required to sustain forward velocity of 29.46m/s exceeds the rpm to sustain 35N of

lift


(I)

Propeller

Dimensions (in)

Rpm to sustain

lift of 35N

Power

(W)


velocity (29.46m/s)

3s 4s 5s 6s 7s 8s 9s

13*6 9,900 668 12,901 60 45 36 30 25 22 20

Table 2.3 shows that by increasing the rpm power consumption increases along with current draw


(I)

Propeller

Dimensions (in)

Rpm to sustain

lift of 35N

Power

(W)


velocity (29.46m/s)

3s 4s 5s 6s 7s 8s 9s

13*6 12,901 1478 12,901 133 99 79 67 57 49 44

Table 3.9 shows updated version of table 2.1

Lithium-Ion Cells (s) & Current Draw (I) Propeller

Dimensions (in)

Rpm to

sustain lift of 35N

Power (W)

Rpm to sustain forward velocity

(29.46m/s)

3 4 5 6 7 8 9

10*8 14,600 1000 9,676 90 67 54 45 38 33 30 10*9 14,100 1013 8,601 91 68 55 46 39 34 30

10*10 13,750 1044 7,740 94 70 56 47 40 35 31 11*6 13,200 811 12,901 73 54 43 36 31 27 24 11*8 14,088 1316 9,676 118 88 71 59 50 44 39 11*9 12,522 1039 8,601 93 70 56 46 40 35 31

11*10 11,700 942 7,740 84 63 50 42 36 31 28 12*6 12,901 1073 12,901 97 72 58 48 41 36 32 12*8 14,088 1864 9,676 167 125 100 83 71 62 55 13*6 12901 1478 12,901 133 99 79 66 57 49 44 13*8 14,088 2567 14088 231 173 138 115 99 86 77

13*10 11,270 1643 11270 148 111 89 74 63 55 49 14*13 8,670 1307 5,954 117 88 70 58 50 44 39 15*6 12901 2621 12,901 236 177 141 118 101 88 79

17*10 11270 4804 11270 432 324 259 216 185 162 144

UAS CHALLENGE 2015


Table 2.4 shows propellers with lowest current draw

Lithium-Ion Cells (s) & Current Draw (I) Propeller

Dimensions (in)

Rpm to

sustain lift of 35N

Power (W)

Rpm to sustain forward velocity

(29.46m/s)

3 4 5 6 7 8 9

10*8 14,600 1000 9,676 90 67 54 45 38 33 30 10*9 14,100 1013 8,601 91 68 55 46 39 34 30

10*10 13,750 1044 7,740 94 70 56 47 40 35 31 11*6 13,200 811 12,901 73 54 43 36 31 27 24 11*9 12,522 1039 8,601 93 70 56 46 40 35 31

11*10 11,700 942 7,740 84 63 50 42 36 31 28 12*6 12,901 1073 12,901 97 72 58 48 41 36 32

H.6. Current drawn

Table 2.5 showing the rc wiring rating and maximum current permissible

AWG Maximum current permissible

8 200

10 140

12 90

14 60

16 35

18 16

24 6

Table 1.7.1 shows updated version of table 1.7 in power supply data


Capacity (mAH)

Coulomb (C)

Weight (g)

Cost (£)

Total Weight

(g)

Total Cost (£)

Turnigy nano-tech 5 4000 25/50 525 31.99 2100 128 Zippy 5 5000 45/55 732 46.57 2196 140

Turnigy power 5 5000 25/30 677 33.12 2031 99 Turnigy power 5 5000 30/40 695 39.53 2085 119

Turnigy nano-tech 5 5000 35/70 659 47.73 1977 143 Zippy 5 5000 20/30 640 29.19 1920 87 Zippy 5 8000 30/40 1054 53.65 2108 107

Turnigy nano-tech 5 8000 25/50 924 63.87 1848 128 Turnigy nano-tech 6 4000 25/50 623 38.36 2492 153

Turnigy power 6 4500 30/40 745 51.84 2980 207

UAS CHALLENGE 2015


Turnigy power 6 5000 35/45 812 44.52 2436 133 Zippy 6 5000 20/30 754 35.60 2262 107 Zippy 6 5000 30/40 784 41.30 2352 124

Turnigy nano-tech 6 5000 25/50 769 51.66 2307 155 Turnigy power 6 5800 25/35 989 55.58 2967 167

Turnigy nano-tech 6 6000 25/50 908 57.48 2724 172 Turnigy nano-tech 6 8000 25/50 1105 78.56 2210 157 Turnigy nano-tech 7 4500 65/130 895 66.99 2685 268 Turnigy nano-tech 7 5000 65/130 978 66.99 2934 201

Zippy 7 5000 25/35 818 46.79 2454 140 Turnigy power 7 5000 60/120 1025 50.46 3075 151

Turnigy nano-tech 8 4400 65/130 1012 90.03 3036 360 Zippy 8 4500 35/45 911 67.19 2733 269 Zippy 8 5000 25/35 937 53.59 2811 161

Turnigy nano-tech 8 5000 65/130 1106 110.03 3318 330 Zippy 8 5800 25/35 1025 60.68 3075 182

Turnigy power 8 5800 25/35 1216 66.99 3648 201 Zippy 9 5000 25/35 1021 69.12 3063 207

Table 2.6 shows section of table 1.7.1 which identifies lowest weight and lowest costing power supplies


Capacity (mAH)

Coulomb (C)

Weight (g)

Cost (£)

Total Weight

(g)

Total Cost (£)

Zippy 5 5000 20/30 640 29.19 1920 87 Turnigy nano-tech 5 8000 25/50 924 63.87 1848 128

H.7. Motor data Table 2.7 showing updated version of table 1.5

Manufacturer/Model

Power (W)

KV (rpm/V)

Max Current

(I)

Working Current

(I)

Power Supply

Cell Range

(s)

Propeller Dimension Range (in)

Weight (g)

Cost (£)

Max RPM

BRC HOBBIES PRODUCTS

EMax GT2826-06

962 710 52 42 4-5 10*5-14-7 175 23.95 12987

Boost 0.50 800 600 55 45 3-5 12-13 295 62.95 8880 Boost 0.60 900 490 60 50 4-6 13-14 345 69.95 8702 Boost 0.80 950 340 60 52 5-7 14-15 395 79.95 7044 Boost 0.90 1000 300 65 55 6-9 16-17 455 79.95 7992

HOBBY KING PRODUCTS

Turnigy G46 925 550 55 46 4-5 12-15 303 34.43 8140 Turnigy

3508-640 550 640 30 25 2-5 10-13 98 20.05 9472

Turnigy 4250

900 540 60 55 3-5 9-12 236 13.80 7992

Scorpion SII- 1000 710 60 55 4-5 12-15 205 83.74 10508

UAS CHALLENGE 2015


3026 Scorpion SII-

3014 550 830 30 25 4-5 10-15 129 61.63 12284

Scorpion SII-3020

780 890 45 40 4-5 10-14 166 73.69 13172

Quanum MT 4010

548.3 580 24.7 20 4-6 9-12 127 23.818 10300

Quanum MT 3510

568.3 630 25.6 18 3-6 9-11 100 16.52 11188

Quanum MT 3510 V2

672.7 700 30.3 24 3-6 10-12 100 16.52 12432

E-FLITE Power 32 800 770 60 45 4-5 11-14 200 50.27 11396 Power 60 1000 470 80 65 5-6 15-17 230 73.73 8347 Power 46 925 670 55 40 5-6 12-14 290 60.32 11899

Table 2.8 shows rc brushless motors that can use 5s lithium-ion power supply

Manufacturer/Model

Power (W)

KV (rpm/V)

Max Current

(I)

Working Current

(I)

Power Supply

Cell Range

(s)

Propeller Dimension Range (in)

Weight (g)

Cost (£)

Max RPM Based on 5s

BRC HOBBIES PRODUCTS

EMax GT2826-06

962 710 52 42 4-5 10*5-14-7 175 23.95 12987

HOBBY KING PRODUCTS

Quanum MT 3510 V2

672.7 700 30.3 24 3-6 10-12 100 16.52 10360

E-FLITE Power 46 925 670 55 40 5-6 12-14 290 60.32 9916

Table 2.9 shows maximum thrust, maximum velocity, current draw, power consumption and maximum flight

time based on 83% thrust setting and 11168rpm

Maximum Thrust per motor (N)

Maximum angle (degrees)

Maximum Velocity

(m/s)

Power (W)

Current Draw (I)

Flight Time (Minutes)

34 43 20.9 800 43 5.6

UAS CHALLENGE 2015


H.8. ESC data Table 4.0 shows comparison between different esc’s

Manufacturer/Model Constant Current (I) Burst Current (I) Weight (g) Cost (£)

Platinum Pro 60 90 68 47.95

Robotbirds Pro 60 80 63 33.95

Turnigy Super Brain 60 70 50 33.49

4 max 60A 60 70 62 44.95

SimonK 60 80 63 17.49

Hobbywing 60 80 60 39.97

Table 3.0 show a section of appendix L that has two esc’s that is lowest in weight and lowest in cost

Manufacturer/Model Constant Current (I) Burst Current (I) Weight (g) Cost (£)

Robotbirds Pro 60 80 63 33.95

Hobbywing 60 80 60 39.97

H.9. Cost and weights In this section the total cost and weight of the four different components will be gathered and transferred on

table 3.1.

Table 3.1 showing total cost and weight of each component

Component Quantity Cost (£) Weight (g)

Propeller 4 15.8 100

Power supply 2 128 1848

Motor 4 95.8 700

ESC 4 135.8 252

Total 375.4 2900

At the start of the project certain limitation as stated below were set. By looking at table 3.1it can be seen

that these initial limitation has been met and even exceeded resulting in extra cost saving of £174.6 and total

weight saving of 800g.

“Initial cost limitation of £550 after taking into account structural other electrical components

Initial propulsion and power supply weight limitation of 3.7kg was set after taking into account structural,

payload and electrical components weights

UAS CHALLENGE 2015


H.10. Testing data Test Rigs & Data

In section: D table 2.4 identified seven different propellers that required further testing to identify the most

efficient propeller that would achieve the velocity and lift required for this project. In this section two test rigs

will be presented one for single motor thrust and torque testing and the other wind tunnel test rig used for

testing the velocity of the Quad-rotor. These test rigs were constructed so that an accurate value for current

draw

The first test rig was used to measure six different parameters which are all essential for this project. The

parameters that were measured are listed below:

1. Current draw

2. Power required

3. Thrust

4. Torque

5. RPM

6. Temperature

Figure: 1.7 shows the full set-up of the first test rig

The test rig in figure 1.7 works by applying throttle from the radio controller to the receiver, this then engages

the esc, which then controls the rc motor rotational speed and ultimately the rotation of the propeller. The rc

motor is attached to an aluminium tube and the aluminium tube at the bottom end is wired onto the lift scale.

When the throttle setting is increased the lift also increased which pulled the lift scale and displayed the

result on the digital read out in grams. When the motor and propeller is rotated there is also a torque

component that occurs in the same direction as the rotation, this is measures by the torque scale. The

torque values were particularly important to the stability section when it came to using the Quad-rotor

simulation model. For this testing an ammeter is used for measuring the current draw and power required

illustrated in figure 1.8. These two values were essential as the rc brushless motor and the esc both have a

specific power and current draw values that cannot be exceeded, if these values are exceeded has the

potential to damage the motor and the esc. Current draw was particularly important because this is the

deciding factor for flight time, the higher the current draw the lower the flight time will be for a given power

supply capacity. Another important factor that was tested involved using an infrared temperature sensor, this

is illustrated in figure 1.9. RPM (revolutions per minute) was investigated using an optical rpm reader

UAS CHALLENGE 2015


illustrated in figure 2.0, rpm values were particularly important in calculating the maximum velocity of the

Quad-rotor. Figure 1.8 shows ammeter used for testing Figure: 1.9 shows infrared temperature sensor

Figure 2.0 shows the rpm reader used for testing

Table 5.3 is used to log the data obtained for propeller with dimensions 11*8

Propeller Dimensions 11*8 Throttle Setting Percentage (%)

0 16.6 33.2 49.8 66.4 83 99.6

Current drawn (A) 0 2.35 6.20 11.26 19.51 37.20 49.63

Power required (W) 0 49.2 128.2 232.0 398.7 747.8 918.1

Lift (N) 0 0.29 0.70 2.71 8.45 22.68 23.86

Torque (Q) 0 0.082 0.193 0.292 0.427 0.794 0.915

RPM 0 1020 3510 5124 7514 9741 1302

Temperature (0C) 37 39 43 45 52 68 72

From this data the total flight time can be obtained using equation 3.2. In this case we have 4 motors and

reach one can draw up to 49.63A, therefore 198.52A in total. Under these values the Quad-rotor can have a

flight time of up to 4.83 minutes. The Quad-rotor can hover for much longer time than 4.83minutes as the

current draw reduces to around 11.26A, this give a total hover time of up to 21minutes. Take the

temperature of the motor as particularly important because over heating could lead to motor failure during

the competition day. Temperature results was conducted after running the motors at each motor setting to

1minute which allowed for them to heat up to a certain degree presented in table 5.3. Similar analysis as the

one in table 5.3 was conducted for the remaining six propellers and presented below.

