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The University of AdelaideSchool of Mechanical Engineering
859 Design and build of a UAV with
morphing configuration
MORPHEUS - Final Report
Kevin Chan 1132668
Crystal Forrester 1118686
Ian Lomas 1132921
Simon Mitchell 1132439
Carlee Stacey 1132235
Supervisor:
Dr. Maziar Arjomandi
Executive summary
This report details the design and development of a morphing Unmanned Aerial Vehicle
(UAV) by a group of five undergraduate engineering students from the School of Mechani-
cal Engineering at the University of Adelaide during 2009. Sharing a common background
in Aerospace Engineering, the students aimed to develop a remotely piloted UAV capable
of morphing between two different configurations. Dubbed ‘The Morpheus Project,’ the
aircraft design was driven towards a multi-mission platform which reduces the need for
performance compromise during different flight phases.
The conceptual design of the airframe was derived using a classical approach, based on an
extensive feasibility study and statistical analysis of the global UAV and morphing tech-
nology industries. Motivated by aerodynamic, structural and manufacturing limitations,
a telescoping wing and tail concept was developed based on a conventional aircraft config-
uration. The aircraft platform features non-tapered outboard wing sections which extend
and retract from a tapered inboard wing section. To control the longitudinal stability
of the aircraft during flight, a telescoping tail boom extends and retracts from the rear
of the fuselage. While this design presents numerous challenges, particularly in terms of
stability and manufacturing, the overall airframe demonstrates an innovative and creative
approach to engineering design.
The aircraft is to be primarily constructed from composite materials to provide struc-
tural strength and rigidity whilst minimising weight. The use of an electric propulsion
system consisting of a brushless motor and lithium-polymer battery technology allowed
for a reduction in aircraft complexity and development time. Stable and sustained flights
were achieved in all possible aircraft configurations, and morphing during flight was also
demonstrated. The aircraft has a theoretical maximum speed of 147km/h in the ex-
tended configuration and 166 km/h in the retracted configuration. The aircraft has also
demonstrated the capability of 700g of payload, and has a theoretical endurance of 36
minutes.
From the beginning, the project objectives were deemed ambitious due to the difficulty in
developing and manufacturing the morphing mechanisms, and the reliance of all project
goals on successful test flights. The resourcefulness of the group provided a strong founda-
tion from which the majority of the primary goals were achieved. Several extended goals
were also specified to provide the group with additional challenges to an already ambitious
project. Theoretical calculations were performed toward the achievement of these goals;
however there was insufficient time available for flight testing. The work undertaken by
the project group provides a solid basis for further development of the Morpheus UAV.
Disclaimer
We, the authors, declare that the material contained within this report is entirely our
own, unless otherwise specified.
Kevin Chan
Date:
Crystal Forrester
Date:
Ian Lomas
Date:
Simon Mitchell
Date:
Carlee Stacey
Date:
Acknowledgements
We, the authors, would like to acknowledge the contributions made by many people
throughout the course of the project; without their support and guidance, the project
would not have been successfully completed. The authors would like to thank the project
supervisor, Dr Maziar Arjomandi, who has provided the group with invaluable guidance
and technical knowledge. Thanks must also go to the Electronics Workshop, and the
technicians at the Mechanical Workshop, particularly Mr Bill Finch and Mr Richard
Pateman, for their consultation and technical expertise.
The authors are sincerely thankful to the main sponsors of the project; Aeronautical En-
gineers Australia and Babcock Integrated Technology Australia. Without their financial
support, the project would not have been possible. The authors would also like to thank
Australian Aerospace Limited, who has provided the group with technical assistance and
in-kind support.
Finally, the authors would like to thank their families and friends for their support over
the duration of the project.
Contents
Executive summary i
Disclaimer ii
Acknowledgements iii
Glossary xx
1 Introduction 1
1.1 Motivation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1
1.2 Aims and objectives . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2
1.2.1 Primary project goals . . . . . . . . . . . . . . . . . . . . . . . . . . 2
1.2.2 Extended project goals . . . . . . . . . . . . . . . . . . . . . . . . . 3
1.3 Scope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4
2 Literature review and feasibility study 5
2.1 Literature review . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5
2.2 Market evaluation and benchmarking . . . . . . . . . . . . . . . . . . . . . 6
2.2.1 Virginia Tech BetaMax Morphing Wing Project . . . . . . . . . . . 6
2.2.2 Delft University of Technology Roboswift . . . . . . . . . . . . . . . 7
2.2.3 Lockheed Martin Skunk Works Morphing UAV Concept . . . . . . 8
2.2.4 NextGen Aeronautics MFX-2 . . . . . . . . . . . . . . . . . . . . . 8
2.3 Analysis of morphing methods . . . . . . . . . . . . . . . . . . . . . . . . . 9
2.3.1 Wing morphing methods . . . . . . . . . . . . . . . . . . . . . . . . 9
2.3.2 Tail morphing methods . . . . . . . . . . . . . . . . . . . . . . . . . 10
2.4 Technical task . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10
2.4.1 Standard Requirements . . . . . . . . . . . . . . . . . . . . . . . . . 10
2.4.2 Special systems and miscellaneous . . . . . . . . . . . . . . . . . . . 11
2.4.3 Performance parameters . . . . . . . . . . . . . . . . . . . . . . . . 12
2.4.4 Technical level of product . . . . . . . . . . . . . . . . . . . . . . . 14
2.4.5 Economical parameters . . . . . . . . . . . . . . . . . . . . . . . . . 14
2.4.6 Power plant type and requirements . . . . . . . . . . . . . . . . . . 14
2.4.7 Main system parameter requirements . . . . . . . . . . . . . . . . . 15
2.4.8 Reliability and maintenance . . . . . . . . . . . . . . . . . . . . . . 16
iv
2.4.9 Unification level . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
2.5 Mission profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 16
2.6 Summary of design requirements and feasibility . . . . . . . . . . . . . . . 17
3 Conceptual design 18
3.1 Aircraft configuration design . . . . . . . . . . . . . . . . . . . . . . . . . . 18
3.2 Initial aircraft concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
3.2.1 Delta wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 18
3.2.2 Lifting body . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19
3.2.3 Telescopic wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
3.2.4 Folding wing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
3.2.5 Selected concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . 21
3.3 Wing morphing concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22
3.3.1 Telescopic wing configuration . . . . . . . . . . . . . . . . . . . . . 22
3.3.2 Wing mechanism selection . . . . . . . . . . . . . . . . . . . . . . . 24
3.4 Empennage morphing concepts . . . . . . . . . . . . . . . . . . . . . . . . 25
3.4.1 Telescopic fuselage . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
3.4.2 Sliding tail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
3.4.3 Boom-mounted tail . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
3.5 Mechanism actuator concepts . . . . . . . . . . . . . . . . . . . . . . . . . 27
3.5.1 Rack and pinion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
3.5.2 Winch . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
3.5.3 Pneumatics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
3.5.4 Threaded rod actuator . . . . . . . . . . . . . . . . . . . . . . . . . 30
3.5.5 Morphing mechanism actuator selection . . . . . . . . . . . . . . . 30
3.6 Aircraft Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 31
3.6.1 Statistical Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
3.6.2 Preliminary design parameters . . . . . . . . . . . . . . . . . . . . . 35
3.6.3 Sizing criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38
3.6.4 Matching diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . 39
3.6.5 Aileron sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41
3.7 Empennage conceptual design . . . . . . . . . . . . . . . . . . . . . . . . . 41
3.7.1 Tail/fuselage configuration analysis . . . . . . . . . . . . . . . . . . 42
3.7.2 Empennage configuration analysis . . . . . . . . . . . . . . . . . . . 42
3.7.3 Empennage sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . 43
3.7.4 Ruddervator sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . 43
3.7.5 Tail geometry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44
3.8 Propulsion system design . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45
3.8.1 Propulsion type selection . . . . . . . . . . . . . . . . . . . . . . . . 45
3.8.2 Electric motor selection . . . . . . . . . . . . . . . . . . . . . . . . 46
3.8.3 ESC selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47
3.8.4 Battery selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47
3.8.5 Energy requirements . . . . . . . . . . . . . . . . . . . . . . . . . . 47
3.8.6 Propeller solutions . . . . . . . . . . . . . . . . . . . . . . . . . . . 48
3.8.7 Propeller selection . . . . . . . . . . . . . . . . . . . . . . . . . . . 49
3.9 Landing gear configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . 50
3.10 Fuselage sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 51
3.11 Conceptual design summary . . . . . . . . . . . . . . . . . . . . . . . . . . 51
4 Preliminary and Detailed Design 54
4.1 Wing design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 54
4.1.1 Airfoil selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55
4.1.2 Installed incidence angles . . . . . . . . . . . . . . . . . . . . . . . . 57
4.1.3 Wing loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 58
4.1.4 Wing structural layout . . . . . . . . . . . . . . . . . . . . . . . . . 63
4.1.5 Structural analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . 68
4.1.6 Wing design summary . . . . . . . . . . . . . . . . . . . . . . . . . 75
4.2 Empennage design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75
4.2.1 Airfoil selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 75
4.2.2 Stall recovery and installed incidence angle . . . . . . . . . . . . . . 76
4.2.3 Empennage loading . . . . . . . . . . . . . . . . . . . . . . . . . . . 76
4.2.4 Structural layout . . . . . . . . . . . . . . . . . . . . . . . . . . . . 78
4.2.5 Structural analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . 80
4.2.6 Tail rail shear analysis . . . . . . . . . . . . . . . . . . . . . . . . . 82
4.2.7 Tail rail bending analysis . . . . . . . . . . . . . . . . . . . . . . . . 82
4.2.8 Empennage design summary . . . . . . . . . . . . . . . . . . . . . . 82
4.3 Morphing mechanism design . . . . . . . . . . . . . . . . . . . . . . . . . . 82
4.3.1 Morphing mechanism loads . . . . . . . . . . . . . . . . . . . . . . 82
4.3.2 Threaded rod design . . . . . . . . . . . . . . . . . . . . . . . . . . 83
4.3.3 Motor design and selection . . . . . . . . . . . . . . . . . . . . . . . 84
4.3.4 Roller design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86
4.4 Control system design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 87
4.4.1 Thrust subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . 87
4.4.2 Control surfaces subsystem . . . . . . . . . . . . . . . . . . . . . . . 88
4.4.3 Morphing subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . 89
4.5 Fuselage design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90
4.5.1 Component layout . . . . . . . . . . . . . . . . . . . . . . . . . . . 90
4.5.2 Structural layout . . . . . . . . . . . . . . . . . . . . . . . . . . . . 91
4.5.3 Weight distribution and centre of gravity . . . . . . . . . . . . . . . 92
4.5.4 Landing gear . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 94
4.5.5 Fuselage loads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 95
4.5.6 Structural analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . 97
4.5.7 Fuselage design summary . . . . . . . . . . . . . . . . . . . . . . . . 100
4.6 Flight performance analysis . . . . . . . . . . . . . . . . . . . . . . . . . . 100
4.6.1 Longitudinal stability analysis . . . . . . . . . . . . . . . . . . . . . 100
4.6.2 Theoretical performance . . . . . . . . . . . . . . . . . . . . . . . . 103
4.6.3 Differential telescoping analysis . . . . . . . . . . . . . . . . . . . . 104
4.6.4 Optimal configurations for various flight phases . . . . . . . . . . . 105
4.7 Preliminary and Detailed Design Summary . . . . . . . . . . . . . . . . . . 106
5 Manufacturing 108
5.1 Available manufacturing methods . . . . . . . . . . . . . . . . . . . . . . . 108
5.1.1 Foam cutting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 108
5.1.2 Composite layup . . . . . . . . . . . . . . . . . . . . . . . . . . . . 109
5.2 Common components found in the Morpheus UAV . . . . . . . . . . . . . 110
5.2.1 Ribs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 110
5.2.2 Leading and trailing edges . . . . . . . . . . . . . . . . . . . . . . . 110
5.2.3 Control surfaces . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 110
5.2.4 Carbon fibre components . . . . . . . . . . . . . . . . . . . . . . . . 111
5.3 Inboard wing construction . . . . . . . . . . . . . . . . . . . . . . . . . . . 111
5.3.1 Foam core . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 111
5.3.2 Ribs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112
5.3.3 Spars . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112
5.3.4 Fibreglass skin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112
5.4 Outboard wing construction . . . . . . . . . . . . . . . . . . . . . . . . . . 113
5.4.1 Foam core . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 113
5.4.2 Ribs . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 113
5.4.3 Carbon fibre components . . . . . . . . . . . . . . . . . . . . . . . . 113
5.4.4 Fibreglass skin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114
5.5 Fuselage construction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114
5.5.1 Plug . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 114
5.5.2 Skin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115
5.5.3 Fuselage internal structure . . . . . . . . . . . . . . . . . . . . . . . 115
5.6 Empennage construction . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115
5.6.1 Foam core . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115
5.6.2 Fibreglass skin . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 116
5.6.3 Tail sliding block . . . . . . . . . . . . . . . . . . . . . . . . . . . . 116
5.7 Aircraft assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 116
5.7.1 Fuselage internal structure installation . . . . . . . . . . . . . . . . 117
5.7.2 Outboard wing and wing sliding block installation . . . . . . . . . . 117
5.7.3 Inboard wing installation . . . . . . . . . . . . . . . . . . . . . . . . 118
5.7.4 Empennage installation . . . . . . . . . . . . . . . . . . . . . . . . . 118
5.7.5 Undercarriage installation . . . . . . . . . . . . . . . . . . . . . . . 118
5.8 Electronics installation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 119
5.8.1 Propulsion system installation . . . . . . . . . . . . . . . . . . . . . 119
5.8.2 Morphing system installation . . . . . . . . . . . . . . . . . . . . . 120
5.8.3 Radio control system installation . . . . . . . . . . . . . . . . . . . 121
5.9 Painting and finishing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 122
5.9.1 Two-pack paint . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 122
5.9.2 Solartrim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 122
5.10 The completed Morpheus UAV . . . . . . . . . . . . . . . . . . . . . . . . 122
6 Testing 123
6.1 Component tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123
6.1.1 Propulsion - Static Thrust Test . . . . . . . . . . . . . . . . . . . . 123
6.1.2 Morphing Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . 124
6.1.3 Wing - Structural Test . . . . . . . . . . . . . . . . . . . . . . . . . 126
6.1.4 Assembled Electronics, Morphing and Control Systems . . . . . . . 127
6.2 Flight testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127
6.2.1 Heavy model certification . . . . . . . . . . . . . . . . . . . . . . . 128
6.2.2 Balance & stability . . . . . . . . . . . . . . . . . . . . . . . . . . . 128
6.2.3 Ground test - range checks . . . . . . . . . . . . . . . . . . . . . . . 128
6.2.4 Ground handling tests . . . . . . . . . . . . . . . . . . . . . . . . . 129
6.2.5 Stability Test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 130
6.2.6 Airworthiness test . . . . . . . . . . . . . . . . . . . . . . . . . . . . 133
6.2.7 Morphing test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134
6.2.8 Endurance test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 135
6.2.9 Flight parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . . 136
6.3 Evaluation of airframe and flight performance . . . . . . . . . . . . . . . . 136
6.3.1 Flight Performance . . . . . . . . . . . . . . . . . . . . . . . . . . . 136
6.3.2 Stability and Controllability . . . . . . . . . . . . . . . . . . . . . . 137
6.3.3 Morphing Mechanism Performance . . . . . . . . . . . . . . . . . . 137
6.3.4 RC System Performance . . . . . . . . . . . . . . . . . . . . . . . . 138
7 Management 139
7.1 Management structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 139
7.1.1 Technical coordinator . . . . . . . . . . . . . . . . . . . . . . . . . . 139
7.1.2 Logistics coordinator . . . . . . . . . . . . . . . . . . . . . . . . . . 140
7.1.3 CAD officer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140
7.1.4 Manufacturing coordinator . . . . . . . . . . . . . . . . . . . . . . . 140
7.1.5 Procurements and assemblies coordinator . . . . . . . . . . . . . . . 141
7.1.6 Quality assurance officer . . . . . . . . . . . . . . . . . . . . . . . . 141
7.1.7 Test coordinator . . . . . . . . . . . . . . . . . . . . . . . . . . . . 141
7.1.8 Safety officer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 141
7.2 Risk management . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142
7.3 Resource Management . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142
7.3.1 Project meetings . . . . . . . . . . . . . . . . . . . . . . . . . . . . 142
7.3.2 Scheduling . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143
7.3.3 Labour . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143
7.3.4 Finances . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 144
8 Conclusion 146
8.1 Project Achievements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 146
8.1.1 Primary Project Goals . . . . . . . . . . . . . . . . . . . . . . . . . 146
8.1.2 Extended project goals . . . . . . . . . . . . . . . . . . . . . . . . . 147
8.1.3 Additional achievements . . . . . . . . . . . . . . . . . . . . . . . . 148
8.2 Issues and setbacks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 149
8.3 Future work and recommendations . . . . . . . . . . . . . . . . . . . . . . 149
8.4 Project summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 150
Reference List 151
Appendices 152
A Electronics subsystems specification and design 153
A.1 Specifications of electronic components . . . . . . . . . . . . . . . . . . . . 153
A.1.1 Battery specifications . . . . . . . . . . . . . . . . . . . . . . . . . . 153
A.1.2 Radio control specifications . . . . . . . . . . . . . . . . . . . . . . 153
A.1.3 Motor and ESC specifications . . . . . . . . . . . . . . . . . . . . . 153
A.2 Wiring diagram - Thrust subsystem . . . . . . . . . . . . . . . . . . . . . . 154
A.3 Wiring diagram - Control surfaces subsystem . . . . . . . . . . . . . . . . . 154
A.4 Wiring diagram - Morphing subsystem . . . . . . . . . . . . . . . . . . . . 154
B Landing gear positioning 155
C Fuselage load calculation 157
C.1 Empennage aerodynamic loads . . . . . . . . . . . . . . . . . . . . . . . . 157
C.2 Inertial loads . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 158
C.3 Wing aerodynamic loads . . . . . . . . . . . . . . . . . . . . . . . . . . . . 158
C.4 Shear and bending moment diagrams . . . . . . . . . . . . . . . . . . . . . 159
C.5 Full aileron roll torsion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 159
C.6 Static thrust . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 160
D Theoretical performance calculations 161
D.1 Wing and power loading . . . . . . . . . . . . . . . . . . . . . . . . . . . . 161
D.2 Stall speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 161
D.3 Takeoff distance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 162
D.4 Drag polar . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 162
D.5 Maximum speed . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 162
D.6 Endurance . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 163
D.7 Rate of climb . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 163
D.8 Performance summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 163
E Manufacturing photos 165
F Component test procedures 167
F.1 Propulsion System Static Thrust Test . . . . . . . . . . . . . . . . . . . . . 167
F.1.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 167
F.1.2 Intended results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 167
F.1.3 SOP required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 167
F.1.4 Related/required tests . . . . . . . . . . . . . . . . . . . . . . . . . 167
F.1.5 Apparatus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 168
F.1.6 Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 169
F.1.7 Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 169
F.1.8 To do . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 172
F.1.9 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173
F.2 Mechanism motor test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173
F.2.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173
F.2.2 Intended results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173
F.2.3 Project phase . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173
F.2.4 SOP required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173
F.2.5 Other/related tests required . . . . . . . . . . . . . . . . . . . . . . 173
F.2.6 Apparatus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 174
F.2.7 Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 174
F.2.8 To Do . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175
F.2.9 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 175
F.3 Wing Structural Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 177
F.3.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 177
F.3.2 Intended results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 177
F.3.3 SOP required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 177
F.3.4 Related/required tests: . . . . . . . . . . . . . . . . . . . . . . . . . 177
F.3.5 Apparatus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 178
F.3.6 Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 178
F.3.7 Loading conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . 180
F.3.8 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 181
F.3.9 Assembly electronics, morphing and control test checklist . . . . . . 183
G Heavy model certification 184
H Heavy model requirements 186
I Flight test procedures 188
I.1 Pre-flight ground checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . 188
I.1.1 Things to Check . . . . . . . . . . . . . . . . . . . . . . . . . . . . 188
I.1.2 Actual Tests . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 188
I.1.3 Electronics start-up procedure . . . . . . . . . . . . . . . . . . . . . 189
I.1.4 Inboard wing installation . . . . . . . . . . . . . . . . . . . . . . . . 190
I.1.5 Outboard wing Installation . . . . . . . . . . . . . . . . . . . . . . 190
I.1.6 Tail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 190
I.1.7 Ready to fly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 191
I.1.8 Landing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 191
I.1.9 End of consecutive flights . . . . . . . . . . . . . . . . . . . . . . . 192
I.1.10 End of flight/day . . . . . . . . . . . . . . . . . . . . . . . . . . . . 192
I.1.11 To Change thrust batteries . . . . . . . . . . . . . . . . . . . . . . . 192
I.1.12 Trouble shooting Ground Checks . . . . . . . . . . . . . . . . . . . 192
I.1.13 Top 10 trouble shooting . . . . . . . . . . . . . . . . . . . . . . . . 193
I.2 Range Checks . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 194
I.3 PF1 - Ground handling - Taxi test . . . . . . . . . . . . . . . . . . . . . . 195
I.3.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 195
I.3.2 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 195
I.3.3 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 196
I.4 PF2 - Ground Run . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 196
I.4.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 196
I.4.2 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 197
I.4.3 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 197
I.5 F1 - Stability test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 198
I.5.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 198
I.5.2 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 198
I.5.3 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 199
I.6 Propulsion System Static Motor Test . . . . . . . . . . . . . . . . . . . . . 199
I.6.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 199
I.6.2 Intended results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 200
I.6.3 SOP required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 200
I.6.4 Related/required tests . . . . . . . . . . . . . . . . . . . . . . . . . 200
I.6.5 Apparatus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 201
I.6.6 Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202
I.6.7 Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 202
I.6.8 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205
I.7 Propulsion System Static Motor Test . . . . . . . . . . . . . . . . . . . . . 205
I.7.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205
I.7.2 Intended results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205
I.7.3 SOP required . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205
I.7.4 Related/required tests . . . . . . . . . . . . . . . . . . . . . . . . . 206
I.7.5 Apparatus . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 206
I.7.6 Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207
I.7.7 Method: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 207
I.7.8 TO DO: . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209
I.7.9 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209
I.7.10 Conclusion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 209
I.8 F2 - Airworthiness test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210
I.8.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210
I.8.2 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 210
I.8.3 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211
I.9 F3 - Morphing mechanism test . . . . . . . . . . . . . . . . . . . . . . . . . 211
I.9.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 211
I.9.2 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 212
I.9.3 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 214
I.9.4 Weather conditions . . . . . . . . . . . . . . . . . . . . . . . . . . . 216
I.10 F4 - Endurance test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216
I.10.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216
I.10.2 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216
I.11 F5 - Performance parameter tests . . . . . . . . . . . . . . . . . . . . . . . 217
I.11.1 Ext Goal 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217
I.11.2 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 217
I.11.3 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218
I.12 F6 - Differential span roll control test . . . . . . . . . . . . . . . . . . . . . 218
I.12.1 Aim . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218
I.12.2 Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 218
I.12.3 Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 220
J Risk management Plan 221
K Meeting minutes 233
L Gantt Charts 305
M Labour 306
N Documents used in obtaining sponsorship 307
O Business plan 310
List of Figures
1 Coordinate System Designation . . . . . . . . . . . . . . . . . . . . . . . . xxvi
2.1 Virginia Tech BetaMax Morphing Wing Project (Tech 2004) . . . . . . . . 7
2.2 Delft University of Technology Roboswift (Roboswift 2009) . . . . . . . . . 7
2.3 Lockheed Martin Skunk Works Morphing UAV (Martin 2009) . . . . . . . 8
2.4 NextGen Aeronautics MFX-2 (2009 2009) . . . . . . . . . . . . . . . . . . 9
2.5 Mission profile . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
3.1 Sliding plates concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 19
3.2 Sketch of the lifting body morphing concept . . . . . . . . . . . . . . . . . 19
3.3 Telescopic aircraft sketch . . . . . . . . . . . . . . . . . . . . . . . . . . . . 20
