imeche report final_fixed
TRANSCRIPT
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UAS CHALLENGE 2015
CRITICAL DESIGN REVIEW
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Table of Contents 1 Project management ...................................................................................................................... 1
1.1 The Project Organization Structure .......................................................................................... 1
1.2 Project Planning ........................................................................................................................ 2
1.2.1 Milestones ......................................................................................................................... 3
1.3 Team Communication .............................................................................................................. 3
1.4 Project Budgeting ..................................................................................................................... 4
1.4.1 Summary of Project Budget .............................................................................................. 5
2 Quad-Rotor Design .......................................................................................................................... 5
2.1 Design Rationale - Quad-Rotor ................................................................................................. 6
3 UAV Mass Breakdown ..................................................................................................................... 6
4 UAV Cost Breakdown ...................................................................................................................... 7
5 Structural Analysis........................................................................................................................... 7
5.1 Load Case Definition and Free Body Diagrams......................................................................... 7
6 UAV Stress Analysis ......................................................................................................................... 9
6.1 Pressure Loading on Plates ....................................................................................................... 9
6.2 Load Transfer ............................................................................................................................ 9
Fixed and Movable Arm Stress Maximum ...................................................................................... 9
6.3 Simplified Plate Deflection ....................................................................................................... 9
6.4 Plate Deflection - Assembly Contact Model as Built .............................................................. 10
6.5 Undercarriage Buckling Calculation ....................................................................................... 10
6.6 Undercarriage Bending ........................................................................................................... 10
6.7 Undercarriage Bending - Assembly Contact Model ............................................................... 11
6.8 Undercarriage Torsion ............................................................................................................ 11
6.9 Undercarriage Combined Loading - Torsion and Bending ..................................................... 11
6.10 Undercarriage Combined Loading - Assembly Contact Model ........................................... 12
6.11 Simplified Analytical Modal Analysis – Fixed Arm .............................................................. 12
6.12 Simplified FEA Modal Analysis – Fixed Arm ........................................................................ 12
6.13 As Built FEA Modal Analysis – Fixed Arm ............................................................................ 13
6.14 Summarised Margin of Safety Table ................................................................................... 13
7 Performance & Propulsion ............................................................................................................ 14
7.1 Introduction ............................................................................................................................ 14
7.2 Take-Off Velocity .................................................................................................................... 14
7.3 Time To Reach Cruise Altitude ............................................................................................... 15
7.4 Max Velocity ........................................................................................................................... 15
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7.5 Stall ......................................................................................................................................... 16
7.6 Range, Power Consumption, Battery Life ............................................................................... 16
8 Rationale for Design Specification/Selection ................................................................................ 17
8.1 Autopilot ................................................................................................................................. 17
8.1.1 Specification (3dr, 2014): ................................................................................................ 17
8.2 On screen display board (OSD) ............................................................................................... 17
8.2.1 Specifications: ................................................................................................................. 18
8.3 GPS system ............................................................................................................................. 18
8.3.1 Specification: ................................................................................................................... 18
8.4 Telemetry kit .......................................................................................................................... 19
8.4.1 Specification: ................................................................................................................... 19
8.5 Camera ................................................................................................................................... 19
8.5.1 Specification .................................................................................................................... 19
8.6 Servo ....................................................................................................................................... 20
8.6.1 Specifications /Rational ................................................................................................... 20
8.7 BEC .......................................................................................................................................... 20
8.7.1 Specification / Rational ................................................................................................... 20
9 UAS Sub-systems ........................................................................................................................... 21
9.1 Navigation System .................................................................................................................. 21
9.2 Flight Control System ............................................................................................................. 21
9.3 Communication System .......................................................................................................... 23
9.3.1 Serial Connection ............................................................................................................ 23
9.3.2 Telemetry Kit Connection................................................................................................ 24
9.3.3 Radio Connection ............................................................................................................ 24
9.4 System Schematics ................................................................................................................. 24
10 Payload box mechanism ............................................................................................................ 25
10.1 System of the payload box .................................................................................................. 25
10.1.1 Controlling the servo as a servo .................................................................................. 25
10.1.2 Testing with the Mission Planner ................................................................................ 26
11 Manufacturing ............................................................................................................................ 27
11.1 Design Phases ..................................................................................................................... 27
11.2 Materials Selection ............................................................................................................. 27
11.2.1 Nylon .................................................................................................................................. 27
11.2 .2 Nylon 6.6 (PA 6.6 Black Cast Sheet) and Nylon 6.6 Rod (PA 6.6 Dia 25mm Rod) ............. 27
11.2.3 Nylon 6 (PA 6 Extruded Sheet) ........................................................................................... 27
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11.2.4 PVC (Polyvinyl chloride), (Black hard plastic tube/Rigid angle section) ............................. 28
11.2.5 Aluminium Alloy (AL-2024-T6) ........................................................................................... 28
11.3 Machining Selection ............................................................................................................ 28
11.3.1 Machines ............................................................................................................................ 28
11.3.2 Tools ............................................................................................................................ 28
11.4 Manufacturing process of Quad-rotors components ......................................................... 29
11.4.1 Fixed Bracket ...................................................................................................................... 30
11.4.2 Motor arm end bracket ...................................................................................................... 30
11.4.3 Movable arm vertical fixed bracket /support bracket ....................................................... 30
11.4.4 Landing gear top support bracket ...................................................................................... 30
11.4.5 Main Body Plate ................................................................................................................. 30
11.5 Overview of Machining ....................................................................................................... 31
11.5.1 Milling Machines (Bridgestone Series 2) ............................................................................ 31
11.5.2 XYZ 1330 Lathe ................................................................................................................... 31
11.5.3 Tortec Laser cutter ............................................................................................................. 31
11.5.4 Vertical Bandsaws Machine ............................................................................................... 31
11.5.5 CNC Machines (Router 2600 Pro and VMC 1300) .............................................................. 31
12 Testing ........................................................................................................................................ 31
1.1. Octagonal gimbal test rig ....................................................................................................... 31
12.1 Weight estimation for octagonal test rig ............................................................................ 32
12.1.1 Cost breakdown for octagonal test rig .............................................................................. 32
12.2.2 Structural testing ............................................................................................................... 32
12.2 Material Testing .................................................................................................................. 32
12.3 Component Testing ............................................................................................................. 33
12.4 Payload drop testing ........................................................................................................... 33
12.5 Qualification test plan ......................................................................................................... 34
12.5.1 Electrical Performance Tests (Initial, In-Process, Final) ............................................... 34
12.5.2 Storage Temperature Cycling ...................................................................................... 34
12.5.3 Thermal Shock ............................................................................................................. 34
12.5.4 Random/Sine Vibration ............................................................................................... 34
12.5.5 Operational Temperature Cycling ............................................................................... 34
12.6 Verification and Validation Test matrix .............................................................................. 35
13 Safety Case ................................................................................................................................. 35
13.1 Overview ............................................................................................................................. 35
13.2 Hazardous Components ...................................................................................................... 35
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13.3 Flight Controller safety mechanism .................................................................................... 36
13.1 Safety measures for flight testing ....................................................................................... 36
13.1 Description of functionality for flight termination cases .................................................... 36
13.1 GPS loss ............................................................................................................................... 36
14 UAV Technical Specifications ..................................................................................................... 40
REFERENCES ........................................................................................................................................ 161
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Table of Figures Figure 1 - Project Organization Chart ..................................................................................................... 1
Figure 2 - Show progress to date of the project ..................................................................................... 3
Figure 3 - Quad-rotor design ................................................................................................................... 5
Figure 4 - Stowage Instructions .............................................................................................................. 6
Figure 5 - Quad-rotor in Stowed Configuration ...................................................................................... 6
Figure 6 – Free Body Diagram - Flight and Landing Cases ...................................................................... 7
Figure 7 - Free Body Diagram - Landing Cases ........................................................................................ 8
Figure 8 - Free Body Diagram - Flight and Gust Load Cases.................................................................... 8
Figure 9 – Fixed Arm Cross Section – See also Appendix H.5 ................................................................. 9
Figure 10 - Flight and Gust condition of Main Body with 0.13mm Deflection ..................................... 10
Figure 11 - Lateral Impact Case on Single Leg - 60.6MPa Stress ........................................................... 11
Figure 12 – Stress Element A with Principal Stress for - Analytical – Undercarriage Combined Loading
– Bending, Buckling and Torsion (H.11) ................................................................................................ 12
Figure 13 – Simplified Fixed-arm modal analysis with 1st Natural Frequency at 19.6 Hz .................... 12
Figure 14 – As-Built Fixed-arm modal analysis with 1st Natural Frequency at 451 Hz ........................ 13
Figure 15 – Pixhawk (3DR, 2014) .......................................................................................................... 17
Figure 16 - MinimOSD v2 (APM, 2014) ................................................................................................. 17
Figure 17 - 3DR uBlox GPS (3drobotics, 2014) ...................................................................................... 18
Figure 18 – 3DR telemetry kit V.2 ......................................................................................................... 19
Figure 19 - Mobius Action Cam (UNMANNEDtech, 2014) .................................................................... 19
Figure 20 – MG90S Servo ...................................................................................................................... 20
Figure 21 – SBEC26 - Turnigy ................................................................................................................ 20
Figure 22- Waypoint Command File ..................................................................................................... 21
Figure 23- Telemetry Information transmitted to ground control station ........................................... 23
Figure 24 - Transmission Link Statistics (Serial Connection) ................................................................. 24
Figure 25- Transmission Link Statistics (Telemetry Kit) ........................................................................ 24
Figure 26 - Configuration of the servo on Pixhawk ............................................................................... 25
Figure 27- Verification of the performance of the Servo ...................................................................... 26
Figure 28 -Updated Octagonal Test Rig Assembly ................................................................................ 32
Figure 32 - Payload drop test ................................................................................................................ 33
Figure 33 - Overall View of Quad-Rotor ................................................................................................ 49
Figure 34 - Motor Mount Design (Left) & Undercarriage T-Joint (Right) .............................................. 49
Figure 35 - Undercarriage Pivot Design (Left) & Main Body Sandwich Design (Right) ......................... 49
Figure 36 - Movable Arm Pivot Design ................................................................................................. 50
Figure 37 - Project Main Body Area ..................................................................................................... 88
Figure 38 – SOLID187 Element (Ansys, November 2013c) ................................................................... 92
Figure 39 – PLANE182 Element (Ansys, November 2013c) .................................................................. 92
Figure 40 - Fixed-arm Cross-section ...................................................................................................... 96
Figure 41 - Arm Cross-section for Stress Calculation ............................................................................ 96
Figure 42 - Tension & Compression Stress in Arm ................................................................................ 96
Figure 43 – Mesh for Fixed-arm Assembly – Values as per Appendix H.4 ............................................ 97
Figure 44 - Deflection of Fixed-arm Assembly (Flight Loads) with 7.6mm Deflection .......................... 97
Figure 45 - Stress of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa
(Peak) .................................................................................................................................................... 98
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Figure 46 – Stress (Close-up) of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and
20MPa (Peak) ........................................................................................................................................ 98