UAS CHALLENGE 2015



𝑇𝑜𝑡𝑎𝑙 𝐶𝑢𝑟𝑟𝑒𝑛𝑡 𝐷𝑟𝑎𝑤 (𝐴)∗ 60 = 𝑇𝑜𝑡𝑎𝑙 𝑓𝑙𝑖𝑔ℎ𝑡 𝑚𝑖𝑛𝑢𝑡𝑒𝑠 Equation 3.2

16 (𝐴ℎ)

49.63 (𝐴) ∗ 4∗ 60 = 4.83𝑚𝑖𝑛𝑢𝑡𝑒𝑠

Table 5.3.1 is used to log the data obtained for propeller with dimensions 10*8


0 16.6 33.2 49.8 66.4 83 99.6

Current drawn (A) 0 2.42 7.21 14.7 22.1 36.4 48.2

Power required (W) 0 44.7 133.3 271.9 408.8 673.4 891.7

Lift (N) 0 0.24 0.57 3.71 7.1 17.1 20.2

Torque (N.m) 0 0.074 0.187 0.274 0.421 0.697 0.845

RPM 0 795 3005 4901 7521 8964 12940

Temperature (0C) 37 40 42 44 54 65 70



0 16.6 33.2 49.8 66.4 83 99.6

Current drawn (A) 0 2.43 7.24 14.9 22.5 37.1 48.9

Power required (W) 0 44.9 133.9 275.6 416.2 676.3 904.6

Lift (N) 0 0.22 0.54 3.45 6.9 17.0 19.7

Torque (N.m) 0 0.087 0.201 0.312 0.511 0.721 0.874

RPM 0 841 3117 5101 7045 9521 12972

Temperature (0C) 37 41 43 47 55 67 71

UAS CHALLENGE 2015




0 16.6 33.2 49.8 66.4 83 99.6

Current drawn (A) 0 2.47 8.2 15.4 24.1 38.2 49.4

Power required (W) 0 45.7 151.7 284.9 445.9 706.7 913.9

Lift (N) 0 0.25 0.64 5.41 9.41 21.12 22.45

Torque (N.m) 0 0.094 0.297 0.387 0.547 0.799 0.944

RPM 0 940 3201 4987 7012 9624 12984

Temperature (0C) 37 43 45 48 57 68 74



0 16.6 33.2 49.8 66.4 83 99.6

Current drawn (A) 0 3.4 12.0 22.1 34.1 47.2 52.1

Power required (W) 0 62.9 222.0 408.9 630.8 873.2 963.8

Lift (N) 0 0.27 0.75 5.87 8.98 21.47 23.1

Torque (N.m) 0 0.97 0.301 0.421 0.687 0.822 0.972

RPM 0 1074 3521 6210 7742 1045 1308

Temperature (0C) 37 46 49 54 67 72 78



0 16.6 33.2 49.8 66.4 83 99.6

Current drawn (A) 0 2.2 7.1 12.42 31.74 40.3 45.7

Power required (W) 0 40.7 131.4 229.7 587.2 745.5 845.5

Lift (N) 0 7.1 12.9 18.2 21.7 29.2 31.2

Torque (N.m) 0 0.31 0.68 0.94 1.09 1.20 1.31

RPM 0 1558 3766 6623 8051 11168 12987

Temperature (0C) 37 38 30 42 47 55 67

Now that the propeller that will be used on the Quad-rotor has been identified as 12*6, its time to analyse the

performance further. The second test rig was built with the intention to calculating the maximum velocity the

Quad-rotor can achieve under different wind speeds and also the current draw increase due to increased

headwind velocity.

UAS CHALLENGE 2015


Figures 2.1 & 2.2 illustrate the full test rig set up in the wind tunnel.

Figure 2.1 shows test rig inside wind tunnel Figure 2.2 shows test rig structure in wind tunnel

For this test

rig to

operate a

prototype

Quad-rotor

was built as

shown in

figure 2.3

Figure 2.3 shows the prototype Quad-rotor built for testing purposes.

Figure 2.3 shows the prototype Quad-rotor was built for the purpose of initial propulsion system integration

and performance testing. System integration involved all the components, ecs’s, motors, and power system

working together harmoniously, this also helped to eliminate any problems before it was wired onto the main

Quad-rotor. Some of the problems encountered involved esc throttle recalibration and esc programming, all

of which would have been difficult to accomplish once the system gets wired onto the main Quad-rotor and

the esc wires will be installed inside the arms of the copter. The prototype also allowed for the positive,

negative and signal wires to be accurately cut to side ready for installation. From the performance side of

things the KK 2 board was an important and a necessary piece of hardware to positioning the Quad-rotor in

correct pitch angle via digital read out.

Operating the test rig

The prototype Quad-rotor is set in place as shown in figure 2.1, pitch angle is set using KK 2 board and

clamped into place. The stopping pin is inserted into the rear of the test rig, this will allow for the Quad-rotor

to be powered up to 83% throttle setting without moving forward. After 5 seconds at 83% throttle setting the

stopping pin is removed. Removing the pin allows for the copter to travel in the x-axis very rapidly. To

calculate forward speed distance travelled and time is required, in figure 2.4 a pre-set distance of 0.185m is

identified. To measure time taken to cover 0.185m a standard camera was used to record the event and

using software called videopad video editor, the recording was slowed down so that time taken can be

KK 2 Board

UAS CHALLENGE 2015


calculated to cover 0.185m. Figure 2.5 to 271 shows how the events video recording was captured. This test

was conducted under different wind speeds from 0m/s to 25.6m/s which was the limit of the wind tunnel.

Table 5.3.6 identifies the velocity that the Quad-rotor can travel at under different headwind velocities.

Figure 2.4 shows pre-set distance of 0.185m Figures 2.5 shows start of recording

Figure: 2.6 shows motion capture midway at 0.0925m Figure: 2.7 shows full distance 0.185m

Table 5.3.6 shows wind tunnel velocity and corresponding Quad-rotor velocity

Wind Tunnel Velocity m/s Quad-rotor velocity achieved

m/s

0 20.2

10.28 10.4

12.86 8.4

15 5.7

20 1.2

25 Went Backwards

Effect of current draw due to headwind

From the use of the first test rig it was identified that propeller dimension of 12*6 will be used on the Quad-

rotor therefore further current draw testing was conducted using the wind tunnel. One APC 12*6 propeller

was put into the wind tunnel to investigate the aeroelastic effect on current draw. The testing was conducted

at different wind tunnel head speed of 25.6m/s. The results are shown in table 5.4

Table 5.4 shows the effect of propeller current draw increase at head speed of 25.6m/s


0 16.6 33.2 49.8 66.4 83 99.6

Current drawn (A) 0 2.23 7.5 12.8 32.4 42.2 47

Power required (W) 0 41.2 138.7 236.8 599.4 780.7 869.5

Temperature (0C) 37 38 30 41 43 51 64

UAS CHALLENGE 2015


From table 5.4 and table 5.3.5 the same propeller can be compared to each other, as stated earlier one of

the propeller is tested in ideal conditions with 0m/s head wind and the other in the wind tunnel with 25.6m/s

head wind. The results indicated that the maximum current draw increases from 45.7A to 47A but the overall

temperature of the motors decreased as a result of headwind cooling down the motors. The rpm of the

propellers in this case could not be obtained because the door to the wind tunnel at 25.6m/s was very hard

to open due to suction cause by the wind turbine.

H.11. Prop Performance Propeller Efficiency

Propeller performance is required to be recalculated to improve accuracy of the calculations. From the

University of Illinois at Urbana-Champaign (UIUC) students performed propeller analysis such as propeller

efficiency for every rc propeller that is available for use as an rc propeller, in this section their data and

graphs will be used to improve the accuracy of the calculations performed in this report.

The aim of this section is to calculate propeller thrust taking into account propeller efficiency obtained from

data that is presented by UIUC students

Engine power = Torque * rpm * 2𝜋

60 Equation 1.8

Rc motor torque can be calculated using “equation 1.9 (rc groups, 2008)”

𝑇𝑜𝑟𝑞𝑢𝑒 (𝑙𝑏 − 𝑓𝑡) = 𝐻𝑃

𝑟𝑝𝑚∗ 5252 Equation 1.9

Where

HP = horsepower of motor 1.289HP, as 1HP = 746W and the motor that is used has 962W

Revs Per Minute = 10508 rpm, this was calculated earlier

𝑇𝑜𝑟𝑞𝑢𝑒 (𝑙𝑏 − 𝑓𝑡) = 1.289

10508∗ 5252

Theoretical Torque Value = 0.6442lb-ft = 0.873N.m

Calculated Experimental Torque Value = 1.21N.m obtained from table 5.3.5

Now the engine power can be calculated using “equation 1.8 (Hart, 2013)” and data obtained from

experimental data, table 5.3.5 based on 83% thrust setting

Engine power = 1.21 * 11168 * 2𝜋

60

Engine power (N) = 1415N

Advance Ratio (J) is calculated using “equation 2.0 (mit education, 2012)”

J = 𝑉0

𝑛𝐷 Equation 2.0

Where

V0 = Forward velocity = 20.2m/s obtained from wind tunnel test

D = Propeller diameter = 0.3048m

n = 𝑟𝑝𝑚

60 = 186rev/s

Advanced ratio J can now be calculated using equation 2.0

J = 20.2

186∗0.3048 = 0.356

UAS CHALLENGE 2015


Now by using figure below propeller efficiency Vs advanced j ratio provided by UIUC propeller efficiency can

be obtained. Which in this case is 42%.

“Figure 3.1 (Ananda, 2015)” shows graph of propeller efficiency Vs advanced J ratio for propeller dimension

12*6

To calculate propeller power engine power is required as it can be seen from “equation 2.2 (Hart, 2013)”

Propeller power = Engine power * Propeller efficiency Equation 2.2

Propeller power = 1415N * 0.42

Propeller power = 594N

Therefore propeller actual thrust = 𝑝𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑝𝑜𝑤𝑒𝑟 𝑁

𝑉0 =

594

20.2= 29.5𝑁

Theoretical thrust was calculated using equation 1.3 as 34N per motor at thrust setting of 83% and the

results were shown in table 2.9, appendix G.

Experimental data using the test rig shows that thrust obtained at 83% thrust setting is equal to 29.5N.

As the propeller thrust has changed from 34N as stated in table 5.3.5 to 29.5N so will the maximum angle,

maximum velocity, power current draw and flight time. The new changes has been calculated and stated in

table 3.2

Table 3.2 shows the updated Quad-rotor performance

Maximum Thrust per

motor

(N)

Maximum angle

(degrees)

Maximum Velocity

(m/s)

Power

(W)

Current Draw

(I)

Flight Time

(Minutes)

29.5 32 20.2 800 43 5.6

UAS CHALLENGE 2015


H.12. Flight dynamics In this section “Flight Dynamics” is referenced to a research paper that was written to investigate a Quad-

rotors flight ability” (Khan, 2014), therefore flight dynamics will be investigated to further deepen the

capabilities of this copter. The Quad-rotor can be defined under two frame’s one which moves with its body

called the body frame and the other which is defined with respects to the ground, the layouts of these can be

seen in figure 1.5

Figure 1.5 (Khan, 2014)” showing body frame and inertial frame

Each of the four motors operate

and produce thrust independently for example if the Quad-rotor is required to hover at any height the thrust

must equal to its MTOW which in this case is 68.67N and reach motor will work independently to produce

one fourth of this value. If the Quad-rotor is to perform any type of manoeuvre such as pitch to move forward

then again each motor would work independently to produce the thrust required but in this case each motors

thrust value will vary. Manoeuvres such as, pitch, roll and climb. “Equation 2.3 (Khan, 2014) “ shows total

thrust required during certain pitch angle and roll angle.

𝑇 = 𝑚𝑔

𝐶𝑜𝑠(𝛳)∗𝐶𝑜𝑠(𝜑) Equation 2.3

Where

T = Thrust (N)

mg = MTOW (N)

cos(𝜃) = pitch angle

𝑐𝑜𝑠(𝜑) = roll angle

If the Quad-rotor climbs and performs any of the same manoeuvres then “equation 2.4 (Khan, 2014) “ is

considered to calculate the total thrust required.

𝑇 = 𝜌∗𝐴∗4∗𝑔∗(ℎ𝑓−ℎ)+𝑚𝑔

𝐶𝑜𝑠(𝛳)∗𝐶𝑜𝑠(𝜑) Equation 2.4

Where

ρ = Density kg/m3

A = Total propeller area m2

g = 9.81m/s2

hf = FInal altitude (m)

h = Initial altitude (m)

UAS CHALLENGE 2015


Magnitude of component vectors within the x-y and z axis are calculated using “equations 2.5, 2.6 and 2.7

(Khan, 2014)”.

𝑇𝑥 = √−𝑇2 ∗ cos(𝜃)2 ∗ (1 −1

cos(𝜃)2) Equation 2.5

𝑇𝑦 = 𝑇 ∗ 𝑐𝑜𝑠(𝜃) ∗ sin(𝜑) Equation 2.6

𝑇𝑧 = 𝑇 ∗ cos(𝜃) ∗ cos(𝜑) Equation 2.7

“Table 3.2 (Khan, 2014)” shows the pitch and roll angle that the Quad-rotor can operate within

Manoeuvre Pitch (𝜃) Roll (𝜑) 𝑇𝑥 𝑇𝑦 𝑇𝑧

Hover 0 0 0 0 + or -

Pitch Forward 0-90 0 + 0 + or -

Pitch Backward -90-0 0 - 0 + or -

Roll Left 0 0-90 0 + + or -

Roll Right 0 -90-0 0 - + or -

Pitch Forward and Roll left 0-90 0-90 + + + or -

Pitch Backward and Roll Right -90-0 -90-0 - - + or -

The manoeuvre that is required ultimately depends on the angle setting as it can be seen form table 3.2. For

example if the Quad-rotor was required to hover then both pitch and roll is required to be zero as it can be

seen in table 3.2. Another example can be considered when the Quad-rotor moves forward therefore pitch is

required, this results in a pitch forward manoeuvre with angles between 0 and 90, which requires zero roll

angle.

Table 3.3 shows thrust required to sustain hover

x-axis y-axis z-axis Total Thrust

Thrust (N) 0 0 68.6 68.8

“Table 3.4 (Khan, 2014)” shows thrust that is required by each propeller to perform a certain manoeuvre

Propeller Hover Pitch Forward/Pitch

Backward

Roll left/Roll

right

Pitch Forward and roll left/Pitch backward and roll

right

1 +

𝑍

4 +

𝑍

4 +

𝑍

4 +

𝑍

4

2 −

𝑍

4 −

𝑍

4 −(𝑇 − 𝑍) +

𝑍

4 −(𝑇 −

𝑍

2) +

𝑍

4

UAS CHALLENGE 2015


3 −

𝑍

4 −

𝑍

4 −

𝑍

4

𝑍

4

4 +

𝑍

4 +(𝑇 − 𝑍) +

𝑍

4 +

𝑍

4 +(𝑇 −

𝑍

2) +

𝑍

4

Table 3.4 is used to calculate thrust required by individual propellers example for hover can be seen in table

3.5

For hover the very first column will be used and it can be seen that each individual propeller is divided by 4

as it’s a Quad-rotor. All the values within the x and y axis is identified as zero because no pitching or roll is

required.

Table 3.5 shows thrust required by individual propeller

Propeller x-axis y-axis z-axis Total thrust

1 0 0 17.15 17.15

2 0 0 17.15 17.15

3 0 0 17.15 17.15

4 0 0 17.15 17.15

0 0 68.6 68.6

Another example can be identified in table 3.6 and 3.7 which shows forward flight with an angle setting of

320.

Table 3.6 shows thrust required by each axis and

total thrust at forward flight with angle setting of 320

From table 3.6 it can be seen that there is a variation of thrust required by different propellers, which is quite

different from hover, table 3.7 also confirms this.