3.4 Folding wing concept in folded configuration . . . . . . . . . . . . . . . . . 21
3.5 External telescoping wing section with rectangular planform . . . . . . . . 23
3.6 Internal telescoping section without taper . . . . . . . . . . . . . . . . . . . 23
3.7 Internal telescoping with taper . . . . . . . . . . . . . . . . . . . . . . . . . 24
3.8 Twin rail concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 25
3.9 Roller concepts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
3.10 Telescopic fuselage concept . . . . . . . . . . . . . . . . . . . . . . . . . . . 26
3.11 Sliding tail concept . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 27
3.12 Boom-mounted tail concept . . . . . . . . . . . . . . . . . . . . . . . . . . 27
3.13 Wing rack and pinion sketch . . . . . . . . . . . . . . . . . . . . . . . . . . 28
3.14 Tail rack and pinion sketch . . . . . . . . . . . . . . . . . . . . . . . . . . . 28
3.15 Winch mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 29
3.16 Wing pneumatic mechanism sketch . . . . . . . . . . . . . . . . . . . . . . 30
3.17 Tail pneumatic mechanism sketch . . . . . . . . . . . . . . . . . . . . . . . 30
3.18 Wing threaded rod actuator sketch . . . . . . . . . . . . . . . . . . . . . . 31
3.19 Tail threaded rod actuator sketch . . . . . . . . . . . . . . . . . . . . . . . 31
3.20 Technology diagram for UAVs with a takeoff weight between 1.8kg and 28.1kg. . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 32
3.21 Stall and cruise speeds versus takeoff weight. . . . . . . . . . . . . . . . . . 33
3.22 Wing span versus takeoff weight for electric UAVs . . . . . . . . . . . . . . 34
3.23 TO distance versus takeoff weight for four electric UAVs . . . . . . . . . . 35
3.24 Matching diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 40
3.25 Final empennage configuration . . . . . . . . . . . . . . . . . . . . . . . . . 42
xv
3.26 Preliminary tail geometry . . . . . . . . . . . . . . . . . . . . . . . . . . . 44
3.27 Recommended propeller pitch . . . . . . . . . . . . . . . . . . . . . . . . . 50
4.1 Lift to drag ratio of candidate inboard wing airfoils . . . . . . . . . . . . . 56
4.2 Lift to drag ratio of candidate outboard wing airfoils . . . . . . . . . . . . 57
4.3 V-n diagram for the retracted wing configuration . . . . . . . . . . . . . . 59
4.4 V-n diagram for the extended wing configuration . . . . . . . . . . . . . . 59
4.5 Spanwise lift distribution for both wing configurations . . . . . . . . . . . . 60
4.6 Extended wing configuration load distribution . . . . . . . . . . . . . . . . 61
4.7 Retracted wing configuration load distribution . . . . . . . . . . . . . . . . 61
4.8 Wing shear diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 62
4.9 Wing bending moment diagram . . . . . . . . . . . . . . . . . . . . . . . . 62
4.10 Torque as a function of angle of attack . . . . . . . . . . . . . . . . . . . . 63
4.11 Schematic of the outboard wing structural layout . . . . . . . . . . . . . . 65
4.12 Schematic of the inboard wing structural layout . . . . . . . . . . . . . . . 66
4.13 Removeable tip rib . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 67
4.14 Schematic of the wing block structural layout . . . . . . . . . . . . . . . . 68
4.15 Fuselage attachment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 68
4.16 Wing tongue brackets . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 69
4.17 Wing shear stress distribution for carbon-fibre members . . . . . . . . . . . 70
4.18 Wing bending stress distribution . . . . . . . . . . . . . . . . . . . . . . . 71
4.19 Bending stress in the leading rail, tongue and reinforcement tubes . . . . . 72
4.20 Wing skin torsional stress for both wing configurations . . . . . . . . . . . 74
4.21 Shear diagram for the fuselage and empennage . . . . . . . . . . . . . . . . 77
4.22 Bending moment diagram for the fuselage and empennage . . . . . . . . . 77
4.23 V-tail and boom structural layout . . . . . . . . . . . . . . . . . . . . . . . 79
4.24 Boom and V-tail mounted to the tail rails by the tail block . . . . . . . . . 80
4.25 Roller model . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 86
4.26 Control surfaces subsystem electronic components . . . . . . . . . . . . . . 88
4.27 Fuselage structural layout . . . . . . . . . . . . . . . . . . . . . . . . . . . 92
4.28 Landing gear mounting layout . . . . . . . . . . . . . . . . . . . . . . . . . 93
4.29 Centre of gravity envelope . . . . . . . . . . . . . . . . . . . . . . . . . . . 93
4.30 Selected main landing gear (Pilot-RC Inc. 2009) . . . . . . . . . . . . . . . 94
4.31 Maximum fuselage shear stress . . . . . . . . . . . . . . . . . . . . . . . . . 98
4.32 Bending stress in the upper longeron . . . . . . . . . . . . . . . . . . . . . 99
4.33 Static margin envelope . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 102
5.1 Inboard wing assembly drawing . . . . . . . . . . . . . . . . . . . . . . . . 111
5.2 Outboard wing and block assembly drawing . . . . . . . . . . . . . . . . . 113
5.3 Fuselage assembly drawing . . . . . . . . . . . . . . . . . . . . . . . . . . . 114
5.4 Empennage assembly drawing . . . . . . . . . . . . . . . . . . . . . . . . . 116
5.5 Aircraft assembly drawing . . . . . . . . . . . . . . . . . . . . . . . . . . . 117
6.1 Static thrust set-up . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124
6.2 Static thrust curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 124
6.3 Second morphing test with 3.8G (6.95kg) loading . . . . . . . . . . . . . . 125
6.4 Wing structural test . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 127
6.5 Attempt 1 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 131
6.6 Attempt 1 - GPS output . . . . . . . . . . . . . . . . . . . . . . . . . . . . 131
6.7 Attempt 2 . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 132
6.8 Attempt 2 - GPS output . . . . . . . . . . . . . . . . . . . . . . . . . . . . 133
6.9 Airworthiness test flight images . . . . . . . . . . . . . . . . . . . . . . . . 134
6.10 Morphing test flight images . . . . . . . . . . . . . . . . . . . . . . . . . . 135
7.1 Management structure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 140
7.2 Labour distribution between tasks and members . . . . . . . . . . . . . . . 144
7.3 Usage of project funds . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 145
C.1 Fuselage and tail boom shear diagram . . . . . . . . . . . . . . . . . . . . . 159
C.2 Bending moment diagram for the fuselage and empennage boom . . . . . . 160
List of Tables
2.1 Mission Profile Segments . . . . . . . . . . . . . . . . . . . . . . . . . . . . 17
3.1 Weight Budget Breakdown . . . . . . . . . . . . . . . . . . . . . . . . . . . 33
3.2 Oswald’s efficiency factor for the retracted and extended configurations. . . 36
3.3 Sizing requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 38
3.4 Matching diagram conclusions . . . . . . . . . . . . . . . . . . . . . . . . . 41
3.5 Tail sizing results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44
3.6 Tail geometry . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 45
3.7 Energy requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 48
4.1 Airfoil Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 55
4.2 Inboard wing candidate airfoils . . . . . . . . . . . . . . . . . . . . . . . . 55
4.3 Outboard wing candidate airfoils . . . . . . . . . . . . . . . . . . . . . . . 56
4.4 Wing componenet materials . . . . . . . . . . . . . . . . . . . . . . . . . . 65
4.5 Deflection results for individual wing sections . . . . . . . . . . . . . . . . 73
4.6 Candidate empennage materials . . . . . . . . . . . . . . . . . . . . . . . . 78
4.7 Maximum loads on the tail boom . . . . . . . . . . . . . . . . . . . . . . . 80
4.8 Requirements for logic circuitry . . . . . . . . . . . . . . . . . . . . . . . . 90
4.9 Candidate materials for the fuselage structure . . . . . . . . . . . . . . . . 91
4.10 Aircraft weight breakdown summary . . . . . . . . . . . . . . . . . . . . . 93
4.11 Main landing gear requirements and specifications of the selected gear . . . 94
4.12 Neutral axes and Moment of intertia for various fuselage sections . . . . . . 97
4.13 Torsional shear stress at former locations with a safety factor of 2.25 . . . . 100
4.14 Morpheus UAV longitidinal stability . . . . . . . . . . . . . . . . . . . . . 102
4.15 Morpheus UAV Performance . . . . . . . . . . . . . . . . . . . . . . . . . . 103
6.1 Piecewise wing load distribution up to 3G total load . . . . . . . . . . . . . 126
6.2 Wing deflection under load . . . . . . . . . . . . . . . . . . . . . . . . . . . 127
6.3 Static margin for each configuration obtained during the morphing test . . 128
6.4 Ground handling - Control surfaces . . . . . . . . . . . . . . . . . . . . . . 129
8.1 Morpheus UAV Performance . . . . . . . . . . . . . . . . . . . . . . . . . . 148
8.2 Optimal configurations for a reconaissance mission . . . . . . . . . . . . . . 148
B.1 Parameters used for the tip-back angle calculation . . . . . . . . . . . . . . 156
xviii
B.2 Landing gear positioning criteria . . . . . . . . . . . . . . . . . . . . . . . . 156
C.1 Aircraft weight breakdown summary . . . . . . . . . . . . . . . . . . . . . 158
D.1 Morpheus UAV parameters . . . . . . . . . . . . . . . . . . . . . . . . . . . 161
D.2 Morpheus UAV Performance . . . . . . . . . . . . . . . . . . . . . . . . . . 164
M.1 Labour contributions by each group member . . . . . . . . . . . . . . . . . 306
Glossary
Nomenclature
Acronyms and initialisms
AR Aspect Ratio
AC Aerodynamic Centre
CAO Civil Aviation Orders
CASA Civil Aviation Safety Authority
CASR Civil Aviation Safety Regulations
CFD Computational Fluid Dynamics
CG Centre of Gravity
CGR Climb Gradient Ratio
DARPA Defense Advanced Research Projects Agency
ESC Electronic Speed Controller
FEA Finite Element Analysis
GPS Global Positioning System
LiPO Lithium Polymer
MAAA Model Aeronautical Association of Australia
MAC Mean Aerodynamic Chord
MAV Micro Aerial Vehicle
NASA National Aeronautics and Space Administration
NiCd Nickel Cadmium
RC Radio Control
RF Radio Frequency
RPM Revolutions Per Minute
SM Static Margin
UAV Unmanned Aerial Vehicle
Symbols
A Aspect ratio, cross-sectional area
b Wing span
C Chord, cruise distance
c Mean aerodynamic chord
ct Tip chord
c0 Root chord
CD Drag coefficient
CD0 Drag coefficient of airfoil
Cfe Skin friction drag coefficient
cHT Horizontal tail volume ratio
CL Lift coefficient of wing
Cl Lift coefficient of airfoil
CLα Lift curve slope
CLTO Takeoff lift coefficient
CLmax Maximum lift coefficient
CLmaxTO Maximum takeoff lift coefficient
CM Wing moment coefficient (quarter chord)
Cm Airfoil section moment coefficient (quarter chord)
CM0 Moment coefficient of airfoil
cV T Vertical tail volume ratio
D Drag, Diameter
d Diameter
DC Direct current
E Endurance, modulus of elasticity
e Oswald efficiency factor
g Acceleration due to gravity
h Height
I Moment of inertia
i Installed angle of incidence
J Energy consumed
K Drag-due-to-lift factor, gust alleviation factor
kts Knots
Kv Motor constant
L Lift, length
La Spanwise lift coefficient
Lb Spanwise lift coefficient
LH , LV Distance between the wing and tail quarter chord points
L/D Lift-drag ratio
M Bending moment
n Load factor
nm Nautical mile
P Power
p Pitch
Pcr Critical bucking load
Q First moment of area
q Dynamic pressure
r Radius
Re Reynolds number
RPM Revolutions per minute
S Area, planform area, surface area
SG Ground roll distance
SH Horizontal stabiliser planform area
Sref Reference area
STOG Takeoff distance
SV Vertical stabiliser planform area
Swet Wetted area
SM Static margin
T Torque, thrust
t Thickness
TOP23 FAR23 takeoff parameter
U Gust velocity
V Voltage, volume, velocity, shear load
Vclimb Climb velocity
VNE Never exceed velocity
W Weight
WTO Takeoff weight
WP
Power loading
W/S Wing loading
xac Aerodynamic chord position
xcg Centre of gravity position
y Distance from the neutral axis
zh Distance between horizontal and tail planes
Greek Symbols
α Angle of attack
αinstalled Installed angle of wing relative to longitudinal axis of aircraft
αstall Stall angle of attack of the main wing
γ Climb angle, dihedral angle
∆αOL Change in zero-lift angle of atatck of main wing dur to flap deflection
∆D0gear Change in zero-lift drag coefficient due to landing gear
∂ε∂α
Downwash derivative
ε Twist angle of wing
η Efficiency
ηp Propeller efficiency
ΛLE Sweep angle
λ Taper ratio
µ Mass ratio
ρ Free stream air density
σ Stress, density ratio
τ Shear stress
ω Motor speed
Subscripts
A Aircraft
airfoil Airfoil
air speed Air speed
aileron Aileron
cl Climb
climb Climb
cr Cruise
cruise Cruise
des Descent
de Design
e,E Empty
extended Extended
f Final
h, hori Horizontal
i Initial
induced Induced
loiter Loiter
max Maximum
motor Motor
O Zero angle of attack
OL Zero lift
p, prop Propeller
payload Payload
retracted Retracted
root Root
stall Stall
static Static
tip Tip
TO Takeoff
w Wing
wet Wetted area
wind Wind gust
wing Wing section
v,vert Vertical
x X axis (with respect to)
y Y axis (with respect to]
z Z axis (with respect to)
Coordinate frame
The coordinate frame used throughout this report is shown in the Figure 1 below.
Figure 1: Coordinate System Designation
1. IntroductionA morphing Unmanned Aerial Vehicle (UAV) is a high-performance aircraft that can
operate efficiently in multiple flight regimes by changing its external shape. Morphing
is generally achieved using either smart materials (materials which have one or more
properties that can be significantly changed, in a controlled manner, by external stimuli),
or structural morphing. Morphing can encompass many aspects of the aircraft design,
including the location, shape, area and angle of the wings, tail or fuselage. The Morpheus
UAV was designed to morph between at least two configurations during flight using a
combination of wing and tail structural morphing mechanisms.
1.1 Motivation
Unmanned Aerial Vehicle (UAV) technology is currently one of the fastest growing sectors
of the international aerospace industry. Rapid technological advances in both materials
science and electronics have recently accelerated the development of UAV design (Sarris
2001). UAVs can be utilised for a diverse range of applications in both the civilian and
military sectors. Such applications include surveillance, reconnaissance, search and rescue,
bushfire monitoring, mapping, surveying, remote sensing, transport, scientific research
and precision attacks. Using UAVs in place of human-occupied vehicles eliminates the
danger to human life by permitting hazardous tasks to be undertaken with reduced risk.
Aircraft design generally involves compromise between different requirements. A morph-
ing aircraft can overcome the need for compromise, allowing multiple and often contradic-
tory aircraft configurations to be incorporated into a single platform. This allows a much
wider range of mission tasks to be performed efficiently by the aircraft. Currently NASA,
DARPA, Lockheed Martin, Boeing and many other aerospace and defence companies are
studying morphing technology with the aim of exploring alternate UAV designs which
are more versatile, efficient and reduce the need for performance compromise in aircraft
design.
A particular application in which morphing UAVs are of interest, could include reconnais-
sance missions, where combinations of long, low speed endurance and high cruise/dash
speeds are desirable. On a normal aircraft, a compromise between speed and low speed
endurance is required, however a morphing aircraft removes this need for compromise. An
example of a military application for such a mission would be to gain extended intelligence
via surveillance (requiring a long, potentially slow speed loiter), where the aircraft is re-
1
1.2. AIMS AND OBJECTIVES 2
quired to arrive at the desired location as soon as possible (in which case high cruise/dash
speed capabilities are required). A civilian application could be as an emergency services
response aircraft, where real time data is required over extended periods, requiring long
loiter capabilities. In emergency services activities such as search and rescue, or fire fight-
ing, it is often important to obtain this information as soon as possible, or for the target
area to change rapidly, requiring the aircraft to also change location rapidly. This would
require an aircraft that would also be capable of fast cruise/dash.
1.2 Aims and objectives
The aim of the Morpheus project was to design, build and test a remotely piloted UAV
with a morphing configuration, as a test bed for morphing technology. The project focus
was on varying the UAV wing span to change the aerodynamic properties of the aircraft,
and changing the location of the tail to control longitudinal stability during flight. The
UAV was to be designed and developed using existing techniques and readily available
materials and components. The project included the design of the airframe and the
morphing mechanisms to extend and retract the wings and tail, the manufacturing, and
testing of the aircraft to validate the success of the airframe and morphing mechanism
designs. For the purpose of this project, the UAV was to be flown using a standard radio
control system. The inclusion of an automatic control system for either the aircraft or
the morphing mechanisms deemed to be beyond the scope of this project.
The project objectives consists of both primary and extended goals. The primary goals
focus on demonstrating the aircrafts flight capabilities, including takeoff, landing, cruise,
and payload and morphing capabilities. The extended goals focused on aircraft perfor-
mance and the effects of morphing.
1.2.1 Primary project goals
1. The UAV shall have a normal takeoff and landing method.
This goal is achieved if the UAV can demonstrate a normal takeoff and landing
method. A normal takeoff and landing method is defined as the use of landing gear
on a runway rather than hand launching or landing without an undercarriage.
2. The UAV shall be capable of having a loiter time of at least 30 minutes.
This goal is achieved if the UAV can demonstrate a loiter time of at least 30 minutes.
For this goal to be achieved, it is not a requirement for the entire 30 minutes to be
spent in flight, but can be proven by testing the loiter time for a shorter period, and
determining the total loiter time from the remaining battery power. The 30 minutes
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3 CHAPTER 1. INTRODUCTION
loiter time need only occur within a simple mission profile involving takeoff, climb,
loiter, descent and landing.
3. The UAV shall be capable of cruising within line of sight.
This goal is achieved if the UAV can demonstrate a short cruise segment. The line of
sight restriction is due to CASA regulations and the inability to control the aircraft
via remote control if the UAV is out of sight.
4. The UAV shall be capable of carrying a 500g payload.
This goal is achieved if the UAV can demonstrate takeoff, 30 minute loiter and
landing with a 500g payload onboard. This ensures that the UAV is capable of
achieving a purposeful mission.
5. The UAV shall morph the wing to achieve a wing span increase of at least 50% of
the original wing span during flight.
This goal is achieved if the UAV can demonstrate the operation of the morphing
mechanism to achieve a 50% minimum increase in wing span during flight This
should be achieved without major loss of control of the aircraft.
6. The UAV shall change the tail position to control the longitudinal stability during
flight.
This goal is achieved if the UAV can demonstrate the operation of the tail morphing
mechanism during flight without major loss of control of the aircraft. It should also
be theoretically shown that this operation affects the static margin of the UAV.
1.2.2 Extended project goals
1. Measure the performance of the aircraft in different configurations during flight.
This goal is achieved if at least 4 performance parameters are measured in at least
2 UAV configurations. Performance parameters may include, but are not restricted
to, takeoff distance, cruise speed, endurance, landing distance, dash speed, range or
turn rate.
2. Theoretically optimise the morphing parameters for a predetermined mission.
This goal is achieved if optimal wing spans and tail positions are calculated for a
predetermined mission.
3. Achieve roll control through differential span morphing.
This goal is achieved if one circuit of flight is completed using only differential span
morphing to control the roll angle.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
1.3. SCOPE 4
1.3 Scope
The Morpheus project includes the design, manufacture, and basic testing of the UAV.
The design includes both the initial investigation into possible morphing methods, con-
cept design, and detailed design. The design of the aircraft will only consider longitudinal
stability and will not involve extensive use of computational fluid dynamics. The man-
ufacturing of the aircraft was limited by the project budget, as well as the capabilities
of the students and the engineering workshop. Owing to time restrictions, and the cost
and expertise of acquiring and utilising test equipment, only basic testing of the aircraft
was included in the scope of the project. The use of smart materials, and any form of
automation were also deemed to be beyond the scope of the project due to the lack of
available time and knowledge in these specific areas. The scope is further refined in the
technical task (Section 2.4).
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
2. Literature review and feasibility
studyIn determining the feasibility of designing, building and testing a morphing UAV, an
extensive investigation was conducted. Existing morphing aircraft were investigated to
provide the group with an understanding of the level of complexity of these aircraft, and
for the selection of prototypes to be used for benchmarking purposes. An investigation
into the effects of varying aircraft parameters was also conducted as a means to determine
the effectiveness and feasibility of different morphing methods. Small scale UAVs were also
investigated to perform a statistical analysis to determine the performance parameters to
which the aircraft would be sized. It was necessary to investigate these non-morphing
UAVs, to form a base design from which to work. This was necessary as morphing
UAVs vary significantly in their methods of morphing, making comparison difficult, and
information on the performance of these UAVs is often limited, and does not provide the
information required for sizing.
2.1 Literature review
To determine the feasibility of this project, a review of available literature was conducted.
This literature review allowed for the determination of suitable prototypes for benchmark-
ing (Section 2.2), an analysis of possible aircraft parameters to be morphed (Section 2.3)
and a statistical analysis of existing aircraft to gain information pertaining to aircraft
sizing.
When investigating morphing aircraft, numerous websites, and research papers were con-
sulted, providing valuable information regarding types of morphing previously attempted
and achieved. Of particular note, are the website for the Virginia Tech Morphing Wing
Project (Tech 2004), and the thesis database of the University of Maryland (University of
Maryland 2008), which provided information about realistic student morphing projects.
The literature used to determine the effectiveness of changing specific aircraft parame-
ters included a range of texts pertaining to different aspects of aircraft design. These
include the books Aircraft Design: A Conceptual Approach (Raymer 2006) and Volumes
1-7 of Airplane Design (Roskam 1989). Although these books are aimed toward the de-
sign of large scale, non-morphing aircraft, the general information and basic equations
provided a means by which the effectiveness of changing certain aircraft parameters could
5
2.2. MARKET EVALUATION AND BENCHMARKING 6
be determined.
A number of databases containing information on small scale UAVs were used to ob-
tain statistical data pertaining to the dimensions, speeds and capabilities of small scale
aircraft. Two databases of particular note, were Jane’s unmanned Aerial Vehicles and
targets (Vehicles & Targets 2002), and 2007 UAV World Roundup (American Institute of
Aeronautics and Astronautics 2007), which provided an extensive list of UAVs of all sizes
for use in the statistical anlysis.
2.2 Market evaluation and benchmarking
A market evaluation of existing aircraft revealed several morphing UAVs which have
been successfully designed, built and tested. The market evaluation was conducted in
parallel with the statistical analysis (section 3.6.1), and provided a means against which
the Morpheus project aims and objectives could be compared and evaluated to ensure a
feasible, yet worthwhile project. Four morphing aircraft were selected as prototypes for
benchmarking, based on the following criteria:
• Type of morphing
• Physical size
• Weight
• Mission requirements and application
2.2.1 Virginia Tech BetaMax Morphing Wing Project
Virginia Tech has been extensively involved with the design and development of morphing
UAVs, especially those with telescopic wings. The most relevant of these, the BetaMax,
is shown in Figure 2.1. This aircraft successfully demonstrated the use of differential
telescoping for roll control, which is an extended goal for the Morpheus project. BetaMax
uses a glow plug engine for propulsion and a rack and pinion morphing mechanism to
extend and retract the wings. BetaMax was fully instrumented to provide information
about the aircraft’s flight characteristics. Recorded data included flight speed, roll and
pitch rates, accelerations and control surface deflections. This data demonstrated that
retracting the wings enabled the aircraft to be quick and manoeuvrable, while extending
the wings provides the aircraft with more lift, resulting in improved fuel economy at
slow speeds (Tech 2004). This aircraft was selected for benchmarking purposes as it is a
student project which focuses on morphing structures, and utilised differential roll control.
In these ways, it is similar in scope to the Morpheus Project.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
7 CHAPTER 2. LITERATURE REVIEW AND FEASIBILITY STUDY
Figure 2.1: Virginia Tech BetaMax Morphing Wing Project (Tech 2004)
2.2.2 Delft University of Technology Roboswift
The Roboswift, shown in Figure 2.2, is a Micro Aerial Vehicle (MAV) which morphs by
changing the area, sweep, slenderness and camber of it’s wings. It was designed and
built by a student team from the Delft Univeristy of Technology in the Netherlands, who
used nature mimicry to imitate the appearance and flight characteristics of a swift. This
aircraft enhances the performance envelope of the aircraft allowing for efficient flight at
both low and high speeds (Roboswift 2009). This aircraft was chosen for benchmarking
purposes as it is a student project capable of morphing several aircraft parameters.
Figure 2.2: Delft University of Technology Roboswift (Roboswift 2009)
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
2.2. MARKET EVALUATION AND BENCHMARKING 8
2.2.3 Lockheed Martin Skunk Works Morphing UAV Concept
The Lockheed Martin Skunk Works Morphing UAV morphs by folding its wings to achieve
a change in wing span, wing area, wing shape and wing location. The aircraft is shown
in Figure 2.3. The UAV is designed to perform both long endurance loiter surveillance
and high speed, short dash attack missions. Morphing is achieved using shape changing
actuation systems located within the wing skin of the UAV, which relax and contract
when energised by an electrical current. Morphing between the two configurations occurs
in 25 seconds and results in a 71% change in wing area. Onboard flight control systems
manages the vehicle dynamics as the aerodynamics and centre of gravity of the aircraft
changes during morphing (Vehicles & Targets 2002). Although the UAV utilises smart
materials, the geometry of the morphing was considered relevant for benchmarking.
Figure 2.3: Lockheed Martin Skunk Works Morphing UAV (Martin 2009)
2.2.4 NextGen Aeronautics MFX-2
The MFX-2 is a jet powered morphing UAV capable of independently varying wing area
and wing sweep. The aircraft is shown in Figure 2.4. It achieves a 40% change in
wing area, 73% change in wing span and 177% change in aspect ratio. The MFX-2
can switch between autonomous and radio control modes during flight, and features a
unique autopilot system which utilises a variable stability and control scheme. The MFX-
2 performed five successful flights of approximately 10 minutes duration, demonstrating
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
9 CHAPTER 2. LITERATURE REVIEW AND FEASIBILITY STUDY
autonomous morphing in approximately 10 seconds (Aeronautics 2007). This UAV was
selected for benchmarking purposes as it varies several of its external parameters..
Figure 2.4: NextGen Aeronautics MFX-2 (2009 2009)
2.3 Analysis of morphing methods
An investigation into possible morphing geometries was conducted to determine the use-
fulness, and applicability of changing different aircraft parameters. This reseach indicated
that the most effective use of morphing would be to design an aircraft which could change
it’s wing and tail parameters. The effects of varying different aspects of wing and tail ge-
ometry were then considered to determine the most effective and applicable parameter(s)
to vary.