Figure 47 – Mesh for Arm Assembly (With additional Tab) – Mesh Values as per H.4 ........................ 99
Figure 48 - Modified FB-002 for reduction in point contact stress concentration ............................. 100
Figure 49 - Stress Concentration at Arm (without addition) Contact (a) & Close-up (b) .................... 100
Figure 50 - Deflection of Modified Movable Motor Arm of 7.88mm for flight loads with SF ............ 101
Figure 51 - Stress of Modified Movable Motor Arm of 20.8MPa for flight loads with SF .................. 101
Figure 52 - Modified Movable Motor Arm with Stress of 20.8MPa for flight loads with SF (a) & Close-
up (b) ................................................................................................................................................... 101
Figure 53 - Load on the Lug (Niu, 1988) .............................................................................................. 102
Figure 54 - Components of the Load (Niu, 1988) ................................................................................ 102
Figure 55 - Areas on the Lug ............................................................................................................... 102
Figure 56 - Lug Bracket Without Flange (Left) & with additional Flange (Right) ................................ 103
Figure 57 – Lateral Unit Load Deflection (Left) & Stress (Right) of Lug Bracket Without Flange ....... 104
Figure 58 – Lateral Unit Load Deflection (Left) & Stress (Right) of Lug Bracket With Flange ............. 104
Figure 59 - Mesh for MP-001 (Appendix C.7) with values as per Appendix H.4 ................................. 105
Figure 60 – Motor Plate Deflection (0.038 mm) and Stress (41.7 MPa) for flight case with SF at start-
up ........................................................................................................................................................ 105
Figure 61 - Error Elements in Model - Due to Separation at FB-001 and EB-001 ............................... 105
Figure 62 - Simplified Plate Representations ...................................................................................... 106
Figure 63 - Simple Plate Deflection Carried out on CATIA structural analysis .................................... 107
Figure 64 - Mesh of Main Body Plate - Values as per Appendix H.4 ................................................... 108
Figure 65 – Single Main Body Plate Analysis – with 17.8MPa Stress at contact holes for flight case
with pressure load .............................................................................................................................. 108
Figure 66 – Mass Representation of components and payloads as per Appendix D ......................... 109
Figure 67 - Mesh of Main body assembly with Values as per Appendix H.4 ...................................... 109
Figure 68 – Contact model Flight Case for Main body assembly Deflection (left) and Equivalent Stress
(right) .................................................................................................................................................. 109
Figure 69 - Contact model Flight Case for Main body assembly - Equivalent Stress with predicted
locations .............................................................................................................................................. 110
Figure 70 - Resolving Component to Determine Vertical Load .......................................................... 111
Figure 71 - Undercarriage Leg Under Pure Bending ........................................................................... 111
Figure 72 - Undercarriage Leg Under Pure Torsion ............................................................................ 112
Figure 73 - Stress Element A (Warren C. Young) ................................................................................ 113
Figure 74 - Plan View of Stress Element A .......................................................................................... 113
Figure 75 - Stress Element A with Principle Stresses .......................................................................... 114
Figure 76 - Undercarriage Mesh for Contact Model with values as per H.4 ...................................... 115
Figure 77 – Lateral Landing on Single Undercarriage Leg with 53.6mm Deflection ........................... 115
Figure 78 - Lateral Landing on Single Undercarriage Leg with 60MPa Bending Stress ...................... 116
Figure 79 - Lateral Landing on Single Undercarriage Leg with 60MPa Bending Stress (Close-up) ..... 116
Figure 80 - Tip Landing on Single Undercarriage Leg with 60MPa Bending Stress ............................. 117
Figure 81 - Tip Landing on Single Undercarriage Leg with 66mm Combined bending and torsion
deflection ............................................................................................................................................ 117
Figure 82 - Tip Landing on Single Undercarriage Leg with 71MPa Combined bending and torsion
stress ................................................................................................................................................... 118
Figure 83 – Entire Quad-Rotor Flight Deflection of 7.9mm at Motor Arm Tips ................................. 119
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Figure 84 - Entire Quad-Rotor Flight Deflection of 7.9mm at Motor Arm Tips (Close-up) ................. 119
Figure 85 - Entire Quad-Rotor Flight Stress of 28.8 MPa at Motor mount plates .............................. 119
Figure 86 - Entire Quad-Rotor Flight Stress with Plate Stress peak at 14.42Mpa .............................. 120
Figure 87 – Downward Load - 1kg Payload and 10N Additional Load onto PB-005 Plate .................. 121
Figure 88 - Side Load - 1kg Payload and 10N Additional Load onto Hinge Plate at 45deg to horizontal
............................................................................................................................................................ 121
Figure 89 - Side Load - 1kg Payload and 10N Additional Load onto short edge 45deg to horizontal 122
Figure 90 - Side Load as per Figure 89 - Showing Pre-mature Release due to global deflection ....... 122
Figure 91 – Downward Load as per Figure 87 with new design showing 0.73mm Deflection ........... 123
Figure 92 - Side Load as per Figure 88 – with new rigid design and Deflection of 1.56mm ............... 123
Figure 93 – Side Load as per Figure 89 and Figure 90 – with new design and deflection of 0.41mm*
............................................................................................................................................................ 123
Figure 94 - Arm and Mass for Rayleigh Method ................................................................................. 124
Figure 95 – Mass Representation of Motors, Blocks, Plates, Fasteners and ESC ............................... 125
Figure 96 – Simplified FE analysis with 1st Nat freq as 19.64Hz – 69.3mm Deflection (Left) and
164MPa Stress (Right) ......................................................................................................................... 125
Figure 97 – Simplified FE with 2nd Nat freq as 20.06 Hz (Left) and 3rd Nat freq as 134.6 Hz (Right) ... 125
Figure 98 – Simplified FE with 4th Nat freq as 224.1 Hz (Left) and 5th Nat freq as 411.9 Hz (Right) ... 125
Figure 99 – As Built FE Analysis - Mass Representation of Motors, Fasteners, Cables and ESC ......... 126
Figure 100 – As Built FE analysis with 1st Nat freq as 451 Hz – 69.0mm Deflection (Left) and Stress
(Right) .................................................................................................................................................. 126
Figure 101 - As Built FE analysis with 2nd Nat freq as 736 Hz (Left) and 3rd Nat freq as 1707 Hz (Right)
............................................................................................................................................................ 126
Figure 102 - As Built FE analysis with 4th Nat freq as 2 KHz (Left) and 5th Nat freq as 4.1 KHz (Right)126
Figure 103 - Overall System Hardware Block Diagram Video graphics processing unit (VGPU) ........ 136
Figure 104 - Overall Software Block Diagram ..................................................................................... 137
Figure 105- Pixhawk hardware connections ....................................................................................... 138
Figure 106- Quadcopter Propulsion setup .......................................................................................... 138
Figure 107 - Transmitter and Receiver with Video Graphics Processing Unit (VGPU) the MinimOSD
............................................................................................................................................................ 139
Figure 108 - Servo and motor control schematics .............................................................................. 140
Figure 109: Nylon 6.6 Rod Figure 110: Nylon 6.6 Sheet ............................................................ 147
Figure 111: Nylon 6 sheet ................................................................................................................... 147
Figure 112: PVC rigid angle section……………………………………………………………………………………………… 147
Figure 113: PVC Hard Plastic Tube…………………………………………………………………………………………………147
Figure 114: AL-2024-T6 Sheet ............................................................................................................. 147
Figure 115: Orientation of brackets in Quad copter ........................................................................... 148
Figure 116: Machined fixed brackets by ............................................................................................. 148
Figure 118: Machined end bracket ..................................................................................................... 148
Figure 119: Machined movable arm support bracket………………........................................................148
Figure 120: Machined movable arm vertical fixed bracket ……………………………………………………………148
Figure 121: Machined landing gear top support bracket ……………………………………………………………….149
Figure 122: Machined arm pivot for movable arm…………………………………………………………………………149
Figure 123: Main body plate after laser cutting …………………………………………………………………………..149
Figure 124: Laser cutting of Nylon 6 extrude……………………………………………………………………………….149
Figure 125 - Melted edges on main body plate after laser cutting .................................................... 149
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Figure 126: Milling arm pivot ……………………………………………………………………………………………...150
Figure 127: Drilling centre hole in fixed bracket ................................................................................. 150
Figure 128: Chamfering of movable arm support bracket……………………………………………………………..150
Figure 129: Smoothing surface by fly cutter………………………………………………………………………………….150
Figure 130.1-2: Drilling using slot drills ............................................................................................... 151
Figure 131: High speed steel tool ....................................................................................................... 151
Figure 132.1-2: Machining arm pivot on lathe ................................................................................... 151
Figure 133.1-2 Laser Cutting of Nylon 6 sheet for main body plate ................................................... 152
Figure 134:Cutting Nylon 6.6 cast block in vertical band saw machine ............................................. 152
Figure 135.1-3: Practising samples on CNC machine .......................................................................... 152
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Table of Equations Equation 1- Take off velocity ................................................................................................................ 14
Equation 2- Vertical distance travelled ................................................................................................. 15
Equation 3- Pitch angle ......................................................................................................................... 15
Equation 4 - Max velocity at straight level flight .................................................................................. 16
Equation 5 - Max pitch velocity ............................................................................................................ 16
Equation 6 - Projected Area .................................................................................................................. 88
Equation 7 - Thrust at 54 Degrees ........................................................................................................ 88
Equation 8 - Drag Equation (R. H. Barnard, 2010) ................................................................................ 88
Equation 9 – Working out Moment ...................................................................................................... 96
Equation 210 - Moment for Fixed-arm ................................................................................................. 96
Equation 11 - Stress in a Cylindrical Pipe (Warren C. Young) ............................................................... 96
Equation 12 - Moment for Movable-arm ............................................................................................. 99
Equation 13 - Area A1 on Lug (Niu, 1988) ........................................................................................... 102
Equation 14 – Area A2 on Lug (Niu, 1988) .......................................................................................... 102
Equation 15 - Area A3 on Lug (Niu, 1988) ........................................................................................... 102
Equation 16 - Area A4 on Lug (Niu, 1988) ........................................................................................... 102
Equation 17 - Average Area of Lug (Niu, 1988) ................................................................................... 102
Equation 18 - Bearing Area on Lug (Niu, 1988) ................................................................................... 103
Equation 19 - Flexural Rigidity of the Plate (Ventsel and Krauthammer, 2001) ................................. 106
Equation 20 – Navier solution (Ventsel and Krauthammer, 2001) ..................................................... 106
Equation 21 - Navier stokes coefficient 1 (Ventsel and Krauthammer, 2001) .................................. 106
Equation 22 - Navier Stokes coefficient 2(Ventsel and Krauthammer, 2001) .................................... 106
Equation 23 – Slenderness Ratio (Warren C. Young) .......................................................................... 110
Equation 24 - Radius of Gyration (Warren C. Young) ......................................................................... 110
Equation 25 - Critical Load to Cause Buckling (Warren C. Young) ...................................................... 110
Equation 26 - Critical Stress to Cause Buckling (Warren C. Young) .................................................... 110
Equation 27 - Angle of Twist (Warren C. Young) ................................................................................ 112
Equation 28 - Polar Moment (Warren C. Young) ................................................................................ 112
Equation 29 - Shear Stress (Warren C. Young) .................................................................................... 113
Equation 30 - Compression Stress on Pipe (Warren C. Young) ........................................................... 114
Equation 31 - Principle Stress 1 and 2 (Warren C. Young) .................................................................. 114
Equation 32 - Principle Stress Angles (Warren C. Young) ................................................................... 114
Equation 33 - Shear Due to Combined Loadings ................................................................................ 115
Equation 34 -Static Deflection Curve (MEGSON, 1999) ...................................................................... 124
Equation 35 - Rayleigh's Natural Frequency Equation (MEGSON, 1999) ........................................... 124
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Table of Tables Table 1- Work Breakdown Outline.......................................................................................................... 2
Table 2 - Forms of communication used in project ................................................................................ 4
Table 3 - UAS Challenge 2015 Budget ..................................................................................................... 4
Table 4 - Summarised Margin of Safety Table ...................................................................................... 14
Table 5-shows selected propellers, brushless motor, esc’s and power supply .................................... 14
Table 6- Input Parameters for Quad-rotor Velocity Calculations ......................................................... 15
Table 7- Data for vertical distance travelled ......................................................................................... 15
Table 8 – Weight Variables ................................................................................................................... 15
Table 9 – Propeller data ........................................................................................................................ 16
Table 10- Speed at straight level flight ................................................................................................. 16
Table 11- Effects on the close loop response from PID (University of Michigan, 1996) ...................... 22
Table 12- List of Machines .................................................................................................................... 28
Table 13 - List of tools and their functions ........................................................................................... 28
Table 14 - Bill of Material for manufacturing........................................................................................ 29
Table 15 Qualification Test Plan ........................................................................................................... 34
Table 16 – Itemised Mass Breakdown of all Structural UAV Components ........................................... 82
Table 17 - Electronics and Misc Component Masses ........................................................................... 83
Table 18 – UAV Itemised Cost Breakdown ........................................................................................... 85
Table 19 – Mesh Attributes for Components ....................................................................................... 95
Table 20 – Comparison of Simplified Plate Deflection for Model Substantiation .............................. 108
Table 21 – Summary of Modal Frequencies for Fixed Motor Arm ..................................................... 127
Table 22 - Overall system hardware definitions ................................................................................. 136
Table 23 - Overall software definition ................................................................................................ 137
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1 Project management To achieve the project objectives, effective organisation, planning budgeting and management styles
were adopted. This chapter describes the organisational structure of the project. It describes the
project management, organisation structure, project planning, budgeting and risk management.
1.1 The Project Organization Structure Project organisation needs to be a structure 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. Figure 1 shows
an organisational structure to highlight each person’s responsibility.
Figure 1 - 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. 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.
Alfred Dzadey
Project Manager
Zuber Khan
Structural Team Leader
Structural / Cost / Weights / Assembly Engineer
Osman Sibanda
Bussiness Case
Mozammel
Manufacturing Engineer
Amit Ramji
Structural / Stress / Design & Assembly Engineer
Mohammed Mohinuddin
Test Rig 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
Propulsion / Testing / Assembly Engineer
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1.2 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 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 Figure 2. This is a simplified
overview of the progress made so far which is detailed in the project plan shown in Appendix A. The
progress made so far and completion of tasks can be seen in more detail in the project plan.