Table 3.7 shows thrust required by individual propeller for forward flight.

Propeller x-axis y-axis z-axis Total thrust

1 9.09 0 14.54 17.15

2 9.09 0 14.54 17.15

3 9.09 0 14.54 17.15

4 15.61 0 24.97 29.45

Thrust (N) 42.9 0 68.6 80.9

x-axis y-axis z-axis Total Thrust

Thrust (N) 42.91 0 68.67 80.9

UAS CHALLENGE 2015


Calculating variation in thrust for forward flight

To calculate variation in thrust for forward flight table 3.6 must be obtained using equations 2.4, 2.5, 2.6, and

2.7. By using table 3.4 each propeller total thrust is obtained

Calculating Total thrust column

Propeller: 1 = 𝑍

4

Propeller: 2 = 𝑍

4

Propeller: 3 = 𝑍

4

Propeller: 4 = +(𝑇 − 𝑍) +𝑍

4

Where

Z = Total thrust in the z-axis which in this case is 68.6N

T = Total thrust which in this case is 80.8N

Now that the total thrust column has been calculated z-axis column can be calculated.

It is known that total z-axis thrust is 68.6N and total thrust is 80.9N, therefore 68.6𝑁

80.9𝑁 = 0.848. Individual thrust

for each propeller in the z-axis be calculated.

From table 3.7 it can be seen that propeller 1 has total thrust of 17.15N, therefore propeller 1 in the z-axis

can be calculated as 17.1N * 0.848 = 14.54N

Similarly the same method can be applied to propeller 4 in the z- axis, where 29.45N * 0.848 = 24.97N

The same method can be used to calculate thrust required by individual propeller in the x-axis. X-axis total

thrust is calculated as 34.76N.

42.9𝑁

80.9𝑁 = 0.530

Propeller 1 = 17.15N * 0.530 = 9.09N

Propeller 4 = 29.45N * 0.530= 15.61N

Now that the thrust for individual propeller has been obtained for performing a certain manoeuver, velocity of

each propeller and hence the voltage reduction can be calculated using “equations 2.8 and 2.9 (Khan,

2014)”.

𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦𝑝𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑁𝑜. = √𝑇𝑜𝑡𝑎𝑙 𝑡ℎ𝑟𝑢𝑠𝑡 𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑁𝑜.

𝜌∗𝐴 Equation 2.8

𝑉𝑜𝑙𝑡𝑎𝑔𝑒𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑁𝑜. = 1

𝑀𝑜𝑡𝑜𝑟 𝐾𝑉*

60∗𝑣𝑒𝑙𝑜𝑐𝑖𝑡𝑦𝑝𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝑁𝑜.

2∗𝜋∗𝑟 Equation 2.9

Taking the example used for forward flight each propellers thrust, velocity and corresponding voltage is

shown in table 3.8

Table 3.8 shows thrust, velocity and voltage for each propeller

UAS CHALLENGE 2015


Propeller No Propeller Thrust (N) Velocity using equation 2.8 (m/s) Voltage using equation 2.9 (V) Voltage required per

motor (V)

1 17.15 13.85 1.22 18.5

2 17.15 13.85 1.22 18.5

3 17.15 13.85 1.22 18.5

4 29.45 18.15 1.60 20.9

From table 3.8 it can be seen that for the Quad-rotor to attain an angle of 32 degrees the voltage required

from the batteries is 18.5V for three motors and 20.9V for one motor which can be achieved by the lithium-

ion batteries chosen for this project.

H.13. Velocity of Quad rotor Take of velocity for a Quad-rotor can be calculated based on the velocity of the air while the free stream of

the Quad-rotor is equal to zero.

𝑉ℎ = √𝑇

2𝜌𝐴 Equation 3.0

Where:

T = Thrust

𝜌 = Density kg/𝑚3

𝐴 = 𝑃𝑟𝑜𝑝𝑒𝑙𝑙𝑒𝑟 𝐴𝑟𝑒𝑎 𝑚2

Table 4.0 shows different density setting at certain altitude coupled with Quad-rotor thrust required with and

without payload

Using table 4.0 and equation 3.0 take-off velocity can be calculate:

Take-Off Velocity with payload to 30.48m = 9.9m/s

Take-Off Velocity without payload to 30.48m = 9.1m/s

Take-Off Velocity with payload to 121.92m = 10m/s

Take-Off Velocity without payload to 121.92 = 9.2m/s

Time to reach cruise altitude

Time to cruise altitude of between 100ft and 400ft can now be calculated using “equation 3.1 (Physics,

2014)”

𝑑 (𝑚) = 𝑉𝑖 ∗ 𝑡𝑖 + 1

2∗ 𝑎 (

𝑚

𝑠2) ∗ 𝑡2 Equation 3.1

Table 4.0 Density at

30.48m

kg/𝑚3

Density at

121.92m

kg/𝑚3

Mass with

Payload

(kg)

Thrust required

with Payload

(N)

Mass without

Payload

(kg)

Thrust required

without

payload

(N)

Propeller

Area

(𝑚2)

1.192 1.179 7 68.67 6 58.86 0.292

UAS CHALLENGE 2015


Where:

d = Distance m

𝑉𝑖 = Initial velocity m/s

𝑡𝑖 = Initial time s

𝑎 = 𝐹𝑜𝑐𝑒 (𝑁)

𝑀𝑎𝑠𝑠 (𝑘𝑔)= Acceleration m/𝑠2

t = time taken s

Table 4.1 Distance

(m)

Distance

(m)

Initial

velocity

(m/s)

Initial

time (s)

Force

(N)

Mass

with

payload

(kg)

Mass without

payload (kg)

Acceleration

with payload

(m/𝑠2)

Accelerati

on

without

payload

(m/𝑠2)

Time

taken (s)

30.48 121.92 0 0 86.33 7 6 12.33 14.38 6

Using table 4.1 and equation 3.1 can be rearranged to

𝑑 (𝑚) =1

2∗ 𝑎 (

𝑚

𝑠2) ∗ 𝑡2 or to calculate time to height t = √𝑑 (𝑚)∗2

𝑎(𝑚

𝑠2)

Time to height of 30.48m with payload = 2.2s

Time to height of 30.48m without payload = 2.0s

Time to height of 121.92m with payload = 4.45s

Time to height of 121.92 without payload = 4.1s

Stall

Stall for a Quad-rotor that weights 7kg will stall if the maximum tilt angle of 320 is exceeded

H.14. Flight performance Flight performance below is calculated based the flight path that would be undertaken during the

competition. Each leg of the flight path is identify by a number (i.e. [1]). The distance of each leg is

calculated using google earth.

Important Data

UAS CHALLENGE 2015


1. Flight performance calculations will be based on a worst case scenario were the Quad-rotor has a

mass 7kg and in full gust conditions throughout the flight path.

2. Having constructed a test rig and performed analysis on the propeller/brushless motor combination it

was obtained that a current draw per motor is identified as 47Amps and power required as

829Watts.

“Figure 1.6 (google, 2015)” shows example flight course provided by IMECH

Initial starting point: On the runway with no power

Table: 4.2 shows initial starting state of the Quad-rotor


[Runway] 0 0 100 0

First leg: Quad-rotor will take-off to its cruise altitude of 100ft ready to transit to its maximum pitch angle of

26.870. 100ft is used as the cruise altitude so that when it approaches the drop box it can perform a quicker

drop of time. Also a Quad-rotor cannot tilt immediately from the runway position as the propeller will make

contact with the asphalt, therefore it would require a certain height before a manoeuver is performed.

As the time is known to vertically climb to height of 30.46m and also the current draw of 37.26A per motor is

obtained from test rig based on 75% throttle setting, the battery percentage can therefore be calculated

using equation 3.2


𝑇𝑜𝑡𝑎𝑙 𝐶𝑢𝑟𝑟𝑒𝑛𝑡 𝐷𝑟𝑎𝑤 (𝐴)∗ 60 = 𝑇𝑜𝑡𝑎𝑙 𝑓𝑙𝑖𝑔ℎ𝑡 𝑚𝑖𝑛𝑢𝑡𝑒𝑠 Equation 3.2


37.26(𝐴) ∗ 4∗ 60 = 2.2 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)

Battery Capacity remaining can be calculated

𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 (𝐴ℎ) 𝑈𝑠𝑒𝑑 =2.2 ∗ 0.0167

60∗ (37.26 ∗ 4)

𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 (𝐴ℎ) 𝑈𝑠𝑒𝑑 = 0.091

Therefore

16𝐴ℎ − 0.091𝐴ℎ = 15.91𝐴ℎ

UAS CHALLENGE 2015


Battery Status %

15.91𝐴ℎ

16𝐴ℎ∗ 100 = 99.4%

Table 4.3 shows time taken and battery state from runway to cruise altitude


[Runway] to [30.46m] 0 2.2 99.4% 0

Second Leg: Now that the Quad-rotor is at a safe altitude maximum tilt angle of 26.870 can be applied. Also

using google earth the distance from runway to point [1] is calculated as 282m. Earlier it was calculated that

the Quad-rotor can achieve maximum velocity of 20.2m/s. Taking into account wind condition of 25knots

(12.86m/s) then the Quad-rotor can travel at a maximum velocity of 7.51m/s. Current draw of 47A per motor

was obtained again from the sophisticated test rig

Using the data above calculations for time and battery status can be calculated

𝑇𝑖𝑚𝑒 (𝑠) = 𝐷𝑖𝑠𝑡𝑎𝑛𝑐𝑒 (𝑚)

𝑆𝑝𝑒𝑒𝑑 (𝑚𝑠

)

𝑇𝑖𝑚𝑒 (𝑠) = 282𝑚

7.51𝑚/𝑠

𝑇𝑖𝑚𝑒 (𝑠) = 37.55𝑠


𝑇𝑜𝑡𝑎𝑙 𝐶𝑢𝑟𝑟𝑒𝑛𝑡 𝐷𝑟𝑎𝑤 (𝐴)∗ 60 = 𝑇𝑜𝑡𝑎𝑙 𝑓𝑙𝑖𝑔ℎ𝑡 𝑚𝑖𝑛𝑢𝑡𝑒𝑠


47(𝐴) ∗ 4∗ 60 = 37.55 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)



60∗ (47 ∗ 4)


Therefore

15.91𝐴ℎ − 1.96𝐴ℎ = 13.95𝐴ℎ

Battery Status %

13.95𝐴ℎ

16𝐴ℎ∗ 100 = 87.2%

Table 4.4 shows time taken and battery state from cruise altitude to point [1]


[30.5m]-[1] 7.51 37.55 87.2 282

Third Leg: After a quick turn at point [1] the Quad-rotor will travel another 842m which again was measured

from google earth. Again the velocity will be taken as 7.51m/s and current draw of 47A per motor.


7.51𝑚/𝑠

𝑇𝑖𝑚𝑒 (𝑠) = 112.12𝑠



UAS CHALLENGE 2015



47(𝐴)∗4∗ 60 = 112.12 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)


𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 (𝐴ℎ) 𝑈𝑠𝑒𝑑 =112.12∗0.0167

60∗ (47 ∗ 4)


Therefore

13.95𝐴ℎ − 5.87𝐴ℎ = 8.08𝐴ℎ

Battery Status %

8.08𝐴ℎ

16𝐴ℎ∗ 100 = 50.5%

Table 4.5 shows time taken and battery state from [1] to point [2]


[1]-[2] 7.51 112.12 50.5 842

Fourth Leg: Again at point [2] the Quad-rotor will perform a sharp turn to align itself with point [3] which 418m

away from point [2]. With velocity of 7.51m/s and current draw of 47A per motor


7.51𝑚/𝑠

𝑇𝑖𝑚𝑒 (𝑠) = 55.7𝑠




47(𝐴)∗4∗ 60 = 55.7 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)



60∗ (47 ∗ 4)


Therefore

8.08𝐴ℎ − 2.91𝐴ℎ = 5.17𝐴ℎ

Battery Status %

5.17𝐴ℎ

16𝐴ℎ∗ 100 = 32.3%

Table 4.6 shows time taken and battery state from [2] to point [3]


[2]-[3] 7.51 55.7 32.3 418

Fifth Leg: Again for this section same performance criteria can be assumed


7.51𝑚/𝑠

𝑇𝑖𝑚𝑒 (𝑠) = 44.47𝑠



UAS CHALLENGE 2015



47(𝐴)∗4∗ 60 = 44.47 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)



60∗ (47 ∗ 4)


Therefore

5.17𝐴ℎ − 2.33ℎ = 2.84𝐴ℎ

Battery Status %

2.84𝐴ℎ

16𝐴ℎ∗ 100 = 17.8%

Table 4.7 shows time taken and battery state from [3] to [Target]


[3]-[Target] 7.51 44.47 17.8 334

Hover and Navigation Leg: A this point the Quad-rotor will be hovering over the top of the target but also

navigating so that is can precisely on top of the 2x2 red square. It is estimated that it would take 20 seconds

for this to occur with current draw of 17.3A per motor

𝑇𝑖𝑚𝑒 (𝑠) = 20𝑠




17.3(𝐴)∗4∗ 60 = 20 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)


𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 (𝐴ℎ) 𝑈𝑠𝑒𝑑 =20∗0.0167

60∗ (17.3 ∗ 4)

𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 (𝐴ℎ) 𝑈𝑠𝑒𝑑 = 0.39Ah

Therefore

2.84𝐴ℎ − 0.39𝐴ℎ = 2.45𝐴ℎ

Battery Status %

2.45𝐴ℎ

16𝐴ℎ∗ 100 = 15.3%

Table 4.8 shows time taken and battery state at [Target]


[Target] 0 20 15.3 0

Final Leg: Final leg of the mission is to return from the target drop off point back to the runway


7.51𝑚/𝑠

𝑇𝑖𝑚𝑒 (𝑠) = 16.51𝑠




47(𝐴)∗4∗ 60 = 16.51 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)

UAS CHALLENGE 2015




60∗ (47 ∗ 4)


Therefore

2.45𝐴ℎ − 0.86𝐴ℎ = 1.59𝐴ℎ

Battery Status %

1.59𝐴ℎ

16𝐴ℎ∗ 100 = 9.93%

Table 4.8 shows time taken and battery state from [Target] to [Runway]


[Target]-[Runway] 7.51 16.51 9.93 124

Time Taken up to this point 288.55s (4.82minutes)

Reload Leg: At this point the Quad-rotor will be on the runway, power supply (8Ah) and

the second payload will be replaced ready for flight the estimated time for this will be 30

seconds. After the reload the same performance criteria as the first leg and can used.