2.3.1 Wing morphing methods
Wing geometric parameters considered for morphing include:
• Airfoil profile or chord
• Wing position
• Twist
• Dihedral angle
• Angle of incidence
• Sweep
• Area, aspect ratio and/or span
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
2.4. TECHNICAL TASK 10
Each of the possible wing morphing methods was analysed for usefulness, to determine
the project focus. The option to morph the airfoil profile or chord was immediately
eliminated as this form of morphing is commonly used on existing aircraft in the form of
flaps. Changing the wing position, twist, dihedral angle, or angle of incidence was also
eliminated as analysis determined that morphing the wing area, aspect ratio or sweep
would result in the UAV being able to satisfy a larger range of conflicting requirements
for different flight phases. Variable sweep is mainly of use for supersonic applications. As
the Morpheus aircraft is restricted to subsonic speeds due to the size of the aircraft and
availability of technology, this was deemed unsuitable unless used as a means to morph
the wing area. The most viable options for morphing the wings of the Morpheus aircraft
were therefore determined to be the area, aspect ratio and/or span of the aircraft
2.3.2 Tail morphing methods
The main aim of morphing the aircraft tail is to affect the longitudinal stability of the
aircraft. The two main ways of doing this include morphing the tail area or location.
These two options were compared to determine the most feasible method. The compar-
ison considered stability effects and the location of the morphing mechanisms required.
Varying the tail location could operate using a single mechanism housed in the fuselage,
possibly allowing the entire tail to be moved as one. Varying the tail area would require
multiple mechanisms (minimum of one per horizontal tail surface) housed in the tail itself.
Morphing the tail position of the aircraft was therefore considered to be the most effective
and feasible parameter due to simplicity.
2.4 Technical task
The technical task utilises the project aims and objectives, along with information dis-
covered during the initial market research and benchmarking, to provide a comprehensive
set of specifications to which the Morpheus aircraft should be designed. These specifica-
tions cover relevant standards, aircraft system requirements, performance requirements,
technical level and economic requirements.
2.4.1 Standard Requirements
The UAV design shall be compliant with the associated CASA 101, CASA UA25, MAAA
Manual of Procedures and CAO 95.21 regulations.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
11 CHAPTER 2. LITERATURE REVIEW AND FEASIBILITY STUDY
2.4.2 Special systems and miscellaneous
Overall UAV
1. The UAV and morphing system shall be designed, built and tested within one year.
Rationale 1: This is a given requirement of the project.
2. The UAV shall be able to fit within the back seat of a four-wheel drive. This may
involve dismantling and reassembling the aircraft.
Rationale 2: This allows easy transportation without a trailer to and from the
takeoff and landing sites.
3. The UAV shall be designed to morph during flight.
Rationale 3: This is part of the project definition.
4. The UAV shall be tested and flown in an approved area.
Rationale 4: More stringent and limiting regulations apply for UAVs flown outside
approved areas. CASA 101.240 defines an approved area as ’an area approved under
regulation 101.030 as an area for the operation of UAVs’ (CASA 2007).
5. The UAV shall be able to be flown by a single pilot with Gold Wing experience level
during testing. NB: a second control mechanism may be used for morphing.
Rationale 5: Pilots with higher qualfications and experience are more difficult and
expensive to hire for testing purposes. As the UAV will not have an automatic
control system, a second controller may be required for the morphing mechanism
control during testing.
Morphing system
6. The morphing system shall allow for morphing of the UAV between at least 2
configurations.
Rationale 6: This is taken from the project definition.
7. The morphing system should interface with the onboard electronics and controls.
Rationale 7: The morphing system is required to interface with the electronics and
controls so that the morphing mechanism can be integrated into the automatic
controller or controlled remotely.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
2.4. TECHNICAL TASK 12
Control authority system
8. The UAV should be designed to accommodate an automatic control system.
Rationale 8: The final/future UAV must be able to fly a specified mission out of
line of sight capabilities and with minimal ground input.
9. The UAV shall be controlled by means of a radio controller during the design and
testing phases.
Rationale 9: Cost and time constraints prevent an automatic control system from
being developed and integrated into the UAV for the design and testing phases.
10. The UAV shall morph under its own power in flight via a simple input signal from
the controller(s) i.e. pilot or automatic control system.
Rationale 10: This is required to make the aircraft easily controlled by the pilot.
2.4.3 Performance parameters
Weight
11. The UAV maximum takeoff weight shall be less than 7 kg.
Rationale 11: MAAA heavy model aircraft rules, guidelines and procedures require
that all model aircraft having a dry mass (including batteries if electric powered)
greater than 7kg and less than 25kg must be inspected by an MAAA Heavy Model
Inspector prior to the first flight. To avoid heavy model inspection, the UAV shall
have a maximum takeoff weight less than 7kg.
12. The UAV shall be capable of carrying a payload of 500g mass. The payload will be
completely independent of all aircraft systems.
Rationale 12: A payload capability will provide the UAV with a functional purpose,
and increase marketability. The payload must be independent from all aircraft sys-
tems to provide greater flexibility in payload type and requirements and to prevent
aircraft malfunction due to a malfunction of the payload or its associated power
system.
Takeoff and Landing
13. The UAV shall be designed for standard take off and landing on a prepared runway
(including firm grass strips).
Rationale 13: This allows a broad range of mission capabilities both in rural and
city areas.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
13 CHAPTER 2. LITERATURE REVIEW AND FEASIBILITY STUDY
Loiter/Endurance
14. The UAV shall be capable of achieving a minimum of 30 minutes loiter.
Rationale 14: This is one of the primary project goals which must be achieved.
Altitude
15. The UAV shall have an altitude of less than 400ft during testing.
Rationale 15: M.A.A.A. Manual of procedures state that a pilot can only fly a model
aircraft up to 400ft unless allowed under civil aviation requirements.
16. The UAV shall have a maximum altitude determined by radio controller operational
range and pilot visibility .
Rationale 16: For safety and operational purposes the UAV should be operated
within the pilot’s skill and visibility.
Operational radius
17. The UAV shall have a maximum operational radius during remote control testing
of line-of-sight or radio control, whichever is the smaller.
Rationale 17: Due to the time and overall budget allocations of this project it is
infeasible to design and install an onboard automated system on the UAV and thus
the only control will be via a pilot using a radio control system. CASA 101.385 also
requires that ’a model aircraft be operated only if the visibility at the time is good
enough for the person operating the model to be able to see it continuously’ (CASA
2007).
Operating conditions
18. The UAV shall be designed to operate at temperatures between 10 and 40 degrees
Celsius.
Rationale 18: For temperatures between 10 and 40 degrees no heat shielding or
specialised components are required. This allows readily available, off the shelf
components to be used.
19. The UAV shall be designed to fly from calm conditions up to a gentle breeze as
defined by the Beaufort wind scale [wind speeds up to 18.5 kph (10kts)]
Rationale 19: On the Beaufort wind scale a gentle breeze is defined as ‘Leaves and
twigs in constant motion, wind extends a light flag’ (Bureau of Meteorology 2009).
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
2.4. TECHNICAL TASK 14
For the UAV with no automatic controller, a gentle breeze at ground level was
deemed to be the maximum allowable conditions for safe flight.
2.4.4 Technical level of product
20. The UAV shall use only currently available technologies
Rationale 20: The UAV must be built within the year and thus all parts and
components must be able to be acquired in 2009 or be manufactured using existing
and available technology.
21. The UAV should use off-the-shelf products wherever possible.
Rationale 21: This ensures that parts are readily available should replacement be
required. This ensures that the cost of parts and time to acquire or manufacture
parts is kept to a minimum.
22. The UAV shall be able to be setup and operated (once fully operational) by a single
person with basic knowledge of the UAV.
Rationale 22: Once fully operational with installed automatic control the UAV
should be easily setup and operated by trained personnel.
2.4.5 Economical parameters
23. The UAV shall remain within the allocated budget.
Rationale 23: Limited funds are available for the project and these should not be
exceeded.
2.4.6 Power plant type and requirements
Whole system
24. The UAV shall have two isolated power sources one each for the payload and the
platform.
Rationale 24: This ensures that the UAV flight performance and operation is inde-
pendent of the payload.
25. The platform power system shall be able to provide continuous power to the engine,
morphing mechanisms and control system for the duration of the mission.
Rationale 25: This is required for a successful and safe flight.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
15 CHAPTER 2. LITERATURE REVIEW AND FEASIBILITY STUDY
26. The UAVs should include a mechanism which allows the batteries to be easily
charged or changed.
Rationale 26: The UAV will need to be tested and flown numerous times and thus
the batteries must be able to be recharged for each flight. The batteries also need
to be easily replaced in the event of damage or end of life.
2.4.7 Main system parameter requirements
Structure
27. The UAV structure shall be designed to withstand a manoeuvring load factor of at
least a 3.8G loading.
Rationale 27: This reduces the probability of major damage occurring during an
emergency landing and is a minimum requirement of CASA regulations (CASA
2000).
28. The metallic and wooden structures of the UAV must have a safety factor of 1.5.
Rationale 28: This reduces the possibility of failure and is a minimum CASA re-
quirement (CASA 2000).
29. Fibre reinforced primary composite structures must have a safety factor of at least
2.25.
Rationale 29: This reduces the possibility of failure and is a minimum CASA re-
quirement (CASA 2000).
30. Non-critical, non-structural components should be designed to be sufficient for the
required task.
Rationale 30: Non-critical, non-structural components are not essential to the air-
craft operation. A safety factor is therefore not necessary.
Landing gear
31. The UAV landing gear shall be designed to have an impact vertical load factor of
1.33 in the event of an emergency landing.
Rationale 31: This prevents major damage occurring in the majority of emergency
landings and is a CASA requirement (CASA 2000)
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
2.5. MISSION PROFILE 16
2.4.8 Reliability and maintenance
32. The UAV structure and external material shall withstand a reasonable amount of
wear and tear due to normal/mission usage, transportation, setup and recovery.
Rationale 32: To reduce maintenance time and improve reliability the aircraft should
be able to withstand reasonable wear and tear before it requires repairs.
33. The UAV shall be designed for easy maintenance on the airfield.
Rationale 33: This may involve detaching different sections of the UAV to provide
easy access to all internal sections of the aircraft as required.
34. The UAV should be repairable within a fortnight from any minor damage in a cost
effective manner.
Rationale 34: This ensures that all minor repairs can be covered within budget
without any major project delays.
35. The UAV shall be easily repaired by field personnel using on-hand tools, excluding
major damage to primary structure and mechanisms
Rationale 35: Any minor damage sustained during testing and operation must be
easily repaired onsite to prevent delays to the project.
36. The UAV shall not require more than 1 hour of standard maintenance per 10 hours
of flight.
Rationale 36: This ensures that maintenance time does not impinge on testing
periods. This includes any joints and mechanisms that may require greasing prior
to flight. Assembly, repairs and initial calibration are not included in this time.
2.4.9 Unification level
No consideration is required in regard to using existing technology or designs from previous
projects. Consideration is to be given to the use of a fuselage plug from a previous Adelaide
University UAV project, as a means of reducing the cost, however this is not essential.
2.5 Mission profile
The mission profile for the Morpheus UAV, was designed to incorporate loiter, cruise, and
dash phases, to demonstrate the UAVs diverse capabilities. The mission profile selected
as a guide for this project consists of a taxi, takeoff, climb, cruise, loiter, dash, decent,
landing and final taxi phases. This mission profile will demonstrate the UAVs performance
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17 CHAPTER 2. LITERATURE REVIEW AND FEASIBILITY STUDY
in multiple flight phases and in each configuration. The mission profile is defined in Table
2.5 and Figure 2.5.
Table 2.1: Mission Profile SegmentsMission phase Phase requirements
TaxiTakeoffClimbCruise 2.5km at 80 km/h and 400 ftLoiter 30 minutes at 1.4Vstall and 400 ftDash 2.5 km at 120 km/h and 400 ft
DescentLanding
Taxi
Figure 2.5: Mission profile
2.6 Summary of design requirements and feasibility
The key design requirements follow. The aircraft:
• Must be capable of increasing it’s wing span by 50% and capable of moving its tail
to affect stability
• Must follow CASA design regulations
• Must be remote controlled
• Should be less than 7kg
• Should be capable of 30 min loiter, with a 500g payload
The project is considered feasible for a final year honours project group of five. Similar
small scale morphing aircraft projects have been successful at other universities and three
previous UAV projects have been completed at The University of Adelaide as honours
projects.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3. Conceptual designThe conceptual design process aimed to generate, select and develop the most feasible
concepts that could meet all the design requirements. This process was conducted using
a classical conceptual design approach involving multiple design iterations. Each iter-
ation led to further development of the concepts until design decisions could be made.
The following section outlines the conceptual design process. This includes the aircraft
configuration selection, initial platform concepts, wing morphing concepts, empennage
morphing concepts and morphing actuator concepts. This is followed by weight estima-
tion, aircraft sizing, empennage conceptual design, propulsion system selection, landing
gear configuration design and fuselage sizing.
3.1 Aircraft configuration design
In order to begin the conceptual design phase of the project, the aircraft configuration
had to be selected. Choosing an unconventional morphing aircraft configuration would
introduce unneccessary complexities to an already challenging project. To ensure that
the aircraft possessed inherent stability, and to simplify the design and manufacture of
the airframe, a conventional aircraft configuration was selected.
3.2 Initial aircraft concepts
The first stage of the concept design process involved generating ideas for possible plat-
forms that could morph to achieve a span increase. Many concepts were initially gener-
ated, four of which were considered for further development. The four morphing platforms
considered were the delta-wing concept, the lifting-body concept, the telescoping-wing
concept and the folding-wing concept.
3.2.1 Delta wing
One concept to be considered was a flying wing aircraft with a single boom tail that
could morph between a conventional configuration and a delta-wing configuration, shown
in Figure 3.1. The wing morphing was achieved through a series of sliding plates pivoting
about the point where the wing leading edge meets the fuselage. The tail was also required
to morph to maintain the stability of the aircraft and to provide the delta-wing shape.
18
19 CHAPTER 3. CONCEPTUAL DESIGN
This was achieved through a telescoping boom tail, allowing the tail moment arm to
change as required.
Figure 3.1: The sliding plates concept with wing and tail morphing mechanisms shown
There were many advantages to the sliding plates design, as the span, area and sweep
angle were changed considerably between the two extreme configurations. This allowed
morphing from a conventional configuration with high span, large area and low sweep
wing, to a delta-wing configuration with low span, reduced area and high sweep. However,
a smooth aerofoil shape could not be achieved due to the tail being directly behind the
wing.
3.2.2 Lifting body
The concept involves a lifting body aircraft with internally stored auxiliary wings, as
shown in Figure 3.2. The UAV morphs between two configurations by sweeping the
auxiliary wings from inside the fuselage to a neutral sweep position outside the fuselage.
A boom mounted tail moves longitudinally to vary the longitudinal stability of the UAV.
Figure 3.2: Sketch of the lifting body morphing concept
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.2. INITIAL AIRCRAFT CONCEPTS 20
The swept configuration is suited to high-speed flight phases. The reduced frontal and
wetted area decreases profile and skin friction drag. This configuration can achieve greater
roll manoeuvrability than the unswept configuration through the combination of a lower
roll moment of inertia and the use of both elevators and differential sweeping of the
auxiliary wings to achieve lateral control authority. The unswept configuration is best
suited for slow speed flight phases. The increased lifting area reduces the stall speed of the
UAV, thus enabling the UAV to takeoff and land in shorter distances, as well as achieving
greater loiter endurance. The pitch manoeuvrability of both configurations is dominated
by the tail position.
3.2.3 Telescopic wing
The telescoping concept involves outboard wing sections extending and retracting from
inboard wing sections. Figure 3.3 shows a sketch of a telescopic morphing aircraft.
Figure 3.3: Telescopic aircraft sketch
The retracted configuration is best suited to high-speed flight phases. The reduced frontal
area and wetted area reduces profile drag and skin friction drag. The extended config-
uration is best suited to slow speed flight phases. The increased area reduces the stall
speed of the UAV, thus enabling the UAV to take-off and land in shorter distances and
achieve greater loiter endurance. Both the longitudinal and lateral stability of the UAV
will be greatly affected by the extension and retraction of the wings. In the retracted
configuration, the UAV will have decreased lateral stability due to a reduced roll mo-
ment of inertia. The extension and retraction of the wings will alter the position of the
aerodynamic centre, which can be compensated for by extending or retracting the tail.
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21 CHAPTER 3. CONCEPTUAL DESIGN
3.2.4 Folding wing
The folding wing concept shown in Figure 3.4 involves a conventional configuration UAV
that morphs by folding its wings and translating a boom-mounted tail. The wing contains
a root pivot and a mid-span pivot, which allows the UAV to fold its wings either above or
below the fuselage and thereby vary the wing planform area. The longitudinal translation
of the boom-mounted tail alters the longitudinal and directional stability of the UAV.
Figure 3.4: Folding wing concept in folded configuration
The folded configuration is best suited to high-speed missions where it benefits from
reduced drag due to a lower frontal and wetted area. The unfolded configuration is best
suited to slow speed mission phases where the larger planform area and aspect ratio
generate increased lift. This would decrease the stall speed of the UAV and enable the
aircraft to takeoff and land in shorter distances and achieve greater loiter endurance.
Both the longitudinal and lateral stability of the UAV will be greatly affected by the
folding wing morphing. In the folded configuration, the UAV will have decreased lateral
stability, due to a reduced roll moment of inertia, and increased longitudinal stability
as the neutral point of the UAV moves aft. The opposite will occur in the unfolded
configuration.
3.2.5 Selected concept
Four feasible platform solutions have been developed, each with varying degrees of tech-
nical and manufacturing difficulties. The UAV platform solutions were assessed using
six selection criteria, including structure and mechanism feasibility, stability and flight
predictability, morphing effectiveness, liklihood of completion, manufacturability and aes-
thetics.
The telescoping concept was selected based on good performance against all selection
criteria. The telescoping wing concept will limit the concentrated loads experienced by
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3.3. WING MORPHING CONCEPTS 22
each of the other morphing platform solutions. However, the requirement of hollow wing
sections may present some difficulty. The morphing of this concept involves minimal
deviation from a conventional aircraft configuration, and hence, the stability of the concept
should be predictable using theoretical methods. This concept involves the alteration of
wing area and wetted area with minimal detrimental effects on aircraft performance. The
simplicity of the fuselage and wing profiles required, as well as the performance of the
concept against the other selection criteria, suggests that the likelihood of project success
is high.
3.3 Wing morphing concepts
The development of wing morphing concepts considered the use of internal or external
telescoping and the appropriate mechanism support structures.
3.3.1 Telescopic wing configuration
Three feasible planform shapes incorporating a telescoping wing section were devised.
Two primary distinctions exist between each of the concepts: the external or internal
telescoping of the wing section and the use of wing taper. No feasible design was found
which incorporated both external telescoping and taper. Tapering of both the inboard
and outboard wing sections was considered in combination with the internal telescoping
method. It was concluded that a tapered outboard section would result in a large chord
length discontinuity at the junction between the inboard and outboard sections, when the
outboard section is in an intermediate position. Consequently, only the tapering of the
inboard planform was considered. These initial choices resulted in three main planform
shape concepts: external telescoping wing sections with a rectangular planform, internal
telescoping wing sections with a rectagular planform, and tapered inboard wing sections
with internal telescoping rectangular wing sections.
Concept 1: External telescoping wing section with rectangular planform
This concept involves rectangular inboard and outboard wing sections as shown in Fig-
ure 3.5, allowing for uniform cross sections within each wing segment. The outboard
section must have a hollow cross section to allow the outboard section to slide over the
inboard section. This will reduce the wing structural weight in the outboard section, but
will also result in the outboard section having a greater chord than the inboard section.
Consequently, the taper ratio for the entire wing would be greater than one, resulting in
increased lift generation at the wingtip.
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23 CHAPTER 3. CONCEPTUAL DESIGN
Figure 3.5: External telescoping wing section with rectangular planform
Concept 2: Internal telescoping wing section with rectangular planform
This concept involves rectangular inboard and outboard wing sections shown in Figure
3.6, allowing for uniform cross sections within each wing segment. The inboard section
must have a hollow cross section for the majority, if not the entire, inboard span. This
arrangement allows the outboard section to retract within the inboard section and gives
the overall wing planform a taper ratio of less than one due to the reduction of chord
between the inboard and outboard sections required for structural supports. The hollow
cross section of the inboard wing will result in reduced structural integrity.
Figure 3.6: Internal telescoping section without taper
Concept 3: Tapered inboard planform with internal telescoping rectangular wing tip
This concept involves a tapered inboard section and a rectangular outboard wing section
as shown in Figure 3.7, requiring varying cross sections within the inboard wing segment.
The inboard section must have a hollow cross section for the majority, if not the entire,
inboard span. This arrangement allows the outboard section to retract within the inboard
section and gives an overall wing planform taper ratio of less than one. The hollow cross
section of the inboard wing will result in reduced structural integrity. However, the
increased root chord will improve the structural integrity of the wing. This wing will not
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3.3. WING MORPHING CONCEPTS 24
benefit from the usual structural benefit of reducing weight towards the wing tip due to
the structural reinforcement required for the telescoping outboard section.
Figure 3.7: Internal telescoping with taper
Configuration selection
The telescoping layouts were assessed using selection criteria such as aerodynamic perfor-
mance, structural integrity, effect of morphing, manufacturability and cost, and aesthetics.
The selection criteria allowed concept 3 to be identified and selected as the most favourable
of the three concepts analysed, due to its favourable aerodynamic performance, structural
integrity and aesthetics.
3.3.2 Wing mechanism selection
The wing mechanism conceptual design involved the development of the support structure
for the outboard wing which involved the use of guide rails and rollers.
Rails
The choice of a mechanism that extends and retracts the wings and tail requires the use
of a set of guide rails. Both square cross-section rails and circular cross-section rails were
investigated. Square cross section rails provided an increased likelihood of the rails seizing
under load if the rails were slightly misaligned. Additionally, it was found that square
cross-section material was more difficult to source, which would make the procurement
of the components more difficult. Hence, two circular cross section rails were chosen for
the design, as this configuration uses readily-available components and has the highest
probability of success. The twin rail design is shown in Figure 3.8.
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25 CHAPTER 3. CONCEPTUAL DESIGN
Figure 3.8: Twin rail concept
Rollers
It was determined that rollers would be required to guide the wings and tail during the
morphing process. Two roller configurations were investigated throughout the design of
the morphing mechanism. The first configuration involved two sets of rollers on each
wing, as shown in Figure (a) in 3.9.. The first set of rollers was positioned on the inboard
wing tip rib and the second set of rollers was positioned on a rib further inboard.
Although the second set of rollers would guide the outboard wing more accurately then
one set of rollers, the design posed several challenges due to the position of the rollers
within the wing. Firstly, their position increased the difficulty of installation, as there
would be no direct access to the rollers during the assembly of the aircraft. Secondly, if at
any stage the rollers required maintenance or repairs, a lack of direct access would make
this nearly impossible. It was also shown that a second set of rollers was not required
for the morphing mechanism to work successfully, and would have been a redundant
system adding unnecessary weight and complexity to the aircraft. Hence, the second
roller configuration, using only one set of rollers on each inboard wing tip rib, was chosen
for the final design for simplicity, ease of access and reduced weight. This design can be
viewed in Figure (b) in 3.9.
3.4 Empennage morphing concepts
Three empennage morphing concepts were considered to vary the tail position. These
involves a telescopic fuselage, a sliding tail or the boom mounted tail.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.4. EMPENNAGE MORPHING CONCEPTS 26
(a) Double set of rollers (b) Single set of rollers
Figure 3.9: Roller concepts
3.4.1 Telescopic fuselage
One possible method of moving the tail position is a telescopic fuselage, as shown in Figure
3.10. This method involves the fuselage extending and retracting to vary the longitudinal
position of the tail. This concept presents a number of challenges. A complex mechanism
with high tolerances would be required to ensure that each section of the telescoping
fuselage extends and retracts as designed. The mechanism would need to be powerful, as
it would have to move the large amount of weight associated with the telescoping fuselage
sections. Additionally, a telescopic fuselage would encroach on space within the rear of the
fuselage, which would decrease the amount of available space for electronics and payload.
Figure 3.10: Telescopic fuselage concept
3.4.2 Sliding tail
The sliding tail concept is shown in Figure 3.11. This method involves translating the tail
along the top of the fuselage to vary the longitudinal position of the tail with respect to the
wings. The mechanism for this concept would be simple to manufacture and implement,
and would provide adequate space for electronics and payload within the fuselage. This
concept, however, has excess wetter area when in the retracted tail configuration.
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27 CHAPTER 3. CONCEPTUAL DESIGN
Figure 3.11: Sliding tail concept
3.4.3 Boom-mounted tail
The third method of moving the tail position was the boom-mounted tail, as shown in
Figure 3.12. This method involves varying the tail position by extending and retracting
the boom. When the tail is retracted, most of the boom length is within the fuselage, and
when the tail is extended, most of the boom length is outside the fuselage. This method
minimises the weight of the empennage, which minimises the weight of the overall aircraft
and reduces the mechanism actuation power. The method also simplifies the mechanisms
and provides adequate space within the rear of the fuselage to mount electronics and
payload. Hence, the boom-mounted tail was selected as the most preferred option for
extending and retracting the tail.
Figure 3.12: Boom-mounted tail concept
3.5 Mechanism actuator concepts
The aircraft requires three morphing mechanisms: one for the port wing, one for the
starboard wing and one for the tail. Four different morphing mechanism concepts were
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.5. MECHANISM ACTUATOR CONCEPTS 28
developed. These were a rack and pinion concept, winch concept, pneumatic concept and
a threaded rod actuator concept.
The system requirements were that linear motion was required for the extension and
retraction of the wings and tail, and a fast response rate was required for differential roll
control.
Each concept was generated by considering the system requirements, researching the types
of mechanisms that could meet the system requirements, and sketching each mechanism
as it would appear in the aircraft.