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Figure 2 - Show progress to date of the project
1.2.1 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
2 July – Completion Day
2 July - End of UAS Challenge
1.3 Team Communication Throughout the project, weekly meetings with supervisors were undertaken to discuss any updates,
complications and actions required. Also Throughout the project, 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 at Appendix B. 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.
0% 10% 20% 30% 40% 50% 60% 70% 80% 90% 100%
Scope
Design Specification/System Requirements
Preliminary Design Review
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
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Communication Aids Type/technique 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.
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
1.4 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 E.
Budget Estimation as of
01/11/2014 Actual as of 1/04/2015
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
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1.4.1 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. Thus allowing 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 C
2 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 D and Appendix E 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 D, 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 3 below. The use of extruded Nylon 6 main body plates (Appendix C.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 C.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 3 through Figure 5 and
Appendix C.7.
Figure 3 - Quad-rotor design
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2.1 Design Rationale - Quad-Rotor
Figure 4 - Stowage Instructions
Figure 5 - Quad-rotor in Stowed Configuration
3 UAV Mass Breakdown Detailed mass (Appendix D) analysis has been carried out to ensure the UAV is within CAA certifiable
weights limits enable flight and to ensure the requirements are met (IMechE, Jan 2015). The total
mass of the Quad-rotor is 6511.8g with an itemised breakdown shown in Appendix D.
UAV Structural Mass The total mass of the structure is calculated to be 1012.5g including all the materials and fixings
depicted in Appendix C.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 D.
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.
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UAV Electrical / Miscellaneous Components Mass The total mass of the Electrical / Misc. components is calculated to be 5499.3g including all the motors, batteries and additional wiring. The itemised breakdown can once again be observed in Appendix D.
4 UAV Cost Breakdown Detailed cost (Appendix E) analysis 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 and invoices provided in
Appendix E. The above cost summary is inclusive of VAT, less delivery and is accurate to retail prices
at the time of purchase.
5 Structural Analysis
5.1 Load Case Definition and Free Body Diagrams
Figure 6 – Free Body Diagram - Flight and Landing Cases
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Figure 7 - Free Body Diagram - Landing Cases
Figure 8 - Free Body Diagram - Flight and Gust Load Cases
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6 UAV Stress Analysis
6.1 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 G.2 comes 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.
6.2 Load Transfer Loads are transferred from the arms to the
Nylon clamps using a moment balance shown
in Figure 41. 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 9 – Fixed Arm Cross Section – See also Appendix H.5
Fixed and Movable Arm Stress Maximum
The maximum bending stress experienced on the Fixed-arm is 14.42MPa as shown in Appendix H.5 and the maximum bending stress experienced by the Movable-arm is 15.26MPa as shown in Appendix H.6. Refer to Appendix C.7 for parts list, Appendix F for material properties, H.2 for boundary conditions, H.3 for Finite Element solver method, H.4 for mesh types and properties and H.5 - H.6 for results of the contact model for bending case of the UAV Arms. A Sample calculation for the Fixed-arm is shown below:
( )
6.3 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 H.9). These results are in the same order of
magnitude and are marginally different; therefore the modelling technique is deemed correct and
usable throughout.
D1 FF
D2
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6.4 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 C.7 for parts list, Appendix F for material properties, H.2 for boundary conditions, H.3 for Finite Element solver method, H.4 for mesh types and properties and H.10 for results of the contact model for in-flight case of the Quad-rotor.
Figure 10 - Flight and Gust condition of Main Body with 0.13mm Deflection
6.5 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 H.11 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 H.11 is shown below:
( )
(
)
6.6 Undercarriage Bending Analysis on pure bending has also been carried out in Appendix H.11, 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 F). However this analysis has assumed 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
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6.7 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 6.6 with a stress of 62.2 MPa is very close to that shown in
Figure 11 (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 C.7 for parts
list, Appendix F for
material properties,
H.2 for boundary
conditions, H.3 for
Finite Element solver
method, H.4 for mesh
types and properties
and H.11 for results of
the contact model for
bending case of the
undercarriage.
Figure 11 - Lateral Impact Case on Single Leg - 60.6MPa Stress
6.8 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 C.7). Appendix
H.11 calculates a pure torsion case to be used for a combined loading effect in section 6.9 and 6.10.
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 F).
6.9 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 H.11.
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 F). 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:
√( )
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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 12 – Stress Element A with Principal Stress for - Analytical – Undercarriage Combined Loading – Bending, Buckling and Torsion (H.11)
6.10 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 H.11 titled “FEA Results – Combined Torsion and Bending – Tip Contact”. Refer to Appendix C.7 for parts list, Appendix F for material properties, H.2 for boundary conditions, H.3 for Finite Element solver method, H.4 for mesh types and properties and H.11 for results of the contact model for combined tip loading of a single undercarriage.
6.11 Simplified Analytical Modal Analysis – Fixed Arm A simple modal analysis was carried out on the UAV arm to ensure the frequencies of the motors
stay away from resonance. A simplified model with all the parts condensed on a point was used
which resulted in a frequency of 34.19Hz (Appendix 124). The justification for carrying out this
calculation is to substantiate following accurate models of modal analysis with “As-built” parts. It is
worth noting the 1st Natural frequency of the simplified cases (34.2Hz Vs 19.6Hz) are of the same
magnitude and very close. The limitations between these models are that the boundary conditions
being slightly different at the clamped ends (See Appendix H.15).
6.12 Simplified FEA Modal Analysis – Fixed Arm The same case as 6.11 has been modelled in Ansys to substantiate the modelling techniques and justify the results for the first natural frequency of a single motor arm. Refer to Appendix C.7 for parts list, Appendix F for material properties, H.2 for boundary conditions, H.3 for Finite Element solver method, H.4 for mesh types and properties and H.15 for results of the contact model for in flight loading of a single simplified arm. Masses for the motor, fasteners, ESC’s, cables and end brackets has been input into the model as a point-mass as carried out in the analytical solution shown in 6.11 (supplemented by Appendix H.15).
Figure 13 – Simplified Fixed-arm modal analysis with 1st Natural Frequency at 19.6 Hz
𝜏𝑥𝑦
𝜎𝑥
𝜎𝑦
𝜏𝑥𝑦
𝜏𝑦𝑥 𝜎𝑦
𝜎𝑥
𝜏𝑦𝑥
A 𝜃
𝜃
27.09
-34.495
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6.13 As Built FEA Modal Analysis – Fixed Arm As mentioned in section 6.12 the stiffness limitations of point masses are not considered in the simplified cases, hence an accurate model of a single arm assembly is generated. Refer to Appendix C.7 for parts list, Appendix F for material properties, H.2 for boundary conditions, H.3 for Finite Element solver method, H.4 for mesh types and properties and H.16 for results of the contact model for in flight loading of a single arm assembly. This method produces a much more accurate method of analysing the actual structure as the stiffness contribution of fasteners, inertia of offset motors and fasteners are considered. The results show a higher natural frequency due to the tip structure being a much higher stiffness. The modelling techniques have been demonstrated in 6.11 and 6.12,
where the order of frequency magnitude is the same and difference in frequencies is minimal. As the same modelling techniques have been performed in Appendix H.16 as carried out in H.15, the analysis can be deemed as correct with the only difference being the inclusion of actual parts as built (Appendix C.7).
Figure 14 – As-Built Fixed-arm modal analysis with 1st Natural Frequency at 451 Hz
6.14 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 C.7)
Case / Calculation /
Section
Loading Description
Maximum Applied
Load/Stress
Maximum Allowable
Load/Stress Appendix F
Safety Factor, SF= Allowable
/Applied
FA-001 Case 1 (H.5)
Maximum Thrust from
Motors 14.42MPa 55MPa 3.81
MA-001 Case 1 (H.6)
Maximum Thrust from
Motors 15.26MPa 55MPa 3.60
UV-001 Case 2 (H.11)
Undercarriage Pipe Under
Buckling 10.5Kg 56.76Kg 5.41
Case 4 (H.11)
Undercarriage Pipe Under
Torsion 30.57MPa 1099.3MPa 35.96
LB-003 (H.7) Undercarriage Lug Under Maximum
Loading
72.84N 1765.15N 24.23
UV-001 Analytical – Undercarriage
Combined Loading – Bending,
Combined Loading on
Undercarriage Vertical Leg
27.09MPa 55MPa 2.03
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Buckling and Torsion (H.11)
Combined Loading on
Undercarriage Vertical Leg
34.5MPa 55MPa 1.59
BP-001 & BP-002
Assembly.
Appendix H.10
Main Body Deflection
due to Maximum Thrust and
Gusts
5.83MPa 55MPa 9.43
Table 4 - Summarised Margin of Safety Table
7 Performance & Propulsion
7.1 Introduction It is the performance and propulsion engineer’s role to investigate the possible in-runner and out-runner electric motors, propellers, and power sources that are capable of producing the thrust required. This thrust firstly includes lift of the Quad-rotor and secondly to attain the velocity required to complete the mission on time prior to excessively draining of batteries.
1. To calculate Quad-rotor performance, the MTOW was one of most vital piece of information that was required, where 7kg has been used.
2. Identified Hover thrust – Using MTOW of 7kg it was identified that for the Quad-rotor to hover it would require each of the four motors to produced 1.75kg of lift to hover
3. Identified thrust for manoeuvrability – Using an equation provided by leading multicopter developers such as DJI, thrust required for improved manoeuvrability was calculated
(
)
4. From the thrust value above, propellers of dimensions 11” x 8” would be adequate to produce the thrust required.
5. From the 11” x 8” propeller, a specific brushless motor can be identified due to very few motors being able to perform with efficiency.
6. The electronic speed controllers were selected based on the maximum current draw that can be obtained from the brushless motor. In this case 47A, therefore an ESC of 60A was appropriate.
Propeller Brushless motor
ESC Power supply
11” x 8” EMax 2826-06 Robotbirds Pro-60A
5s Turnigy nano-tech with 8000mah capacity (x2)
Table 5-shows selected propellers, brushless motor, esc’s and power supply
7.2 Take-Off Velocity Take off 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.
√
Where: T = Thrust N Density kg/ =
Equation 1- Take off velocity
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Density at 30.48m kg/
Density at 121.92m
kg/
Mass with Payload
(kg)
Thrust required
with Payload
(N)
Mass without Payload
(kg)
Thrust required without payload
(N)
Propeller Area ( )
1.192 1.179 7 68.67 6 58.86 0.2452 Table 6- Input Parameters for Quad-rotor Velocity Calculations
Using Table 6: Take-Off Velocity with payload to 30.48m = 10.8m/s
Take-Off Velocity without payload to 30.48m = 10.0m/s
Take-Off Velocity with payload to 121.92m = 10.9m/s
Take-Off Velocity without payload to 121.92 = 10.1m/s
7.3 Time To Reach Cruise Altitude Time to cruise altitude of between 100ft and 400ft can now be calculated using the equation below.
( )
(
)
Equation 2- Vertical distance travelled
Where: d = Distance m Initial velocity m/s = Initial time s
( )
( )= Acceleration m/ t = Time taken s
Using Table 7 and Equation 2 can be simplified to:
( )
(
) or to calculate time to height t = √
( )
(
)
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
7.4 Max Velocity To calculate the maximum velocity attainable by the Quad-rotor requires propeller diameter and pitch angle, maximum motor RPM and also maximum pitch angle that can be achieved by the Quad-rotor without instability.
As the propeller data is known the maximum tilt angle can be calculated F*cos( )=Mass
Equation 3- Pitch angle
Where F = Force N M = Quad-rotor mass kg = Maximum tilt angle
Table 7- Data for vertical distance travelled
Altitude (m)
Altitude (m)
Initial velocity
(m/s)
Initial time (s)
Force (N)
Mass with
payload (kg)
Mass without payload
(kg)
Acceleration with payload
(m/ )
Acceleration without payload (m/ )
Time taken
(s)
30.48 121.92 0 0 86.33 7 6 12.33 14.38 6
Force (N) Weight with payload (N) Weight without payload (N)
86.33 68.67 58.86
Table 8 – Weight Variables
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Using Equation 3 and input variables in Table 8, the maximum tilt angle can be calculated
(
( )
( ))
Therefore: Maximum tilt angle with payload box = 37.30 Maximum tilt angle without payload boa = 470
Quad-Rotor maximum speed in straight and level flight can be calculated using equation 4 *0.44704
Equation 4 - Max velocity at straight level flight
Propeller Pitch
11164.75 8 Table 9 – Propeller data
Using Equation 4 and Table 9
= 9851.25*8*0.000954*0.44704 = 33.61m/s
For the case of a Quad-Rotor straight and level flight velocity cannot be used as the equation
assumes that the flight path perpendicular to the x-axis, therefore the has to be modified
to take into account Quad-rotor maximum tilt angle.