37.26(𝐴)∗4∗ 60 = 2.2 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)



60∗ (37.26 ∗ 4)


Therefore

8𝐴ℎ − 0.091𝐴ℎ = 7.91𝐴ℎ

Battery Status %

7.91𝐴ℎ

8𝐴ℎ∗ 100 = 98.8%

Table 4.9 shows time taken and battery state from [Runway] to [30.46m]


[Runway] to [30.46m] 0 2.2 98.8% 0

Final Leg: The Quad-rotor will be at a height of 30.46m and will head towards the target.


7.51𝑚/𝑠

𝑇𝑖𝑚𝑒 (𝑠) = 16.51𝑠

UAS CHALLENGE 2015





47(𝐴)∗4∗ 60 = 16.51 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)



60∗ (47 ∗ 4)


Therefore

7.91𝐴ℎ − 0.86𝐴ℎ = 7.1𝐴ℎ

Battery Status %

7.1𝐴ℎ

8𝐴ℎ∗ 100 = 88.12%

Table 5.0 shows time taken and battery state from [Runway] to [Target]


[Runway]-[Target] 7.51 16.51 88.12 124

Again we will have the Hover and Navigation Leg: A this point the Quad-rotor will be hovering over the top of

the target but also navigating so that is can precisely on top of the 2x2 red square. It is estimated that it

would take 20 seconds for this to occur with current draw of 17.3A per motor

𝑇𝑖𝑚𝑒 (𝑠) = 20𝑠




17.3(𝐴)∗4∗ 60 = 20 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)


𝐵𝑎𝑡𝑡𝑒𝑟𝑦 𝐶𝑎𝑝𝑎𝑐𝑖𝑡𝑦 (𝐴ℎ) 𝑈𝑠𝑒𝑑 =20∗0.0167

60∗ (17.3 ∗ 4)


Therefore

7.1𝐴ℎ − 0.39𝐴ℎ = 6.71𝐴ℎ

Battery Status %

6.71𝐴ℎ

8𝐴ℎ∗ 100 = 83.8%

Table 5.1 shows time taken and battery state at [Target]


[Target] 0 20 83.8 0

Lastly we have final leg again: Final leg of the mission is to return from the target drop off point back to the

runway

UAS CHALLENGE 2015


Table 5.2 shows time taken and battery state from [Target] to [Runway]


[Target]-[Runway] 7.51 16.51 73.1 124


7.51𝑚/𝑠

𝑇𝑖𝑚𝑒 (𝑠) = 16.51𝑠




47(𝐴)∗4∗ 60 = 16.51 ∗ 0.0167(𝑚𝑖𝑛𝑢𝑡𝑒𝑠)



60∗ (47 ∗ 4)


Therefore

6.71𝐴ℎ − 0.86𝐴ℎ = 5.85𝐴ℎ

Battery Status %

5.86𝐴ℎ

8𝐴ℎ∗ 100 = 73.12%

Time taken from reload to landing = 55.22s (0.922 minutes)

Total Time Taken: 5.74minutes

H.15. Target recognition Target Identification

Due to the usual GPS inaccuracies it was identified that in order for the Quad-rotor to accurately locate the

2x2 square drop zone point that a system had to be implemented. In this section two different target

identification methods are discussed, first method is target identification using a camera which is

programmed to identity target and a motor to crab the Quad-rotor to the target. The second method identifies

GPS inaccuracies and corrects the latitude and longitude coordinate in accordance with the location of the

2x2 target.

Method 1 works on the bases of having two servos attached to each other as shown in fig 1. Each servo has

the capability of rotating within the range of 90 degrees for tracking purposes. To track an object a webcam

is required to optically identify the 2x2 target box, in figure 2 shows the assembly between two servo’s and a

camera. The camera will identity the 2x2 target and the servos will be used to centre the camera with the

target. These three components will identify and track the target but without propulsion the Quad-rotor will

not move anywhere. Figure 3 shows the final tracking system with a propulsion system for manoeuvrability

The propulsion system has been attached on top of the camera so that the Quad-rotor would move in the

direction that the camera is looking at. When the Quad-rotor is directly underneath the target it would not

move in any direction because the camera’s angle will be 90 degrees facing towards the 2x2 target box and

the height will be controlled by pixhawk.

This system will work independently from pixhawk and will require its own software, firmware and hardware.

Some of the hardware has already been discussed such as servos, camera and propulsion system, but

UAS CHALLENGE 2015


there are others such as the Arduino Uno were it enables the two servos and the camera to be integrated

together.

To ensure that all the hardware works with each other certain software’s programs are required to ensure

that the firmware is implemented correctly. Codes are available on request.

How the whole system will function

Method 2

The second idea involves calculating the error between google maps and the pixhawk. Firstly a location is

chosen in google maps as seen in figure 2.8 and figure 2.9

“Figure 2.8 (google, 2015)” shows an easily identifiable location for easy GPS extraction

UAS CHALLENGE 2015


“Figure 2.9 (google, 2015)” show street view of the location in figure 2.8

By identifying a specific point on google earth the coordinates can then be identified pretty accurately, e.g.

the location above has coordinates of 51038’47.13”N; 0

004’29.80”W. The next process was to take pixhawk

to that exact same location which showed coordinates of 51038’47.28”N and 0

004’29.88N which is about 3.5

meters in difference. This process is repeated over several days at different locations so that a data base is

built up and the error between the two coordinates can be calculated and inputted back into pixhawk so that

it would have the same coordinate points as google earth.

Competition day scenario

On the competition day when the target location in given, at that point it would be logged into google maps

and the location identified via surrounding structures e.g. runway. From there the coordinates will be

identified and inputted into pixhawk after taking into account the error which has been calculated. From

figure 3.0 google maps identifies the target coordinates as 53055’128.83”N; 0

058’27.94”W

“Figure 3.0 (google, 2015)” shows the target drop point zone

UAS CHALLENGE 2015

261 UAS System Set Up MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Johnathan Appendix. I UAS System Set Up H.1. Control Systems Standard Operating Procedures These procedures are step by step instructions on how to operate, maintain and test control systems components without the need of the control systems manager. These instructions would cover the following operations:

Installing the Autopilot firmware and configuring the autopilot main sensors

Implementing a flight plan

Configuring all other necessary parameters for the autopilot system

H.2. Connecting the Autopilot Controller (PIXHAWK) There are two main ways of connecting Pixhawk to a ground control station (Laptop with a Mission Planner Software) which are Serial connection and Telemetry kit. The steps are shown below:

Connect Pixhawk to the laptop, the driver necessary to make it connect should be

downloaded automatically or you can download from the PX4 official website. The first form

of connection to a GCS should be done through a USB port.

When Pixhawk is connected to the GCS, the buzzer would play a musical note to alert the

programmer that it has been connected to a power source.

After the first connection, a dedicated port is chosen by the laptop for the serial connection

and this port would show as COM5 on the GCS whenever Pixhawk is connected.

To connect Pixhawk to a GCS through a telemetry kit, the port to be used is COM3. An

easier method is to choose the option of AUTO whenever connecting Pixhawk to the GCS,

it would automatically find the correct port for whatever connection is chosen.

After the correct port is chosen and Pixhawk can then be fully connected. The following

screen would show if Pixhawk is connected properly.

Figure 161: Proof of Connection

H.3. Configuring the Autopilot Before the autopilot system can be programmed, the firmware must be downloaded to Pixhawk and this can be done through the Wizard function on Mission Planner. Another way to do this is to manually create a model of the UAV you want to build by specifying the UAV’s properties. The steps to do this are show below:

Click on the Initial Setup Button on the top menu of mission planner

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Figure 162: Mission Planner top menu

The Initial Setup menu is the environment where the firmware and all hardware are

configured whether they are mandatory or optional.

Figure 163: Initial Setup for all components

Every single component to be connected to Pixhawk is to be configured from here including all the primary on-board sensors such as the accelerometer and compass. To configure a specific one, click on it and follow the instructions.

H.4. Implementing a Flight Plan To implement a flight plan (especially for autonomous flight), GPS coordinates are required to be inputted into Mission Planner.

Figure 164: Mission Planner Waypoint Entry Point

Before any waypoints are entered, the GPS should be left for a minute or two in order to

get satellite fix and then that first fixed position would be the home or launch position.

At the above shown environment in mission planner, the longitude, the latitude and altitude

of the waypoints for a flight plan. The waypoint radius, loiter radius and default altitude are

set at this environment.

UAS CHALLENGE 2015


Other commands for secondary mission purposes such as servo commands or camera

commands for image processing or payload deployment. Flight modes such as loiter is also

set in this environment.

The commands would work at the waypoint altitude before them and as such it is not

necessary to set waypoints for the secondary commands.

Figure 165: Secondary Commands

After all the necessary commands for the flight plan has been inputted into Mission

Planner, the next step is to write it to Pixhawk Memory and start the mission; the

environment to write the at the right side of the mission planner

Figure 166: Area for writing flight plans into Pixhawk's Memory

Another method to enter in commands is to load waypoint files that have been saved in the

form of text files.

The speed of the Quad-rotor flight to waypoints can be also programmed at this

environment with a DO_CHANGE_SPEED command or at the configuration setting area

where PID values are set.

H.5. Stability and Control Procedures for Quad-rotor For the Quad-rotor to be fully stable in flight, Pixhawk programming has in built controllers that control the Quad-rotor and stabilises it if it any external disturbances are encountered in the Quad-rotor flight path. Although Pixhawk has in built controllers, the controllers need to be programmed for the different motions. The PID numbers have to be set and there are a variety of methods to get it.

To input the PID numbers, click on the CONFIG/TUNING button at the top of the Mission

Planner Software.

Click on Extended tuning; in the environment, all PID numbers can be set as well.

Waypoint Speed, radius, ascending and descending speed.

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Figure 167: Stability Tuning for Quad-rotor Control

H.6. Configuring other Autopilot Parameters The Quad-rotor flight parameters can be changed or fine-tuned in order to improve flight performance and stability. Such parameters include control systems fail-safes, sensor settings, monitoring systems settings, radio controller settings, etc.

To get to this area in mission planner, click on the CONFIG/TUNING button at the top

menu on Mission Planner.

After that, click on standard parameters button on the side menu.

A list of parameters would appear and most of them would have been set to default or

disable in order to prevent systems from malfunctioning.

For changed parameters to have effect they have to be written to Pixhawk memory, to do

that, after changing the parameter click write and to be sure it has saved, a progress bar

would appear and then disappear when the parameter is being written.

Figure 168: Mission Plannner environment for changing parameters

UAS CHALLENGE 2015


There are a lot of parameters that can be changed and as such in this manual; only a

necessary few would be shown such as GPS failsafe’s, change of speed and acceleration,

flight modes.

Instead of scrolling to find parameters, click on the find button and type in the parameter

you are looking for and it would show only that parameter and the others would disappear.

Figure 169: Fail Safe parameters

To change any fail safe parameters, click find and type fail safe, the screen should show

parameters that look like those in figure 9. They all have different options of what to do

when the fail safes are activated.

The fail safe parameters in figure 9 are all set to disabled as a default value as they can

vary depending on the type of UAS

UAS CHALLENGE 2015


Figure 170: Typical Set Fail Safe Values

Figure 10 shows the appropriate fail safe values to be set for a Quad-rotor; after these

values have been chosen click on the write button to copy them to Pixhawk memory.

These parameters would hold even after a reboot but they would be set to their default

values if Pixhawk is reset (firmware deleted or overwritten on the board).

The monitoring system parameters are shown at their defaults value in figure 11.

The monitoring system parameters include the arming check for the Quad-rotor, GPS

failsafe, throttle failsafe enable.

To be flight ready, the monitoring system should be set as shown in figure 12.

UAS CHALLENGE 2015


Figure 171: Monitoring System Values

Figure 172: Flight ready monitoring system

To change the different velocity and different acceleration for different flight modes such as

take-off and land speed, waypoint speed and also respective acceleration.

UAS CHALLENGE 2015


Figure 173: Quad-rotor Acceleration and Velocity parameters

The units of velocity and acceleration are in cm/s and cm/s2.

H.7. General remarks and safety warnings To configure the autopilot system for flight, the most important configurations are all shown above. To make sure that all configurations are done in a safe way in order to prevent damage, the following notes should be taken into consideration.

Whenever Pixhawk is to be used, the buzzer and the safety switch must always be plugged

in. The musical tones would sound when Pixhawk is booted up or when important

parameters are saved on Pixhawk hard drive.

Do not change the transmission rate for any connection methods.

Do not turn off Pixhawk without using the safety switch.

Only use Mission Planner to program the Pixhawk.

Before any flight, pre-arm check must be carried out on the components of the control

system.

Do not arm the motors when the batteries are low, if the propellers aren’t screwed on

properly and if the pre-arm check is failed.

UAS CHALLENGE 2015

269 Systems MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Malwenna Malwenna

Appendix. J Systems

J.1. AltHold Mode Tuning

Test Changes to PID values

Expectations Changes required

1 Altitude Hold P set to zero

PID at standard 3DR Quad-rotor levels

P -0.7500

I- 1.500

D- 0.000

Keeping the throttle stick in the middle will still cause a climb with very little stability

This would be done to primarily observe the effect of altitude hold P gain on the system and purpose of AltHold mode being maintaining the altitude to hover, Altitude Hold gain P will be never zero.

1.1 Altitude Hold P slightly increased


P -0.7500

I- 1.500

D- 0.000

The copter will start responding to the throttle command and will be able to maintain altitude at 40% -60% throttle

Increment in Altitude Hold P is needed for a better desired climb or desired decent rate to maintain the altitude

1.2 Altitude Hold P increased more


P -0.7500

I- 1.500

D- 0.000

The copter will start to aggressively maintain altitude while having considerable amount of oscillations

Altitude Hold P gain will better the response time to correct however may need to change by further increasing to determine the best response time. PID values will need changing since current is optimum values for the 3DR Quad-rotor which is fairly small compared to ours.

1.3 Altitude Hold P increased to a higher number


P -0.7500

I- 1.500

D- 0.000

Controlling the throttle will result in abrupt stops and starts of motors

Altitude Hold P gain should always be kept below this value since it can cause mechanical failure.

2 Changing The Throttle Rate P, D

Has to be observed since the expectation is not

As advised in the Ardupilot tuning guide, The Throttle Rate gains would not require tuning and therefor will be

UAS CHALLENGE 2015


gains


P -0.7500

I- 1.500

D- 0.000

clear kept at standard values;

P – 6.0000

D – 0.0000

3 PI at standard 3DR Quad-rotor levels

P -0.7500

I- 1.500

Increase D gain

The quad is expected to be less responsive to throttle command

D gain is usually used to damp out or limit accelerations towards desired output and since the throttle is required to archives accelerations, this should always be kept at zero

3.1 Decrease throttle Accel PI gains

Quad should be more stable with less oscillations while correcting altitude

For powerful Quad-rotors like ours, decreasing these values will better the performance. To maximize the performance, further changing will be required. However, while changing gains, P: I ratio of 1:2 will be maintained.