3.5.1 Rack and pinion
The rack and pinion concept for the wing can be seen in Figure 3.13, and the rack and
pinion concept for the tail can be seen in Figure 3.14. A rack and pinion meets the
system requirements and requires low maintenance. However, the mechanism is heavy,
and procurement of the materials and components required to manufacture a custom
mechanism would be difficult.
Figure 3.13: Wing rack and pinion sketch
Figure 3.14: Tail rack and pinion sketch
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29 CHAPTER 3. CONCEPTUAL DESIGN
3.5.2 Winch
A winch is a mechanical device that is used to extend, retract or adjust the tension of
a rope, wire or cable. The winch concept for the wing can be seen in Figure (a), and
the winch concept for the tail can be seen in Figure (b) in 3.15 below. A winch is cheap
to manufacture, meets system requirements, utilises components and materials that are
readily available, is easy to maintain and is simple. However, a winch system is heavy,
as it requires a large rope, wire or cable running the full span of each wing and the full
length of the fuselage.
(a) Wing winch sketch
(b) Tail winch sketch
Figure 3.15: Winch mechanism
3.5.3 Pneumatics
Pneumatics involves the use of pressurized gas to create mechanical motion. The pneu-
matic concept for the wing can be seen in Figure 3.16, and the pneumatic concept for
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.5. MECHANISM ACTUATOR CONCEPTS 30
the tail can be seen in Figure 3.17. A pneumatic system meets the system requirements,
requires minimal maintenance and is reliable. However, a pneumatic system is expensive,
difficult to integrate, exceedingly heavy and complex to operate.
Figure 3.16: Wing pneumatic mechanism sketch
Figure 3.17: Tail pneumatic mechanism sketch
3.5.4 Threaded rod actuator
The threaded rod actuator concept for the wing can be seen in Figure 3.18, and the
threaded rod actuator mechanism for the tail can be seen in Figure 3.19.
A threaded rod actuator meets the system requirements, is easy to integrate, has a low
weight, is easy to maintain and has a high reliability. No major disadvantages were noted
in comparison to the other morphing mechanisms. This mechanism does not require a
locking mechanism to maintain its position.
3.5.5 Morphing mechanism actuator selection
Based on the selection criteria, it was determined that a threaded-rod actuator is the
most suitable morphing mechanism to use for the aircraft, as it is easy to integrate, has
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31 CHAPTER 3. CONCEPTUAL DESIGN
Figure 3.18: Wing threaded rod actuator sketch
Figure 3.19: Tail threaded rod actuator sketch
a low weight, is easy to maintain and has a high reliability.
3.6 Aircraft Sizing
The preliminary sizing was conducted to determine the power and wing area of the Mor-
pheus UAV required to meet the specifications outlined in the technical task. This was
conducted using a matching diagram, however reasonable performance and geometric pa-
rameters were first required for sizing purposes. To determine the design weight of the
aircraft, a technology diagram was used. Statistical data was then used to determine the
aircraft performance parameters, and an investigation into the wing design variables was
used to determine the remaining parameters.
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3.6. AIRCRAFT SIZING 32
3.6.1 Statistical Analysis
A statistical analysis of existing aircraft was carried out to determine appropriate cruise
and stall speeds, wingspan and takeoff distance for the initial sizing of the Morpheus
aircraft. These parameters were not identified in the technical task. A statistical analysis
relevant to a morphing UAV proved to be a challenge, as the different morphing capabili-
ties varied significantly between aircraft, making comparison difficult. In addition to this,
performance information pertaining to morphing UAV’s is generally not readily available.
A statical analysis of standard aircraft was therefore conducted to determine a base design
from which the morphing aircraft parameters could be determined.
Weight estimate
To determine the design takeoff weight of the Morpheus aircraft, a statistical approach was
used to generate a technology diagram, shown in figure 3.20. To develop this technology
diagram, 11 electric UAVs with takeoff weights in the range of 1.8 to 28.1kg were analysed.
Figure 3.20: Technology diagram for UAVs with a takeoff weight between 1.8kg and 28.1kg.
The project goals specify that the Morpheus UAV must be capable of carrying a payload
of at least 500g as outlined in the primary project goals in Section 1.2.1. Due to the
limited data available on morphing aircraft, the technology diagram was generated using
non-morphing UAVs. To accommodate for the additional weight of the morphing mech-
anisms, 0.5kg of extra payload was included to account for the morphing mechanisms.
The technology diagram indicates that for a payload of 1.0 kg, an empty weight of 5kg is
required. This results in a design takeoff weight of 6kg.
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33 CHAPTER 3. CONCEPTUAL DESIGN
The weight budget breakdown is defined in Table 3.1. This weight estimate allows for the
design weight to be exceeded by 1 kg before the UAV is classed as a heavy model aircraft
and certification of the aircraft is required.
Propulsion and other electronics 2 kgWings (including morphing mechanism) 2 kgFuselage and Tail (including morphing mechanism 1.5 kgPayload 0.5 kgTotal 6 kg
Table 3.1: Weight Budget Breakdown
Cruise and stall speeds
For the purpose of determining the design cruise and loiter speeds, the type of fuel and
morphing capabilities of the aircraft were deemed irrelevant for this analysis. The main
reason for this is that the Morpheus aircraft must have comparable performance to all
other, similarly sized aircraft. This analysis therefore included statistical data on both
electric and fuel aircraft. To account for the change in fuel weight, where applicable, the
average of the takeoff and empty weights was used. For the stall speed analysis, 4 UAVs
were used with weights varying between 8.6kg and 45.0kg. The cruise speed analysis
utilised 10 UAVs with weights ranging from 4.5 to 60 kg. The results of this analysis can
be seen in figure 3.6.1.
Figure 3.21: Stall and cruise speeds versus takeoff weight.
Analysis of this data indicates that the Morpheus UAV should be sized to a stall speed
of 55kph, and a cruise speed of 75kph. Consultation of experienced model aircraft pilots
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.6. AIRCRAFT SIZING 34
indicated that a lower stall speed of approximately 45 km/h would be more appropriate
for a 6 kg aircraft to minimise approach speed and reduce the required pilot skill for
landing. A cruise speed of 80 km/h was selected for the extended configuration to provide
performance slightly superior to the average.
The statistical analysis also indicated that the average maximum speed for similar sized
aircraft was 110 km/h. A retracted cruise speed in excess of this maximum speed of
120 km/h was selected to give the Morpheus UAV far superior speed performance in the
retracted configuration compared to similar sized aircraft.
Wing span
Ordinarily, the wing aspect ratio would be determined using statistical methods. As the
primary objective for the project is to increase the wing span by a minimum of 50%, and
since the geometry of the wing, is largely dictated by the morphing mechanism, the wing
span was considered instead. In addition to this, the unique shape of the wing, with a
tapered inner section, and rectangular outer section makes comparison with other UAV
aspect ratios difficult. An estimate of wing span was obtained using 5 electric UAVs. This
analysis is shown in 3.22
Figure 3.22: Wing span versus takeoff weight for electric UAVs
The wing span determined by the statistical analysis was found to be 1.6m. This value
was used in initial estimates of aspect ratio and to verify the feasibility of the sized wing
spans. The retracted wing span should be below this value, whilst the extended wing
span should exceed this value.
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35 CHAPTER 3. CONCEPTUAL DESIGN
Takeoff distance
A statistical analysis was performed on four UAVs with electric propulsion systems. The
results of this analysis can be seen in figure 3.23.
Figure 3.23: TO distance versus takeoff weight for four electric UAVs
For a design takeoff weight of 6 kg, a takeoff distance of 42.m was obtained. A 40m takeoff
distance will therefore be used for preliminary sizing purposes.
3.6.2 Preliminary design parameters
The selection of the aspect ratio, Oswald’s efficiency factor, taper ratio, twist, dihedral
angle, wing position and sweep are necessary to allow the preliminary sizing of the wings.
It is also necessary to determine an estimate for propeller efficiency and the zero lift drag
coefficient.
Aspect ratio, A.
Higher aspect ratio wings have greater aerodynamic efficiency as less of the wing is af-
fected by three dimensional airflow. This results in higher lift (through a higher wing lift
coefficient) and decreased induced drag. Lower aspect ratio wings have greater structural
integrity due to the reduced moment arm of the lift forces, resulting in decreased weight.
The selection of an internal telescoping wing section will reduce the structural integrity
and increase the weight of the inboard wing. Hence, a low inboard aspect ratio of 3 was
selected for both structural and weight advantages. The decreased aerodynamic efficiency
of this aspect ratio can be compensated for by taper and wing tip devices if necessary.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.6. AIRCRAFT SIZING 36
Preliminary calculations indicated that the 50% extension of a single stage telescopic wing
could approximately double the aspect ratio of the inboard wing. Hence an aspect ratio
of 6 was selected for the extended configuration.
Wing sweep, ΛLE
Sweep is used to reduce compressibility effects and wave drag during supersonic flight.
Wing sweep also delays the onset of transonic flight. In subsonic flight, however, wing
sweep decreases the aerodynamic efficiency of the aircraft. Hence, for the design speeds
of the UAV, an unswept wing would be required. For structural reasons, the wing taper
is to be distributed between the leading and tailing edges of the wing. This will result in
some indirect sweep angle.
Oswald’s efficiency factor, e.
The Oswald’s efficiency factor is a measure of the efficiency of a ’non-elliptical’ spanwise
lift distribution. The estimated values are included in Table 3.6.2 and were calculated
using Equation 3.1 which is valid for aircraft with leading edge wing sweep less than 30◦.
Configuration Aspect Ratio Oswald’s efficiency factorRetracted 3 0.9709Extended 6 0.8691
Table 3.2: Oswald’s efficiency factor for the retracted and extended configurations.
e = 1.78(1− 0.045A0.68)− 0.64 (3.1)
Taper ratio, λ
Tapered wings have greater aerodynamic and structural efficiency, however, excessive
taper will lead to wing tip stall. Raymer (2006) states that an elliptical wing is the most
aerodynamically efficient planform, but is not commonly used due to high cost. A wing
with taper ratio λ = 0.45 provides the same reduction in induced drag as an elliptical
wing to within 1% (Raymer 2006). Hence, a taper ratio of λ = 0.45 was selected for
aerodynamic and structural advantages.
Twist, ε
Negative wing twist is used to reduce the effective angle of attack of the wing tip and
thereby ensure root stall before tip stall. Negative twist also reduces structural weight
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
37 CHAPTER 3. CONCEPTUAL DESIGN
by reducing the lift at the wing tip, and hence, the bending moment on the wing. Twist
significantly increases the complexity and cost of the wing. Additionally, wing twist will
increase the difficulty in designing a feasible telescoping mechanism. Hence, an untwisted
wing was selected for both the inboard and outboard wing sections.
Dihedral angle
Positive dihedral angle increases the lateral stability of the UAV. Negative dihedral angle
(anhedral angle) decreases the lateral stability of the UAV. For simplicity, a dihedral angle
of 0◦ was selected.
Wing position
An aircraft may use a low, high or mid-wing configuration. Low wing aircraft have
decreased lateral stability and high wing aircraft have increased lateral stability. A neutral
wing position was unsuitable as it divided the volume for payload and systems within the
fuselage. A low wing configuration is preferable to a high wing configuration as it will
decrease the lateral stability of the UAV, and thereby increase the possibility of achieving
roll control through differential telescoping.
Propeller efficiency, ηp.
The propeller efficiency is a measure of the ratio of the power produced from the propeller
to the motor shaft power. Raymer (2006) suggests that the propeller efficiency of a fixed
pitch propeller during loiter is 0.7. As loiter forms the major portion of the mission profile
it seems reasonable to use ηp = 0.7 as an initial estimate.
Zero lift drag coefficient, CD0
The zero-lift drag coefficient was estimated from the wetted area ratio (SwetSref
)and the
equivalent skin friction drag coefficient Cfe. Raymer (2006) suggests that the equivalent
skin friction drag coefficient of a single engine light aircraft can be estimated to be Cfe =
0.0055. The wetted area ratio was estimated for the extended configuration from sketches
to be SwetSref
= 3.3355. Roskam (1985) suggests that landing gear can add between 0.015
and 0.025 to the zero-lift drag coefficient. As the UAV will have fixed landing gear this
effect was estimated to be 0.025. Hence the zero-lift drag coefficient may be estimated as
CD0 = 0.0433 using Equation 3.2.
CD0 =SwetSref
∗ Cfe + ∆CD0gear (3.2)
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.6. AIRCRAFT SIZING 38
These preliminary design parameters can now be used in the preliminary sizing of the
UAV.
3.6.3 Sizing criteria
The UAV was sized to stall speed, takeoff distance and climb requirements which are
critical to the achievement of project goals. The UAV was also sized to different cruise
speeds for the extended and retracted configurations in order to demonstrate the effec-
tiveness of morphing. The sizing requirements are given in Table 3.3. All requirements
were converted to imperial units, with the results converted back to SI.
Table 3.3: Sizing requirementsCriterion ValueTakeoff distance [m] 40Stall speed [km/h] 45Climb Gradient [%] 16.66Extended wing cruise speed [km/h] 80Retracted wing cruise speed [km/h] 120
Sizing to takeoff distance
The UAV was sized to FAR23 takeoff distance requirements, which included a 50 ft
obstable. It was also assumed that takeoff occurs at approximately 1.1Vstall. Equations
3.3 to 3.5 were used to determine the relationship between wing and power loading.
CL,TO =CL,maxTO
1.21(3.3)
STOG = 4.9TOP23 + 0.009TOP232 (3.4)
W
P=
TOP23σCL,TOW/S
(3.5)
Sizing to stall speed requirements
The relationship between wing loading and stall speed was described by equation 3.6.
W
S=
1
2ρVstall
2CL,max (3.6)
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39 CHAPTER 3. CONCEPTUAL DESIGN
Sizing to climb requirements
CASA requires a minimum climb gradient of 8.33% at 1.3Vstall. A safety factor of two was
applied to this requirement to give a climb gradient of CGR=16.66%. The aircraft was
sized to this requirement using the FAR23 method. The climb coefficient of lift was cal-
culated with equation 3.7 and the climb speed requirement. The drag coefficient for climb
was then calculated from equation 3.8 and the climb gradient parameter determined from
equation 3.9. Equation 3.10 described the limiting relationship between power loading
and wing loading.
CL,climb = 1.2
(1
1.3
)2
(3.7)
CD,climb = CD,0 +CL,climbπAe
(3.8)
CGRP =CGR +
(CL,climbCD,climb
)−1
CL,climb0.5 (3.9)
W
P=
18.97ηpσ0.5
CGRP (W/S)0.5(3.10)
Sizing to cruise requirements
Equation 3.11 was derived by equating thrust to drag for straight and level flight at 75%
power, which is the common cruise setting for propeller aircraft (Roskam 1989).
P
W=
ρcruiseVcruise3CD0
0.75× 2× 550ηp(W/S)+
2(W/S)
550ηpρcruiseVcruiseπAe(3.11)
3.6.4 Matching diagram
The matching diagram method is traditionally used to determine a design point for an air-
craft which must meet several requirements simultaneously. A morphing aircraft, however,
is able to change configuration in order to meet different requirements. The Morpheus
UAV, in altering its wing span and area, was designed to meet low speed and high speed
requirements in separate configurations. To overcome the problem of sizing for multiple
configurations, the use of a design line was implemented and all requirements were plotted
on the same matching diagram, seen in figure 3.24.
Area 1 and 2 are the met areas for low and high speeds requirements respectively. Area 3
is the met area for all requirements and would be used for a conventional aircraft design.
The design line method uses the following process. The length of the design line is
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.6. AIRCRAFT SIZING 40
Figure 3.24: Matching diagram
determined from the maximum achievable change in wing loading and was determined
using an iterative process to be 2.69kg/m2. The design line is horizontal as the aircraft
has a constant power loading. The design line was positioned such that its endpoints were
situated in areas 1 and 2. Area 3 could have been used: however, this would imply that
a morphing aircraft was not required. The endpoints of the design line gave the design
points for the extreme configurations of the aircraft. The design line method could also
be used to meet the requirements of more than two met areas provided that the design
line can be placed such that it passes through each met area.
The design line location in Figure 3.24 was selected as it required the minimum power
to meet all design requirements. This design line gives the sizing results in Table 3.4(a).
The extended configuration, with its lower wing loading, will have superior slow speed
performance whilst the retracted wing configuration, with its higher wing loading, will
perform better at high speeds.
The results of the aircraft geometry and the selected preliminary design parameters are
met by the wing geometry in Table 3.4(b). Numbers have been rounded up to simplify
the drawing and manufacturing process.
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41 CHAPTER 3. CONCEPTUAL DESIGN
Table 3.4: Matching diagram conclusions
(a) Aircraft sizing results
Extended wing loading [kg/m2] 10.89Retracted wing loading [kg/m2] 13.58Extended wing area [m2] 0.5510Retracted wing area [m2] 0.4417Power loading [kg/kW] 6.83Power[W] 878.5
(b) Required wing geometry
Root chord [mm] 530Inboard tip chord [mm] 240Outboard chord [mm] 160Retracted wing span [mm] 1150Extended wing span [mm] 1840Retracted wing area [m2] 0.44275Extended wing area [m2] 0.55315
Required wing geometry
3.6.5 Aileron sizing
Ailerons provide roll control for conventional aircraft. Whilst this project aims to achieve
roll control through differential telescoping, conventional ailerons are necessary as a re-
dundancy. Ailerons must be appropriatly sized and located to ensure sufficient roll control
authority. Eger (1983) suggests that the area of a single aileron should be approximately
7% of a single wing as described by Equation 3.12.
2SaileronS
= 0.07 (3.12)
The aileron must provide sufficient roll control for both the extended and retracted con-
figurations. Hence, the aileron area was sized for the extended wing area of 0.55315m2,
resulting in a required aileron area of 0.0194m2. Raymer (2006) suggests that ailerons
should be positioned between 50-90% of the span and are typically 15-25% of the chord.
Due to the volume required by the wing morphing mechanism, the chord length of the
ailerons was limited to the lower end of the range suggested by Raymer (2006). Conse-
quently, it was necessary to increase the span of the ailerons to obtain the required aileron
area. The aileron was positioned between 20% and 76% of the half span of the wing with
aileron chords of 15% and 17% respectively. This produced an area of 0.019513, which
equates to 7.05 % of the half wing area.
3.7 Empennage conceptual design
The final empennage configuration was chosen to be a single telescoping boom tail with
a V-tail configuration as seen in Figure 3.25. There were two sections to the empennage
selection; the first section was the configuration analysis and the second section was the
morphing mechanism analysis.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.7. EMPENNAGE CONCEPTUAL DESIGN 42
Figure 3.25: Final empennage configuration
The empennage configuration selection was divided into two sub-sections, the overall
tail/fuselage configuration and the specific tail configuration. This sub-sectional analysis
was performed to increase the ease of analysis and to reduce the number of possible
configurations that needed to be analysed.
3.7.1 Tail/fuselage configuration analysis
The first section was the analysis of the overall tail/fuselage configuration, which involved
comparing the advantages and disadvantages of fuselage mounted, boom tail and twin
boom tail configurations. These configurations were assessed against selection criteria
such as weight, aerodynamic efficiency, control and positioning of the mechanism and
aesthetics, cost, structural stability and ease of manufacture.
Analysis outcome
The single boom tail is the best choice for the UAV. This result was primarily due to
the low weight, low drag and high aerodynamic efficiency, low cost, high control and
positioning of the mechanism and good aesthetics of the boom tail in comparison to the
other configurations.
3.7.2 Empennage configuration analysis
The second analysis was used to analyse the empennage configuration, which involved
comparing the advantages and disadvantages of conventional, cruciform, T-tail, H-Tail,
V-tail, inverted V-tail (single and double boom) and Y-tail configurations. These con-
figurations were assessed against weight, mechanical control, aerodynamic efficiency, aes-
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
43 CHAPTER 3. CONCEPTUAL DESIGN
thetics, ease of manufacture and life/fatigue.
Analysis outcome
The final result of the analysis was that a V-tail configuration was the preferred choice for
the UAV. This result was primarily due to the low weight, low drag and high aerodynamic
efficiency, low cost and high ease of manufacture, and good aesthetics of a V-tail compared
with the other configurations.
3.7.3 Empennage sizing
The preliminary empennage sizing was carried out using horizontal and vertical volume
ratios and typical statistical values. Due to the need for control authority, and the V-tail
configuration selected, it was determined that the use of volume coefficients corresponding
to a jet fighter would be appropriate.
The horizontal and vertical tail volume ratios can be used to calculate the corresponding
areas using Equations 3.14 and 3.14. From initial sketches, the tail position for the
retracted configuration was determined and the required horizontal and vertical areas
calculated. This tail size will be used for both configurations, but the tail position will be
varied. As suggested by Raymer (2006), the area of a V-tail should be such that the tail
has equivalent surface area i.e. the horizontal and vertical areas must be added together
to determine the V-tail area. The dihedral angle of the surfaces should be approximately
equal to the angle calculated using Equation 3.15. The results of these calculations are
presented in Table 3.5 and will be used as initial estimates for tail arm and tail area.
Final values for tail arm will be determined from a stability analysis.
SV T =cV T bwSwLV T
(3.13)
SHT =cHTCwSwLHT
(3.14)
θ = tan−1(√SV T/SHT ) (3.15)
3.7.4 Ruddervator sizing
Ruddervators are used to control both the pitch and yaw of the aircraft by combining the
effects of an elevator and a rudder. The ruddervators were sized according to the process
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.7. EMPENNAGE CONCEPTUAL DESIGN 44
Table 3.5: Tail sizing resultsLV T , LHT 0.97mSV T 0.039m2
SHT 0.074m2
ST 0.11m2
Dihedral angle 35.2◦
outlined in Simons (2002). It is stated that the chord of the empennage control surfaces
should be between 20% and 30% of the stabiliser average chord. Hence, it was decided to
have a ruddervator chord 30% of the stabiliser chord to gain adaquete control authority.
The average stabiliser chord is 161mm, so the ruddervator chord was calculated to be
48mm.
Raymer (2006) suggests that the empennage control surfaces either extend to the tip of
the stabiliser or extend to approximately 90% of the stabiliser span. To simplify manu-
facturing, it was decided that the ruddervators would extend to the tip of the stabiliser.
Figure 3.26 shows the preliminary tail geometry.
Figure 3.26: Preliminary tail geometry
3.7.5 Tail geometry
Table 3.6 shows the tail geometry requirements, and the associated chosen values.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
45 CHAPTER 3. CONCEPTUAL DESIGN
Table 3.6: Tail geometryParameter Requirement (Raymer 2006) Value
Total area (m2) 0.11Length of one tail half (m) 0.35
V-tail angle (degrees) 110.00Span (m) 0.57
Root chord (m) 0.22Tip chord (m) 0.10
Thickness-to-chord ratio Similar to wing 0.16Taper ratio Horizontal: 0.3-0.6, Vertical: 0.3-0.6 0.45
Leading edge sweep angle (degrees) 5 degrees greater than wing 19.14Trailing edge sweep angle (degrees) Straight hinge line 0.00
Aspect ratio Horizontal: 3-6, Vertical: 1.3-2 2.95
3.8 Propulsion system design
The propulsion system selection considerd the type of propulsion to use and then selected
the appropriate components.
3.8.1 Propulsion type selection
Propulsion for a UAV can be provided by several different types of engines. The statistical
analysis showed that two of the most common types of propulsion for a UAV or model
aircraft are an internal combustion engine or an electric motor. The statistical analysis
also showed that both internal combustion engines and electric motors have been success-
fully used to power UAVs up to and over 10kg. Hence, internal combustion engines were
compared against electric propulsion systems.
A suitable propulsion system was selected based upon selection criteria such as system
requirements, power-to-weight ratio, cost, reliability and complexity.
The power-to-weight ratio of the propulsion system was considered to be one of the most
important criteria as the aircraft requires the highest power from the lowest propulsion
weight. Cost, integration into the airframe, reliability, complexity and availability were
all considered to be of intermediate importance.
Internal combustion engines
Glow plug engines are unreliable, require high maintenance, have high operational costs
and also produce environmental emissions and noise.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
3.8. PROPULSION SYSTEM DESIGN 46
Electric motors
Electric motors offer several advantages in comparison to internal combustion engines,
including reduced noise, greater power consistency, greater reliability, ease of maintenance,
reduced vibrations, greater versatility and environmentally friendliness.
Electric motors can be used to drive a propeller (for a conventional aircraft) or an impeller
(for a ducted fan aircraft). From extensive research and manufacturer’s data, electric
ducted fans appear to offer additional performance, but are more expensive.
Propulsion system solution selection
Based on the selection criteria, it was determined that an electric motor driving a propeller
is the most suitable propulsion system to use for the aircraft, as this propuslion system
provides a high power-to-weight ratio.
Propulsion system location
The placement of the engine on the aircraft can affect the stability, performance, and
efficiency of the aircraft. The tractor configuration and the pusher configuration were
both considered.
A pusher configuration reduces drag and increases the efficiency of the wing due to the
absence of prop wash over the wing. A tractor configuration allows greater ground clear-
ance, improved cooling, reduced vibrations, and can be readily designed from existing
resources.
Despite the drag and efficiency benefits, a pusher configuration is not preferred due to
stability concerns and the issues caused by the limitation of the take-off rotation an-
gle. For these reasons it was recommended that a tractor configuration be used for the
aircraft. Secondary reasons for choosing a tractor configuration include the improved
crashworthiness, decreased vibration, and the engine cooling advantages.
3.8.2 Electric motor selection
An electric motor driving a propeller was selected for the aircraft propulsion system.
From propulsion system calculations, it was determined that the aircraft would require a
1050W motor to adequately perform in all flight phases. Hence, electric motors with a
power greater than 1050W were sourced to determine the most appropriate motor which
would meet the requirements and specifications.