= * Sin( ) Equation 5 - Max pitch velocity
(m/s)
with payload
(degrees)
without payload
(degrees)
33.61 37.3 47 Table 10- Speed at straight level flight
At maximum tilt angle of 37.30 = 20.37 m/s IAS
At maximum tilt angle of 470 = 24.58 m/s IAS
7.5 Stall Stall for a Quad-rotor with a mass of 7kg will stall if the maximum tilt angle of 37.30 is exceeded Stall for a Quad-rotor with a mass of 6kg will stall if the maximum tilt angle of 470 is exceeded
7.6 Range, Power Consumption, Battery Life Quad-rotor range, power consumption and battery life is analysed in the Flight performance section
in detail as per Appendix I
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8 Rationale for Design Specification/Selection
8.1 Autopilot The need of an autopilot system was required to control the Quad-rotor as well as its systems. The criteria for choosing the autopilot system are:
On board processing power Interference from external devices Capability of use with different types of aerial view Programmable Firmware and software Cost Fail Safe Systems Support for autonomous flight Product support for troubleshooting
The final two systems that meet the above requirements were APM and Pixhawk. While APM was inexpensive and extensively user tested, it was decided that Pixhawk (Figure 15) shall be used due to its on-board failsafe processor and increased processing power able to handle additional devices without lag.
Figure 15 – Pixhawk (3DR, 2014)
8.1.1 Specification (3dr, 2014): 168MHz/252MIPS Cortex-M4F processor 14 PWM (Pulse-width modulator)/servo outputs (8 with failsafe and manual override, 6
auxiliary, high-power compatible) Abundant connectivity options for additional peripherals (UART, I2C, CAN) Integrated backup system for in-flight recovery and manual override with dedicated
processor and stand-alone power supply Backup system integrates mixing, providing consistent autopilot and manual override mixing
modes). Redundant power supply inputs and automatic failover External safety switch Multicolor LED main visual indicator High-power, multi-tone Piezo-audio indicator microSD card for high-rate logging over extended periods of time
8.2 On screen display board (OSD) For such a complex UAS, it is required for a method of viewing the telemetry data interlaced with a live-feed; an on screen display board was hence selected. Overall, it was decided that the MinimOSD (Figure 16) would be best suited for the project. The criteria used for the selection is as follows:
Figure 16 - MinimOSD v2 (APM, 2014)
Compatibility for PIXHAWK control board Number of telemetry data outputs Configuration ease Cost Power consumption Size Error indication and warning system (lost GPF fix, stall, over speed, battery voltage and
percentage and received signal strength indication)
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8.2.1 Specifications:
ATmega328P
MAX7456 monochrome on-screen display
FTDI cable compatible pin-out
Standard 6-pin ISP header
Two independent power sections with an LED indicator on each
Solder jumpers for combining the power supply 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
8.3 GPS system As the Quad-rotor is designed with autonomy in mind, Pixhawk would require a method of
knowing its position and therefore a GPS is required.
The criteria for selection were as follows:
Compatibility with Pixhawk control board
GPS accuracy
Configuration ease
Cost
Power consumption
Battery life
Battery recharge ability
Protectiveness Figure 17 - 3DR uBlox GPS (3drobotics, 2014)
After careful consideration, the 3DR uBlox GPS was selected as the suitable GPS + Compass
module for the project. The uBlox GPS outperforms most other GPS modules due to the larger
antenna and next-gen chipset. It has an expected usable time of 180-200 hours on a full charge
with a rechargeable backup battery for warm starts. It allows for up to 1-2o degree compass
accuracy, protective casing and its usability within a strong magnetic field environment.
8.3.1 Specification:
uBlox LEA-6H module
3-axis digital compass IC HMC5883L
5Hz update rate
25 x 25 x 4mm ceramic patch antenna
LNA 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 pan
38 x 38 x 5.8mm total size, 16.8grams
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8.4 Telemetry kit A telemetry kit combined with a radio controller, allows for two-way communication with the
ground station. There are mainly two types; radio set kit and Bluetooth data link set. While
Bluetooth is inexpensive, it is only intended for pre-flight ground-use and not a replacement for
an RC transmitter and receiver. The main disadvantage of Bluetooth is that it has a very limited
range making it an unrealistic option for this purpose.
It was decided that the 3DR telemetry kit V.2 is suitable, allowing for greater range and
maximum compatibility with Pixhawk.
8.4.1 Specification:
100mW maximum output power
-117dBm receive sensitivity
RP-SMA connector
2-way full-duplex communication
UART interface
Transparent serial link
MAVlink protocol framing
Error correction up to 25% Figure 18 – 3DR telemetry kit V.2
8.5 Camera As GPS cannot be relied upon for accurate reading of position for payload deployment, it was
decided to make use of a camera which would allow for a live feed of the target area as a first
person view (FPV). The Mobius Action Camera as shown in Figure 19 for its extensive features and
lightweight.
8.5.1 Specification
1080p-30fps, 720p-60fps and 720p-30fps functionality
Adjustable movie quality setting
Recording cycle time setting of 3, 5, 10, 15 minutes and
max
Loop recording
Movie flip (180o and rotation)
Variable resolution photo mode Figure 19 - Mobius Action Cam (UNMANNEDtech, 2014)
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8.6 Servo Deployment of the payload as per the mission requirements requires the use of a servo to release
the payload door as depicted within Appendix C.7. It was decided to choose the MG90S servo or
equivalent. The servo can rotate approximately 180° (90° in each direction), is controllable with any
code, hardware or library to control.
8.6.1 Specifications /Rational
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
Dead band width: 5 µs
Figure 20 – MG90S Servo
The MG90S servo uses a switched mode power system that is considerably more efficient than an
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. The MG90S servo
has a small dead band, faster response, smoother acceleration, and improved positional holding.
8.7 BEC
The BEC requires +5V to power the opto-isolator and while the Pixhawk can be powered from the
servo rail, it does not support +5V output to the servo rail hence the reasoning for a BEC. The
Turnigy SBEC26 is an advanced switching DC-DC regulator which will supply a constant 5A. It works
with 2s – 7s Li-po and supplies a constant 5 - 6v to the receiver and is interference-free and perfect
for confined spaces
8.7.1 Specification / Rational
Type: Switching
Input protection: Reverse polarity protection
Output (Constant): 5v/5A or 6v/5A
Input: 8v-26v (2-7cell lipo)
Weight: 18g
Figure 21 – SBEC26 - Turnigy
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9 UAS Sub-systems Efficient systems architecture involves the discretisation of a larger system into sub-systems. This
section describes the systems on board the UAS such as navigation, communications as well as
schematics showing detailed information on how the system components are integrated. A detailed
specification is included herein along with detailed system schematics as depicted in Appendix J.
9.1 Navigation System 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 locked on the waypoint and target. The on board compass, gyroscope and GPS coordinates work together via Pixhawk to determine the motion on the relevant UAV axis and control the motors through the ESCs. The GPS coordinates are programmed into the navigation system with the use of waypoint files as shown in Figure 22.
Figure 22- Waypoint Command File
The ground control station is used to input the GPS coordinates and payload release commands in
the form of a mission plan. The ground station is also used to monitor the data generated by the
sensors on-board the UAS and is transmitted back via the telemetry kit. The ground control station
consists of a laptop, telemetry antenna and mission planner software.
9.2 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, altitude and position. 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, a command is sent from the ground
control station to the autopilot or by on-board pre-programming. The autopilot 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.
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The flight control system also carries out the stability and control function for the UAV. The
autopilot system has a in-built controller, reprogrammed to correct errors and make
adjustments in flight control. The controller is a PID (Proportional Integral and Derivative) variant
and can auto-tuned in-flight with the use of the radio controller. The PID numbers are to be
calculated before being input into the UAV prior to initial flight where the methods used to
obtain the PID’s are:
Creating a MATLAB model for theoretical values of PID a. Understanding and selecting the right PID controller within the Simulink
environment for a Quad-rotor
b. Configure the Simulink model to fit the UAV, such as ideal conditions as well as
strong wind and light rain
c. Tune the gain values until satisfied with the results from the graphs, for still
conditions, Table 11 describes the effect of PID gains on a closed loop response
d. Test PID values under disturbances to check if the Quad-rotor can stabilise
e. Further tune the controller gain values to meet all conditions
Testing the UAV system using a test rig a. Set the Quad-rotor inside the test rig and ensure rigid fixing onto the test frame.
b. Input the PID values from MATLAB
c. Test the Quad-rotor under multiple conditions
d. Use a high airflow fan to replicate strong gusts to observe how well the Quad-rotor
can stabilise or fly under wind conditions
e. Fine-tune the PID values if results not satisfactory
Test fly the quad a. Take the quad to an open area for test flying while adhering to CAA requirements
b. Start with simple manoeuvres before moving onto more extreme manoeuvres
c. Further fine-tune if refinement is required
CL 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 11- 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 torque,
forces produced by the motors and the Quad-rotor’s inertial frame in relation to non-linear
dynamics. With the above information, equations of motion can be generated using a rotation
matrix to simulate the motion of the Quad-rotor. An appropriate controller can be designed to
reduce any error produced by the Quad-rotor system. The model is not an entire 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. The test rig will be
used to fine-tune our close-to-final PID values before actual test flight of the Quad-rotor.
An integral part of the flight control system is the autopilot system. The autopilot system comprises
of three layers:
Hardware
Firmware
Software
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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 is capable of carrying out functions such as autonomous flight, computer vision operations and robotic functions. The 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 provides information about different systems on board the UAV and data about flight performance, this information 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 can be analysed to make adjustments to any system to raise the performance of the UAS.
Figure 23- Telemetry Information transmitted to ground control station
9.3 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.
9.3.1 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 connection is used to load the firmware and software needed to run the autopilot system and to calibrate on-board sensors. 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.
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Figure 24 - Transmission Link Statistics (Serial Connection)
9.3.2 Telemetry Kit Connection
The telemetry kit is used to connect the Pixhawk 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.
It 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 not advisable. 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; ranging from servo activation to camera functions. The strength in telemetry
connection would decrease as the UAV moves further away from the ground control station.
Figure 25- Transmission Link Statistics (Telemetry Kit)
9.3.3 Radio Connection
The radio controller is used to connect to the Pixhawk autopilot and the UAV through a frequency of
2.4 GHz (Section 14). 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 auto-tune methods.
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 measure if the UAV flies out of
range of telemetry range or outside the airspace of the ground control station.
9.4 System Schematics Details of all system schematics and functionality can be found in Appendix J.
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10 Payload box mechanism As part of the UAS design challenge, there is a task to design a payload delivery system and its
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 copter 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 as shown in Appendix
C.7.
10.1 System of the payload box 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. 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.
10.1.1 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 26 - Configuration of the servo on Pixhawk
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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.
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.
10.1.2 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 27- Verification of the performance of the Servo
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11 Manufacturing In order to achieve an efficient structure, manufacturing methods were identified at the primary
stage of the project. The manufacturing plan included materials to be used, joining methods,
machines and the best possible way to carry out the task on time. Considering the weight and cost
restriction of the project the most attractive and ideal materials are polyamide (nylon 6, nylon 6.6)
and carbon composite. Hence a systematic progress is done to design the parts of the UAV to make
machining achievable.
11.1 Design Phases 11.1.1 Phase 1 (Preliminary Design Review)
At the primary stage of the design, due to high strength-to-volume ratio composite materials were
considered for top and bottom plates and the tubes (arms and landing gear strut and stabilizer) in
the UAV; and Nylon was considered for the support clamps and spacer material to minimise the
weight and strengthened the structure for its effective compression ratio (3.0:4.0)
11.1.2 Phase 2 (Critical Design Review)
As the project progressed, after weight breakdown and cost estimation, the materials were
reconsidered. Albeit composite material are the most desirable material for the UAV’s structure to
accomplish efficient strength to withstand the load but after detailed stress and deflection analysis
nylon and PVCs are finalised for the plates and tubes respectively as they are inexpensive and have
low density.
11.2 Materials Selection Decisions concerning materials selection were mainly on mechanical properties as the aim is to
withstand unexpected impact and damp the vibrations from the motors at the arms. Machining of
the materials was also considered while designing the Quad-rotor’s components.
11.2.1 Nylon
Nylon was chosen because of its low density and high strength properties and hence chosen for
manufacturing. Nylon is one of the best materials, where milling and lathe are fairly straight forward
and does not leave any dull cut or scratches on the surface. It also has great impact strength or
shock resistance compare to metallic counterparts. As the UAV is likely to have vibration from the
motors, vibration or shock will be absorbed by the nylon and integrated system components will not
be affected.