3.2 Decrease throttle Accel PI gains by 50%

Should be very stable while maintaining altitude and will have a good response to throttle command.

As mentioned in the Ardupilot tuning guide, the best response for a powerful quad can be achieved by this and might need slight changes approximating around these values to obtain the optimum performance.

UAS CHALLENGE 2015


J.2. Auto Tuning

Ensure the quad is properly tuned at AltHold mode prior Auto Tuning.

Procedure

1. Change one flight mode switch to AltHold.

2. Change Ch7 Opt or Ch8 Opt at Extended parameters to Auto tune.

3. Keep the Auto tune set switch in Low position.

4. Take the cat in a large open area away from the crowd to be tested.

5. Ensure there is no trim been set up in Radio controller.

6. Arm motors and take off to a desired altitude (not too high) and switch to AltHold mode to

hover.

7. Put Auto tune set switch to high position to engage auto tuning.

8. Input roll, pitch and yaw if quad starts drifting away.

9. Use switch to abandon Auto tuning if it seems too destabilized.

10. At the end of the tuning, PID gains will be changed back to original and can be monitored

through mission planner.

11. Switch Auto tune set switch to low position and back to high position to test the new PIDs.

12. Land and disarm the motors while at high position to save the new PIDs.

13. Land and disarm motors at low position to return to original PIDs

J.3. Risk Assessment

Likelihood definition

1 – 0-10% probability / Rare

2 – 11-40% probability / Unlikely

3 – 41-60% probability / Moderate

4 – 61-90% probability / Likely

5 – 91-100% probability / Very likely

UAS CHALLENGE 2015


Severity definition

1 – There will be little or no impact and need to review quarterly.

2 – There will be a nominal impact associated with small budgets and lateness impacts and

unlikely to require monitoring.

3 – There will be significant effects on the project exceeding the budget by at least 10% with at

least a 10% lateness impact.

4 – There will be a significant impact on the outcome of the project exceeding the budget by at

least 25% with at least a 25% lateness impact.

5 – The project is likely to fail exceeding the budget by at least 50% with at least a 50%

lateness impact.

Lik

elih

oo

d

5 5 10 15 20 25

4 4 8 12 16 20

3 3 6 9 12 15

2 2 4 6 8 10

1 1 2 3 4 5

1 2 3 4 5

Severity

UAS CHALLENGE 2015


I.D Risk Likeli

hood

Seve

rity

Ris

k

Lev

el

Control/Mitigation

1 Bird Strikes 1 5 5 Cannot be managed.

2 One motor failure 1 5 5 power of the motor in front of the failed to

counter the rotation about yaw axis and

guide the copter to safety.

3 Adverse weather conditions 2 4 8 Monitor weather forecast and avoid flying

in hazardous weather conditions.

4 Take-off and Landing failure 1 4 4 Use a checklist to ensure equipment are

working properly prior to take off.

5 Incorrect assembly of UAS

components

1 3 3 Use a checklist to be used prior every

flight, use setup guides and manuals

provided by equipment manufacturers.

6 Radio frequency interference 3 2 6 Keep wire/cable away from transmitters

and antennas, Use of shielding for your

wiring runs, Keep antennas as far apart as

possible, Monitor RC Channel interference

in between flights.

7 Propeller Injuries 1 5 5 Operate away from congested areas, 50m

away from all personals and structures.

8 Battery detachment 2 4 8 Use a Velcro Strap to hold the batteries.

9 Battery combustion 1 5 5 Monitor their temperature and regulate

their charging and discharging.

10 Systems compatibility issues 2 4 5 Research on compatibility and use same

suppliers

11 CAD and analysis work lost 2 2 4 Keep multiple backups

12 Suppliers delaying the

delivery of components/

material

3 3 9 Plan ahead and include a contingency in

time plan

13 Run out of budget 2 5 10 Accurate cost analysis and good planning

UAS CHALLENGE 2015


14 Insufficient time for testing 2 4 8 Stage testing earlier and include a

contingency in time plan

15 Manufacturing lab and

equipment unavailable

2 4 8 Book in advance

16 University procurement

process delays

2 3 6 Finalizing required materials and

components early and communicate with

procurement early

17 UAS overheats 2 3 6 Check for any malfunctions before running

and do not exhaust the system

18 Wind tunnel unavailable 2 3 6 Book sessions in advance, design

alternative testing methods

19 Stability and control

algorithms fail

3 5 15 Use MATLAB to validate obtained PID

values through testing

20 Project delays 3 5 15 Good planning and including a

contingency time

21 Structural failure 1 5 5 Perform FEA test and revalidate

22 Autonomy fails 2 3 6 Designed to be able to manually control

UAS CHALLENGE 2015


J.4. Diposal methods

Component The process used Environmental Hazards

Circuit Boards Open burning

Acid baths

Hazardous gas

emissions

Pollutants such as tin,

lead, glass powder

(brominated dioxin,

beryllium cadmium and

mercury) discharge into

rivers.

Gold plated components Chemical stripping using nitric

and hydrochloric acid

Burning of chips

Tin and lead discharged

directly into rivers

acidifying fish and flora.

Air emissions of

brominated dioxins,

heavy metals and

hydrocarbons can be

poisonous

Wires Stripping to remove copper

Open burning

Hydrocarbon residues,

free into water, air and

soil.

Table 29 system component disposal method

UAS CHALLENGE 2015


J.5. OSD Specification

This section is used to justify the purchase of the autopilot equipment explained below.

Scope

The UAS Equipment is the on screen display board (OSD) to view the telemetry data.

Purpose

The On Screen Display is the video output of telemetry data of the UAS and will be

connected to the autopilot control board. The OSD transmits the telemetry ground data to the

ground control station. The module chosen for purchase is the MINIM OSD V1.1.

System Description

Overview

Figure 174 Minim OSD V2.1 (unmannedtechshop, 2015)

Part Name/Number The UAS Equipment is the on-screen display board (OSD) to view the telemetry data. The

module chosen for purchase is the MINIM OSD V2.

Criteria for Selection

A number of OSD modules were evaluated under the following checklist:

Compatibility for PIXHAWK control board.

Number of telemetry data outputs

Configuration ease

Cost

Power consumption

Size

Error indication and warning system (Lost GPS Fix, stall, over speed, battery voltage

and percentage, RSSI)

UAS CHALLENGE 2015


List of Other Choices

1) MAX7456 On Screen Display OSD

2) Remzibi OSD 3DR

3) DJI iOSD Mark II

4) OSD Pro Pkg

5) Skylark Dianmu OSD

Conclusions

After comparing with other OSD modules the MEng control system group have decided that

this is the OSD module that should be bought.

Specifications

ATmega328P with Arduino bootloader

MAX7456 monochrome on-screen display

FTDI cable compatible pinout

Standard 6-pin ISP header

Two independent power sections with an LED indicator on each

Solder jumpers for combining the power sections

+5V 500mA regulator for up to +12V supply input

Solder jumper for PAL video option

Exposed test points for HSYNC and LOS

Dimensions: 0.7"W x 1.7"L (2.4" w/ pins as shown) x 0.3"H

Suppliers

The following are links where the Minim OSD Rev 1.1 can be bought.

http://www.hobbyking.co.uk/hobbyking/store/__36844__Minim_OSD_v1_1.html

http://store.3drobotics.com/products/apm-minimosd-rev-1-1

http://www.buildyourowndrone.co.uk/ardupilot-mega-minim-osd-rev-1-1.html

http://www.unmannedtechshop.co.uk/sample-marc-retro-style-summer-mid-dress/

Prices

Buidyourowndrone - £ 45.98

Hobby King - £ 13.63

Unmanned tech shop - £44.95

http://www.hobbyking.co.uk/hobbyking/store/__36844__Minim_OSD_v1_1.html

http://store.3drobotics.com/products/apm-minimosd-rev-1-1

http://www.buildyourowndrone.co.uk/ardupilot-mega-minim-osd-rev-1-1.html

http://www.unmannedtechshop.co.uk/sample-marc-retro-style-summer-mid-dress/

UAS CHALLENGE 2015


J.6. GPS Specification

This section is used to justify the purchase of the autopilot equipment explained below.

Scope

The UAS Equipment is the external GPS with compass. The GPS with Compass chosen for

purchase is the 3DR uBlox GPS with Compass Kit,

Purpose

The GPS+compass unit will be the primary means navigation and tracking of the UAS and

will be connected to the autopilot control board.

System Description

Overview

Figure 175: 3DR uBlox GPS with Compass Kit (unmannedtechshop, 2015)

Part Name/Number The UAS Equipment is the external GPS with Compass. The GPS with Compass chosen for

purchase is the 3DR uBlox GPS with Compass Kit.

Criteria for Selection

A number of GPS + Compass modules were evaluated under the following checklist:

Compatibility for PIXHAWK control board.

GPS accuracy

Configuration ease

Cost

UAS CHALLENGE 2015


Power consumption

Battery life

Battery rechargerbility

Protectiveness

List of Other Choices

1) Zubax GNSS

2) UBLOX NEO-M8N GPS GNSS

3) MediaTek MT3329 GPS V2.0

4) I2C GPS Shield Rev 2.0

Features and Specifications

ublox LEA-6H module

3-Axis Digital Compass IC HMC5883L

5 Hz update rate

25 x 25 x 4 mm ceramic patch antenna

LNA and SAW filter

Rechargeable 3V lithium backup battery

Low noise 3.3V regulator

I2C EEPROM for configuration storage

Power and fix indicator LEDs

Protective case

APM compatible 6-pin DF13 connector

Exposed RX, TX, 5V and GND pad

38 x 38 x 8.5 mm total size, 16.8 grams.

Conclusion After comparing with other GPS modules and also considering the recommendation to use

3DR uBlox GPS with Compass Kit on Pixhawk by Pixhawk manufacturer, the MEng control

system group have decided that this is the GPS module that should be bought. However,

given the unavailability, , GPS Crius CN-06 v2 was purchased instead.

Suppliers

The following are links where the uBlox GPS with Compass Kit can be bought.

http://www.buildyourowndrone.co.uk/3dr-ublox-gps-with-compass-lea-6h.html

http://www.hobbyking.co.uk/hobbyking/store/__42833__UBLOX_LEA_6H_GPS_Mod

ule_w_Built_in_Antenna_2_5m_Accuracy_V1_01.html

http://store.3drobotics.com/products/3dr-gps-ublox-with-compass

Prices Buidyourowndrone - £54.16

Also includes following

DF13 6 Position connector cable

http://www.buildyourowndrone.co.uk/3dr-ublox-gps-with-compass-lea-6h.html

http://www.hobbyking.co.uk/hobbyking/store/__42833__UBLOX_LEA_6H_GPS_Module_w_Built_in_Antenna_2_5m_Accuracy_V1_01.html

http://www.hobbyking.co.uk/hobbyking/store/__42833__UBLOX_LEA_6H_GPS_Module_w_Built_in_Antenna_2_5m_Accuracy_V1_01.html

http://store.3drobotics.com/products/3dr-gps-ublox-with-compass

UAS CHALLENGE 2015


DF 13 4 Position connector cable.

Hobby King - £35.13

3D Robotics - £ 57.24

Also includes following

Four-position cable (compass)

Five-position-to-six-position cable (GPS) for APM or PX4

UAS CHALLENGE 2015

281 Altitude control MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Reyad Malwenna Appendix. K Altitude control

Figure 176 CG calculations for the x and y-axis

Figure 177 CG calculations for z-axis

Assuming one motor is doing all the work, it requires 180g of additional thrust to compensate for CG that is off-centre by 2cm. As thrust/power ratio is not linear, increasing thrust will reduce motor efficiency.

With 1kg of payload, CG is at 0.88 cm and -1.67 cm with no payload

UAS CHALLENGE 2015


J.1. Altitude control on MATLAB PID Values Effect on Quad-rotor Desired effect Mitigation

P = 0 I = 0 D = 0

Quad-rotor does not move No expectations As this is a trial and error method, the PID values were set to zero to see their effect, before adjustments can start being made

P = 1 I = 0 D = 0

Quad-rotor reaches 100ft at just under 3 seconds but continuously oscillates between 250ft and ground level due to lack of dampening

Quad-rotor to gradually reach 100ft, from a low P value, although this may be due to the high power of the motors, with a much lower levels of oscillation

Increase proportional value to see if a higher value may increase the time to 100ft and possibly reduce the oscillation

P = 2 I = 0 D = 0

Time to 100ft reduced further to approximately 2 seconds, with a peak altitude of hardly changed while the oscillation amplitude increased with the peak just over 320ft and no sign of damping. Throttle continuously hits 100% for extended periods of time.

Small change in going to 100ft Decrease the P value which will also reduce the voltage draw

P = 0.5 I = 0 D = 0

Time to reach 100ft increased to a much more reasonable level of 5 seconds while the peak altitude reduced to just above 220ft while oscillations also reduced

Amplitude to go down alongside the number of oscillations

Reduce the P value further, but this time by a much smaller amount. However, increasing the P value increases the sensitivity, but oscillations also increase. Derivative value will need to be increased to reduce oscillation

P = 0.4 I = 0 D = 0

Very little change from above Maximum height to fall 0.4 for the P value is far too low, even if the overshoot is very high. Reducing the P value from 0.5 to 0.4 reduced has very little effect which means we’re already at the point of diminishing returns. The P value will remain at 0.5 for the meantime while the Derivative value will be dampen the oscillations and overshoot

P = 0.5 I = 0 D = 0.1

Maximum height reduced to just over 200ft once the derivative value was included. However, while the oscillations are being damped, the rate of damping is very low, taking over 250 seconds to stabilise while the error margin was about 100% above the ideal position at 201ft.

Basic assumption that if the Proportional value is low, the Derivative value may also need to be fairly low to be effective which is not the case

D value must be increased by a fairly large amount to stabilise the Quad-rotor much quicker otherwise we’ll reduce forward speed whilst continuously trying to stabilise.