After thorough research, three electric motors were sourced and compared using a decision
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
47 CHAPTER 3. CONCEPTUAL DESIGN
matrix. The suitability of the propulsion system was determined based upon cost, weight,
power, and volume. Based on a decision matrix, it was determined that the Dualsky
1650W motor is the most suitable electric motor to use for the aircraft. The Dualsky
1650W motor has the lowest cost, lowest weight, highest power-to-weight ratio and lowest
volume. Appendix A.1 shows the data for the selected motor.
3.8.3 ESC selection
An ESC (electronic speed controller) is a device which varies the motor speed or direction
depending on user inputs from a transmitter. Using an ESC which is recommended by
the manufacturer is preferred. The manufacturer of the electric motor recommended the
Dualsky 90 Amp ESC for use with the Dualsky 1650W motor. Hence, the recommended
ESC was chosen for the aircraft. Appendix A.1 shows the data for the selected ESC.
3.8.4 Battery selection
Using batteries which are recommended by the supplier is preferred. The electric motor
supplier recommended the use of two Flight Power Evolution 5000 mAh 5S 18.5V Lithium
Polymer (Li-Po) batteries for use with the Dualsky 1650W motor. Two Flight Power Evo-
Lite 5350 mAh 5S 18.5V Li-Po batteries were sourced and borrowed from the University
of Adelaide and these batteries are be placed in series to provide the required voltage.
Appendix A.1 shows the data for the selected batteries.
3.8.5 Energy requirements
For each phase of the mission profile (see Section 2.5 ) the energy was calculated. The
energy requirements of the cruise and loiter phases were calculated using the equations
derived from Roskam (1989). The loiter equation is included in Equation 3.16. The data
for these equations were obtained from the technical task (see Section 2.4).
Jloiter =E
η
[1
2ρV 3SCD0 +
2W 2TO
ρV SπARe
](3.16)
Where:
• Jloiter is the energy consumed in the segment [J]
• E is the time the aircraft loiters [secs]
Where the energy requirement data could not be obtained from direct calculations, the
data was constructed by multiplying the estimated time that the aircraft spends in each
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3.8. PROPULSION SYSTEM DESIGN 48
mission section by the throttle that the pilot was typically providing for the aircraft.
These data values are included in Table 3.7.
Table 3.7: Energy requirementsMission Phase Time Throttle Setting Energy Required
Taxi 36 secs 50% 15812 JTakeoff 11 secs 100% 10278 JClimb 30 secs 100% 26354 JCruise 76 secs 75% 55685 JLoiter 1800 secs 65% 435522 J
Descent 30 secs 25% 6588.6 JLanding 20 secs 15% 2635 J
Taxi 36 secs 50% 15813 JTotal Energy Required: 625558 J
The number of batteries required for this mission profile is calculated in Equation 3.17.
In this equation the batteries are placed in series.
mAh required =Energy required
Volts
1000
3600
=625558
18.5 + 18.5
1000
3600(3.17)
= 0.939 banks of two batteries
Thus, two batteries in series will provide sufficient power for this mission profile.
3.8.6 Propeller solutions
This section involves the sizing and selection of appropriate propellers for the design
conditions. This will involve the sizing of propeller diameter and pitch, as well as the
sourcing of a propeller and spinner.
Propeller sizing
The sizing of the propeller is based on three factors: number of blades, propeller diameter,
and propeller pitch. The selection of these values is outlined below. Considerations such as
the shape of the propeller blade and twist were not considered as an off-the-shelf propeller
was purchased. As a result, it is not possible to dictate the propeller shape.
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49 CHAPTER 3. CONCEPTUAL DESIGN
3.8.7 Propeller selection
The selection of the propeller for the aircraft was based upon the following constraints:
• The propeller, when coupled with the motor, must produce sufficient thrust for the
aircraft to perform adequately in all flight phases.
• The propeller tips must not reach sonic flow.
• The propeller must be lightweight and readily available.
Diameter selection
To determine the diameter of the propeller, there are several considerations. In general,
larger propellers are more efficient, and it is important to determine that the propeller
tips remain in sonic flow. Additionally, the practical diameter, the motor manufacturer’s
recommendations and the aircraft speed must also be considered.
The recommended diameter can be determined using Equation 3.18 from Simons (2002).
d = 24500× 4
√motor power [kW]
(n[RPM ])
2
× Vairspeed × 24.8 (3.18)
The actual values for the recommended propeller diameter were calculated to be 15.6”,
17.2” and 19.9” for cruise in the retracted configuration (120kph), cruise in the extended
configuration (80kph), and in the loiter configuration (45kph) respectively.
The propeller selected for use on the aircraft is to have a propeller diameter of 16 inches.
This diameter was selected as it meets all the requirements and represents a compromise
between the recommended propeller size for cruise and loiter.
Pitch selection
The propeller pitch is dependent on the engine speed, aircraft speed, mission, availability
and the pitch recommended by the motor manufacturer.
The selection of pitch for our propeller was based on Equation 3.19 (Simons 2002).
Pitch =V × 1000
n× 0.6× 0.393700787 [inches] (3.19)
The pitch values recommended by the motor manufacturer are greater than the values
calculated at any of the speeds our UAV shall achieve during flight. Hence, a 10” pitch
propeller was selected, as this is the lowest recommended pitch by the motor manufacturer,
and hence the closest to the pitch recommended by the equations.
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3.9. LANDING GEAR CONFIGURATION 50
Recommended Pitch vs. Velocity
0
2
4
6
8
10
12
14
0 20 40 60 80 100 120 140 160 180 200
Velocity (kph)
Pit
ch (
inch
es)
Range of values recommended by the motor
manufacturer.Range of values for the velocity of
the aircraft.
Figure 3.27: Recommended propeller pitch
Propeller selection
For testing purposes, we required two propellers of the same brand and diameter to provide
an appropraite comparison. An investigation of available propellers revealed that within
the diameter range of 16 to 18 inches, it was only possible to acquire a 16 inch propeller
which met the requirements. The Graupner 16 inch propeller was available with a pitch
of 8, 10 or 12. For testing purposes, the Graupner 16 inch × 8 inch and the 16 inch × 12
inch were selected.
3.9 Landing gear configuration
For a conventional takeoff and landing as required for our aircraft (defined in the section
2.4), there are six main types of landing gear configurations. The UAV is a small scale
aircraft, and as such, the multiple-bogey, quadricycle, single main and bicycle configu-
rations are inappropriate due to the excessive number of wheels which contribute to the
weight and drag of the aircraft. The tail-dragger configuration can be eliminated as it
interferes with the morphing of the tail, or a long tail wheel would be needed to prevent
the tail boom from striking the ground during landing. The alternative is to mount the
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
51 CHAPTER 3. CONCEPTUAL DESIGN
tail wheel on the tail boom. This is undesirable as it would require the tail boom to
withstand a considerable load during landing. The most appropriate configuration for
use on the aircraft is therefore the tricycle configuration.
3.10 Fuselage sizing
A minimum fuselage diameter of 0.155 m was required to fit several internal components
and provide sufficient room for handling of components. Fuselage friction drag is min-
imised with an overall fineness ratio of between 6 - 9 (Roskam 1989). It is desirable to
have the minimum fuselage length to reduce weight and to enable the Morpheus UAV to
change between a small and large length configuration. For a fineness ratio of 6 a fuselage
length of 0.93 m was required.
The tapered nose section requires a ratio between 1.2 - 2 according to Roskam (1989).
This gives a required taper length of 0.31m, which is approximately a third of the over-
all fuselage length. The length of the tapered fairing was limited by other geometrical
constraints, in particular the root chord length of the wings and landing gear mounting
requirements. A maximum tapered length of 0.13m for the aft end of the fuselage was
selected.
The tail positions determined in section 3.7.3 results in the tail retracting partway into the
fuselage fairing. The fuselage diameter was increased to 0.16m in order to meet landing
gear mounting requirements. This change in diameter, however, would have minimal
effect on the fuselage drag and hence the fuselage length was not resized due to time
constraints.
3.11 Conceptual design summary
The conceptual design of the Morpheus UAV resulted in a 6 kg aircraft capable of achieving
a 60% increase in wing span and a tail translation of 400 mm. Morphing mechanisms
and actuators for the wings and tail were evaluated and the most appropriate solutions
selected. The wings and propulsion system were sized to meet project, technical task
and statistical analysis requirements. The empennage and fuselage were also sized and a
landing gear configuration selected. Three-view drawings of the Morpheus CAD model
in the extended and retracted configurations along with major design parameters are
included below.
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3.11. CONCEPTUAL DESIGN SUMMARY 52
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53 CHAPTER 3. CONCEPTUAL DESIGN
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4. Preliminary and Detailed DesignThe preliminary and detailed design of the Morpheus UAV considers the aerodynamic
and structural design of the wings, fuselage and empennage. The design of the aircraft
and morphing mechanism control systems was also considered and integrated into the
overall aircraft design. An analysis of the flight behaviour of the aircraft was conducted
to determine the stability of the aircraft and the effects of the wing and tail morphing.
4.1 Wing design
The preliminary and detailed design of the wings was required to meet several operational,
structural and aerodynamic specifications. Aerodynamic specifications relate to the air-
craft achieving stable and efficient flight. Operational specifications involve transport and
maintanence requirements. Structural specifications relate to relevant design standards
and flight loads. The following specifications were addressed and met during the design
process.
• Select suitable airfoils for the inboard and outboards wing to provide CL,max = 1.2
• Determine the required installed incidence angle, iw, to achieve horizontal attitude
during cruise and minimise drag
• Support an 8 kg aircraft (aircraft weighed after first crash repairs and analysis
repeated) during flight with appropriate load and safety factors.
• Withstand normal handling conditions and light impact
• Transfer all loads into the fuselage
• Have removeable wings for transport
• Incorporate the wing morphing mechanism
• Incorporate control surfaces and control surface actuators
• Allow internal access for maintanence
• Be as simple as possible to manufacture
54
55 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
4.1.1 Airfoil selection
Airfoil cross sections from the UIUC airfoil database were analysed to meet the minimum
requirements in Table 4.1. A minimum wing lift coefficient of CL = 1.2 was required to
meet the design points specified in Section 3.6.4. A small negative pitching coefficient is
desirable for longitudinal stability whilst gradual stall characteristics promote safer flight.
After all other requirements have been met, maximum L/D and minimum camber are
required for flight efficiency and manufacturability respectively.
Table 4.1: Airfoil RequirementsLift coefficient (CL) > 1.2Pitching moment (Cm) < 0 and smallStall characteristics GradualL/D MaximumManufacturability Minimum camber
Inboard wing airfoil
The inboard wing is required to have a thick airfoil cross-section to provide maximum
internal volume to house the outboard wing. A minimum thickness of 16% was found to be
sufficient from preliminary CAD models. Table 4.2 lists airfoils which met the minimum
thickness requirement and provided a section lift coefficient greater than 1.2. The wing
lift coefficient requirement applies only to the extended wing configuration, which has
an aspect ratio of 6.1. This value was used in the JavaFoil package to determine the
three-dimensional airfoil characteristics. The S8036 and e664 airfoils failed to meet the
wing lift coefficient requirements and the e1098 airfoil was eliminated due to its pitching
moment coefficient and stall behaviour.
Table 4.2: Inboard wing candidate airfoilsAirfoil Stall characteristic Pitching moment CLmax
S8036 Gentle -0.033 to -0.066 1.18S8037 Gentle -0.036 to -0.071 1.22e1098 Moderate -0.125 to -0.1573 1.27e664 Moderate -0.06575 to -0.12725 1.17NACA 2416 Gentle -0.06 to -0.083 1.27NACA 4416 Gentle -0.111 to -0.136 1.47
The remaining airfoils (S8037, NACA2416 and NACA4416) were assessed on the basis
of L/D and camber. Figure 4.1 shows the L/D of the S8037 and NACA2416 airfoils
are significantly greater than that of the NACA4416 airfoil. In comparison to the S8037
airfoil, the NACA2416 airfoil has superior L/D performace at cruise angles of attack. The
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.1. WING DESIGN 56
NACA2416 airfoil profile was selected to maximise L/D performance during cruise whilst
maintaining good performance at other angles of attack. NACA airfoil cross sections are
also simpler than modern airfoils such as the S8037, which will simplify the manufacturing
process.
Figure 4.1: Lift to drag ratio of candidate inboard wing airfoils
Outboard wing airfoil
The outboard wing is required to have a thin airfoil cross-section to allow it to be stored
within the inboard wing. A maximum thickness of 12% was found to be sufficient from
preliminary CAD models. Table 4.3 lists the airfoils which meet this maximum thickness
requirement and provided a section lift coefficient greater than 1.2. The three-dimensional
lift coefficient was determined using the JavaFoil package with an aspect ratio of 6.1.
The sg6042 and NACA 4412 airfoil sections were the only airfoils to meet the wing lift
coefficient requirements.
Table 4.3: Outboard wing candidate airfoilsAirfoil Stall characteristic Pitching moment CLmax
e214 Moderate -0.068 to -0.161 1.17s2091 Moderate -0.048 to -0.082 1.14s4310 Severe -0.041 to -0.089 1.11s4320 Severe -0.044 to -0.096 1.10sd7032 Moderate -0.044 to -0.099 1.12sg6042 Moderate -0.068 to -0.133 1.21sd7034 Moderate -0.048 to -0.82 1.15NACA 2412 Moderate -0.045 to -0.073 1.06NACA 4412 Moderate -0.078 to -0.125 1.25
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
57 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
Figure 4.2: Lift to drag ratio of candidate outboard wing airfoils
The sg6042 airfoil provided superior L/D performance at lower angles of attack than the
NACA 4412 as seen in figure 4.2. The NACA 4412, however, was selected instead as the
greater wing lift coefficient of this section allows for reductions in the manufacturing pro-
cess. The NACA 4412 section also has more favourable pitching moment characteristics
and provides superior L/D performance at α > 5◦.
4.1.2 Installed incidence angles
The installed incidence angle of the inboard wing was specified to reduce drag during
cruise in the retracted configuration. For a design weight of 6 kg, cruising at a speed
of 120 km/h at the maximum altitude permitteed by MAAA (122 m), a required lift
coefficient of 0.198 is given by equation 4.1. For a Reynolds number range of 5.4× 105 to
1.2× 106 and A = 3, this corresponds to an angle of attack of 0.33◦.
CL =2W
ρV 2S(4.1)
Initially the outboard wing was positioned at an incidence angle of −3◦ to enable both the
inboard and outboard wings to attain their maximum lift coefficient at the same aircraft
angle of attack. The outboard wing incidence angle, however, was modified to zero to
solve otherwise unresolvable inboard wing tip layout issues. The outboard wing stalls at
αaircraft = 12◦ whilst the inboard wing will stall at αaircraft = 15◦. This will limit the
inboard wing to a maximum lift coefficient of 1.2 without causing stall on the outboard
wing. This result is acceptable as both the inboard and outboard wings still achieve the
required wing lift coefficient of 1.2 within the operational aircraft angle of attack range.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.1. WING DESIGN 58
4.1.3 Wing loading
Aircraft wings experience shear, bending, torsional and axial loads resutling from aero-
dynamic and weight forces. Shear, bending and torsional loads were calculated from a
maximum load factor, wing lift distribution and wing weight distribution. Axial loads
were assumed to be negligible for a small UAV.
Maximum load factor
The maximum load factor was determined from a combination of the following require-
ments:
• CASA UA25.337 requires a limit maneuver load factor of +3.8 and -1.5
• Section 2.4 specifies operation in wind speeds up to 18.5 km/h
The maneuver V-n diagram is defined by the stall curve (Equation 4.2) and the limit
maneuver load factor.
n =CLρSV
2
2W(4.2)
Deviation from the nominal load factor of one, due to gusts, is given by equation 4.3.
Equation 4.4 is the modified gust velocity which accounts for the gust alleviation factor
given by equation 4.5. The mass ratio, given in Equation 4.6, accounts for the influence
of aircraft weight on the effect of the gust.
∆n =ρUV CLα2((W/S)
(4.3)
U = KUde (4.4)
K =µ1.03
6.95 + µ1.03(4.5)
µ =2(W/S)
ρgcCLα(4.6)
The combined maneuver and gust V-n diagram is bound by the VNE for each configuration.
VNE was estimated to be 50% greater than the cruise speed (Raymer 2006). The V-n
diagram was determined for an aircraft weight of 8 kg and a motor power of 1.65 kW.
Increases in the weight of the aircraft will reduce the gust load factor and do not need to
be considered in this section.
Figure 4.3 shows a maximum load factor of n = 3.8 for the retracted wing configuration.
Figure 4.4 shows a maximum load factor of n = 5. The Morpheus UAV, however, is not
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
59 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
designed to fly at high speeds in the extended configuration and hence it is reasonable to
assume that the UAV would be in the retracted configuration whilst flying at high speed.
Based on this assumption the extended VNE was re-specified as 147km/h, which reduces
the load factor experienced to n = 3.8. This compromise allows for reduced structural
weight whilst still meeting CASA requirements.
Figure 4.3: V-n diagram for the retracted wing configuration
Figure 4.4: V-n diagram for the extended wing configuration
Wing lift and weight distribution
The wing lift distribution was calculated using the Lifting Line Method documented in
Abbott (1959). Schrenk’s lift distribution approximation was considered to be inappror-
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.1. WING DESIGN 60
piate due to the complex geometry and discontinuous geometry of the extended wing
configuration. The section lift coefficient, given in Equation 4.7, is comprised of the
’basic’ and ’additional’ components given in Equations 4.8 and 4.9 respectively. For an
untwisted wing clb = 0. The values of La were tabulated in Abbott (1959) for various
taper and aspect ratios.
cl = clb + CLcla1 (4.7)
clb =εαeSLbcb
(4.8)
cla1 =SLacb
(4.9)
The extended wing configuration was modelled as a single wing with A = 6 and a taper
ratio described by Equations 4.10 and 4.11. The retracted wing configuration was mod-
elled as a single wing with A = 3 and λ = 0.45. The resulting lift distributions are shown
in Figure 4.5.
λ = 0.45 fory
b/2< 0.625 (4.10)
λ = 1.0 fory
b/2≥ 0.625 (4.11)
Figure 4.5: Spanwise lift distribution for both wing configurations
Wing weight was distributed according to chord length as suggested by Raymer (2006).
The discretised wing load distribution was found according to equation 4.12 which includes
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
61 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
a load factor of 3.8. The discretised lift, wing weight and load distributions for the
extended and retracted configurations are shown in Figures 4.6 and 4.7 respectively.
Pi = n× (Li −Wwing,i) (4.12)
Figure 4.6: Extended wing configuration load distribution
Figure 4.7: Retracted wing configuration load distribution
Wing loads
The net load distribution was used to calculate the shear, bending and torsional loads
on the wings. Spanwise axial loads and drag loads are negligible and have not been
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.1. WING DESIGN 62
considered. The extended load distribution is the critical load case due to the increased
moment arm. Due to time constraints, only this critical load case was analysed.
The wing spars were designed to carry all shear and bending loads. The shear and bending
moment diagrams in Figures 4.8 and 4.9 account for the bracket supports which fix the
wing tongues to the fuselage formers.
Figure 4.8: Wing shear diagram
Figure 4.9: Wing bending moment diagram
The wing torsion was calculated as the moment produced by the lift force around the
shear centre of the foam cross section, where the lift force was placed at the centre of
pressure. The shear centre was assumed to coincide with the centroid of the foam cross
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
63 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
section. The wing weight acts through the centriod and hence has no contribution to the
moment.
The shear centre varies with halfspan location due to the tapered geometry of the wings
and the untapered geometry of the hollow section. The position of the inboard wing
shear centre was calculated numerically for the root and tip sections and interpolated for
other halfspan locations. The position of the outboard shear centre was also calculated
numerically. The calculated shear centres are constant for all angles of attack.
The variation of the centre of pressure with various parameters was considered using the
JavaFoil package. The centre of pressure was constant with halfspan location, but moved
forwards with an increasing angle of attack. Variation with Reynolds number was also
considered, but was found to be negligible for Reynolds numbers up to 1.2× 106 up to a
stall angle of α = 15◦.
The difference between the shear centre and the centre of pressure was used to calculate
the torque generated at each of the discretised halfspan locations. The torque produced
in the wings was plotted as a function of angle of attack in Figure 4.10. This shows that
maximum wing torque was generated at the stall angle of α = 15◦.
Figure 4.10: Torque as a function of angle of attack
4.1.4 Wing structural layout
The wing structural layout was designed to incorporate the morphing mechanism, struc-
tural members and internal access for maintanence. The wing structure consists of the
outboard wing, wing block, inbound wing and fuselage attachment. Each sub-structure
was designed to include structural members which carried torsional, bending and shear
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.1. WING DESIGN 64
loads.
Internal structure type selection
Built up and foam core structures were considered as two layout solutions which could
meet the structural design requirements. A built up structure uses a framework of spars,
ribs and stringers which enclosed by a skin. Built up structures are more weight efficient,
but are more difficult and costly to manufacture. A foam structure uses a foam core
which is shaped to the required wing geometry. Spars, end ribs or a skin may be added as
reinforcements. A foam structure is heavier, but is simpler to manufacture. The reduced
requirement for precision cut components in a foam core wing should also reduce cost.
A foam core structure was selected for manufacturing simplicity and lower cost. Foam core
structures easily allow for the integration of additional structural or functional components
as these items may be directly bonded to the foam core. This structure type has been used
extensively in previous UAV projets at the University of Adelaide and can be considered
to be a proven option.
Material selection
Materials for the wing structure were selected on the basis of specific strength, manufac-
turability, availability and cost. Additional consideration was required to avoid excessive
use of carbon-fibre which may result in radio frequnecy interference. Table 4.4 lists mate-
rials used and the associated componeets. Components which are specific to the morphing
mechanism are considered in 4.3.
Outboard wing
The outboard wing is required to carry inertial and aerodynamic loads when in the ex-
tended configuration and transfer these loads to the fuselage. The outboard wing is also
required to contain roller strips and a threaded rod sleeve to protect the outboard wing
from roller and threaded rod impact damage.
The outboard wing structural layout is shown in Figure 4.11. The foam core supports the
airfoil shape and locates other components. The foam in reinforced at the root and tip
with plywood ribs which provide further support to structural components and protection
from impact damage. Four 10 x 1.5 mm unidirectional carbon strips provide roller impact
protection as well as acting as carrying bending and shear loads. These strips are located
at positions which correspond to the chordwise position of the rollers. The 12 mm outer
diameter carbon tube acts as a sleeve for the threaded rod and carries outboard wing
bending and shear loads. Three layers of 85 gsm fibreglass in a ±45◦ / 0◦/90◦ / ±45◦
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
65 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
Table 4.4: Wing componenet materialsMaterial Possible useExtruded polystyrene foam Wing coreE glass, Fibreglass SkinsCarbon fibre Guide rails
SparsRoller stripsThreaded rod sleveBracketsWing tongue
Plywood RibsSparsServo hatches
Aluminium RibsSparsBracketsWing tongue
Hardwood Servo mountsEpoxy-resin Used in conjunction with fibreglass and carbon fibre
Used as a bonding agent between components
orientation form the wing skin. The two ±45◦ layers carry torsional loads whilst the
0◦/90◦ layer provides additional handling strength. This layup was incorporates an odd
number of plies to reduce warpage and a symmetrical layer orientation as in common
practice in composite skins (Raymer 2006).
Figure 4.11: Schematic of the outboard wing structural layout
Bending and shear loads are transfered to the wing block through the 12 mm carbon tube
which extends into the wing block. Torsional loads are transfered to from the fibreglass
skin to the plywood root rib which is bonded to the wing block. Torsional loads are also
transfered to the inboard wing through the mechanism rollers.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.1. WING DESIGN 66
Inboard wing
The inboard wing is required to carry intertial and aerodynamic loads and transfer these
loads to the fuselage. The inboard wing must contain a hollow section for the retracted
outboard wing, contain guide rails and other mechanism components at the tip rib. The
inboard wing design must also allow for internal access.
The inboard wing structural layout is shown in Figure 4.12. The foam core maintains
the airfoil shape and wing geometry whilst also providing mounting points for the servos
and ailerons. Balsa leading and trailing edges were included in the design as an option
if the desired surface finish was not obtainable from the foam core. Two 10 mm carbon
tube spars carry bending and shear loads whilst also acting as the mechanism guide rails.
The wing skin consists of three layers of 85 gsm fibreglass in a ±45◦ / 0◦/90◦ / ±45◦
orientation. The fibreglass skin carries torsional loads and provides handling strength.
The root and tip ribs transfer loads from the skin and foam core to the spars and to the
fuselage. The root rib is fibreglassed to the foam core to facilitate improved load transfer.
Figure 4.12: Schematic of the inboard wing structural layout
Internal access to the inboard wing was facilitated by removeable tip rib, which is shown
in Figure 4.13. The tip rib is bolted to four carbon-fibre brackets which are fibreglassed
into the inboard wing skin. The bolts are epoxyed into the brackets to prevent movement
during tip rib installation. The brackets were positioned to provide a rigid joint without
interfering with other tip rib components. Two 12 mm carbon-fibre tubes, bonded to the
tip rib, act as support sleeves and locate the spars within the tip rib. This enables the
tip rib to support the wing spars whilst still being removable. Rollers were positioned on
the tip rib to provide torsional support to the outboard wing whilst not interfering with
other tip rib components.
Geometrical constraints on the tip rib required small edge distances around the bolt
holes for the brackets and rollers. Aluminium, consequently, was considered to be a more
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
67 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
appropriate material than plywood. Five series, 6 mm thick aluminium sheet was selected
for the tip rib as this material was commonly available from the Mechanical Engineering
Workshop.