11.2 .2 Nylon 6.6 (PA 6.6 Black Cast Sheet) and Nylon 6.6 Rod (PA 6.6 Dia 25mm Rod)
Due its attractive compressive characteristics nylon 6.6 is used to manufacture the supporting
brackets which are mounted in the sandwiched structure in the UAV. It is important to ensure the
arm pivot is efficient to hold the arms and hence Nylon 6.6 rod is the most suitable material for this
particular function.
11.2.3 Nylon 6 (PA 6 Extruded Sheet)
Since Nylon 6 is more compressive and inexpensive than Nylon 6.6 so it is used to manufacture the
main plates and the spacer in the structure. Alongside the compressive characteristic of nylon it is
also desirable when compared to brittle acrylic for this purpose.
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11.2.4 PVC (Polyvinyl chloride), (Black hard plastic tube/Rigid angle section)
PVC is very durable and can typically withstand impacts for reasonable times when compared with
other plastic. It is also a very good electrical insulator and has a low melting point which is highly
desirable for the arms. Hence using PVC, the manufacturing is more affordable and correspondingly
helped to minimise the overall weight remarkably.
11.2.5 Aluminium Alloy (AL-2024-T6)
The motor mount plates are prone to heat rise and vibration from the motor during operation.
Therefore considering the fact aluminium alloy is the best material for this purpose as it has
reasonable density and high melting point; it is also a good environmental resistant material which
minimises any probable risk involved.
11.3 Machining Selection Acknowledging the weight and budget limit for the project, the manufacturing process includes
milling, lathe, laser cut and CNC machining which are available within the lab facility of the university
11.3.1 Machines
The following machines are used or practised to manufacture the parts depending on their operating
functions.
Machine Type Functions
Bridgeport Series 2 Milling machine
Use end mills to obtain precise dimension 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 center 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 12- List of Machines
11.3.2 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
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 13 - List of tools and their functions
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11.4 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 components are
marginally modified. All the sharp edges will be machined (preferably by grinding) before assembling
all the components to ensure that cracks are not anticipated during dynamic motion of the Quad-
rotor.
Material Type Components Quantity
MAIN STRUCTURE
Nylon 6.6(Black) 10 mm thick cast sheet Fixed bracket 16
Motor arm end bracket 4
Movable arm vertical fixed bracket 2
Movable arm support bracket 2
Nylon 6.6(Black) 16mm thick cast sheet Landing gear top support bracket 2
Landing gear bottom support bracket 2
T-joint top half 2
T-joint bottom half 2
Nylon 6.6(Black) 30 mm thick cast sheet Landing gear lug bracket 2
Landing gear pivot 2
Nylon 6.6 (Black) Dia 25 mm solid rod Arm pivot 2
Nylon 6 (Black) 2 mm thick extruded sheet Main body plate (top & bottom) 2
PVCs Movable arm 2
Fixed arm 2
Landing gear strut 2
Landing gear stabilizer 2
Aluminium Alloy 2024-T6 Motor mount plate 8
PAYLOAD
Nylon 6.6(Black) 10 mm thick cast sheet Slot bracket 1
Turn button 1
Nylon 6 (Black) 2 mm thick extruded sheet Triangle payload support 1
Nylon 6 (TECAMID 60 MO FILLED) 1 mm cast sheet
Thick bonded corner supports 8
1PVC HARD PLASTIC RIGID ANGLE SECTION H707 (BSA10) 16 x 16 x 1.5
Angle section for edges (long, short and vertical)
11
Angle section for solenoid support 1 Table 14 - Bill of Material for manufacturing
Machined Components
To be machined (purchased )
To be machined (awaiting delivery)
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11.4.1 Fixed Bracket
The fixed brackets are made of nylon 6.6, which has the favourable characteristics to hold the fixed
arms in place. To achieve good finishing qualities a milling machine will be used to manufacture the
fixed brackets. Due to some limitation in CNC machining, smooth edges were not obtained as shown
in Figure 116 in Appendix L
11.4.2 Motor arm end bracket
Due to compressive characteristics of nylon 6.6, the brackets at the end of the each arm will securely
hold the motors mounted onto the motor mount plates. Refer Appendix L for machined bracket. The
brackets are drilled by 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.
11.4.3 Movable arm vertical fixed bracket /support bracket
From nylon 6.6, the brackets are machined in milling machine as per the technical drawings. Refer
Appendix L for machined components.
Figures in Appendix L show the machined brackets that are obtained by milling machine. All
dimensions are carefully machined according to the technical drawings but the edge fillet of radius
of 12.5 mm was comprised due to complexity and availability of appropriate tools.
11.4.4 Landing gear top 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. Refer Appendix L for machined
component. The versatile properties of nylon 6.6 make the arm pivot very efficient to withstand
bending moment and compressive force and since it has a high compressive ratio which ensures it
will not crack unexpectedly like materials ABS plastic (Acrylonitrile Butadiene Styrene) or PVCs
(Polyvinyl Chloride) Refer for machined part.
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 desired length and diameter was cut by high speed steel
tooling; then the rod was bored to 22mm inside and followed by milling the 5.5 mm from the both
sides on the other end. The remaining flat part was then drilled to make it suitable for pivotal
function for the movable arm.
11.4.5 Main Body Plate
The top and bottom plates are manufactured from nylon 6 extruded sheets which ensure consistent
thickness throughout the material. Therefore load of the components are evenly spread on the
plates. The supporting brackets also strengthen the structure and make it efficient as desired. Refer
Appendix L for figures
Due to simplicity in Trotec Laser machine setup, laser cutting was chosen and figures in Appendix L
shows the plate after machining. Even though accurate dimensions were obtained but during
machining the heat of the laser melted the edges clear smoke was observed as shown in Figure 122
in Appendix L. Figures in Appendix L shows the laser burned the surface of the plate and which
anticipates the risk of cracks at the edges for further use. To overcome the challenge, milling would
be done using Denford Router 2600 Pro. The advantage of such machining would give accurate
dimensions and the material properties would not be affected by any thermal energy.
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11.5 Overview of Machining
11.5.1 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 L for figures of some machined components by milling.
11.5.2 XYZ 1330 Lathe
The arm pivot was machined in the lathe; 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 L for figures regarding lathe
11.5.3 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 L for figures on laser machining set up
11.5.4 Vertical Bandsaws Machine
Vertical bandsaws machine is used to cut the purchased block or sheet (aluminium alloy) into
required dimensions for the components. Refer Appendix L for figure
11.5.5 CNC Machines (Router 2600 Pro and VMC 1300)
Simplest form of machining has been chosen all along the manufacturing process considering the
budget limit in the competition. Utilising the machining facility of university, all the components
would be manufactured. Unfortunately, due to lack in technical facilities (tools constraints) on CNC
machines (Denford Router 2600 Pro and Denford VMC 1300 Milling Machines), machining is mostly
dependent on manual milling and drilling for simplicity. Refer Appendix L for CNC practised samples
12 Testing 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 shown below. 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.
1.1. Octagonal gimbal test rig In order to overcome a few challenges encountered through initial test rig design shown above, a
more compact and robust design has been established as shown below. 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.
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Figure 28 -Updated Octagonal Test Rig Assembly
12.1 Weight estimation for octagonal test rig The spreadsheet in Appendix M 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.
12.1.1 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 a generous sponsorship. The spreadsheet detailing the costing can be found in Appendix M.
12.2.2 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. Although the 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 complex. Therefore a safe testing method would have to be implemented to avoid damage on such a costly design.
12.2 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 (Figure 30) and the manufactured base plate after cut-outs (Figure 29).
Figure 29 Manufactured Base Plate
Figure 30 - Nylon Specimen
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A compression test, using the Hounsfield 1kN
Tensometer, was carried out on the base plate
material as seen in figure 32. 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.
12.3 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. 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.
12.4 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.
Figure 32 - Payload drop test
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.
Figure 31 - Compression Test conducted on Hounsfield Tensometer
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12.5 Qualification test plan Upon completion of all manufacturing activities and processes, all space qualification hardware shall be tested according to and following the order of the qualification test plan found in Table 15.
Test Description
Initial Electrical Performance Test
Storage Temperature Cycling
Electrical Performance Test
Thermal Shock
Electrical Performance Test
Sine Vibration
Electrical Performance Test
Random Vibration
Electrical Performance Test
Operational Temperature Cycling
Final Electrical Performance Test Table 15 Qualification Test Plan
12.5.1 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.
12.5.2 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.
12.5.3 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.
12.5.4 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 quadcopter without any degradation to its coatings,
surfaces, or electrical performance.
12.5.5 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.
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The aforementioned qualification tests must be performed at regular intervals to ensure worthy
performance of the quadcopter. The results obtained shall also be verified and validated to meet
certain conformances. The results from various tests should also be recorded and checked.
12.6 Verification and Validation Test matrix In Appendix K numerous tests that need to be conducted in order to achieve successful
accomplishment of this project.
13 Safety Case
13.1 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. Since the UAS is less than 7Kg MTOM, it 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. (Civil Aviation Authority, 2012; Corbett, 2014)
The main requirements extracted from those articles are as below;
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 take-off and
landing
Apart from this, it is essential that team is referring to the UAS Challenge 2015 competition rule
book while designing, manufacturing and testing and the demonstration of UAS.
13.2 Hazardous Components 1. High speed propellers – detachment of propellers in flight can cause serious injuries to
people and animals. Therefore 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 miss use, which can
cause serious structural damage to the aircraft. Uses of high quality batteries are considered
and monitoring their charging and temperature can avoid such failure
(Rogershobbycenter.com, 2014). Purchased batteries will be made brighter in colour to
identify them in a crash and they will be mounted using Velcro Straps for easy removal.
36 | P a g e
13.3 Flight Controller safety mechanism The flight controller Pixhawk has a number of safety mechanisms. It includes a motor arming safety
feature when manually controlling the Quad rotor. 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.
13.1 Safety measures for flight testing Ensuring no personnel is 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.
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.
13.1 Description of functionality for flight termination cases The core safety functionality of the Pixhawk software “APM:Plane” is for RTL, if the UAV loses
contact with the ground station or during manual control. If more advanced options are required
than Pixhawk has an on-board Advanced Failsafe system (AFS). The autopilot can be set up for
failsafe conditions so that the multi-copter can loiter for a short period of time before RTL
automatically lands or terminates flight (APM Plane, 2014).
13.1 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 is initiated (APM
Plane, 2014). This involves the system looking at the data input into the “AFS_WP_GPS_LOSS”
parameter which instructs the multi-copter on its next action, ranging from loiter for a period of
time or reducing motor RPM to land the craft. If the GPS regains positioning then the multi-copter
will continue its mission from where it left off.
GROUND STATION COMMUNICATION LOSS
The AFS system constantly monitors the strength of the data-link between the multi-copter and the
ground station using “HEARTBEAT MAVlink” (APM Plane, 2014) messages that are transmitted by
the ground station. If for a period of 10 seconds or greater the multi-copter does not receive a
HEARTBEAT (similar to handshake ping in IT systems) message, then it enters AFS state. During AFS
state, the system looks for the “AFS_WP_COMMS” parameter, a waypoint number to which the UAV
should navigate to when communication loss has occurred (APM Plane, 2014). If this is a non-zero
value the UAV will change its target waypoint given by the “AFS_WP_COMMS”, otherwise it shall
RTL.
37 | P a g e
If the craft loses GPS positioning and connection with the ground station then this is considered
‘dual loss’ and the craft will carry out pre-programmed commands. However the user can override
Pixhawk and enable manual mode and take control regardless of loss of GPS or connection between
ground - stations. If all connections are lost, including manual control, then the multi-copter will
terminate flight after a specified time in milliseconds, in this case for 30,000 milliseconds (30
seconds).
GEOFENCE BREACH
Geofence allows the user to set the 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 (APM Plane, 2014).
MAXIMUM PRESSURE ALTITUDE BREACH
When the airspace is being shared by multiple UAVs, the flight altitude will be measured by a
common reference pressure, typically the QNH, a Q-code. The AFS system can be used to set a
pressure altitude limit, a value in millibars into the “AFS_AMSL_PRESSURE” parameter, which the
craft will not exceed. Similarly the pressure altitude limit can be set in the “AFS_AMSL_LIMIT” (APM
Plane, 2014) parameter in meters. If both parameters are set and are exceeded, then the AFS will
initiate a pre-programmed termination process.
The AFS system will also monitor the barometer, and if it shows to be malfunctioning for 5 seconds, then the AFS system will look at the “AFS_AMSL_ERR_GPS” (APM Plane, 2014) parameter. The multi-copter will enter flight termination immediately if it is set at the default value of -1. If not, it will continue flight and use the value as a margin to add to the GPS height. This allows 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 (APM Plane, 2014).
38 | P a g e
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
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.