P = 0.5 I = 0

Quad-rotor now stabilises at 45 seconds, a far cry from the previous 250+ seconds

Rapidly damped oscillations D value will be further increased to reduce the time to settle

UAS CHALLENGE 2015


D = 0.5 while the max altitude is at 155ft

P = 0.5 I = 0 D = 1

Max amplitude at approximate 120 seconds with oscillations fully damped by 25 seconds

Improved damping D value needs to be increased further to reduce damping time

P = 0.5 I = 0 D = 2

No overshoot. Settles at 100ft by about 15 seconds

Small overshoot D needs to be increased as settling time needs to reduced further but there is an issue here where the P value is too low and can cause the settling time to increase as D is increased

P = 0.5 I = 0 D = 4

Just as before, no overshoot but settling time is now 40 seconds

No specific expectation, simply curiosity The current D value is too high for the P value, although both are below what is ideal. D value will be returned to 2 while P will be increased

P = 1 I = 0 D = 2

Small overshoot, 7% over which is quite ideal while settling time is about 15 seconds

Settling time to reduce with a small overshoot which was close to what was desired

D value increased to see its effect as they are both assumed to be lower than ideal

P = 1 I = 0 D = 4

No overshoot while settling time did not change

Settling time to increase With the D value increased with very little change in settling time, it seems like the D value can be increased further. Although not without the P value first which will most create an overshoot but reduce the settling time

P = 2 I = 0 D = 4

No overshoot with settling time at approximately 7.5 seconds

Settling time reduced with a small overshoot

Seems like the P and D values are close and can be worked with. However, the P value will be increased to see for any positive changes

P = 4 I = 0 D = 4

Very large overshoot, with a maximum altitude of just over 120ft with no change in settling time

Overshoot with a reduction in settling time

P value will be reduced slightly to reduce the overshoot and settling time

P = 3 I = 0 D = 4

Overshoot reduced to a much more tolerable level while settling time reduce to just about 6 seconds

Reduced overshoot and settling time Seems close to where the PID values should be, therefore no more changes will be made as of yet. However, no disturbances were included where the I value will be needed

Table 30 Attitude control on MATLAB

UAS CHALLENGE 2015


J.2. Roll and Pitch control on MATLAB PID Values Effect on Quad-rotor Desired effect Mitigation

P = 1 I = 0 D = 0

Very unstable, but cannot stabilise and falls out of the sky in less than 20 seconds

Gradually pitching down before becoming unstable

D value will be increased to see if It can stabilise

P = 1 I = 0 D = 1

Stabilises fairly quickly and continues to fly

Some form of stabilisation Double values to see if there are any changes

P = 2 I = 0 D = 2

Flight quite stable Some possible small changes Seems like any PID numbers in the PD numbers will not result in a change. Disturbances will be introduced to see what effect it has

P = 2 I = 0 D = 2

With oncoming wind, the quad gets thrown off initially before flying relatively well

A small struggle with the oncoming wind

Increase the I value

P = 2 I = 1 D = 2

Flight far less smooth and corrects itself a few times

A smoother flight Increase the I value further

P = 2 I = 2 D = 2

Flight path not followed correctly whatsoever with very sharp turns which are also inaccurate

Better correction While it seems like reducing the Integral value will allow for better flight path, the P and D values will be increased to check for any improvements

P = 4 I = 2 D = 4

Flight path much more smoother, small corrections need to be made every now and then

Smoother flight While the flight path is now much smoother, it still made one incorrect adjustment. The P and D values will be increased to see if they can rectify this

P = 6 I = 2 D = 6

Smooth flight path without any need for adjustments

Flight path to be the same as if there were no disturbances

No changes required for the PID values but may require further fine-tuning with the physical test rig

Table 31 Pitch and Roll control on MATLAB

UAS CHALLENGE 2015


Test Changes to PID values

Expectations Changes required

1 P minimised I set to zero D set to zero

Very little sensitivity, gradually moving in the axis with little recovery from stabilise mode

While it may not be the case, the P will most likely required to be higher as this Quad-rotor will be quite heavy, including payload. Depending on the level of sensitivity, the P value may be increased by a large number or a low number, therefore either 2.1, 2.2 or 2.3 will be followed

2.1 P slightly increased I at zero D at zero

Much better levels of sensitivity

If a low value of P allows for a good level of control then further increasing P may allow for tighter sensitivity

2.2 P moderately increased I at zero D at zero

Improved responsiveness or possibly oscillations

If the P value required a moderate increase, then this should either greatly improve the level of control and a slight improvements required or cause the Quad-rotor to oscillate as the P value may have been set too high

2.3 P value greatly increased I at zero D at zero

Very high level of sensitivity or possibly oscillations which may be needed

As the P value had a major change, the sensitivity may be there but there is a likely chance that it will oscillate.

3 P value in/decreased I at zero D at zero

Quad-rotor should show very low levels of oscillations

If the P value will be further increased or decreased to get the Quad-rotor to show very low levels of oscillations. At this point, the P value will remain the way it is and the D value will be increased

4 P set as before I at zero D slightly increased

Low levels of recovery when pitched/rolled and lower to no oscillations

The D value at this point will most have a very small impact on the flight mode, although that will depend on how high or low the P value is. However, it will allow the quad to gradually return to zero and smoothen the flight

5 P value from previous I at zero D slightly increased

Quad-rotor should have a smoother flight once the stick has been removed to neutral and ideally all forms of oscillations gone

D may require a further slight increase to further improve the stick free movement although it may reduce the speed of response

6 P value from previous I at zero D slightly in/decreased

Quad-rotor may once again start to oscillate but at a lower level than when P was on its own

The D may be too high and oscillate and therefore require a slight decrease or slight increase of it hasn’t yet started to oscillate. Once it starts to oscillate, this will be the maximum value for D and P must be further increased to remove the oscillation

7 P increased I at zero D set as before

Quad-rotor should move a little faster with reduced oscillations

Increase the P value further

8 P value increased I at zero D as before

As the P value is increased, the responsiveness should increase until it starts to oscillate again

Once the Quad-rotor starts to oscillate, the P value needs to be increased to just below where the oscillations starts which should be where it will be set. Now the I value needs to be increased

9 P as before I slight increased D as before

The quad should no longer pull back as quickly once stick is no longer pushed

I being quite low will have very little visible effect on the Quad-rotor, which should be further hampered by the fact that it is on a test rig. The I value should still be increased a little more

10 P as before I slightly increased D as before

The quad should take a little longer to pull back

Now the quad should have a much smoother flight, if it was free of the test rig but a further increase might be beneficial

11 P as before The quad should take a At this stage the quad may stay in that

UAS CHALLENGE 2015


I slight increased D as before

while longer to stabilise once again

position for too long which means reducing it.

12 P as before I slightly decreased D as before

The Quad-rotor should fly as intended for its task

Now that the PID values have been gathered and are effective, they need to be tested under disturbances

13 P as before I as before D as before weight added to one motor

With an added weight added to one motor, it should no longer fly as smoothly as before since the quad should be trying to rectify the change in CG and moment

If the weights are causing too much of an issue then some of the PID numbers may require adjustment which will depend on the level of change in flight the extra weight causes

14 P as before I as before D as before High airflow fan used for controlled wind conditions

The high airflow should buffet the Quad-rotor but it should be able to stabilise relatively quickly

It if struggles then the PID numbers may require some adjustment

15 P as before I as before D as before

Now that the all the PID values have been obtained, they need to reduced slightly (10-20%) as during flight they will no longer be attached to the test rig and therefore the extra force will no longer be needed and could cause oscillations during flight

16 P as before I as before D as before Flown outside

Smooth movements and relatively speedy reactions when stick is pulled back

If it doesn’t fly as well as intended then the PID may need to be changed. However, this will always depend on the pilot as not all pilots prefer one type of setup over another

Table 32 Pitch and Roll tuning on the test rig

Figure 178 Overview of the Simulink model

UAS CHALLENGE 2015


Figure 179 Section to change PID values

Figure 180 Quad-rotor control mixing

UAS CHALLENGE 2015


Figure 181 Quad-rotor dynamics

Figure 182 GUI of the Quad-rotor general parameters

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289 Verification and validation MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Tarek Appendix. L Verification and validation

K.1. Verfication Matrix Requirements Verification

Requiremen

t number

Requirement Inspection Analysi

s

Demo Test Comments

2.3 The UAV shall be capable of being controlled manually via radio

control however autonomous control is preferred.

Manual control can be demonstrated

by pilot and autonomous flight will be

demonstrated by setting the flight

conditions using pixhawk.

2.4 GPS waypoint locations and delivery will be provided on the

event day hence the UAS shall be programmable in the field.

Demonstration of the programming of

the flight setting can be shown.

2.5 The UAV shall be designed to remain within the range of 1km of

the ground station

An analysis will be required to be

carried out before testing and

demonstrating the

2.6 The UAV shall be visible at a distance of 500m* from the ground

station safety pilot (0) within the operating altitudes.

Simple visual assessment of the

Quad-rotor from a set distance.

2.7 UAV shall take off from the designated take-off and landing area

(APPENDIX C), remain in steady controlled flight from take-off to

an altitude between 100-400ft AGL (Above Ground Level).

Element of ground-based assistance for take-off and landing is

permitted, with transition to automatic control subject to point

penalties (APPENDIX A).

Demonstration of the Quad-rotor

operating at the required conditions

that are taking off from a set take of

space and land in the same space.

Also fly at altitude of up to 400ft AGL.

2.8 The UAS shall be controllable in forward speed together with 3 Pixhawk will be able to achieve this

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axis control (Roll, Pitch and Yaw). and it will be demonstrated by setting

the Quad-rotor for autonomous flight.

2.9 The Ground Control Station shall display the following

information and be visible to the Operators, Flight Safety Officer

and Judges:

Current UAV position on a moving map.

Local Airspace including any No Fly Zones.

Search Area Boundaries.

Height AGL.

Indicated Airspeed (kts).

Information on UAV Health such as remaining fuel/battery,

engine/motor RPM and Orientation.

This can be demonetarised by the

use of the ground station and the

minimOSD which will show the

required information on screen.

3.1 The UAS shall have a Maximum Take-Off Mass (MTOM) of 7kg. Analysis of Quad-rotors weight can

be carried out using a scale.

3.2 The UAV control system shall have adequate sensitivity for

corrections during take-off and landing in conditions ranging

from 0kts up to winds of 5kts and gusts of 8kts.

The testing and demonstration of this

will be reviewed by using a testing

which will be able to rotate freely

demonstrating flight conditions. Part

of the test a weight will be added to

one of the arms and then time

correction until full stability.

3.3 The UAV must be designed to fly in wind conditions up to 20kts

and gusts of 25kts.

Calculations of wind conditions must

be carried out and demonstrated and

re-evaluated using the wind tunnel.

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3.4 The maximum airspeed of the UAV shall not exceed 60 Kts_IAS)

(60.4Kts_(TAS)).

The testing of the speed of the Quad-

rotor can be tested in the wind tunnel

3.5 The UAS shall operate in temperature range of -10C to 35C

including solar radiation, with an atmospheric humidity of 95%

w/w.

This can be analysed using previous

or experimntal data and running

simulation to reprent wather

conidtions effects on the Quad-rotor.

3.6 The UAV stability* shall be predictable and controllable during

the mission, including during delivery of payload

A pilot with the required license will

shall demonstrate this.

4.1 The UAV shall be able to accomplish a flight path of 2km

considering the local conditions described in 3.2-3.5 and

payload configuration described in 4.5, at a working altitude of

that described in 2.7, in the time frame described in 4.2-4.3.

Wind tunnel can demonstrate the

distance that can be flown, by finding

the flight speed and time which will

allow to determining the distance as it

will also allow simulating different

weather conditions.

4.2 A target time for completion of the mission of 120 seconds is

required for scoring of maximum points. A Penalty (-1

point/5sec) is deductible for the total time of mission going

above 120 seconds.

This can be tested phew times in

different weather conditions using a

stop watch to measure time.

4.3 The UAV must be ready to launch within 5 minutes of the

allocated timeslot.

This can be timed using a stop watch.

4.6 The UAV shall be designed to operate from within a 10m x 30m

box, orientated within 30° of the wind direction and required to

stop within the box. Landing includes touchdown and roll-out,

with the UAV required to stop within the box.

Test will be carried out to determine

the flying ability of the Quad-rotor.

Take-off and landing test at different

weather conditions will need to be

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undertaken to satisfy the requirement.

4.7 Consideration (for maximum accuracy and viability of a system

to autonomously detect an alphanumeric code shall be studied,

to compare in real time against the GPS. The ground marker

position for payload delivery is described. where a 2x2m red box

with an alphanumeric code, all placed within a white 8x8m box

border.

The simulation of the alphanumeric

detection will show the program

ability to process the image. After

processing a simple test can see

weather pixahw triggers the servo for

the payload delivery.

6.1 All radio equipment and data links must comply with EEC

directives, and must be licensed for use in the UK.

When purchasing verification of the

spec met the UK requirement and

specification requirement.

6.2 UAS shall receive (RX) and transmit (TX) data between the

ground station and UAV itself. i.e. Global Positioning System

(GPS) telemetry and health (0) data from a distance of minimum

500m of the control station.

This can be demonetarised by the

use of the ground station to locate the

Quad-rotor and the minimOSD will

show other information on screen.

6.3 The UAS shall autonomously fly around selected GPS

waypoints that shall be provided on the mission day, whilst

remaining inside the designated flying zone, and avoiding no-fly

zones.

Demonstration of the autonomous

flight of the Quad-rotor can be verified

by programing it to fly in a set area.

7.1 Batteries used in the UAV shall have bright coloured casings to

facilitate their location in the event of a crash.

Can visually analyse by seeing the

battery from a distance.

7.4 Batteries used in the UAV shall have bright coloured casings to

facilitate their location in the event of a crash.

Visually inspect the Quad-rotor to see

if the batteries are clearly visible.

7.5 The critical UAV components must be protected for water

ingress by light rain (2mm/hr).

This can be visually assessed by

looking at the parts setup.

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7.6 At least 25% of the upper, lower and each side surfaces of the UAV shall be bright coloured to facilitate visibility (see also 2.2) in the air and in the event of a crash.

Visually inspect the Quad-rotor to see

if the Quad-rotor matches the criteria.

8.1 The ‘Return Home’ command shall be capable of activation by

the safety pilot from the ground station at any time deemed

necessary.

The manual activation can be

demonstrated by manually triggering

the return to home function.

8.2 The UAV shall automatically return to the take-off / landing zone

after loss of data-link of more than 30 seconds.

This can be demonstrated by

deactivating the Tx.

8.3 The UAV shall automatically terminate flight after loss of

controllability (auto & manual TX) signal of more than 3 minutes.

Termination of the flight to return to ground station is preferred if

suitable, however a safe landing is priority allowing landing in an

open remote location away (150m) from people, trees, traffic,

other flying craft, animals and any overhead cable.