Figure 4.13: Removeable tip rib
Wing block
The wing block is required to transfer shear, bending and some torsional loads into the
inboard wing and fuselage structure from the outboard wing. The wing block must also
be able to slide on the mechanism guide rails.
The wing block structural layout is shown in Figure 4.14. The foam core and end ribs
locate structural components. Two 12 mm outer diameter carbon-fibre tubes act as guide
tubes and enable the wing block to slide along the mechanism rails. These two tubes, in
addition to the carbon-fibre tube which extends from the outboard wing, carry shear and
bending loads. Torsional loads are carried by a single layer of ±45◦ 85 gsm fibreglass.
Fuselage attachment
Figure 4.15 shows the fuselage attachment structural layout. The wing tongue consists
of two 12 mm carbon-fibre tubes tubes which are fixed to the fuselage by a total of four
brackets. These tubes transfer shear and bending loads to the fuselage structure. The
brackets were positioned at the closest possible point to the fuselage wall to minimise
the spar and wing tongue bending moment. The spars of each inboard wing are inserted
halfway into the wing tongue tubes. The wing tongue tubes constrains vertical and
horizontal motion whilst axial movement of the wings is constrained by a pair of bolts.
Torsional loads are transfered through the wing tongue and bolts to the fuselage internal
structure.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.1. WING DESIGN 68
Figure 4.14: Schematic of the wing block structural layout
Figure 4.15: Fuselage attachment
The wing tongue brackets, shown in Figure 4.16, were conservatively designed from 6 mm
aluminium plate and fixed to formers by two 3 mm steel bolts.
4.1.5 Structural analysis
The structural analysis of the wing was based on the load carrying capabilities of indi-
vidual components rather than considering the combined load carrying capacity of all
components. This approach greatly simplies the processes, however, wing structural tests
will be required to confirm the integrity of the overall wing structure. All bending and
shear loads were assumed to be carried by the carbon-fibre tubes and strips in the outboard
wings, wing blocks, inboard wings and wing tongue. The fibreglass skin was assumed to
carry all torsional loads. Effects of moisture and temperature were not considered on com-
ponents. The effects of the rollers were not considered in order to simplify the analysis
which results in conservative results.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
69 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
Figure 4.16: Wing tongue brackets
The structural analysis used the following safety factors as specified by CASA UA25.303:
• 2.25 for all composite components where moister and temperature are not considered
• 1.5 for all metal and wood components
Neutral axis and moment of interia
The neutral axis of the outwing wing spars was determined using Equation 4.13. The
centroids of the wing block tubes, inboard wing spars and wing tongues were situated in
the same plane and hence the neutral axis was located through the centroids.
yNA =
∑Aiyi∑Ai
(4.13)
Moments of interia about the respective neutral axes were calculated using the centroid
moments of inertia of each spar and the parallel-axis theorem (Equation 4.14). The
centroid moment of inertia for the strips and tubes were calculated using Equations 4.15
and 4.16 respectively. The spars were then considered as a single structure with the
equivalent moment of interia.
INA = Ic + Adc (4.14)
Ic,rectangle =bh3
12(4.15)
Ic,cylinder =π
64
(d4o − d4
i
)(4.16)
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4.1. WING DESIGN 70
Shear analysis
Shear stresses within the carbon-fibre members were calculated using Equation 4.17, where
V is the shear load, Q is the first moment of area, I is the moment of inertia and b is
the width at the point of interested. Due to the spanwise variation of the shear load and
the cross-sectional properties of the spars, it was necessary to determine the shear stress
over the entire wing structure. This analysis assumed a uniform distribution of shear
loads between the various carbon-fibre structural members. The results of this analysis
are shown in Figure 4.17. The highest shear stress occured at the inboard wing root rib
location as a result of the decrease in carbon-fibre tube cross-section. This maximum
shear stress of 7.94 MPa, including a safety factor of 2.25, gives a reserve factor of of 12.6.
The large reserve factor indicates that it is not necessary to consider the actual shear load
distribution between the two inboard wing carbon-fibre tubes.
τ =V Q
Ib(4.17)
Figure 4.17: Wing shear stress distribution for carbon-fibre members
Bending analysis
Bending stresses within the carbon-fibre structural members were calculated using Equa-
tion 4.18, where M is the bending moment, y is the distance between the neutral axis and
the point of interest and I is the cross-sectional moment of inertia. The bending stress
distribution for carbon-fibre members was calculated over the entire wing half-span as
shown in Figure 4.18. Only components which were furthest from the respective neutral
axis were considered in this analysis. This analysis assumed a uniform load distribution
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
71 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
between components within each wing section. For an ultimate compressive stress of 910.1
MPa all components, except for the wing tongues and rails, have reserve factors greater
than 13. The large reserve factor indicates that the this simplified analysis is sufficient for
these components. A more detailed analysis, however, was required for the wing tongue
and rail components.
σ =−My
I(4.18)
Figure 4.18: Wing bending stress distribution
The load distribution between the two rails was calculated by assuming the lift force acts
at the aerodynamic centre and using Equation 4.19. The leading rail was calculated to
carry 69.1% and 92.9% of the outboard wing and inboard wing loads respectively. The
load distribution was used to determine the total bending moments carried by the leading
rail. The maximum bending stress was found to be 1133 MPa at the wing root, which
was above the ultimate compressive stress of 910.1 MPa.
%Mleading spar =xtrailing spar − xac wing
xtrailing spar − xleading spar(4.19)
The bending stress in the leading rail was reduced through the addition of an 8 mm
OD carbon-fibre tube inside the rail. The reinforcement was extended into the wing rail
until the maximum compressive stress was reduced below 910.1 MPa. A reinforcement
extension of 100 mm was required to reduce the maximum compressive stress to 833.8
MPa, which gives a minimum reserve factor of 1.1. Figure 4.19 shows the bending stress
distribution in the leading rail, tongue and reinforcement tubes.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.1. WING DESIGN 72
Figure 4.19: Bending stress in the leading rail, tongue and reinforcement tubes
Deflection analysis
The deflection of the outboard wing tip was calculated to determine the wings structural
behaviour during flight. The deflection analysis assumed that non-carbon components
made no contribution to the rigidity of the wing structure. The outboard wing and wing
block were analysed seperatly from the inboard wing. The results from each individual
analysis were combined to determine the overall deflection of the outboard wing tip.
The outboard wing was assumed to be cantilevered from the wing block. The deflection
due to each discritised load on the outboard wing was calculated using Equation 4.20 and
the results superposed to determine the cantilevered outboard wing deflection.
δ =Pa2
6EI(3L− a) (4.20)
The inboard wing was assumed to be cantilevered from its root rib. The deflection due
to each discritised load was calculated using Equation 4.20 for loads on the inboard wing
and Equation 4.21 for moments due to loads on the outboard wing. The total inboard
deflection was found through superposition. The angular deflection at the inboard wingtip
was calculated in a similar manner using Equations 4.22 and 4.23.
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73 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
δ =ML2
2EI(4.21)
θ =Pa2
2EI(4.22)
θ =ML
EI(4.23)
The results of the individual inboard and outboard deflection analyses are shown in Table
4.5. The deflection in the ouboard wingtip, due to the angular deflection of the inboard
wing, was determined using the small angle approximation and then combined with the
deflections of the outbord and inboard tip to give the overall outboard wingtip deflection of
79.3 mm. An elastic modulus for carbon-fibre of 217 GPa was assumed for this calculation
(GraphiteStore.com 2009).
Table 4.5: Deflection results for individual wing sectionsOutboard deflection 2.12× 10−6 mInboard deflection 28× 10−3 mInboard angular deflection 0.108 radians
Torsional analysis
The torsional stress analysis of the wings assumed that all torsional loads were carried by
the fibreglass skin. Using an angle of attack of α = 15◦ the torsional stress in the skin was
calculated as a function of halfspan location using Equation 4.24. Figure 4.20 shows the
skin torsional stress in the extended and retracted wing configurations. The maximum
skin stresses in the extended and retracted wing configurations occured at the outboard
and inboard wing roots repsectively. These maximum stresses of 495 kPa and 401 kPa
are significantly below 54.5 MPa, which is the utlimate shear stress of E-glass/epoxy with
a 45% fibre fraction (Raymer 2006). Hence the wing skin will be capable of carrying the
torsional loads.
τ =T
2At(4.24)
Wing tongue bracket
A pair of wing tongue brackets are subjected to a load of 100.7 N. The bracket on the
leading tongue will be subjected to 85 N of the total load, which corresponds to 92.9%
of the inboard and 69.1% of the outboard wing loads. This load produces a moment
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.1. WING DESIGN 74
Figure 4.20: Wing skin torsional stress for both wing configurations
of 0.51 Nm about the bracket root. Due to time constraints the bracket geometry was
simplified in a conservative manner and analysed with hand calculations. The rounded
section of the bracket was replaced with a rectangular section with a wall thickness of 4
mm. The maximum bending stress and shear stress are expected to occur at the root of
the bracket and were calculated using Equations 4.25 and 4.26. The calculated stresses
are 16.6 kPa and 88.5 kPa for bending and shear stresses respectively. These values are
significantly below the yield stress for aluminium. Whilst over designed, the aluminium
thickness was required to distribute the load applied to the carbon wing tongues and to
provide sufficient area for bonding.
σ =My
I(4.25)
τ =V Q
Ib(4.26)
Each of the 3 mm diameter bolts may be assumed to carry half of the shear load of 85
N. The maximum shear stress in the circular bolts was calculated to be 8.02 MPa using
Equation 4.27. This stress is significantly below the yield stress of steel.
τmax =4V
3A(4.27)
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75 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
4.1.6 Wing design summary
The aerodynamic, structural and operational specifications of the wing were met during
the design process. NACA 2416 and NACA 4412 airfoil sections, with a 0.33◦ zero angle
of incidence, were selected to meet the aerodynamic design requirements. Carbon-fibre
tubes and strips, also used by the wing mechanism, were design to carry all bending and
shear loads with the appropriate load and safety factors. Torsional, handling and light
impact loads were carried by three layers of 85 gsm fibreglass. The inboard wing was
designed with a removeable wing tip to allow internal access and the control surfaces
were succesfully incorporated. The wings are mounted to the fuselage using concentric
carbon-fibre tube wing tongues and are easily removed by undoing a total of four nylon
bolts.
4.2 Empennage design
The preliminary and detailed design of the empennage considers aerodynamic and struc-
tural requirements. Aerodynamic requirements were used to determine the appropriate
airfoil section and the installed incidence angle for a Reynolds number up to 7 × 105.
Structural requirements were considered in conjunction with the empennage morphing
concept developed in section 3.4 to design a suitable structural layout. A structural anal-
ysis of the empennage also ensured that the empennage met CASA structural design
requirements.
4.2.1 Airfoil selection
The V-tail airfoil was selected to provide the required balancing moments in the longi-
tudinal and lateral directions. This function requires the selected airfoil section to be
able to generate approximately equal moments in either direction and consequently a
symmetrical airfoil was required. A tail thickness ratio similar to that of the wings was
also desireable (Raymer 2006). The variable location of the empennage would make it
difficult to mount servos in the fuselage and hence sufficient airfoil thickness was required
to mount the servos within the tail itself.
A NACA 0016 airfoil was selected on the the above requirements. A detailed airfoil
analysis for the tail airfoil was not possible given the project time constraints, however,
the NACA profile will make the tail simplier to manufacture than other airfoils.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.2. EMPENNAGE DESIGN 76
4.2.2 Stall recovery and installed incidence angle
The effective angle of attack of the V-tail may be approximated using Equation 4.28
(Raymer 2006) which describes the effective angle of attack of a horizontal tail surface.
This equation can be solved for the installed angle ih required for the tail to stall at the
same time as the wing. The NACA 0016 airfoil selected for the tail stalls at αh = 15◦ in
the given Reynolds number range. Using Equation 4.28 the maximum premissable tail
installed incidence angle is ih = 6◦. Installed incidence angles less than this value will
benefit from enhanced stall recovery behaviour.
αh = (α + iw)(1− δe
δα) + (ih − iw) (4.28)
An neutral installed incidence angle of zero degrees was selected for improved stall recovery
behaviour and reduced drag. A neutral V-tail installed incidence angle also simplifies the
manufacturing process. All trim forces must be generated by the ruddervators due to the
neutral angle of incidence. This is desirable as it allows the pilot to have complete control
over the trim of the aircraft in the air.
4.2.3 Empennage loading
The V-tail and boom experiences bending, shear and torsional loads. The highest loads
in each case correspond to a full ruddervator deflection with the aircraft travelling at its
maximum speed of 165 km/h. The greatest bending and shear loads correspond to a
pull-up maneuver with full ruddervator deflection. The greatest torsional loads occur at
the instant of a full aileron deflection. Both load cases occur with the tail in its extended
position.
Pull-up maneuver loads
A maximum downforce of 128 N was calculated in Appendix C. The resulting shear and
bending moment diagrams are shown in Figures 4.21 and 4.22 respectively.
Full aileron roll
The maximum torque generated by a full aileron roll was calculated in Appendix C to be
67.72 Nm. At the instant of the aileron deflection this torque will be transmitted throught
the boom to the V-tail.
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77 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
Figure 4.21: Shear diagram for the fuselage and empennage
Figure 4.22: Bending moment diagram for the fuselage and empennage
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4.2. EMPENNAGE DESIGN 78
Tail rail loads
During a pull up maneuver the tail generates a maximum downforce of 128 N and the
empennage weight generates a downwards load of 6.4 N. The moment produced by these
forces is reacted at the tail block by the tail rails. To simplify the analysis it was assumed
that the rail reaction loads act as point loads at the foremost and aftmost locations of
the tail block. This results in reactions forces of 712.4N and 846.6N. These loads were
found to produce the greatest shear loads and bending moments when the tail was in its
fully retracted position. The maximum shear load was calculated to be 680.6 N and the
maximum bending moment was 68.1 Nm.
4.2.4 Structural layout
The empennage structural layout is required to provide paths for bending, shear and tor-
sional load transfer. The layout must also enable the morphing mechanism to operate and
provide a solution for connecting the receivers leads to the ruddervator servos. Structural
layouts were required for the V-tail, boom and the empennage mounting to the fuselage.
Similar to the wing structure, a foam core structure type was selected for the V-tail. The
materials in Table 4.6 were considered for use in the empennage structure.
Table 4.6: Candidate empennage materialsMaterial Possible useExtruded polystyrene foam Tail coreE glass, Fibreglass SkinsCarbon fibre Guide rails
SparsPlywood Ribs
SparsServo hatches
Hardwood Servo mountsEpoxy-resin Used in conjunction with fibreglass and carbon fibre
Used as a bonding agent between components
V-tail
The V-tail design, seen in Figure 4.23, used a foam core structure similar to that of the
wings, except that the foam and fibreglass carry all shear, bending and torsional loads.
Structural spars and ribs were considered to be unnecessary due to the significantly smaller
shear and bending loads experienced by the V-tail in comparison to the wings. A fibreglass
skin with 0◦/90◦/±45◦/0◦/90◦ was selected. Two 0◦/90◦ layers were used to carry bending
loads in the absence of spars with the single ±45◦ carrying torsional loads in conjunction
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
79 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
with the foam core. The V-tail fins are bonded directly to the carbon boom and have
optional balsa leading and trailing edgees if required during manufacturing.
Figure 4.23: V-tail and boom structural layout
Empennage boom
The empennage boom, also seen in Figure 4.23, was required to contain a threaded rod
and protective sleeves for the ruddervator servo leads. The protective sleeves required
a minimum outer diameter of 5 mm. To house these protective tubes and a threaded
rod, whilst providing sufficient clearance to reduce the probability of the threaded rod
striking the protective tubes, a minimum boom internal diameter of 25 mm was required.
Carbon-fibre booms with 25 mm inner diameters were readily available with a 28 mm
outer diameter. These boom dimensions were selected pending a stress analysis, which
later confirmed the suitability of this component. The protective tubes were located on
either side of the threaded rod to minimise the probability of both servo leads being
damaged simultaneously.
Tail block
The tail block, seen in Figure 4.24, transfers all empennage loads from the boom to the
tail rails. The tail block also contains guides to enable the block to slide smoothly along
the tail rails.
Two 14 mm outer diameter carbon-fibre tubes were selected to act as the tail block guides.
These tubes, along with the carbon-fibre tail boom, carry bending and shear loads within
the block. Torsional loads are carried by the foam and a single layer of ±45◦ 85 gsm
fibreglass. The fibreglass also provides handling protection to the tail block. A triangular
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4.2. EMPENNAGE DESIGN 80
arrangement was selected as this minimised the height and width of the tail block required,
enabling the tapering of the fuselage rear and greater useable internal volume.
Tail rails
The tail rails, also seen in Figure 4.24, transfer all loads from the boom mounted V-tail
into the fuselage, in addition to forming part of the empennage morping mechanism. Two
12 mm outer diameter carbon-fibre tubes were selected for this purpose. The tail rails
are supported in all three directions by fuselage formers.
Figure 4.24: Boom and V-tail mounted to the tail rails by the tail block
4.2.5 Structural analysis
The small bending and shear loads on the tail fins, in combination with the use of three
layers of fibreglass skin and a solid foam core, meant that, due to time constraints, it
was unnecessary to undertake a structural analysis of the tail fins. An analysis of the
tail boom, however, was required. The maximum loads experienced by the tail boom are
given in Table 4.7.
Table 4.7: Maximum loads on the tail boomBending 92.3 NmShear 160.7 NmTorsion 67.72 Nm
Boom shear analysis
The maximum shear stress in the boom was calculated using Equation 4.29 to be 2.57
MPa, which becomes 5.78 MPa after the application of a 2.25 safety factor. This shear
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
81 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
stress gives a reserve factor of 17.9, indicating that the empennage boom is capable of
sustaining flight shear loads.
τmax =4V
3A(r0
2 + r0ri + ri2
r02 + ri2) (4.29)
Boom bending analysis
The maximum bending stress experienced by the boom is at its connection to the tail
block. Using Equation 4.30 stress was calculated to be 117.6 MPa, which becomes 264.5
MPa after a safety factor of 2.25 was applied. This maximum stress results in a reserve
factor of 3.44. Although over engineered, a large reserve factor is desirable as it increases
the likelyhood of the empennage surviving a crash landing.
σmax =−My
I(4.30)
The deflection of the carbon-fibre boom was also considered using the maximum aero-
dynamic force produced by the tail in conjunction with the empennage weight. It was
assumed that the boom was cantilevered from the tail block and that there was no de-
flection in the tail rails. The deflection due to each individual point load was calculated
using Equation 4.31, which gave a maximum deflection of 4.6 mm. It was assumed that
the carbon-fibre had an elastic modulus of 241 GPa (MatWeb 2009) and that 30% of
the fibres, specified by the manufacturer, did not contribute to rigidity in this plane of
motion. This deflection is below the maximum 6 mm permissable before dynamic struc-
tural behaviour must be considered and hence the selected boom dimensions meet the
deflection design requirements.
δ =Pa2
6EI(3L− a) (4.31)
Boom torsional analysis
The torsional stress due to the maximum torsion of 67.72 Nm was calculated using the
approximation for torsional shear stress in a circular tube given in Equation 4.32. The
maximum torsional shear stress was calculated to be 43.1 MPa, which becomes 97 MPa
with the application of a safety factor. This gives a reserve factor of 1.0, assuming an
ultimate torsional strength of 103.4 MPa (GraphiteStore.com 2009). This small reserve
factor is acceptable as the maximum torsion of 67.72 Nm is only applied for an instant.
The maximum continuous torsion, occuring when the empennage resists the roll of the
fuselage, is significantly less than the instantaneous torsion.
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4.3. MORPHING MECHANISM DESIGN 82
τmax =Tr
Ip(4.32)
4.2.6 Tail rail shear analysis
The tail rail shear analysis assumes that each rail carries half of the shear load of 680.1
N. The maximum shear stress was calculated using Equaton 4.29 to be 44 MPa including
the safety factor. This gives a reserve factor of 2.3.
4.2.7 Tail rail bending analysis
The maximum bending stress in the tail rails was calculated using Equation 4.30. This
was determined to be 387.4 MPa, which increased to 871.7 MPa with the inclusion of a
safety factor. This results in a bending reserve factor of 1.04.
4.2.8 Empennage design summary
The preliminary and detailed design of the empennage resulting in aerodynamic and
structural design solutions. A NACA 0016 airfoil was selected with an installed angle of
zero degrees. Empennage loads were determined and applies to the selected structural
layout. All components of the empennage structure analysed met CASA structural design
requirements.
4.3 Morphing mechanism design
The majority of the mechanism structural layout was covered in sections 4.1.4 and 4.2.4.
This section considers the design of the actuation system, the threaded rod and motor,
and the design outboard wing rollers.
4.3.1 Morphing mechanism loads
The reaction forces at the rollers, wing block and tail block may be calculated from statics
in conjunction with the outboard wing and tail loads calculated in sections 4.1.3 and 4.2.3
respectively.
The reaction force between the wing block and the guide rails was calculated to be 67.9
N whilst the reaction force between the outboard wing and the rollers was 96.9 N. The
force between the outboard wing and the rollers becomes 145.4 N with the application
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83 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
of a safety factor of 1.5 as required by CASA UA25.303. These loads were calculated
assuming a point reaction force at the midpoint of the block and at the roller location,
under a load factor of 3.8.
The reaction loads between the tail block and the guide rails was calculated for a pull-up
maneuver with full ruddervator deflection whilst travelling at maximum speed. The tail
block reaction forces were modelled as point loads at the fore and aft ends of the tail
block. Thes reaction forces were calculated to be 712.4N and 846.6N respectively. In
reality the loads will be distributed over the entire length of the tail block.
A friction factor of 0.74 was used to determine the resistance forces which the mechanism
actuator must overcome (Schon 2004). The calculation friction forces were 50.2 N and
626.5 N for the wing block and tail blocks respectively. These forces were used to determine
the required motor torque and minimum threaded rod diameter.
4.3.2 Threaded rod design
The threaded rod must fit within the geometrical constraints imposed by the outboard
wing and the tail boom. The rod must also have the appropriate pitch to enable extension
within approximately two seconds. Structurally the threaded rod must transmit sufficient
torque and not buckle during mechanism extension.
Material selection
Aluminium, steel and nylon threaded rods were considered as candidate materials and
were analysed on the basis of cost, elastic modulus, strength and weight. Nylon was
the lightest material, but was also most likely to buckle under loading. Steel offered the
greatest rigidity, strength and lowest cost, but was also the heaviest. Aluminium was
selected as a compromise between the two materials.
Bucking analysis
The threaded rods are fixed from rotation at the shaft coupler to the motor and at the
threaded insert in the wing or tail blocks. The Euler buckling critical load was calculated
using Equation 4.33 for a variety of threaded rod diameters using a safety factor of 1.5 for
aluminium. A modulus of elasticity of 73 GPa was assumed for this analysis (Gere 2002).
The minimum diameter which gave a critical load above the tail and wing friction forces
were 5.79 mm and 2.87 mm respectively.
Pcr =4π2EI
L2(4.33)
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4.3. MORPHING MECHANISM DESIGN 84
Pitch and diameter selection
The threaded rod was required to have a diameter less than 10 mm to fit within the
outboard wing carbon-fibre sleeve. To increase the interchangability of parts and to
minimise spares the tail and wing threaded rods were required to have the same diameter,
hence the minimum threaded rod diameter was 5.79 mm. It was also desirable to minimise
the threaded rod diameter in order to minimise weight.
Multi-start threaded rods were considered to increase the actuation speed of the mech-
anism. The cost involved in obtaining multi-start threaded rods which met the design
requirements, however, was prohibitive. Manufacturing threaded rods with a custom pitch
was also considered, but again was found to be too expensive.
After an extensive market survey a 6.35 mm diameter rod with a 1.27 mm pitch was
selected. This was the smallest diameter rod which was readily available and the coarsest
pitch available. This diameter threaded rod gives buckling reserve factors of 24.2 and 1.46
for the wing and tail threaded rods respectively.
Maximum transmissable torque
The maximum transmissible torque of the threaded rod is limited by the torsional strength
of the aluminium threaded rod. Assuming a yield stress of 270 MPa (Gere 2002) and
using a safety factor of n = 1.5, the maximum transmissible torque was calculated using
equation 4.34. This was calculated to be 9.05 Nm.
Tmax =τmaxIpnr
(4.34)
4.3.3 Motor design and selection
The morphing mechanism motor was selected based upon theoretical torque and rota-
tional speed requirements and confirmed by test results. Minimising the weight, cost and
dimensions of the motors was also considered. The motor mounting was also considered
in this section.
Design motor power required
The required motor power was calculated from the torque and rotational speed required
to actuate the morphing mechanisms under loading in less than two seconds.
The required motor torque is given by Equation 4.35 which assumes that the linear work
in moving the sliding block is equal to the rotational work in rotating the threaded rod.
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85 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
In this equation p is the threaded rod pitch. The required motor torque was calculated to
be 0.01 Nm and 0.127 Nm for the wing and tail morphing mechanisms respectively. These
values are significantly below the maximum transmissible torque of the selected threaded
rod.
Trequired =pFfriction
2π(4.35)
The rotational speed required to extend the wings and tail in two seconds was calculated
using equation 4.36, where L is the actuation length required and is 0.345 m and 0.4m
for the wings and tail respectively. The required rotational speeds were calculated to be
8150 RPM and 9450 RPM.
RPMrequired =L
tp× 60 (4.36)
The required motor powers were calculated to be 9 W and 125 W for the wing and tail
motors respectively.
Test results and motor selection
The calculated results for the required motor power was used as an initial estimate for
actual morphing motor tests. Motor tests, to determine the required motor power, were
necessary as the theoretical calculations did not account for misalignment in the morphing
system. Details of the morphing motor tests are given in Section 6.1.2. These tests
indicated that a motor power of 33.25 W was required to actuate the wing and tail
morphing mechanisms.