L
ikel
iho
od
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
The above table is the Cause-Consequence Relationship in Risk Management (Weibel and Hansman,
2005)
39 | P a g e
I.D Risks Likelihood Severity Risk Level
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
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
40 | P a g e
14 UAV Technical Specifications
Structural Parameters Maximum Take-off Weight 7Kg (6.5118Kg - Appendix D) Undercarriage arrangement Collapsible legs Payload weight 1kg (x2 with 2nd trip) Control systems and Misc weight 5.4993 Kg (Appendix D) Maximum structural weight 1.0125Kg (Appendix D) Payload mechanism weight 133.67g Control Systems Autopilot Pixhawk Ground Control System Mission Planner GPS unit Ublox GPS module Telemetry kit 3DR 433 MHz Radio kit Electronic Speed Controllers (ESC) Robotbirds Pro 60A ESC Batteries 2x 5s 8,000mAh Turnigy Li-Po Camera Mobius Action Camera On-screen Display Minum OSD Servo MG90S BEC SBEC26 Turnigy Radio controller Receiver
Turnigy 9X 9Ch Transmitter Turnigy 9X 8Ch Receiver
Performance data Motors 4x DC Brushless motors Power 850W Propeller type 2 blade propellers (11” x 8”) Operating Altitude 100-400ft Design Time / Range 5.3 minutes/ 3.48 kilometres UAV Material Quad-rotor arms PVC Flat plate Nylon 6 Brackets Nylon 6.6
41 | P a g e
Appendix A. Project Plan
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 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 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 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 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 usingdesign 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
Integration testing complete 31/05
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 F S25 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
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: UAS Challenge Project PlaDate: Wed 03/12/14
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 Fri 05/12/1420 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 Critical Design Review (CDR)
and Flight Readiness Review (FRR)
71 days Mon09/03/15
Mon 15/06/15
42 Draft CDR report 0 days Mon 09/03/15 Mon 09/03/1543 Deliver CDR report 0 days Wed 01/04/15 Wed 01/04/1544 Draft FRR report 11 days Mon 18/05/15 Sun 31/05/1545 Deliver FRR report 0 days Mon 15/06/15 Mon 15/06/1546 Pre-Competition 15 days Mon 15/06/15 Fri 03/07/1547 Design Presentation 0 days Wed 01/07/15 Wed 01/07/1548 Flight Readiness Review 0 days Wed 01/07/15 Wed 01/07/1549 Certification Test Flight 11 days Mon 15/06/15 Mon 29/06/1550 Competition day 3 days Tue 30/06/15 Thu 02/07/1551 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%
Preliminary Design complete 100%
Deliver PDR to IMeche 100%
100%
100%
100%
100%
Design purchase readyness 100%
100%
100%100%
0%
75%
25%
25%
0%
50%
50%
0%
0%
Integration testing complete 0%
0%
Draft CDR report 100%
Deliver CDR report 100%
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: Wed 01/04/15
44 | P a g e
Appendix B. Minutes and Agendas
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
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!!!
48 | P a g e
Appendix C. UAV Design
C.1. Weight Reduction - Quad-Rotor In order to eliminate excess mass, design considerations such as those discussed in Appendix H.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 C.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 C.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 Appendix G 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 H.9 as an infinite plate assembly. Compared with the analytical technique, the error between results is minimal as is discussed in section 6.3.
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.
C.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 C.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.
49 | P a g e
Figure 33 - Overall View of Quad-Rotor
Figure 34 - Motor Mount Design (Left) & Undercarriage T-Joint (Right)
Figure 35 - Undercarriage Pivot Design (Left) & Main Body Sandwich Design (Right)
50 | P a g e
Figure 36 - Movable Arm Pivot Design
C.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 C.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 H.13 and H.14 show the changes made to PB-006 and PB-008 (Appendix C.7)
to increase the stiffness of the housing during flight conditions to avoid pre-mature deployment of
payloads.
C.4. Supplier Discount and Advertising Opportunities The value of structural components such as raw Nylon (PA6 & PA66) blocks / sheets have been demonstrated in Appendix E 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.
C.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.
51 | P a g e
C.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
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C.7. Production Support and Drawings
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Appendix D. 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.
Quad-rotor Mass Breakdown
Structural Part Name
Part No. (Appendix
C.7)
Material (Appendix F)
Density (g/cm3)
Area cm2
Length/ Thickness
(cm)
Volume cm3
Mass (g) Qty Total
Mass (g) Picture
(Appendix C.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
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Moveable arm pivot
AP-001 Nylon 1.14 6.25 5 13.11 14.95 2 29.90
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
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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
Undercarriage T Joint
TJ-001, TJ-002
Nylon 1.14 13.12 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
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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
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
M5 Nylon Spacer – 5.3mm
Internal, Outer 10mm, length
10mm
M5 Nylon Spacer
Nylon 1.14 10 0.7 2 1.4
M6 Nylon Spacer – 6.4mm
Internal, Outer 12.5mm, length
10mm
M6 Nylon Spacer
Nylon 1.14 10 1 4 4
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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, 18mm
External, c/s 1mm
O-Rings Rubber 30 2
Quick Release Pin
Quick Release
Aluminium 2.7 0.503 4.5 2.26 6.11 2 12.21
Springs Springs Steel 7.2 10 2 20
Total (Structural)
1012.5
(g)
Table 16 – Itemised Mass Breakdown of all Structural UAV Components
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Table 17 - 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
354.6 1 354.6
Lights control board 46 28 13 73.2 1 73.2
Servos 35.5 22.5 12 13.4 1 13.4
Payload 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 297 1 297
Additional Systems battery 102 15 35 118 1 118
ESC’s for servo 50 30 15 55 1 55
Power Rails 87 20 18 52 2 104
Total (Electronics / Misc) 5499.3(g)
From the summation of all the masses for the electronic and miscellaneous components in Table 17, a total mass of 5499.3 grams was calculated. Total Mass of the UAV 6511.8 Grams.
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Appendix E. UAV Detailed Cost Breakdown
Material / Component
Used for (Appendix C.7 for
Structural Parts and Section 8 and 9 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
16mm Nylon 6.6 Block
TJ-001, TJ-002 £4.40 £1.01 £0.51
30mm Nylon 6.6 Block
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 & Board Lights & 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
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M3 Nylon Spacer 3.2mm internal,
outside 6mm, length 25mm
Fasteners £4.09 £0.68 £0.68
M5 Nylon Spacer 5.3mm internal, outside 10mm, length 10mm
Fasteners £3.39 £0.57 £0.57
M6 Nylon Spacer 6.4mm internal, outside 12.5mm,
length 10mm
Fasteners £3.59 £1.20 £1.20
M3 Brass Hexagonal – F/F -
Standoff Fasteners £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 £13.47 £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 £1071.46 £954.80 £906.18
Table 18 – 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
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Appendix F. Material Properties
Material ID Components (Appendix C.7)
Property Value
1 -
Al A
lloy
(MIL
-HD
BK
-5H
)
MP-001 Density ρ= 2770 kg/m3
Young’s Modulus = 7.1E10 Pa = 71 GPa
Poisson’s Ratio = 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,
20
15
)
Brass M3x25 F/F Spacers.
Density ρ= 8450 kg/m3
Young’s Modulus = 1.15E11 Pa = 115 GPa
Poisson’s Ratio = 0.331
Bulk Modulus K= 1.1341E11 Pa = 113.4 GPa
Shear Modulus G= 4.3201E10 Pa = 432 GPa
Tensile Yield Strength σTYS= 160 MPa
Ultimate Tensile Strength σUTS= 270 MPa
3 -
Nyl
on
66
[TEC
AM
ID-6
6-M
O-
Bla
ck]
(En
sin
ger,
20
15
b)
FB, MB, AP, LP, LB, EB & TJ Series. PB-009 & PB-010.
Density ρ= 1150 kg/m3
Modulus of Elasticity (Flexural) = 3100 MPa
Poisson’s Ratio = 0.4
Bulk Modulus K= 4.1667E9 Pa = 4.1667 GPa
Shear Modulus G= 8.9286E10 Pa = 89.28 GPa
Tensile Yield Strength σTYS= 83 MPa
Ultimate Tensile Strength σUTS= 84 MPa
4 -
Nyl
on
6
[TEC
AM
ID-6
-MO
-
Bla
ck]
(En
sin
ger,
20
15
a) BP-001,
BP-002, PB-004 & PB-005
Density ρ= 1140 kg/m3
Modulus of Elasticity (Flexural) = 3100 MPa
Poisson’s Ratio = 0.4
Bulk Modulus K= 4.1667E9 Pa = 4.1667 GPa
Shear Modulus G= 8.9286E10 Pa = 89.28 GPa
Tensile Yield Strength σTYS= 82 MPa
Ultimate Tensile Strength σUTS= 84 MPa
5 -
PV
C H
70
7 E
qu
iv
(Dir
ect_
Pla
stic
s,
20
15
)
FA-001, MA-001, UV-001 & UH-001
Density ρ= 1800 kg/m3
Modulus of Elasticity (Flexural) = 3100 MPa
Poisson’s Ratio = 0.41
Bulk Modulus K= 5.7407E9 Pa = 5.7407 GPa
Shear Modulus G= 1.0993E9 Pa = 1.0993 GPa
Tensile Yield Strength σTYS= 55 MPa
Ultimate Tensile Strength σUTS= 56 MPa
6 -
Au
stin
itic
Sta
inle
ss S
tee
l -
(Cla
ss 7
0, 3
04
gra
de
- co
ld
dra
wn
)
(BSS
A, 2
01
5)
M3 & M5 Fasteners and Nyloc Nuts.
Density ρ= 8030 kg/m3
Modulus of Elasticity = 193 GPa
Poisson’s Ratio = 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
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7 –
PV
C R
igid
An
gle
(Dir
ect_
Pla
stic
s,
20
15
)
PB-001, PB-002, PB-003, PB-006, PB-007 & PB-008
Density ρ= 1800 kg/m3
Modulus of Elasticity (Flexural) = 3100 MPa
Poisson’s Ratio = 0.41
Bulk Modulus K= 5.7407E9 Pa = 5.7407 GPa
Shear Modulus G= 1.0993E9 Pa = 1.0993 GPa
Tensile Yield Strength σTYS= 55 MPa
Ultimate Tensile Strength σUTS= 56 MPa
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Appendix G. Load Cases and Load Transfer (Supplemented by section 5.1)
G.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.
G.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 5.1) to the vertical. Using this along with the total surface area of the main body plates, the Drag force could be calculated.
Figure 37 - Project Main Body Area
( ) ( )
Equation 6 - Projected Area
For steady flight the motors produce enough thrust to balance the weight. Therefore the mass was is
by 4.
. 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.
( ) ( )
Equation 7 - Thrust at 54 Degrees
To calculate the drag force, the following equation is used:
Where: D = Drag Force, p = Density, V = Velocity, S = Area, Cd = Coefficient of Drag
Equation 8 - Drag Equation (R. H. Barnard, 2010)
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).
( )
( )
= 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.
Total Surface Area = 35129.71mm2 54°
Total Projected Area
54°
T
1.75Kg
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Appendix H. Stress Analysis
H.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 C.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 48 through Figure 52), therefore an additional local support (MB-001 – Appendix C.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).
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 already 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.
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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.
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.
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 H.16 and compared to analytical methods as shown in Appendix H.15. 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.
H.2. 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 C.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 H.12) is associated with parts AP-001, LP-001 and FB-002 (Appendix C.7).
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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).
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 C.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 H.12). 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 already been demonstrated in the arm stress analysis (Appendix H.5 and H.6)
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 H.12). 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 already been demonstrated in the arm stress analysis (Appendix H.5 and H.6).
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 C.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.
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H.3. 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.
H.4. Mesh
Element Types Used
SOLID187
The SOLID187 element used as per Table 19 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 38 – SOLID187 Element (Ansys, November 2013c)
PLANE182
The PLANE182 element used as per Table 19 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 39 – PLANE182 Element (Ansys, November 2013c)
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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 19).
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 C.7)
Material ID (Appendix F)
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
94 | P a g e
FB-001 FB-002 MB-001 EB-001 LB-001 LB-002 PB-009 PB-010
3 Element Type
Solid – Tet 10 node
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
LB-003 3 Element Type
Solid – Tet 10 node
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
Solid – Tet 10 node
Types of Mesh
TET10 – SOLID187
Size (Aspect Ratio)
1.28 (Min)
26.4 (Max)
2.82 (Ave)
Refinement None
Inflation None
Pinch Controls
None
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TJ-001 TJ-002
3 Element Type
Solid – Tet 10 node
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
Solid – Tet 10 node
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
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 19 – Mesh Attributes for Components
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H.5. Stationary Motor Arm Stress Analysis Materials: As per Appendix F
D1 = 0.06m & 0.067m D2 = 0.23m & 0.247m Transferring loads from F1 to F2 required a moment transfer using:
Equation 9 – Working out Moment
Figure 40 - Fixed-arm Cross-section
Fixed-arm
Full Arm length of 0.23m ( )
Equation 210 - 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.