This test can be demonstrated by

working a safe location where the Tx

is disconnected allowing for a signal

loss, hence allowing the Quad-rotor

to return to home.

9.1 The UAS shall be able to demonstrate the switch between

manual and autonomous flight and vice versa for CFT

demonstration.

This will require a demonstration of

the required once the Quad-rotor is

built.

9.2 The UAS shall be required to demonstrate manoeuvrability by

flying a figure of eight as a controllability and manoeuvrability

check for CFT.

Demonstration will be required for the

Quad-rotor to pass the certification.

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K.2. Validation Test System Test Procedure Date Result

Telemeter

kit

Establishing

connection

with Tx and

Rx

Plug in the telemetry kit receiver into the laptop using ground

control software (Mission planner) and establish connection with

the transmitter connected to the Pixhawk.

09/03/2015 Setting the ground control station into auto

it will automatically select the operating

frequency allowing for a quick connection.

Transmission

rate

Tilt pixhawk into different orientation and verify the response of

the orientation on the ground control station software (verify

attitude response)

09/03/2015 After carrying out this test an observation

was made that the ground control station

did see a change of attitude but there was

phew occasions where the response

displayed was lagging as the signal was

weakened.

Test the transmission connection in door 09/03/2015 Due to the massive interference of the

indoor testing the data displayed was not

accurate (especially the GPS data)

Place the receiver indoor and the transmitter outdoor and see the

connection response along with signal strength

10/03/2015 This test established the fact of the indoor

interference of the GPS as when the GPS

was placed outdoors it was able establish a

GPS lock on to location. Hence displaying

accurate data.

Transmission

range

Test the distance of transmission at an open field. 16/03/2015 Testing the maximum transmission range

at an open filed, a distance of 450 meters

was recorded but this was limited due to

the size of the field. Therefore the distance

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connection will be further as the spec state

a distance of more than 500meters is

possible.

Video

transmissio

n

Transmission

rate

Transmit the video through pixhawk and receive the live video

feed at the ground station (check if there is a lag in the video

transmission)

11/05/2015

Transmission

range

Test if the distance of transmission is the same as in telemetry kit

test.

11/05/2015

Video display Connect the camera to the Minim OSD and verify if the display

the live feed with the correct information

11/05/2015

Image

processing

Video display Program the Minim OSD to display the battery life, altitude,

attitude and direction.

12/05/2015

Image

processing

The image processing will need to be able to identify

alphanumeric at the location of the target and translate it to a text

file.

19/01/2015 The program has been tested and it

successfully outputted the right results

Pixhawk will able to take picture when triggered, which will then

process the image to output a text file.

14/05/2015

Pixhawk Testing

pixhawk

sensors

The servo test will try to operate four servo channels connected

on the USART2 pins

14/04/2015 The servo was tested and is fully

operational after realisation that the servo

only worked with the use of the BEC.

The tone test will allow to play a tune which will indicate that

pixhawk is ready for use

02/03/2015 The test was a simple test as when power

is supplied to pixhawk it would

automatically alert the user that’s its ready

for use.

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LED test will show the different conditions of the pixhawk (when

on, off, connected)

02/03/2015 When pixhawk is first switched on the kill

switch will indicate a red light however

before disconnecting pixhawk the kill switch

must be toggled to disconnect it and that

will display a flashing light to indicate it’s

safe to disconnect.

Calibrating the RC receiver with pixhawk 16/04/2015 The radio control had been calibrated to fly

the Quad-rotor. An additional joypad was

also calibrated to operate the Quad-rotor.

Navigation

system

Signal

strength

Examine the GPS location test on different weather conditions at

different locations, to test its signal triangulation.

23/03/2015 The GPS was able to achieve its global

positioning lock even when tested inside a

building.

Plan a journey using the ground control, and test the navigation

system GPS response by moving to the destination

19/03/2015 The setting of the way point for the Quad-

rotor is straight forward either by the use of

coordinates or point selection on the map.

Propulsion performance Test the voltage usage at full power and the current drain from

the batteries

19/03/2015

Test the current drain from the engines at different wind speeds

(in wind tunnel).

27/03/2015

Time how long the battery last at full power 25/03/2015

Test the amount of weights the motors are able to carry 20/05/2015

Safety test Drop the final built Quad-rotor from 15 cm from ground to

represent landing

18/05/2015

Test the final model in the wind tunnel to see the structural rigidly 31/03/2015 After testing the Quad-rotor looked staple

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and the arms where able to handle the

conditions

Visually inspect all electrical components and wires to make sure

there is no loose wires or components that may get effected

during flight

16/04/2015 After first initial check phew concerns were

raised and a date is rescheduled for

revisiting the test. New date: 15/05/2015

Physically asses all part and components are secure together 16/04/2015 All components are securely fastened into

the Quad-rotor, however extra mounting

features should be adapted to limit the

change of battery location as CG position

keeps changing every time batteries are

mounted and dismounted.

PID

controller

test

Controllability

of the Quad-

rotor

This is carried by creating a mathematical model of the UAS and

simulating the dynamic behaviour with the use of matlab.

08/04/2015 Test where undertaken on matlab to

simulate the Quad-rotor characteristics.

The values are then applied to Pixhawk.

A test rig is built for the sole purpose of testing the UAS in order

to set its PID numbers before the initial flight.

16/04/2015 Tests have been undertaken using the test

rig. PID values have been narrowed down

and further testing must be carried out from

17/05/2015 - 29/05/2015.

The last method is an auto tune method where the UAS is flown

with a radio controller and the autopilot then auto-tunes the PID

parameters to its final values.

25/05/2015

Payload

deployment

test

Testing the

servo

Connect the pixhak to an oscilloscope and see if there is a signal

transmitted from pixhawk.

20/03/2015 The display on oscilloscope showed that

there was a signal outputted from pixhawk

but was not strong enough to operate the

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servo. Hence we decided to use a BEC as

a signal amplifier.

After connecting the servo to the pixhawk, verify whether the

servo is receiving signal to retract the pin to release the payload

using the BEC as signal amplifier.

13/04/2015 The test was carried and the Mission

Planner was able to control the servo

through pixhawk. Further understanding is

needed on the timing setting of the

deployment. As once the Quad-rotor

reaches target the servo should

automatically deploy the payload.

Deploy the payload at the set destination or target (as the Quad-

rotor would not be built yet an initial test will be carried out as

when the Quad-rotor reached a destination or target it would emit

a signal for the servo to deploy the payload.

24/03/2015 Due to setbacks this was not achieved but

a date has been rescheduled for the

18/05/2015

Flight Test Checking the

operations of

the merged

systems

After integrating all the systems together into

the final product, test the response of the

pixhawk and motors when commanded is

sent by:

Remote

controller

01/04/2015 This test was carried out on the

17/04/2015. There was a good response

from the Quad-rotor but would requires a

bit more work to find the ideal PID values.

Ground control

station

01/04/2015 This test is yet to be carried out,

rescheduled for 19/05/2015.

Test the return home function after signal is lost, during period

more than 30sec.

03/06/2015

Testing maximum flight time with motors running at full speed. 17/04/2015 Rescheduled for 15/05/2015

Testing the

ground

Plan the journey of the Quad-rotor using the ground control

software, with a set coordinates.

01/06/2015

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control

operation

Change flight settings during flight 18/03/2015 This test was done by connecting pixhawk

to a power source and

Deploy the payload at the set destination or target 02/06/2015

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K.3. Schematics Transmitter and Receiver with Video Graphics Processing Unit (VGPU) the MinimOSD

Quad-rotor Propulsion setup

Pixhawk hardware connections

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302 Telemetry kit Specification MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Micky Malwenna

Appendix. M Telemetry kit Specification 1. [3DR Radio telemetry Kit – 433 Mhz (UK)

Price: £82.45 (inc VAT) £68.71 (exc VAT) Weight: 100 Grams

2. 3DR Radio modules V2


3. 5.8 GHz High Gain

Antenna (RP-SMA)

Price: £5.45 (inc VAT) £4.54 (exc VAT) Weight: 30 Grams Specs:

303 | P a g e


Frequency band: 5600-5900MHz,

Gain: 9dbi

Interface: RP- SMA

VSWR: <1.5

Input impedance: 50 ohms

Polarization: vertical

Maximum power: 15 W

Length: 275 mm

Range: 1.2 miles

4. DF13 6 positions connector 15 cm or 30cm

Price: £1.75 (inc VAT) £1.46 (exc VAT) Weight: 2 Grams RADIOS DESCRIPTION

SPECIFICATION Processing

100 mW maximum output power

(adjustable)

-117 dBm receive sensitivity

Based on HopeRF’s HM-TRP module

RP-SMA connector

2-way full-duplex communication

through adaptive TDM

UART interface

Transparent serial link

MAVLink protocol framing

Frequency Hopping Spread Spectrum

(FHSS)

Configurable duty cycle

Error correction corrects up to 25% of

bit error

Configuration through mission planner

and & APM planner

Features 2 Interchangeable air and ground

radio modules

433 mHz since it is in UK

Micro-USB port

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6-position DF13 connector (for

Pixhawk)

Android OTG adapter cable (to

connect radio with your tablet)

Dimensions 26.7cm x 55.5 cm x 13.3 cm (without

antenna)

Power Supply voltage: 3.7-6 VDC (from USB

or DF13)

Transmit current: 100 mA at 20 dBm

Receive current: 25 mA

Serial interface: 3.3 V UART

5. E38 Bluetooth Telmetry Bridge 433mhz


6. DRONCELL –

GSM TELEMTRY

Price: £59 (inc VAT) £49.17 (exc VAT) Weight: 100 Grams The main advantage of this telemetry link is that it potentialy has much greater range, and can also be used to send/receive other information like images, or video from your drone

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FEATURES/SPECS:

LED indication for both network status

and power

Small footprint (5cm x 4.5cm)

Breadboard compatible for easy

prototyping

4.5VDC-16VDC power supply input

3.3V or 5V UART Interface (voltage-

shifting is done on board)

High serial data rate (up to 115200

baud)

GPRS communication rate (86.5 kbps

downlink) - cellular to server

communication

CSD (up to 14.4 kbps) - cellular to

cellular communication

Software configurable baud rate

Works with any SIM card

Quad band cellular connectivity

Internal switch to detect SIM card

presence

Dial and receive phone calls

(however, no microphone or speaker

interface setup)

Send and receive text messages

Send and receive Multimedia

Messages

Send and receive data to any Internet

connected computer

Send and receive data over TCP or

UDP sockets

Super long range (anywhere there is

cell reception)

High altitude (at least 10,000 feet, up

to 30,000 depending on Cell tower)

Phonebook entries and storage

Software libraries for AVR

Real time clock, synced to cellular

tower time

User set alarms

POTENTIAL APPLICATIONS: UAVs and Balloons - live data reporting

- GPS, pressure, altitude, streaming

video

set waypoints, camera commands, etc.

Cars- remote start, car alarm

notification, GPS tracking

Security systems - cars, boards, sheds,

etc.

Home automation - thermostat control,

lighting

Robots - data transfer, remote

commands

Processors and computers- data

transfer, wireless ssh, telnet

Wireless Industrial Systems - reset

computers, activate pumps

Wireless Asset tracking - GPS track

your car, your spouse, your cat

TCP/UDP DATA TRANSFER METHODS:

iPod server/client socket app

HyperTerminal

Custom socket server/client - Perl,

Python,

Previous design before change

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L.1. Payload box

Figure 183: Other CAD views

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307 Servo calculation MEng Team Project Report (7ENT1024) School of Engineering and Technology

M.1. Servo calculation

Calculation for the selection of the servo

Figure 184: schematics for the force calculations

Ffrict-sliding = μfrict-slinding x Fnorm

W = Fnorm + Fhinge

Thus; 1 = Fnorm + Fhinge

Moment about the horn; 0 = 0.0525x1-Fhx0.105 Fh = 0.5N

Fn = 0.5N Ffrict-sliding = μ frict-sliding x Fnorm

Ffrict-sliding = 0.25 x 0.5 = 0.125N

Ffrict sliding =0.125

9.81= 0.013kg. f]

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Section by Mozammel Malwenna

Appendix. N Manufacturing

Figure 185: Machined fixed bracket is CNC Router Pro 2600

Figure 186: Dry assemble of landing gear lug, pivot and the vertical landing strut

Figure 187: Slot bracket Figure 188: Turn button for servo motor

Figure 189: Support corners machined in CNC Figure 190: Triangle payload support glued with hinges

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N.1. Machining by milling machine

Figure 191: Drilling centre hole in fixed bracket Figure 192: Milling arm Pivot

Figure 193: Chamfering of movable arm support Figure 194: Smoothing surface by fly cutter

Figure 195.1-2: Drilling using slot drills

N.2. Machining by XYZ 1330 Lathe

Figure 196: High speed steel tool

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Figure 197.1-2: Machining arm pivot on lathe

N.3. Laser Cutting by Tortec Laser cutter

Figure 198.1-2 Laser Cutting of Nylon 6 sheet for main body plate

N.4. Cutting blocks by vertical bandsaws machine

Figure 199: Cutting Nylon 6.6 cast block in vertical band saw machine

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311 Appendix. O Test Rig MEng Team Project Report (7ENT1024) School of Engineering and Technology

Appendix. O Test Rig

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O.1. Initial Gimbal Test Rig Conceptual Design


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DESIGNED BYMohin

DATE

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CHECKED BY

XXX

DRAWN BYMohinuddin

This drawing is our property.It can't be reproducedor communicated withoutour written agreement.UNIVERSITY OF HERTFORDSHIRE

SCALE 1:1 WEIGHT(kg) XXX SHEET 1/1

SIZE

A4DRAWING NUMBER

ONEREV

X

DRAWING TITLE

Gimbal Frames' Assembly

Gimbal Test Rig

5

36.5

26

822

14

11

27

1314

2647.5

4

Note: All dimensions in mmunless mentioned otherwise

Top view

A

B

Detail AScale: 1:3

Detail BScale: 1:3

864.5

76.59

Side view

101.4 8

8 271.4

2nd Pin

78

8

Stand Pin

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O.2. Updated Octagonal Gimbal Test Rig Assembly


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SCALE 1:1 WEIGHT(kg) XXX SHEET 1/1

SIZE

A4DRAWING NUMBER

ONEREV

X

DRAWING TITLE

Gimbal Frames' Assembly

Gimbal Test Rig

5

36.5

26

822

14

11

27

1314

2647.5

4Note: All dimensions in mmunless mentioned otherwise

121.4 8

2nd Pin8106.4

Mid/Outter Pin

878

Stand Pin

Front View

698.3

41.1

Top View

A

B

Detail AScale: 1:3

Detail BScale: 1:3

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O.3. Octagonal Model Mount Frame Technical Drawing


Note: All dimensions in mmunless stated otherwise.Deburand polish all sharp edges.