A trail and error method of motor selection was adopted due to time constraints and
difficulties in predicting the power requirements of the morphing system as well as the
low cost of the motors. A 160W motor, with a 3.1:1 gear box was selected and tests. The
motor specifications are given in Appendix A.1 and the test results in Section 6.1.2.
Motor mounting
The morphing motor mountings are required to locate the motor and minimise vibration.
The mounting should not interfere with other components and should minimise the po-
tential for alignment issues. The mounting should also allow the removal of the motors
for maintanence.
The wing morphing motors may be mounted to the wing root rib or to the fuselage
internal structure. The fuselage mount will create another point requiring alignment and
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4.3. MORPHING MECHANISM DESIGN 86
will also require the motors to be disconnected from the wing before removal. The wing
root mount will require a larger hole in the fuselage skin, but will not require removal
before the wings can be detached. Wing root mounting was selected as it reduces the
the assembly effort required, particularly alignment, and is a more rigid mount than the
fuselage option. A similar mounting option was also selected for the tail morphing motor,
with the motor mounted directly to a fuselage former.
4.3.4 Roller design
The rollers, seen in Figure 4.25, were designed to be mounted to removeable tip rib. The
mounting locations were selected to maximise the chordwise seperation between the rollers
in order to minimise the magnitude of the point loads in reacting outboard wing torsion.
Figure 4.25: Roller model
Material selection
Delrin was selected for the rolling element material as it is softer than the carbon-fibre
strips on the outboard wing and it is a commonly used bearing material. Steel was
selected for the roller axel as a high strength material due to the small shaft diameter
required to meet the roller geometrical constraints. Aluminium, steel, composite and
plywood were considered as potential roller bracket materials. Composites and plywood
were considered inappropriate due to the small edge distances required by geometrical
constraints. Aluminium was selected over steel for its high specific strength.
Roller Structural Analysis
The most likely mode of failure was determined to be the shear failure of the roller shaft
or bracket in the vicinity of the shaft hole. Equation 4.37 gave a maximum shear stress
in the shaft of 30.8 MPa, which increases to 46.3 MPa with the inclusion of a 1.5 safety
factor. Assuming a steel yield stress of 340 MPa this results in a reserve factor of 7.35
(Gere 2002). Equation 4.38 gave a maximum shear stress in the bracket of 14 MPa, which
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
87 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
increases to 21 MPa with the inclusion of a 1.5 safety factor. Assuming an aluminium
yield stress of 41.4 MPa (MatWeb 2009), the bracket reserve factor is 1.97. The rollers
were therefore deemed sufficient to withstand the possible applied loads.
τmax =4V
3A(4.37)
τmax =QV
Ib(4.38)
4.4 Control system design
The control system electronics were divided into three distinct subsystems: the thrust
subsystem, the control surfaces subsystem and the morphing subsystem. The components
included in each subsystem are as follows:
Subsystem: Included components:
Thrust subsystem Thrust motor
Thrust ESC
Thrust batteries
Control surfaces subsystem Main transmitter
Main receiver and battery
Control surface servo-actuators
Thrust ESC
Charging circuit for the main receiver battery
Morphing subsystem Morphing transmitter and receiver
Morphing motors
Morphing ESCs
Printed circuit board and logic circuitry
Morphing LiPo battery
4.4.1 Thrust subsystem
The selection of the thrust motor is discussed in detail in Section 3.8. The ESC selected
was a Dualsky DSXC9036HV as recommended by the manufacturer for the motor selected.
Specifications of the ESC are included in Appendix A.1. These components were wired
in accordance to the wiring diagram in Appendix A.2
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4.4. CONTROL SYSTEM DESIGN 88
4.4.2 Control surfaces subsystem
It is required in the project specifications that the UAV will be controlled via a radio
control system. A discussion of the selection of components for the electronic system
is outlined below. These components were selected based on system requirements, and
included a transmitter, receiver, receiver battery pack, ESC and servos. These are shown
in Figure 4.26, and the specifications can be found in Appendix A.1.
(a) Transmitter (b) Receiver (c) Battery (d) ESC (e) Servos
Figure 4.26: Control surfaces subsystem electronic components
Two transmitter and receiver sets were available from the university: the Spektrum DX7
and the JR Propo X2610. These components operated on different frequencies (2.4GHz
compared to 36 MHz respectively), and were thus able to be used simultaneously on
the same aircraft. The X2610 was used for the control surface subsystem as there have
been negative reports by aero-model enthusiasts regarding the reliability of the Spektrum
series. These reports are not supported by data, but there was no disadvantage to heeding
the warning of the aero-model enthusiasts and using the Spektrum for the less-essential
morphing subsystem. A dedicated battery was required to power the receiver and the
control surfaces. The use of a battery eliminator circuit (BEC) to power the control
surfaces from the main batteries was considered, but this solution was not feasible due to
the high voltages required for the thrust motor. An external antenna was used for this
receiver in order to reduce the effect of the radio shadow from the carbon tubes inside the
aircraft. Both receivers and antennae were located as far away as possible from the high
voltages and currents used in the thrust electronic subsystem.
Exponential rates were programmed into the main transmitter, to ensure appropriate
sensitivity in different flight phases. Flaps were mixed in to the ailerons and elevators to
assist with lift on take-off and landing. The ruddervators were mixed into the rudder and
elevator channels for full V-tail functionality.
Four JR DS821 digital servos were used to actuate the four control surfaces. The servos
were sized from statistical methods, and the digital servos were selected over analogue
servos on advice from industry experts because of their superior reliability for similar cost
and the same weight.
The components in the control surface subsystem were wired in accordance to the wiring
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89 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
diagram in Appendix A.3
4.4.3 Morphing subsystem
The morphing subsystem consisted of the electronics required for the morphing function-
ality of the aircraft. The subsystem was isolated so that morphing functionality could
easily be switched off if not required or problematic. The morphing transmitter was a
Spektrum DX7, which operated with the Spektrum AR6200 DSM2 receiver. These com-
ponents were provided by the University of Adelaide at no cost. The receiver operated
seven channels, which was more than the four channels required, but there was no weight
penalty for using the more advanced receiver. The receiver did not required its own power
supply, as it was powered by the morphing battery through the logic circuit. Exponential
stick response was programmed into the transmitter to ensure that small adjustments
to the morphing components could be made. The setting of the rates allowed the wing
morphing speed to be synchronised between the wings, despite a difference in friction
conditions.
The morphing battery was sized by calculating the power required to perform the most
demanding test, and requiring that this test be able to be performed twice without charg-
ing. The lightest battery (within reasonable cost) to meet this requirement was selected,
which was the FPEVO25-18002S.
The morphing motors were sourced as 20W 3-phase electric motors. The motors were
relatively cheap and provided sufficient power to morph the wings and tail under design
loading. The motors were compact and light and thus suitable for the aircraft envi-
ronment. Three forward/reverse remote control car ESCs were sourced to control the
morphing motors. The ESCs sourced were EZRun-25A-SL which had suitable continuous
and burst current ratings to match the motor. Forward/reverse ESCs were required in
order to accomodate the extend/retract functionality of the morphing mechanism, and
the ability to control the speed of the actuation.
The morphing mechanism functionality relied on the presence of control logic. This logic
was needed to determine under which conditions the wings and tail should morph. The
logic system needed to allow the morphing mechanism to stop within the physical limits,
and to provide emergency functionality to the aircraft. Limit switches on the extremities of
the wings and tail, and a circuit board with programmable microcontrollers was designed
in order to provide this functionality. The requirements for the logic circuitry are included
in Table 4.8. Pin and wiring diagrams for the morphing subsystem have been included in
Appendix A.4
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4.5. FUSELAGE DESIGN 90
Table 4.8: Requirements for logic circuitryState:
Extended RetractedSignal: Extend 1.55 µs (stop pulse) As received
Retract As received 1.55 µs (stop pulse)Emergency 1.55 µs (stop pulse) 1.90 µs (extend pulse)
4.5 Fuselage design
The fuselage preliminary and detailed design was concerned with meeting the structural
and layout requirements of the fuselage section. The positioning of components within
the fuselage to obtain an appropriate centre of gravit and the positioning of the landing
gear was also analysed.
4.5.1 Component layout
The position and installation of the onboard electronics required consideration in the
design of the internal layout of the fuselage. In order to conform to the strict centre-of-
gravity envelope on which the stability of the aircraft relies, as many electrical components
as possible were located at the front of the fuselage. Exceptions to this were as follows:
In the control surfaces subsystem, the receivers for the radio signal were located as far
away as possible from the high voltage of the thrust motor and the thrust ESC. These
components produce electrical noise that can interfere with the signal from the pilot.
Both receivers and both antennae were located in the aft-most bay of the aircraft for this
reason.
In the morphing subsystem, the printed circuit board for the morphing motor logic cir-
cuitry was located in an area close to the wing morphing motors. This position allowed
easy access to the plugs on the printed circuit board from the morphing motor hatch.
Batteries for the thrust receiver and the morphing motor were located in the aft-most
bay of the fuselage to provide adjustment to the centre of gravity of the aircraft after the
landing gear was placed. This also allowed easy access to these batteries for convenient
replacement between flights.
A design anomaly of the morphing motor logic board was that one particular ESC channel
had to receiver power before the other two ESC channels would become active. The tail
ESC was installed into the “master” channel, and the switch was lengthened and made
accessible through the ESC hatch. A different switch was installed to connect the battery
to the thrust receiver. This allowed the battery to be installed well before flight, and
for control surfaces to be activated easily from one location. This control surfaces switch
was lengthened and made accessible through the ESC hatch. The leads to connect the
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91 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
thrust batteries to the thrust ESC were also made accessible through the ESC hatch.
This allowed all three fuselage electronic subsystems to be switched on through the one
hatch, making pre-flight systems engagement simpler.
4.5.2 Structural layout
A monocoque or built-up internal structure were the two options considered for the fuse-
lage internal structural layout. A moncoque structure would have required an understand-
ing of composite structure beyond that of an undergraduate level and would have been
difficult to modify. A built-up structure was selected instead as it uses well understood
materials and would be easier to modify or repair after initial manufacturing.
High specific strength and specific modulus materials were required for the design of the
fuselage built-up structure in order to minimise weight. The materials should be readily
available and easy to manufacture with. Candidate materials, along with the proposed
uses, are given in Table 4.5.2.
Table 4.9: Candidate materials for the fuselage structureMaterial Proposed usePlywood Formers
LongeronsMounting plates
Aluminium FormersLongeronsMounting platesBrackets
Carbon-fibre Landing gearLongeronsMounting platesReinforcement
Fibreglass Fuselage skinFormer reinforcement
Epoxy resin Used with fibreglassUsed to bond the structure together
The fuselage structural layout selected is shown in Figure 4.27. A total of seven 9 mm
thick plywood formers were positioned as required to mount the motor, nose gear, tail
mechanism, leading and trailing wing spars, tail rails and the fairing. The formers main-
tained the fuselage shape and longeron positions, assisted in carrying torsional loads and
served as a mounting point for other structural components. Four 15 mm x 9 mm ply-
wood longerons carry shear and bending loads. The longerons were positioned to create
the largest moment of inertia about both bending axes whilst maintaining sufficient edge
distances for all cutouts required. The longerons and formers contained cutout slots to
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.5. FUSELAGE DESIGN 92
locate the structure and for ease of manufacture. At cutout locations the longeron cross
section was reduced to 7.5 mm x 9 mm. Three layers of 85 gsm fibreglass were selected to
form the fuselage skin. A ply orientation of 0◦/90◦/± 45◦/0◦/90◦ was selected to provide
handling strength. The 0◦/90◦ are capable of carrying local bending loads which may be
applied directly to the skin whilst the ±45◦ layer carries local torsional loads.
Figure 4.27: Fuselage structural layout
Internal access to the fuselage was obtained hatch cutouts in the fuselage skin. Hatches
were located on the lower side of the fuselage in the nose gear and battery bays and
on the upper side of the fuselage between the spar formers. The empennage morphing
mechanism was accessable through a removeable fairing, which allows the tail rails and
the tail boom to be removed from the fuselage.
Landing gear mounting
The landing gear mounting layout is shown in Figure 4.28. The main landing gear are
bolted to a sacrificial plywood plate which is then bolted to a plywood plate mounted
between the leading and trailing spar formers. Nylon bolts were used at each interface
which are designed to shear during a heavy landing. The sacrificial bolts and plywood
plate protects both the main gear and the fuselage structure.
4.5.3 Weight distribution and centre of gravity
The centre of gravity envelope was calculated using the component list given in Table 4.10.
Minor components and electronics have been included with the fuselage structure weight.
The centre of gravity envelope for the Morpheus aircraft is shown in Figure 4.29, where
the tail leading edge location is plotted on the y-axis. The main landing gear position
listed was the result of an iterative process covered in Section 4.5.4.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
93 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
Figure 4.28: Landing gear mounting layout
Table 4.10: Aircraft weight breakdown summaryComponent Component CG from nose [m] Mass [kg]Motor 0.0315 0.377ESC 0.107 0.125Batteries 0.223 1.295Fuselage structure 0.3743 2.043Payload 0.425 0.5Wings 0.539 2.41Main gear 0.565 0.345Morphing and receiver batteries 0.807 0.226Tail (extended) 1.253 0.654Tail (retracted) 0.853 0.654
Figure 4.29: Centre of gravity envelope. Note that the y-axis is tail position not weight
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.5. FUSELAGE DESIGN 94
4.5.4 Landing gear
The landing gear were designed based upon load bearing and positioning requirements.
Due to project time constraints, the custom design and manufacture of landing gear was
not considered.
Landing gear selection
A steel nose gear, commonly used on similar sized model aircraft, was selected due to a
lack of available alternatives within Australia. Aluminium and composite main landing
gear were considered, with composite gear selected to save weight. The composite gear
shown in Figure 4.30 was selected based on availability and cost. The main landing gear
requirements and the specifications of the selected gear are listed in Table 4.5.4.
Figure 4.30: Selected main landing gear (Pilot-RC Inc. 2009)
Table 4.11: Main landing gear requirements and specifications of the selected gearRequirement Specification
Weight rating Greater than 8 kg 8.1 kgHeight 128-280 mm 195 mmWheel track Greater than or equal to 460 mm 460mmMounting plate width Less than or equal to 160 mm 160 mm
Landing gear positioning
The Morpheus UAV was designed to takeoff in the tail extended configuration as this gives
the greatest elevator control authority. Consequently the landing gear were positioned
only for the tail extended configuration.
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95 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
The nose landing gear was positioned on the second former in order to utilise the existing
fuselage internal structure. The main landing gear was positioned to meet weight distri-
bution, tipback angle and rollover angle requirements. The position calculation used an
iterative process in which a landing gear position with be assumed, the centre of gravity
calculated and the assumed position assessed against the three positioning criteria.
The main landing gear must carry between 80-90% of the aircraft weight. Insufficient
weight on the main gear will make takeoff rotation difficult, whilst excessive weight on
the main gear will result in poor nose gear ground handling qualities. A tipback angle
in excess of αstall = 15◦ was required to ensure the aircraft does not tipback onto its tail
during rotation. A rollover angle less than 63◦ was required to ensure the aircraft does
not rollover during ground maneuvers.
The main landing gear was positioned 565 mm from the aircraft nose based on the cal-
culations shown in Appendix B. This position met all three requirements for the empty
centre of gravity and the payload centre of gravity. This position is also forward of the
aftmost centre of gravity, ensuring that the aircraft will not tip back if the batteries and
payload are removed.
4.5.5 Fuselage loads
Fuselage loads were calculated for the three critical cases of a maximum speed pull-up
maneuver, a full aileron roll and a static thrust case with the tail fixed. These load cases
were calculated in Appendix C.
Pull-up maneuver
A pull-up maneuver at maximum speed and load factor generates the greatest bending and
shear loads in the fuselage. This calculation assumes that the aircraft is in its retracted
wing, extended tail configuration flying at 165 km/h. The shear and bending moment
diagrams are shown in Figures 4.31(a) and 4.31(b) respectively.
Full aileron roll
A full aileron roll at maximum speed generates the highest torsion in the fuselage. This
analysis assumed the critical case of retracted wings and an extended tail. The maximum
instantaneous torque generated by a full aileron deflected was calculated to be 67.72 Nm,
with the maximum continuous torque being 12.75 Nm. The maximum continuous torque
is provided by a full ruddervator deflection to oppose the aileron rolling moment.
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4.5. FUSELAGE DESIGN 96
(a) Fuselage shear diagram
(b) Fuselage bending moment diagram
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97 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
Static thrust
The maximum static thrust was found to be 82.4 N from the static thrust tests. This
axial load is transmitted through the entire fuselage structure assuming that the aircraft
is held at the tail.
4.5.6 Structural analysis
The structural analysis of the fuselage considered the fuselage internal structure and the
skin seperately. The plywood longerons and landing gear plate were assumed to carry all
shear, bending and axial loads. The skin was assumed to carry all torsional loads. The
carbon-fibre rails will reduce the percentage of the load carried by the plywood longerons
and hence were ignored in this analysis.
Neutral axes and moments of inertia
The fuselage internal structure consisted of five sections with different neutral axes and
moments of inertia. The neutral axis was calculated by finding the centroid of the cross
section. The cross section moment of inertia was calculated from the moment of inertia
of individual components about their individual centrelines and the parallel-axis theorem.
The respective neutral axes and moments of inertia are given in Table 4.12. The neutral
axis was measured from the centreline of the lower longerons.
Table 4.12: Neutral axes and Moment of intertia for various fuselage sectionsFuselage position [m] Neutral axis [m] Moment of inertia [m4]
0 < x < 0.472 0.0365 7.23× 10−7
0.472 < x < 0.486 0.0129 1.24× 10−5
0.486 < x < 0.592 0.0102 3.43× 10−5
0.592 < x < 0.615 0.0129 1.24× 10−5
0.615 < x < 0.824 0.0365 7.23× 10−7
Internal structure shear stress
The location of the maximum shear stress in the fuselage internal structure differs with
fuselage station due to changes in the neutral axis of that cross-section. The maximum
stress at each fuselage station was calculated with Equation 4.39 and is shown in Figure
4.31. Here Q is the first moment of area and b is the material thickness at the point of
interest. The maximum shear stress of 179.5 kPa, including a safety factor of 1.5, occurs
slightly forward of the leading spar former. This is due to the reduced cross section at this
point and the local influence of the wing lift. The ultimate shear stress for plywood is 7.93
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
4.5. FUSELAGE DESIGN 98
MPa, which gives a reserve factor of 44.2 (Munitions Board Aircraft Committee 1944).
This indicates that the fuselage structure was overdesigned for shear, but will increase
the chances of the structure withstanding a light crash.
τ =V Q
Ib(4.39)
Figure 4.31: Maximum fuselage shear stress
Internal structure bending stress
The maximum bending stress occurs in the upper longerons at all fuselage stations as
these longerons are located the greatest distance from the neutral axis. The bending stress
distribution in the upper longerons was calculated using Equation 4.40 and is shown in
Figure 4.32. The maximum bending stress of 12.5 MPa, including safety factor, occurs
forward of the leading spar former. This is a result of the lower moment of inertia at this
location. The ultimate stress for plywood is 18 MPa, which gives a reserve factor of 1.44
(Munitions Board Aircraft Committee 1944). This indicates the fuselage structure will
be able to withstand flight bending loads.
σ =My
I(4.40)
Internal structure axial stress
The maximum axial stress in the fuselage structure will occur at the fuselage station with
minimum cross-sectional area. Hence the sections foreward and aft of the wing spars
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99 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
Figure 4.32: Bending stress in the upper longeron
will experience the maximum axial tensile stress, calculated to be 152.6 kPa. The tensile
stress due to static thrust may be added to the maximum bending stress of 12.5 MPa to
obtain the overall maximum tensile stress. The contribution of the static thrust stress,
however, is negligible compared to the bending stresses and may be absorbed into the
reserve factor.
Skin torsional
The torsional loads in the fuselage were carried by the skin and the internal structure. For
the purposes of this analysis it was assumed that the fibreglass skin carried all torsional
loads and that there are no cutouts in the fuselage. The maximum torsion of 67.72 Nm due
to a full aileron roll was assumed to be uniformly applied to the fuselage skin. The skin was
treated as a closed thin-walled section and the torsional shear stresses calculated using
Equation 4.41, where A is the area enclosed by the section and t is the skin thickness
(Megson 2007). The maximum torsional shear stress was calculated to be 35.93 MPa,
including a safety factor of 2.25, which occured at the nose of the aircraft. The torsional
shear stresses calculated at each of the formers are given in Table 4.13. The torsional
shear stress in the sections inbetween the formers may be interpolated from these results.
τ =T
2At(4.41)
Hatch cutout edges were bonded to formers and longerons to reinforce these edges and
transfer shear flow through these structural members.
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4.6. FLIGHT PERFORMANCE ANALYSIS 100
Table 4.13: Torsional shear stress at former locations with a safety factor of 2.25Former Shear stress [MPa]Nose O-ring 38.1Firewall 10Nose gear 6.1Tail motor 5Leading and trailing spar 4.9Rear 6
4.5.7 Fuselage design summary
The structural layout of the fuselage was designed to allow for internal access and to
carry all flight loads, using CASA safety factors, with appropriate reserve factors. The
centre of gravity envelope of the aircraft was calculated and used to position the landing
gear for the normal takeoff configuration. The preliminary componenet layout specified
in this section was flexible and allowed for the movement of the centre of gravity to meet
stability requirements.
4.6 Flight performance analysis
4.6.1 Longitudinal stability analysis
The Morpheus UAV is designed to alter its longitudinal stability by varying its tail posi-
tion. Whilst other components of stability are likely to be affected by the tail morphing,
the calculation of these effects is beyond the scope of the project. The longitudinal sta-
bility of the aircraft was measured by the static margin, defined in Equation 4.42. The
centre of gravity envelope was determined in Section 4.5.3, whilst the neutral point must
be calculated.
SM = xnp − xcg (4.42)
The centre of gravity, neutral point and static margin for each morphed configuration were
calculated as a percentage of the respective mean aerodynamic chord for that configura-
tion. The mean aerodynamic chord for the retracted wing configuration was calculated
for a conventional wing geometry using Equation 4.43. The mean aerodynamic chord for
the extended configuration was calculated as the weighted average of the inboard (i) and
outboard (o) wing mean aerodynamic chords according to Equation 4.44. This method
was selected over the conventional method due to the discontinuous wing geometry and
the different lift-curve slopes of the inboard and outboard wings.
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101 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
cretracted =
(2
3
)c0
(1 + λ+ λ2
1 + λ
)(4.43)
cextended =ciSiCLα,i + coSoCLα,oSiCLα,i + SoCLα,o
(4.44)
The neutral point was determined from Equation 4.45, which considers contributions
from the wing aerodynamic centre, the fuselage and the empennage. Each of these sta-
bility terms are varied by morphing the wings, tail or both and were considered in detail
individually.
xnp = xac −Cmα,fCLα,w
+ VHCLα,tCLα,w
(1− ∂ε
∂α
)(4.45)
The wing contribution to the neutral point is xac. For the retracted wing configuration
this was calculated as the quarter chord location of the mean aerodynamic chord. For
the extended configuration, however, this was calculated as the weighted average of the
locations of the inboard and outboard wing quarter chord points as shown in Equation
4.46.
cextended =xac,iSiCLα,i + xac,oSoCLα,o
SiCLα,i + SoCLα,o(4.46)
The fuselage contribtionCmα,fCLα,w
was calculation from Equation 4.47. Sf is the maximum
cross sectional area, cf is the fuselage length and df is the equivalent fuselage diameter.
The lf term, the distance between the fuselage centre of pressure and the aircraft centre
of gravity, varies with tail location. The wing CLα varies with wing configuration due to
aspect ratio effects and the difference in CLα for the inboard and outboard wings. CLα
for the extended configuration was calculated as a weight average according to equation
4.48.
Cmα,f = −2Sf lfScf
(1− 1.76
(dfcf
)3/2)
(4.47)
CLα,extended =CLα,iSi + CLα,oSo
Si + So(4.48)
The empennage contribution is the most complex of the neutral point terms. The tail
horizontal volume ratio, given in Equation 4.49 (Brandt, Stiles, Bertin & Whitford 2004),
is affected by both wing and tail configurations, CLα,w is affected by wing configuration
as covered previously, and the downwash derivative, equation 4.50 (Brandt et al. 2004),
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4.6. FLIGHT PERFORMANCE ANALYSIS 102
is a function of both tail location and wing configuration.
VH =StltSc
(4.49)
∂ε
∂α=
(21◦CLα,wA0.725
)(cavglh
)0.25(10− 3λ
7
)(1− zh
b
)(4.50)
The calculated static margins for each of the Morpheus UAV tail and wing configurations
are listed in Table 4.14 and shown graphically in Figure 4.33. These show the morphing
the wings and tail both have a significant effect of the longitudinal stability of the aircraft.
Table 4.14: Morpheus UAV longitidinal stabilityWing configuration Tail configuration Static margin Static margin
(empty operational) (with payload)Extended Extended 12.13 13.04Retracted Extended 22.07 22.91Retracted Retracted 18.43 18.71Extended Retracted 13.58 13.88
Figure 4.33: Static margin envelope for the empty operational and operational with pay-load flight configurations
Retracting the wings increases the effectiveness of the empennage due to an increased
tail horizontal volume ratio and an increase the the ratio between tail and wing lift-curve
slopes. There is a reduction in tail effectiveness due to an increased downwash derivative
and a forward movement of the wing aerodynamic centre, however these effects are smaller
than the combination of the other effects. This results in retracted configuration being
more stable than the extended configuration.