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 41 - Arm Cross-section for Stress Calculation
To work out the stress in the arm the following equation was used.
(
)
Equation 11 - 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 42.
Figure 42 - Tension & Compression Stress in Arm
Reaction
Force
0.17m
0.23m
Moment
D1=0.016m
D2=0.0115m
Compression
Tension
F1 D1
F2
D2
97 | P a g e
Stress analysis at Fixed-arm – Analytical Solution
( )
( ) ( )
Stress analysis at Fixed-arm – FEA Method
Mesh: Values as per section H.4
Figure 43 – Mesh for Fixed-arm Assembly – Values as per Appendix H.4
Results:
Figure 44 - Deflection of Fixed-arm Assembly (Flight Loads) with 7.6mm Deflection
98 | P a g e
Figure 45 - Stress of Fixed-arm Assembly (Flight Loads) with Stress 15.8MPa (Contact) and 20MPa (Peak)
Figure 46 – 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 46 (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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H.6. Movable Arm Stress Concentration at Contact Points Materials: As per Appendix F
Movable-arm
Full Arm length of 0.247m ( )
Equation 12 - 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.
Analytical Stress at Movable-arm
( )
( ) ( )
The yield strength of the material used for the arms is 55MPa (Appendix F). 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 F are suitable for the UAV.
FEA Method for Movable-arm
Mesh: Values as per section H.4
Figure 47 – Mesh for Arm Assembly (With additional Tab) – Mesh Values as per H.4
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Results:
mOrange = rV =1.15g / cm3 ´3.142cm3 = 3.6133g
mBlue = rV =1.15g / cm3 ´3.501cm3 = 4.0262g
\Dm = +0.4129g
An additional 0.4129g results in slightly lower stress levels (see Figure 49 & Figure 51). 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 48 - Modified FB-002 for reduction in point contact stress concentration
Figure 49 - 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).
101 | P a g e
Figure 50 - Deflection of Modified Movable Motor Arm of 7.88mm for flight loads with SF
Figure 51 - Stress of Modified Movable Motor Arm of 20.8MPa for flight loads with SF
Figure 52 - Modified Movable Motor Arm with Stress of 20.8MPa for flight loads with SF (a) & Close-up (b)
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H.7. Undercarriage Lug Bracket Flange Addition Materials: As per Appendix F
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 53 - Load on the Lug (Niu, 1988)
To carry out the analysis the load was split into components as shown in Figure 54.
Thickness of Lug (t) = 4.5mm
Width (W) = 25mm
Diameter (D) = 5mm
( )
( )
Ultimate strength of the material (Ftu) = 85MPa
Figure 54 - Components of the Load (Niu, 1988)
The Areas on the lug were determined to be able to calculate the maximum allowable load.
((
) (
))
((
) (
))
Equation 13 - Area A1 on Lug (Niu, 1988)
Figure 55 - Areas on the Lug
(
) (
)
Equation 14 – Area A2 on Lug (Niu, 1988)
Equation 15 - Area A3 on Lug (Niu, 1988)
Equation 16 - Area A4 on Lug (Niu, 1988)
( )
Equation 17 - Average Area of Lug (Niu, 1988)
PA
PT
D W
PT
A2
A1 A3
A4
45'
103 | P a g e
( ) Equation 18 - Bearing Area on Lug (Niu, 1988)
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:
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 H.4
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 56 - Lug Bracket Without Flange (Left) & with additional Flange (Right)
mLeft = rV =1.15g / cm3 ´8.833cm3 =10.1579g
mRight = rV =1.15g / cm3 ´9.415cm3 =10.8272g
\Dm = +0.67g
An additional 0.67g results in lower stress levels (see Figure 57) and is also beneficial for repeated loading and impact consideration during a side impact landing as demonstrated in Appendix H.11.
Although the Lug Bracket (LB-003) in Figure 56 (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.
104 | P a g e
Figure 57 – 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 58 – Lateral Unit Load Deflection (Left) & Stress (Right) of Lug Bracket With Flange
105 | P a g e
H.8. Motor Plate Stress Analysis Materials: As per Appendix F
Mesh: Values as per section H.4
Figure 59 - Mesh for MP-001 (Appendix C.7) with values as per Appendix H.4
Results:
Figure 60 – Motor Plate Deflection (0.038 mm) and Stress (41.7 MPa) for flight case with SF at start-up
Figure 61 - 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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H.9. Main Body Plate Stress Analysis Materials: As per Appendix F
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 62 - 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.
( )
Equation 19 - Flexural Rigidity of the Plate (Ventsel and Krauthammer, 2001)
( ) ∑ ∑
Equation 20 – Navier solution (Ventsel and Krauthammer, 2001)
Equation 21 - Navier stokes coefficient 1 (Ventsel and Krauthammer, 2001)
[(
) (
)]
Equation 22 - Navier Stokes coefficient 2(Ventsel and Krauthammer, 2001)
First the pressure distributed on the whole plate surface was calculated.
Followed by calculating the flexural rigidity
( )
X = a = 315mm
Y =
b =
28
0m
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
107 | P a g e
The Navier coefficients 1 and 2 could be calculated for when mn = 1 1, 1 3, 3 1, 3 3
[(
) (
)]
[(
) (
)]
[(
) (
)]
[(
) (
)]
The coefficients were then input into the Navier solution equation to calculate the deflection at the centre.
( ) (
) (
)
(
) (
) (
)
(
) (
) (
)
( )
( )
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 63 - Simple Plate Deflection Carried out on CATIA structural analysis
108 | P a g e
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 20 – Comparison of Simplified Plate Deflection for Model Substantiation
FEA – As Built (Single Plate) Ansys Results
Mesh: Values as per section H.4
Figure 64 - Mesh of Main Body Plate - Values as per Appendix H.4
Results:
Figure 65 – Single Main Body Plate Analysis – with 17.8MPa Stress at contact holes for flight case with pressure load
109 | P a g e
H.10. Main Body Plate Stress Analysis as Built Ansys Results Materials: As per Appendix F
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 D.
Figure 66 – Mass Representation of components and payloads as per Appendix D
Mesh: Values as per section H.4
Figure 67 - Mesh of Main body assembly with Values as per Appendix H.4
Results:
Figure 68 – Contact model Flight Case for Main body assembly Deflection (left) and Equivalent Stress (right)
110 | P a g e
Figure 69 - Contact model Flight Case for Main body assembly - Equivalent Stress with predicted locations
H.11. Undercarriage Buckling and Torsion Cases Materials: As per Appendix F
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.
√
Equation 23 – Slenderness Ratio (Warren C. Young)
√
Equation 24 - Radius of Gyration (Warren C. Young)
( )
Equation 25 - Critical Load to Cause Buckling (Warren C. Young)
( )
Equation 26 - 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.
111 | P a g e
Working out the second moment of area of the tube.
( )
( )
Cross-sectional area of the tube, ( ) ( )
Radius of Gyration, √
Slenderness Ratio,
√
Therefore the buckling formula can be used for this scenario.
The critical load which would cause the leg to buckle is shown below.
(
)
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.
Figure 70 - Resolving Component to Determine Vertical Load
Analytical – Undercarriage Leg Bending
Figure 71 - 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
112 | P a g e
Equation 11 is used to determine the stress in the tube.
(
)
( )
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 11 the max force could be found out which would cause the undercarriage leg to yield.
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 72 - Undercarriage Leg Under Pure Torsion
Equation 27 - Angle of Twist (Warren C. Young)
G = Shear Modulus = 1.0993 x 109Pa
Equation 28 - Polar Moment (Warren C. Young)
F
113 | P a g e
( )
T=Torque Applied L= Length = 0.175m
Therefore the twist angle was calculated to be:
Equation 29 - Shear Stress (Warren C. Young)
For shear stress to be maximum, r (radius) needs to be maximum.
Yield in shear = 1099.3MPa (Appendix F) 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 73 - Stress Element A (Warren C. Young)
Using the stress element A as
shown in Figure 73, the
equations of combined load can
be used. A plan view of element
A has been shown below.
Figure 74 - Plan View of Stress Element A
T = Torque
Buckling Load
Bending Load
𝜏𝑥𝑦
𝜎𝑥
𝜎𝑦
𝜏𝑥𝑦
𝜏𝑦𝑥
𝜎𝑦
𝜎𝑥
𝜏𝑦𝑥
A
114 | P a g e
The formulas used to work out the combined stress and their principle angles have been shown below.
Equation 30 - Compression Stress on Pipe (Warren C. Young)
(
)
Equation 11 - Stress in a Cylindrical Pipe (Warren C. Young)
Equation 29 - Shear Stress (Warren C. Young)
√( )
Equation 31 - Principle Stress 1 and 2 (Warren C. Young)
The compression stress on the leg using Equation 30:
( )
Bending stress on the leg using Equation 11:
( )
Shear stress on the leg using Equation 29:
Using the above stresses the principle stress could be worked out using Equation 31:
√( )
√( )
Using the stresses above the principle angles were determined to show the direction they were in.
&
Equation 32 - Principle Stress Angles (Warren C. Young)
The principle stresses and their angles
could then be applied to
Figure 74.
Figure 75 - Stress Element A with Principle Stresses
𝜏𝑥𝑦
𝜎𝑥
𝜎𝑦
𝜏𝑥𝑦
𝜏𝑦𝑥 𝜎𝑦
𝜎𝑥
𝜏𝑦𝑥
A 𝜃
𝜃
27.095
-34.495
115 | P a g e
The maximum shear caused by the combined loadings has been calculated below using Equation 33.
√( )
Equation 33 - Shear Due to Combined Loadings
√( ) =
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 H.4
Figure 76 - Undercarriage Mesh for Contact Model with values as per H.4
FEA Results – Bending – Lateral Crash Landing
Figure 77 – Lateral Landing on Single Undercarriage Leg with 53.6mm Deflection
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Figure 78 and Figure 79 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.
Figure 78 - Lateral Landing on Single Undercarriage Leg with 60MPa Bending Stress
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 79 - Lateral Landing on Single Undercarriage Leg with 60MPa Bending Stress (Close-up)
117 | P a g e
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 80 - Tip Landing on Single Undercarriage Leg with 60MPa Bending Stress
FEA Results – Combined Torsion and Bending – Tip Contact
Figure 81 - Tip Landing on Single Undercarriage Leg with 66mm Combined bending and torsion deflection
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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
Figure 82 - 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
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H.12. Entire Quad Non-Liner Contact Model – Flight Case Parts: As per Appendix C.7
Materials: As per Appendix F
Mesh: Values as per section H.4
Results:
Figure 83 – Entire Quad-Rotor Flight Deflection of 7.9mm at Motor Arm Tips
Figure 84 - Entire Quad-Rotor Flight Deflection of 7.9mm at Motor Arm Tips (Close-up)
Figure 85 - Entire Quad-Rotor Flight Stress of 28.8 MPa at Motor mount plates
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Figure 86 - Entire Quad-Rotor Flight Stress with Plate Stress peak at 14.42Mpa
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H.13. Payload Housing Non-Liner Contact Model – Old Design Parts: As per Appendix C.7
Materials: As per Appendix F
Mesh: Values as per section H.4
Results:
Figure 87 – Downward Load - 1kg Payload and 10N Additional Load onto PB-005 Plate
Figure 88 - Side Load - 1kg Payload and 10N Additional Load onto Hinge Plate at 45deg to horizontal
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Figure 89 - Side Load - 1kg Payload and 10N Additional Load onto short edge 45deg to horizontal
Figure 90 - Side Load as per Figure 89 - Showing Pre-mature Release due to global deflection
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H.14. 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 C.7
Materials: As per Appendix F
Mesh: Values as per section H.4
Results:
Figure 91 – Downward Load as per Figure 87 with new design showing 0.73mm Deflection
Figure 92 - Side Load as per Figure 88 – with new rigid design and Deflection of 1.56mm
*No pre-mature
deployment during
manoeuvres as seen in
the previous design
from Figure 89 and
Figure 90
Figure 93 – Side Load as per Figure 89 and Figure 90 – with new design and deflection of 0.41mm*
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H.15. Modal Analysis of Fixed-arm – Simplified Case Parts: As per Appendix C.7
Materials: As per Appendix F
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 94 - 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.
( ) ( )
Equation 34 -Static Deflection Curve (MEGSON, 1999)
∫ (
)
∑ ( )
∫ ∑ ( )
Equation 35 - Rayleigh's Natural Frequency Equation (MEGSON, 1999)
To be able to calculate the natural frequency using Equation 35, the static deflection equation requires to be differentiated twice.