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SCALE 1:1 WEIGHT(kg)2.51

SIZE

A4DRAWING NUMBER

ONEREV

X

DRAWING TITLE

Model Mount Frame Drawing

Gimbal Test Rig

Exact AL Box SectionReq. 5482.4mm

435.3

414.2

1000

127

436.5

Front View

Exploded View

A

B

Detail AScale: 1:4

Detail BScale: 1:4

130

135

1542

20

4TYP

47

Front View

Front View

60

60

4 TYP

10 TYP

25.4

47.3

10

25

34.6

5R

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O.4. Octagonal Mid Frame Technical Drawing


Note: All dimensions in mm unlessstated otherwise.Debur and polishsharp edges.

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DESIGNED BYMohin

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SCALE 1:1 WEIGHT(kg)1.8 Exact AL Box SectionReq. 3946.4mm

SIZE

A4DRAWING NUMBER

ONEREV

X

DRAWING TITLE

Mid Frame Drawing

Gimbal Test Rig

493.2

8246.62

25.4

Mid Frame BoxTop View

493.2

3030

4.9

67.5

Mid Frame BoxSide View

TYP TYP

1139

493.2

472.2 25

.4

Front View Isometric viewScale: 1:14A

Detail AScale: 1:3

130

135

1542

20

4TYP

47

Front view

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O.5. Octagonal Outer Frame Technical Drawing


Note: All dimensions in mm unlessstated otherwise.Debur and polishall sharp edges.

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SCALE 1:1 WEIGHT(kg)1.96

SIZE

A4DRAWING NUMBER

ONEREV

X

DRAWING TITLE

Outer Frame Drawing

Gimbal Test Rig

Exact AL Box SectionReq. 4310.4mm

538.8

30304.9

67.5

Outter Frame BoxSide View

TYP

TYP

8

538.8

269.4

Outter Frame BoxTop View

538.8

517.8

1249

Front View Exploded ViewA

Detail AScale: 1:4

130

135

1542

20

4TYP

47

Front View

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O.6. Octagonal Gimbal Test Rig Stand Technical Drawing


Note: All dimensions in mm unless stated otherwise.Deburand polish all sharp edges. All holes on the bars arepositioned at the center of each part.

AH BG

DE CF BG AH

33

22

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DESIGNED BYMohin

DATE

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DRAWN BYMohinuddin


WEIGHT(kg) 2.7

SIZE

A3

DRAWING NUMBER

Stand Assembly andParts' Drawing REV

X

Gimbal Test Rig

Exact AL Box Section Req.6019.2mm

960

30

30

4.9

45

67.5

Slant Stand BoxSide ViewTYP

960

20

40

Slant Stand BoxTop View

215

30

308

107.5

67.5

Stand Top BoxSide View

Slant Base BracketIsometric View

5R40

20

25.4

Slant Base BracketTop View

1.2

135

80

80

Slant Base BracketSide View

L BracketIsometric View

40

10

5R

L BracketTop View

1.290

60

60

L BracketSide View

Stand Long Leg BoxIsometric View

40

87.3

300

25.4

Stand Long Leg BoxSide View

Stand Short Leg BoxIsometric View

137.3

25.44 0

10

Stand Short Leg BoxSide View

40

31.1

137.3

Stand Short Leg BoxTop View

Stand Isometric View Front View

130

135

1542

20

4TYP

47

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O.7. Gimbal Test Rig Weight / Cost Estimation


Model Mount Frame

COMPONENTS Material Item Exact Length (mm) Qty Cost (£) Ex VAT

Box Auminium Square Tube (1.626mm thickness/ 16swg) 5482.4

(5m lengths) 1 NB: Extra 482.4mm taken from mid

frame left over 16.32

Brackets Auminium Sheet (1.2mm) (80mmx80mm) x 48 (1mx1m) 1 21.1

Pins Stainless Steel Round bar 611.6 (8mm Dia 303) 1m x 1 7.33

Mid Frame

Material Item Exact Length (mm) Qty Cost (£) Ex VAT

Auminium Square Tube 3946 (5m lengths) 1 16.32

Outter Frame


Auminium Square Tube 4310.4 (5m lengths) 1 16.32

Stand


Auminium Square Tube 6019.2 (5m lengths) 1 16.32

Extra Material Auminium Square Tube (5m lengths) 1 16.32

TOTAL REQUIRED LENGTH (mm) 19758

TOTAL NO. OF REQUIRED 5m LENGTH BOXES 3.9516 4 5

Total Cost (£) Ex VAT 110.03

Purchase Cost (£) Inc VAT 132.08

Per person 11.00666667 £11.50

http://www.metals4u.co.uk/stainless-

steel/round/8-mm-diameter-

303/detail.asp?prd_id=1686

http://www.ascmetals.com/downloads/wholebrochure.pdf

http://www.rapidtables.com/calc/wire/swg-to-mm.htm

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O.8. Gimbal Test Rig Manufacturing Cost

O.9. Qualification test plan

On completion of all manufacturing activities and processes, all aerospace qualification hardware shall be tested according to and following the order of the qualification test plan found in Table 1. (Ditom, 2014)

Test Description

Initial Electrical Performance Test

Storage Temperature Cycling

Electrical Performance Test

Thermal Shock


Sine Vibration


Random Vibration


Operational Temperature Cycling

Final Electrical Performance Test

Table 1 Qualification Test Plan

Electrical Performance Tests (Initial, In-Process, Final) To verify electrical performance of the isolator/circulator, electrical performance measurements shall be performed. Measured data displaying insertion loss, Voltage Standing Wave Ratio (VSWR) (every port), and isolation (isolator only) performance over the full operating bandwidth shall be captured for each test. During the initial and final electrical performance tests, RF leakage performance shall also be measured at the center frequency of operation. However all electrical performance tests shall be captured on a calibrated Vector Network Analyser (VNA) given sufficient time to warm up and kept in ambient

conditions (18-26℃) for the entire duration of the test.

Storage Temperature Cycling Non-operational temperature cycling shall be performed to ensure the hardware meets all electrical performance specifications after being exposed to the storage temperature range. The hardware shall be exposed to each temperature extreme for a minimum of 1 hour. The

rate of change between each temperature extreme shall not exceed 20℃/minute. The hardware shall be kept at ambient conditions for no less than 1h after the test is complete prior to electrical performance measurements.

Thermal Shock Thermal shock testing shall be performed to ensure the hardware can survive rapid changes in ambient temperature without any degradation to its coatings, surfaces or electrical performance.

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Random/Sine Vibration Random vibration testing shall be performed to ensure the hardware can survive the vibrations associated with the launch and ascent of the Quad-rotor without any degradation to its coatings, surfaces, or electrical performance.

Operational Temperature Cycling Operational temperature cycling shall be performed to ensure the hardware meets all electrical performance specifications while being exposed to the operational temperature range. All the above mentioned qualification tests have to be performed at regular intervals to ensure worthy performance of the Quad-rotor. The results obtained shall also be verified and validated to meet certain conformances. The results from various tests should also be recorded and checked.

O.10. Initial Involvement in the MEng Team Project I was allocated to a few tasks in the very early stages of this project, the details are as follows:

Write up of the rules and conditions derived from the UAS challenge 2015 handbook

Compiling product design specification document

Research on testing strategies and experiments

Refining design of the gimbal test rig for more than 5 times

The CAD model shown below was the very first concept of the gimbal test rig presented at the PDR. Conversely design iterations had led to changes in to a more robust design discussed in the previous chapters.

The figure on the top right shows an angled bracket that was planned to be used in the updated octagonal gimbal test rig. However lack of facilities present at the university’s fabrication workshop, they had to be subcontracted. But to save finances and quicken the manufacturing stage of the test rig, a tri angular bracket as shown below on the bottom right figure was designed to replace the angled bracket from above. The use of these brackets had tremendously helped in speeding up the fabricating process of the final design as shown on the bottom left figure.

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O.11. Tri Angular Bracket Technical Drawing



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Joint Bracket

Gimbal Test Rig

130

135

15

42

20

4TYP

47

12

24

Front viewScale: 1:1

MATERIAL - MACHINE FROM:ALUMINIUM ALLOY (AL-2024-T3)1.2MM WORK HARDENED AL SHEETSUPPLIER: METALS4UQUATITY REQUIRED: 56

Isometric viewScale: 1:1

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O.12. T-Bracket Technical Drawing



MATERIAL - MACHINE FROM:ALUMINIUM ALLOY (AL-2024-T3)1.2MM WORK HARDENED AL SHEETSUPPLIER: METALS4UQUATITY REQUIRED: 8

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SCALE 1:1 WEIGHT(kg) 0.023 SHEET 1/1

SIZE

A4DRAWING NUMBER

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DRAWING TITLE

T BRACKET

Gimbal Test Rig

60

60

10

25

47.3

4 TYP

10 TYP

5R

34.6

2 5.4

Front viewScale: 1:1

Isometric viewScale: 1:1

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332 Design features for business case MEng Team Project Report (7ENT1024) School of Engineering and Technology

Section by Osman sibanda

Appendix. O Design features for business case

Figure 200 -OXV in storage configuration

Figure 201 - Electro-optic camera on the OXV

Figure 202- Main body of the OXV

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N.1. Risk Assessment for business case The table below quantifies the risks by the probability of the risk occurring and the impact it would have on the business. The table below it numbers these quantities by using the probability x impact ratings. The ratings are explained below.

Impact

Probability

1 2 3 4 5

1 1 2 3 4 5

2 2 4 6 8 10

3 3 6 9 12 15

4 4 8 12 16 20

5 5 10 15 20 25

Where; Impact rating 1 – There is little or no impact at all 2 - Nominal risk 3 - Significant effect on project 4 - Significant impact on outcome 5 – Project may fail and affects organisation function Probability rating 1 0-10% 2 11-40% 3 41-60% 4 61-90% 5 91-100%

#

Impact

Probability

Trivial Minor Moderate Major Extreme

Rare Low Low Low Medium Medium

Unlikely Low Low Medium Medium Medium

Moderate Low Medium Medium Medium High

Likely Medium Medium Medium High High

Very likely Medium Medium High High High

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Potential Failure Mode/Risk

Effect on business

Possible Cause

Impact rating

Likelihood of occurrence

Risk Rating

Remedial Action(s)

Manufacturing

Incorrect assembly of UAS components

Product will not be launch

Lack of executing incorrect procedure

4 2 8

Use a checklist to be used prior every flight, use setup guides and manuals provided by equipment manufacturers.

Suppliers delaying the delivery of components/ material

Delay in manufacturing

Delays in shipping

3 3 9

Plan ahead and include a contingency in time plan

Systems compatibility issues

Product will not operate

Lack of validation

2 4 8

Research on compatibility and use same suppliers

Testing

Legislation changes

Affect and delay project delivery

CAA review 5 1 5

-dedicate a team to follow up and anticipate changes -Get involved with the governing bodies in order to influence changes

Exceeding allowable noise pollution

Product not allowed to be flown

Specifications not well defined

3 2 6

-specifications should be well defined -proper testing should be carried out

Insufficient time for testing

Delay in launch and poor quality

Poor project management

2 4 8

Stage testing earlier and include a contingency in time plan

Stability and control algorithms fail

Product will not meet design specification

Inaccurate stability analysis

5 3 15

Use Matlab to validate obtained PID values through

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testing

Flight Operations

UAS overheats

Product will not sustain in market

Electrical performance test not conducted appropriately

2 3 6

Check for any malfunctions before running and do not exhaust the system

Technology advancement

Business will not sustain longer

Market competition

4 1 4

Allocate Research and Development Team

Bird Strikes Loss of economy

Unexpected encounter

1 5 5 Cannot be managed.

One motor failure

Product will not in operation

Power supply failure

1 5 5

Power of the motor in front of the failed to counter the rotation about yaw axis and guide the copter to safety.

Adverse weather conditions

Reduction in gross sale

N/A 2 4 8

Monitor weather forecast and avoid flying in hazardous weather conditions.

Take-off and Landing failure

Product will be damaged

Testing not conducted appropriately

1 4 4

Use a checklist to ensure equipment are working properly prior to take off.

Autonomy fails

Product may be irresponsive and potential chance of collision

-Error in system code - Power loss in the product’s software

2 3 6

Designed to be able to manually control

Radio frequency interference

Product launch embarrassment and will not operate as desired

Presence of other radio source

3 2 6

Keep wire/cable away from transmitters and antennas, Use of

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shielding for your wiring runs, Keep antennas as far apart as possible, Monitor RC Channel interference in between flights.

Propeller Injuries

Bad reputation Poor quality control

1 5 5

Operate away from congested areas, 50m away from all personals and structures.

Battery detachment

Loss of economy

Business specification was not met

2 4 8 Use a Velcro Strap to hold the batteries.

Battery combustion

Reduction in gross sales due to safety issues

Inappropriate procedure followed during wiring

1 5 5

Monitor their temperature and regulate their charging and discharging.

Hackers

Committed crimes, possible accidents

Intentional suspect/ Pre-planned by criminal

5 2 10

Automatic return home override in case of any control interruptions

UAS theft hacking the system or vandalism

Hacking purposes, selling product parts and criminal activity

1 2 2

- The vehicle must have CCTVs mounted in and around outside -Allow immediate control of vehicle by CT staff -alert CT staff of anything suspicious or out of ordinary

Mechanical Failure

Some system might not function or vehicle may not start

Old parts in the vehicle or not maintained properly

3 1 3

-certify regular maintenance for all vehicle parts and system -adequate

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and professional training provided to maintenance crew

System Failure

product may be irresponsive and potential chance of collision

-Error in system code - Power loss in the product computer

5 1 5

- Substantial product system testing. - Emergency Stop button must be present in the product. - Remotely monitor product - Failsafe system embedded in the product

Staff Changes

Project delay

Retirement, illnesses, demotion, strategic focus

2 3 6

identify skill shortages and act accordingly -sub-contract suitable skilled workers from different organisation

Public damaging the UAS

Loss of business

Poor public awareness for product operation

5 2 10

Regular product usage presentation and advertisements

Disposal

Decomposition of materials

High manufacturing cost

Wrong choice of materials

3 2 6

-thorough analysis before material selection

Carbon foot print

High operational costs

Business expansion

3 2 6

-strategic business plan for product distribution.

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