( )
(
) ( )
( )
The deflection where the concentrated mass is attached:
( ) ( )
Using that and inputting some of the values the equation becomes:
[
]
[
]
( )
√
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.
L = X = 0.234m
m = 0.36846Kg
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Finite Element Modal Analysis – Simplified
Parts: As per Appendix C.7
Materials: As per Appendix F
Mesh: Values as per section H.4
Results:
Figure 95 – Mass Representation of Motors, Blocks, Plates, Fasteners and ESC
Figure 96 – Simplified FE analysis with 1st Nat freq as 19.64Hz – 69.3mm Deflection (Left) and 164MPa Stress (Right)
Figure 97 – Simplified FE with 2nd Nat freq as 20.06 Hz (Left) and 3rd Nat freq as 134.6 Hz (Right)
Figure 98 – Simplified FE with 4th Nat freq as 224.1 Hz (Left) and 5th Nat freq as 411.9 Hz (Right)
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H.16. Modal Analysis of Fixed-arm – Actual Parts (As Built) Parts: As per Appendix C.7
Materials: As per Appendix F
Mesh: Values as per section H.4
Results:
Figure 99 – As Built FE Analysis - Mass Representation of Motors, Fasteners, Cables and ESC
Figure 100 – As Built FE analysis with 1st Nat freq as 451 Hz – 69.0mm Deflection (Left) and Stress (Right)
Figure 101 - As Built FE analysis with 2nd Nat freq as 736 Hz (Left) and 3rd Nat freq as 1707 Hz (Right)
Figure 102 - As Built FE analysis with 4th Nat freq as 2 KHz (Left) and 5th Nat freq as 4.1 KHz (Right)
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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
(H.15)
34.19 N/A N/A N/A N/A
Simplified FEA (H.15)
19.64 20.06 134.6 224.1 411.9
As-built FEA (H.16)
451 736 1707 2000 4100
Table 21 – Summary of Modal Frequencies for Fixed Motor Arm
As predicted from the Rayleigh method in Appendix H.15, 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 C.7 and Appendix F). Aluminium Alloy compared to the cast mild-steel motor brackets which come supplied with the motors are less susceptible to fatigue damage due to repetitive vibration.
H.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 H.2 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.
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Appendix I. 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
1. Flight performance calculations will be based on a worst case scenario were the quadcopter
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 44.77Amps and
power required as 829Watts
Initial starting point: On the runway with no power on
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
[Runway] 0 0 100 0
First leg: Quadcopter will take-off to its cruise altitude of 100ft ready to its maximum angle of 370
based on the quadcopter weighing 7kg. 100ft is used as the cruise so that when it approaches the
drop box it can perform a quicker drop of time. Also a quadcopter 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 manoeuvre is performed
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
[Runway] to [30.46m] 0 2.2 99.4% 0
As the time is known to vertically climb to height of 30.46m and also the current draw of 37.27A per
motor is obtained from practical testing on a sophisticated test rig, the battery percentage can
therefore be calculated
( )
( )
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( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( )
Therefore
Battery Status %
Second Leg: Now that the quadcopter is at a safe altitude maximum tilt angle of 37.30 can be
applied. Also using google earth the distance from runway to point [1] is calculated as 282m. From
section 1.7 it was calculated that the quadcopter can achieve maximum velocity of 20.37m/s. Taking
into account wind condition of 25knots (12.86m/s) then the quadcopter 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
( ) ( )
( )
( )
( )
( )
( )
( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( )
Therefore
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
[30.5m]-[ Waypoint 1] 7.51 37.55 87.2 282
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Battery Status %
Third Leg: After a quick bend at point [1] the quadcopter 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.
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
Waypoint [1]- Waypoint [2] 7.51 112.12 50.5 842
( )
( )
( )
( )
( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( )
Therefore
Battery Status %
Fourth Leg: Again at point [2] the quadcopter 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
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
Waypoint [2]- Waypoint [3] 7.51 55.7 32.3 418
( )
( )
( )
( )
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( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( )
Therefore
Battery Status %
Fifth Leg: Again for this section same performance criteria can be assumed
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
Waypoint [3]-[Target] 7.51 44.47 17.8 334
( )
( )
( )
( )
( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( )
Therefore
Battery Status %
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Hover and Navigation Leg: A this point the quadcopter 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
( )
( )
( )
( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( ) Ah
Therefore
Battery Status %
Final Leg: Final leg of the mission is to return from the target drop off point back to the runway
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
[Target]-[Runway] 7.51 16.51 9.93 124
( )
( )
( )
( )
( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( ) Ah
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Therefore
Battery Status %
Time Taken up to this point 288.55s (4.82minutes)
Reload Leg: At this point the quadcopter 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.
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
[Runway] to [30.46m] 0 2.2 98.8% 0
( )
( )
( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( )
Therefore
Battery Status %
Final Leg: The quadcopter will be at a height of 30.46m and will head towards the target.
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
[Runway]-[Target] 7.51 16.51 88.12 124
( )
( )
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( )
( )
( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( ) Ah
Therefore
Battery Status %
%
Again we will have the Hover and Navigation Leg: A this point the quadcopter 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
( )
( )
( )
( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( ) Ah
Therefore
Battery Status %
Lastly we have final leg again: Final leg of the mission is to return from the target drop off point back
to the runway
Location Velocity (m/s) Time (s) Battery Status (%) Distance Covered (m)
[Target]-[Runway] 7.51 16.51 73.1 124
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( )
( )
( )
( )
( )
( ) ( )
Battery Capacity remaining can be calculated
( )
( )
( ) Ah
Therefore
Battery Status %
Time taken from reload to landing = 55.22s (0.922 minutes)
Total Time Taken: 5.74minutes
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Appendix J. System Schematics
Figure 103 - Overall System Hardware Block Diagram Video graphics processing unit (VGPU)
Table 22 - Overall system hardware definitions
Signal name Description
User input The user input is to turn the quadcopter on/off, toggle on/off video stream, activate/ deactivate the autopilot and arm/disarm the quadcopter. This is achieved through ground control station APM Planner, using the radio controller switches for different controls and by plugging in the batteries to the quadcopter.
Power The power supply of the base station is from the laptop, where it must have the battery fully charged before the mission.
GPS Satellite Signal
The GPS system on the quad copter 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 quadcopter and base station is through the 3DR telemetry kit operating at 433MHz.
Video Data
The video data is sent through the 3DR telemetry 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 quadcopter 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.
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Figure 104 - Overall Software Block Diagram
Table 23 - Overall software definition
Signal name Description
Video
Commands
Video command will make the camera take a picture of the target for it to
then proceeded to read the alphanumeric information at the target.
Video Data The video data is sent from the camera on bored the quadcopter, with
information from the minimOSD. The video will be displayed on the ground
station APM planner.
Telemetry
Command
The telemetry command is sent from the base station APM 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 quadcopter such as the video or the
data from pixhawk. The data from pixhawk includes information such
altitude, attitude, location and speed which are displayed on the APM
planner page.
User Data The user data is the collection of the flight information which is displayed on
the APM 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
quadcopter.
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Figure 105- Pixhawk hardware connections
Figure 106- Quadcopter Propulsion setup
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Figure 107 - Transmitter and Receiver with Video Graphics Processing Unit (VGPU) the MinimOSD
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Figure 108 - Servo and motor control schematics
Flight control
system
(Pixhawk)
RBATTERY 2
(11.1V, 2.2Ah,
2S)
BATTERY
PACK 1
(18.5V, 16Ah,
3S)
PAYLO
AD
SERV
SBECS
M
1
M
2
M
3
M
4
RC
RECEIVER
ESC
ESC
ESC
ESC
+
-
5V – 6V
5A
BEC Out
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Appendix K. Verification and Validation Matrix
Validation Matrix
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 (apm 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 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)
25/03/2015
Transmission range
Test if the distance of transmission is the same as in telemetry kit test. 25/03/2015
Video display Connect the camera to the Minim OSD and verify if the display the live feed with the correct information
26/03/2015
Image processing
Video display Program the Minim OSD to display the battery life, altitude, attitude and direction.
26/03/2015
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.
06/04/2015
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Pixhawk
Testing pixhawk sensors
The servo test will try to operate four servo channels connected on the USART2 pins
25/03/2015
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.
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 27/03/2015
Navigation system
Signal strength Examine the GPS location test on different weather conditions at different locations, to test its signal triangulation.
23/03/2015
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 quadcopter 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
Safety test Drop the final built quad copter from 20 cm from ground to represent landing
03/04/2015
Test the final model in the wind tunnel to see the structural rigidly 31/03/2015
Visually inspect all electrical components and wires to make sure there is no loose wires or components that may get effected during flight
27/03/2015
Physically asses all part and components are secure together 27/03/2015
PID controller
test
This is carried by creating a mathematical model of the UAS and simulating the dynamic behaviour with the use of matlab.
29/03/2015
A test rig is built for the sole purpose of testing the UAS in order to set its PID numbers before the initial flight.
29/03/2015
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.
02/04/2015
Ground
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control station
evaluation
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 ouputed from pixhawk but was not strong enough to operate the 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.
26/03/2015
Deploy the payload at the set destination or target (as the quadcopter would not be built yet an initial test will be carried out as when the quadcopter reached a destination or target it would emit a signal for the servo to deploy the payload.
24/03/2015
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
Ground control station
01/04/2015
Test the return home function after signal is lost, during period more than 30sec.
03/04/2015
Testing maximum flight time with motors running at full speed. 03/04/2015
Testing the ground control
operation
Plan the journey of the quad copter using the ground control software, with a set coordinates.
10/03/2015
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 02/04/2015
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Verification matrix
Requirements Verification
Requirement number
Requirement (IMechE, Jan 2015) Inspection Analysis 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 coordinates 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 copter from a set distance.
2.8 The UAS shall be controllable in forward speed together with 3 axis control (Roll, Pitch and Yaw).
Pixhawk will be able to achieve this and it will be demonstrated by setting the quad copter 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 copters 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.
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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.
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 copter 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 quadcopter.
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.
4.2 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.
4.7 Consideration for maximum 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 in Error! Reference source not found. 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 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 copter 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 copter 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.
7.5 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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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 cables.
This test can be demonstrated by working a safe location where the Tx is disconnected allowing for a signal loss, hence allowing the quadcopter to return to home.
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Appendix L. Manufacturing
Figure 109: Nylon 6.6 Rod Figure 110: Nylon 6.6 Sheet
Figure 111: Nylon 6 sheet
Figure 112: PVC rigid angle section Figure 113: PVC Hard Plastic Tube
Figure 114: AL-2024-T6 Sheet
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Figure 115: Orientation of brackets in Quad copter
Figure 116: Machined fixed brackets by
Figure 117: Machined end bracket
Figure 118: Machined movable arm support bracket Figure 119: Machined movable arm vertical fixed bracket
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Figure 120: Machined landing gear top support bracket Figure 121: Machined arm pivot for movable arm
Figure 122: Main body plate after laser cutting Figure 123: Laser cutting of Nylon 6 extrude
Figure 124 - Melted edges on main body plate after laser cutting
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Figure 125: Milling arm pivot Figure 126: Drilling centre hole in fixed bracket
Figure 127: Chamfering of movable arm support bracket Figure 128: Smoothing surface by fly cutter
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Figure 129.1-2: Drilling using slot drills
Figure 130: High speed steel tool
Figure 131.1-2: Machining arm pivot on lathe
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Figure 132.1-2 Laser Cutting of Nylon 6 sheet for main body plate
Figure 133:Cutting Nylon 6.6 cast block in vertical band saw machine
Figure 134.1-3: Practising samples on CNC machine
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Appendix M. Test Rig
Initial Gimbal Test Rig Conceptual Design
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Updated Octagonal Gimbal Test Rig Assembly
Octagonal Model Mount Frame Technical Drawing
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Octagonal Mid Frame Technical Drawing
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Octagonal Outer Frame Technical Drawing
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Octagonal Gimbal Test Rig Stand Technical Drawing
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Gimbal Test Rig Weight Estimation
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Gimbal Test Rig Manufacturing Cost
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NIU, C. 1988. Airframe structural design: practical design information and data on aircraft structures, Conmilit Press.
PILKEY, W. D. & PILKEY, D. F. 2008. Peterson's stress concentration factors, John Wiley & Sons. R. H. BARNARD, D. R. P. 2010. Aircraft Flight, Pearson Education Limited 1989, 2010. VENTSEL, E. & KRAUTHAMMER, T. 2001. Thin Plates and Shells: Theory: Analysis, and Applications, CRC Press. WARREN C. YOUNG, R. G. B. Roark's Formulas For Stress and Strain. Seventh Edition, 1-855.