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103 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
Moving the tail in the extended wing configuration has minimal effect on longitudinal sta-
bility. Retracting the tail increases stability, indicating the the centre of gravity travel due
to tail morphing is greater than the neutral point travel. In the retracted configuration,
however, the increased effectiveness of the empennage results in the neutral point moving
faster than the centre of gravity as the tail is retracted. This results in the extended tail
being more stable than the retracted tail.
4.6.2 Theoretical performance
The final design of the Morpheus UAV has the properties listed in Table 4.15. The
calculation of these parameters is shown in Appendix D. The extended configuration
has a lower wing loading and higher aspect ratio than the retracted configuration. The
extended configuration, consequently, produces greater lift due to increased wing area and
a higher lift coefficient. The higher lift coefficient results from the reduction of 3D effects
experienced by the inboard wing when the outboard wing is extended. The greater lift of
the extended configuration results in a lower stall speed and takeoff distance and a higher
endurance and rate of climb.
The retracted configuration has a higher wing loading and lower aspect ratio than the
extended configuration. The higher wing loading results in a reduced reference area whilst
the lower aspect ratio reduces the bending moment experienced by the wings. The lower
aspect ratio results in greater induced drag than in the extended configuration, however,
the reduction in planform area has a greater effect in reducing overall drag. The lower
bending moment in the retracted wing configuration also results in a higher velocity
never exceeded for the aircraft. This lower drag and structural loads result in a greater
maximum speed in the retracted configuration.
Table 4.15: Morpheus UAV PerformanceParameter Retracted wings Extended wingsStall speed [km/h] 64.39 48.6Takeoff distance [m] 63.2 35.2Maximum speed [km/h] 165.7 147Endurance [minutes] 22 36Rate of climb [m/s] 12.1 13.3
The interface between the inboard and outboard wings in the retracted configuration will
have a significant influence on the performance of the aircraft. These effects have not
been quantitatively analysed due to time, facility and skill-set constraints. The possible
effects of the interface, however, can be discussed qualitatively.
The geometrical step between the outboard and inboard wing surfaces will act as a fence
for 3D flow around the outboard wing tip. The step may further reduce the 3D effects on
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4.6. FLIGHT PERFORMANCE ANALYSIS 104
the inboard wing, resulting in a more efficient inboard wing with a greater CL,max. This
effect is likely to decrease stall speed and takeoff distances and increase endurance and
rate of climb from the values stated in Table 4.15.
The discontinuity between the outboard and inboard wings will result in turbulance at
the root of the outboard wing. The turbulence will reduce the efficiency and lift generated
by the outboard wing section. This effect will increase stall speed and takeoff distance
whilst decreasing endurance and rate of climb.
The extended configuration has an additional wingtip which will result in an additional
vortex being generated. This will increase the drag of the extended configuration and
would decrease the maximum speed, endurance and rate of climb of the extended config-
uration.
The effects of varying tail position have not been considered in this analysis. Whilst
retracting the tail would reduce wetted area, it may result in increased interference drag
between the fuselage, wings and empennage. The effect of varying tail position would
require a similar analysis to the inboard and outboard wing interface and hence was not
considered. Due to the unknown flow field in the vicinity of the inboard and outboard wing
interface and the empennage, flight testing will be necessary to determine the performance
of the Morpheus UAV in its different configurations.
4.6.3 Differential telescoping analysis
Differential telescoping of the wings will induce a rolling moment which could possibly
be used for roll control during flight. For the purposes of this analysis it will be assumed
that the aircraft is flying at 139 km/h in the retracted configuration and fully extends
one wing to enter a bank.
At 139 km/h the Morpheus UAV would be flying at zero angle of attack. The lift coefficient
of the retracted configuration at this angle is 0.195. The lift coefficients of the extended
configuration at this angle is approximately 0.274 and 0.452 for the inboard and outboard
wings respectively. The lift of the retracted inboard, extended inboard and outboard wings
was assumed to act at the respective mean aerodynamic chord stations. This resulting
in a net rolling moment of 21.04 Nm. An aileron deflection of 6.1◦ gives a similar rolling
moment of 21.2 Nm. This result indicates, that whilst differential telescoping is not
highly effective as a means of roll control, differential telescoping is theoretically capable
of providing a rolling moment to the aircraft and hence it is theoretically possible for
the Morpheus UAV to complete a circuit using only differential telescoping roll control.
The feasibility of this method, however, will require flight testing as this analysis has
not considered the roll rate or response rate required by the pilot. This analysis also
suggests that, should one wing be stuck in the retracted configuration and the other in
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105 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
the extended configuration, the pilot should be able to trim out the effects.
4.6.4 Optimal configurations for various flight phases
A typical reconnaisance mission involves takeoff, climb, cruise, loiter, dash and landing.
The optimal configuration for each of these flight phases may be determined from the
results of Sections 4.6.1 and 4.6.2. Confirmation of the recommended configurations for
each phase of flight will require flight testing.
Takeoff
Takeoff requires high lift generation at low speed and high pitch control authority to
enable rotation. A centre of gravity close to the main gear is also favourable. The
optimal configuration of the Morpheus UAV for takeoff, consequently, is the extended
wing and extended tail configuration.
Climb
The Morpheus UAV is predicted to have the highest rate of climb in the extended wing
configuration. During this stage the empennage high stability would be desirable, however,
the increase in stability by retracting the tail does not justify the resulting reducting
in control authority. Hence the optimal configuration for climb is extended wings and
extended tail.
Cruise
The Morpheus UAV has a higher cruise speed and a higher cruise speed for optimal
range in the retracted wing configuration. During cruise high stability is also desirable to
mimimise pilot or autopilot load. The optimal configuration for cruise is retracted wings
and extended tail.
Loiter
The Morpheus UAV has a significantly higher endurance in the extended wing configura-
tion. High stability is desirable and high control authorityy during loiter is not necessary.
Hence the tail should be retracted. This may also have the additional benefit of reducing
drag. The recommended configuration for loiter is extended wings and retracted tail.
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4.7. PRELIMINARY AND DETAILED DESIGN SUMMARY 106
Dash
A dash is likely to result after being discovered by hostile forces. The retracted wing
configuration provides the highest maximum speed for escape. The retracted tail con-
figuration also provides the most unstable configuration with retracted wings. This will
provide the most maneuverable configuration with some sacrific of control authority. The
balance between reduced stability, reduced control authority and reduced pitch moment
of inertia is beyond the scope of the project and has not been analysed. The recommended
configuration for dash is retracted wings and retracted tail.
Landing
Landing requires the lower stall speed provided by the extended wing configuration. High
longitudinal stability is also desirable during landing. The trade off between increased
stability and loss of control authority, however is beyond the scope of the project. The
extended tail configuration is recommended as it provides greater pitch and yaw control
authority and is likely to provide greater yaw stability.
4.7 Preliminary and Detailed Design Summary
The preliminary and detailed design of the Morpheus UAV resulted in solutions for the
wing, empennage, morphing, control and fuselage systems. Aerodynamic requirements for
airfoils and installed incidence angles were met. All structural components were analysed
using CASA safety and load factor requirements and were determined to be capable of
withstanding flight loads without failure. The morphing mechanisms were successfully
designed and integrated into the wing and empennage structures to save weight. The
aircraft control system was designed to utilise a primary pilot and a co-pilot who was
responsible for the morphing mechanism control. Sufficient access to internal components
was provided through a removeable tip rib, fuselage hatches and a removable fairing.
A weight and balance analysis, updated throughout manufacturing, provided appropriate
landing gear positions, but also indicated that the aircraft weight had increased to 8 kg.
The increase in weight resulted from a lack of understanding of manufacturing methods
and the components required during conceptual design. The aircraft was not resized due
to time constraints.
The static margin of the aircraft was determined in order to ensure sufficient longitudinal
stability and to investigate the effects of morphing on stability. The theoretical perfor-
mance of the Morpheus UAV was also determined to analyse the effects of wing morphing
on performance. A theoretical analysis indicated that roll control through differential
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107 CHAPTER 4. PRELIMINARY AND DETAILED DESIGN
telescoping should be theoretically possible the the optimal morphing configurations for
a typical reconaissance mission were discussed.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
5. ManufacturingThe manufacture of an aircraft focuses primarily on the production of a lightweight air-
frame that can meet all structural and aerodynamic requirements. The choice of the
materials and manufacturing processes used for the Morpheus UAV was largely dictated
by availability of appropriate equipment and advice. Manufacturing was divided into
four major areas: inboard wings, outboard wings and wing sliding blocks, fuselage and
empennage.
Photos of component installation are included in Appendix E, photos of processes are
included in Appendix E, and photos of components are included in Appendix E.
5.1 Available manufacturing methods
Several manufacturing methods were available for constructing the different components
of the Morpheus UAV. These manufacturing methods are outlined below.
5.1.1 Foam cutting
Many components of the Morpheus UAV were manufactured from extruded polystyrene
foam. This foam was cut using three main methods: rig hot-wire cutting, manual hot-wire
cutting and 3D CNC machining.
Rig hot-wire cutting
A hot-wire cutting rig owned and operated by the Mechanical Engineering Workshop
at the University of Adelaide was initially used to cut the foam components. The rig
consisted of a bow, pulley system and power supply. A steel wire drawn taught across
the bow is heated by an electrical current, allowing the wire to cut through the foam.
The pulley system is used to allow the bow to travel at a constant speed through the
foam, so that an improved surface finish can be obtained. In order to achieve the correct
aerofoil shape of the wings, laminex templates were CNC machined. These were secured
to the foam with double sided masking tape, allowing the hot-wire to be drawn across
their surface to achieve the desired shapes. Stations marked on the laminex templates
indicated the speed of the hot-wire through the foam.
The time required to correctly set up the rig, especially for tapered components, greatly
108
109 CHAPTER 5. MANUFACTURING
outweighed the benefits of achieving a high quality surface finish. Temperature fluctua-
tions and environmental effects resulted in vibrations in the bow, which gave the compo-
nents a rippled surface finish. Tapered components such as the inboard wings increased
the depth and frequency these ripples when each side of the hot-wire travelled at different
speeds. The surface finish achieved on many of the components was deemed unacceptable,
and as such, an alternative method of cutting the foam was required.
Manual hot-wire cutting
Manual hot-wire cutting was considered to be a suitable alternative to rig hot-wire cutting.
Both processes are similar, with the exception that the bow is manually moved through
the foam by hand, instead of relying on a pulley to move the bow. Minimal rippling in the
foam was produced through sufficient practice. Hence, the surface finish was considered
acceptable for the remaining components.
A photo of the manual hot-wire process is included in Appendix E
3D CNC machining
Due to the complex geometry of the fuselage, a plug was 3D CNC machined from a solid
block of extruded polystyrene foam. The CAD model of the aircraft was provided to an
external company, who performed the machining at their own facilities.
5.1.2 Composite layup
The main components of the Morpheus UAV were designed to have a composite skin.
The layup of composite material can be achieved by using the hand layup technique, or
by using pre-impregnated cloth. The hand layup technique involves laying the composite
material onto the component prior to applying the resin. Pre-impregnated cloth is pre-
saturated with resin, eliminating the need to apply the resin separately. However, an
autoclave is required if the pre-impregnated cloth method is to be used. The hand layup
technique was used for all components of the Morpheus UAV, as an adequate surface
finish can be achieved with the least amount of time and resources.
Prior to fibreglassing, all components were coated in a thin layer of epoxy resin to minimise
the porosity of the foam. This prevented excess resin seeping into the foam during the
fibreglassing process. Excess resin was removed during the fibreglassing process with a
squeegee to reduce the amount of resin on each component. This translates into significant
weight savings, as excess resin adds unnecessary weight.
The fibreglassing process was repeated for additional layers of fibreglass. A temperature
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5.2. COMMON COMPONENTS FOUND IN THE MORPHEUS UAV 110
control room was use to cure the components overnight. Overhanging fibreglass edges,
particularly at edges or corners, were trimmed with scissors or a stanley knife once the
resin had cured. All components were finished by coating the fibreglass with a thin layer
of resin, and sanding the resin once it had dried. This achieved a high quality surface
finish on all components.
A photo of the wet layup process is included in Appendix E.
5.2 Common components found in the Morpheus UAV
The Morpheus UAV has several subassemblies, many of which have similar components
to other subassemblies. The manufacturing methods for common components found in
the Morpheus UAV are outlined below.
5.2.1 Ribs
Ribs can be cut using either the laser cutting technique or CNC machining. CNC machin-
ing can produce high quality components, but with increased time and cost. Laser cutting
is undesirable for fixed depth cuts or thick sections due to the power of the laser, but has
high precision and a lower cost. The Mechanical Engineering workshop at the University
of Adelaide had a CNC machine capable of meeting all manufacturing requirements for
the aircraft, and as such, was utilised for the manufacture of all ribs.
5.2.2 Leading and trailing edges
All leading edges were initially designed to be balsa wood, so that an adequate leading
edge aerofoil shape could be obtained. However, the hot-wire cutting process resulted
in a flawless leading edge shape, so no balsa wood leading edge was required on any
component. All trailing edges for the wings and tail were initially intended to be foam.
During the manufacturing process, the foam trailing edges were easily damaged during
the handling and construction of the components. Hence, all foam trailing edges were
replaced with balsa wood trailing edges for additional strength and rigidity. The balsa
wood trailing edges were bonded directly to the foam with epoxy resin, and sanded to
shape as required.
5.2.3 Control surfaces
All control surfaces were initially designed to be solid balsa wood. However, it was
discovered that solid balsa wood control surfaces would increase the time of manufacture
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111 CHAPTER 5. MANUFACTURING
and the weight of the control surfaces. Hence, foam control surfaces were constructed by
removing material from the main component with a stanley knife. This resulted in the
control surfaces having the desired shape at a reduced weight. However, balsa wood hinge
lines, trailing edges and end caps were bonded to the control surfaces with epoxy resin
for additional strength and rigidity.
5.2.4 Carbon fibre components
All carbon fibre components were manually cut to length with a hacksaw. Rails and rail
guides for the morphing mechanisms were sanded by hand, until the components slid
freely over each other without resistance. All carbon fibre components were integrated
into the airframe with a mixture of epoxy resin and chopped carbon fibre. The chopped
carbon fibre was added to the epoxy resin to increase the strength of the interface between
the component and the carbon fibre.
5.3 Inboard wing construction
The inboard wings consisted of a foam core, with a plywood root rib, aluminium tip rib,
carbon fibre spars and a fibreglass skin. Figure 5.1 shows an assembly dt of the inboard
wing.
Figure 5.1: Inboard wing assembly drawing
5.3.1 Foam core
Manual hot-wire cutting was used to cut the foam core for the inboard wings. Due to the
taper of the inboard wings, two differing laminex templates were CNC machined. As the
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5.3. INBOARD WING CONSTRUCTION 112
foam was sourced in 50mm thick sheets, and the inboard wings were thicker than 50mm,
an upper wing half and a lower wing half were bonded together with epoxy resin to obtain
an appropriate foam thickness. Manual hot-wire cutting was used to cut the wings, as it
had already been discovered that rig hot-wire cutting did not produce an adequate surface
finish.
A cavity inside the inboard wing was required for the outboard wing to slide into during
the morphing process. This was achieved by threading the hot-wire through a small pilot
hole created prior to bonding the two wing halves together, and hot-wire cutting the
cavity to create a hollow inboard wing. The cavity was created after the wing had been
fibreglassed, as the wing lacked structural integrity at the wing tip prior to fibreglassing.
5.3.2 Ribs
The inner and outer root ribs were manufactured from plywood. The outer root rib was
bonded to the foam with epoxy resin and microballoons prior to fibreglassing, so that the
outer root rib could be fibreglassed into position. The inner root rib was bonded to the
outer root rib after fibreglassing was complete. Composite brackets for the aluminium tip
rib were bonded to the foam and fibreglass with epoxy resin, and were later fibreglassed
in place. The tip rib was fastened to the composite brackets with brass bolts.
5.3.3 Spars
The carbon fibre wing spars were cut to the appropriate length with a hacksaw, and sanded
until the rails in the sliding block slid freely over the spars. The spars were bonded to
the inner root rib with a mixture of epoxy resin and chopped carbon fibre.
5.3.4 Fibreglass skin
The inboard wing was skinned with three multidirectional layers of 85 gsm fibreglass. The
inboard wing tip remained fragile due to the cavity in the inboard wing. Hence, three
additional layers of 85 gsm fibreglass reinforcement were added to the inboard wing tips.
The result was a wing tip with increased strength and resistance to crushing.
A photo of the wing-tip reinforcement can be found in Appendix E.
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113 CHAPTER 5. MANUFACTURING
5.4 Outboard wing construction
The outboard wings and wing sliding blocks consisted of a foam core, plywood ribs, carbon
fibre spars and a fibreglass skin. Figure 5.2 shows an assembly drawing of an outboard
wing and wing sliding block.
Figure 5.2: Outboard wing and block assembly drawing
5.4.1 Foam core
The outboard wings and wing sliding blocks were the first components to be hot-wire
cut. As such, the rig hot-wire cutter was used. Due to the outboard wings and wing
sliding blocks having no taper, the rig provided an adequate surface finish. Laminex
templates used to cut the outboard wings incorporated a hole for the outboard wing spar
and four indentations for the roller strips. The laminex templates for the wing sliding
blocks incorporated holes for the outboard wing spars and rail guides.
5.4.2 Ribs
Plywood root ribs and plywood tip ribs for the outboard wings and wing sliding blocks
were CNC machined and bonded to the foam with a mixture of epoxy resin and microbal-
loons. A thin balsa wood rib was added to the outboard wing tip for aesthetics.
5.4.3 Carbon fibre components
The carbon fibre outboard wing spar, sliding block rail guides and roller strips were cut
to the appropriate length with a hacksaw. The outboard wing spar was bonded into the
outboard wing and sliding block with a mixture of epoxy resin and chopped carbon fibre.
Similarly, a mixture of epoxy resin and chopped carbon fibre was used to bond the rail
guides to the sliding block, and the roller strips to the outboard wing. The same mixture
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5.5. FUSELAGE CONSTRUCTION 114
was used to bond the aluminium insert for the threaded rod into its carbon fibre shroud,
and to bond the carbon fibre shroud into the sliding block.
5.4.4 Fibreglass skin
The outboard wings were skinned with three multidirectional layers of 85 gsm fibreglass,
while the wing sliding blocks were skinned with one layer of fibreglass. No additional
reinforcement was required for these components.
5.5 Fuselage construction
The construction of the fuselage consisted of the manufacture of a foam plug, fibreglass
skin and fuselage internal structure. Figure 5.3 shows an assembly drawing of the fuselage.
Figure 5.3: Fuselage assembly drawing
5.5.1 Plug
Due to the complex geometry of the fuselage, a plug was 3D CNC machined from a block of
extruded polystyrene foam. As the foam was sourced in 50mm thick sheets, several sheets
were bonded together with epoxy resin to create a solid block suitable for machining. The
final product was an exact replica of the fuselage created in CAD. However, some areas
of the plug were pitted due to the CNC machine ripping out areas of foam. These defects
were repaired by bonding the foam back into the plug with epoxy resin. The entire plug
was surfaced with a thin layer of epoxy resin to reduce the porosity of the foam.
A photo of the fuselage plug can be found in Appendix E.
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115 CHAPTER 5. MANUFACTURING
5.5.2 Skin
The fuselage plug can either be laid directly with composite material, or can be used to
create female moulds. Although the latter method allows easy replication of the fuselage
skin, the method is more complex, expensive and labour intensive. As such, the first
option was chosen for the fuselage skin, whereby the fuselage plug was directly laid with
composite material.
The plug was cut into two halves to create a starboard fuselage half and a port fuselage
half. The plug was prepared for fibreglassing by sanding the foam to a high quality surface
finish and applying wax and a PVA release agent. The application of these products
minimises the likelihood of the fibreglass bonding to the plug, as it was necessary for the
fibreglass to be later removed from the plug to create the skin. Three multidirectional
layers of 85 gsm fibreglass were laid onto the plug in a similar procedure to that outlined
previously.
A photo of the fuselage skin can be found in Appendix E.
5.5.3 Fuselage internal structure
Plywood formers, longerons and a landing gear mounting plate for the fuselage internal
structure were manually cut by hand using a jigsaw and scroll saw. Each component was
sanded to fit, and bonded into a fuselage internal structure with a mixture of epoxy resin
and microballoons.
A photo of the fuselage internal structure can be found in Appendix E.
5.6 Empennage construction
The empennage was constructed by cutting a foam core for the tail fins, fibreglassing the
tail fins and manufacturing a tail sliding block. Figure 5.4 shows an assembly drawing of
the empennage.
5.6.1 Foam core
The tapered tail fins were manually hot-wire cut from foam. Two different laminex profiles
were required to obtain the desired taper. The tail fins were bonded to the tail boom
with a mixture of epoxy resin and chopped carbon fibre. A simple rig was constructed to
ensure that the tail fins were bonded at the correct angle.
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5.7. AIRCRAFT ASSEMBLY 116
Figure 5.4: Empennage assembly drawing
5.6.2 Fibreglass skin
The tail fins were skinned with three multidirectional layers of 85 gsm fibreglass. The
connection between the tail fins and tail boom required reinforcement, as the strength
of the foam in this area was inadequate. Hence, a single layer of 300 gsm fibreglass was
added to the tail root on the upper and lower surface of the tail. The result was a tail
root with increased strength.
A photo of the reinforcement at the tail root can be found in Appendix E.
5.6.3 Tail sliding block
The tail sliding block was manually hot-wire cut from foam. The laminex templates used
for the hot-wire cutting incorporated holes for the tail boom and rail guides. Plywood
end caps were CNC machined and bonded to the foam with a mixture of epoxy resin
and microballoons. The carbon fibre rail guides were manually cut to the appropriate
length using a hacksaw, and bonded into the tail block with a mixture of epoxy resin and
chopped carbon fibre. The tail block was skinned with one layer of fibreglass. The sliding
block and aluminium inert for the threaded rod were bonded into the tail boom with a
mixture of epoxy resin and chopped carbon fibre.
5.7 Aircraft assembly
The assembly of the aircraft involved the installation of the fuselage internal structure,
outboard wings and wing sliding blocks, inboard wings, empennage, undercarriage and
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117 CHAPTER 5. MANUFACTURING
electronics. Figure 5.5 shows an assembly drawing of the entire aircraft.
Figure 5.5: Aircraft assembly drawing
5.7.1 Fuselage internal structure installation
The fuselage internal structure was bonded into the starboard fuselage half with epoxy
resin and microballoons. Hatches were cut and removed from the fibreglass skin prior to
the port fuselage half being bonded to the fuselage internal structure. The seam between
the two fuselage halves was covered with a strip of 300 gsm fibreglass. Throughout the
construction of the fuselage, plywood shelves were cut to the required geometry and
integrated into the structure with epoxy resin. These shelves were later used for the
mounting of electronic components such as batteries, speed controllers and receivers. The
motor cowling at the front of the fuselage and the fairing at the rear of the fuselage
were removed from the fuselage skin with a hacksaw and a stanley knife. Balsa wood
and hardwood mounts for the hatches, cowling and fairing were bonded into the fuselage
internal structure with epoxy resin.
A photo of the fuselage internal structure being bonded to the skin is included in Appendix
E.
5.7.2 Outboard wing and wing sliding block installation
The outboard wings and wing sliding blocks were installed by sliding the rail guides of
the sliding block onto the wing spars. The motor mounts were bolted to the inboard wing
root ribs with hex head bolts and blind nuts. The threaded rod was screwed into the
aluminium inserts in the wing sliding blocks. The aluminium tip ribs were bolted onto
the inboard wings.
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5.7. AIRCRAFT ASSEMBLY 118
A photo of the outboard wing and wing sliding block installation is included in Appendix
E.
5.7.3 Inboard wing installation
The wing tongues were cut to the appropriate length with a hacksaw. The inner wing
tongue was bonded into the leading edge wing spar of the port inboard wing with epoxy
resin and chopped carbon fibre. The outer wing tongues were bonded to the wing tongue
brackets with epoxy resin and chopped carbon fibre. The inboard wings were installed
into the fuselage by sliding the inner wing tongue into the outer wing tongues and aligning
the wings spars with the wing tongues. The incidence angle of the wings was adjusted
until the desired angle was achieved.
A photo of the inboard wing installation is included in Appendix E.
5.7.4 Empennage installation
The tail rails were cut to the appropriate length with a hacksaw. These were then inserted
into the formers inside the fuselage. The tail rails can be easily removed to allow easy
access to the fuselage internals. The tail was installed by sliding the rail guides of the
sliding block onto the tail rails. The motor mount was bolted to a former with hex head
bolts and blind nuts. The threaded rod was screwed into the aluminium insert in the tail
boom. The fairing was bolted onto the rear of the fuselage with hex head bolts and blind
nuts.
A photo of the empennage installation is included in Appendix E.
5.7.5 Undercarriage installation
The undercarriage installation involved the installation of the nose landing gear and the
main landing gear into the fuselage.
Nose landing gear
The nose gear was purchased as an off-the-shelf component. The main strut of the nose
landing gear had to be bent to form an axle for the nose wheel. The nose wheel was
attached to the nose gear strut with a brass collar and grub screw, and the nose gear
strut was fastened to a former via a nylon bracket. A servo arm was attached to the nose
gear with a grub screw, and this was linked to a servo via a pushrod. This allowed the
aircraft to have a steerable nose gear for taxiing on the ground.
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119 CHAPTER 5. MANUFACTURING
A photo of the nose gear installation is included in Appendix E.
Main landing gear
The optimal main landing gear location was determined through centre of gravity calcu-
lations. The main landing gear was bolted to a plywood block with 14” nylon bolts and
blind nuts. The plywood block was then bolted to the landing gear mounting plate within
the fuselage with 14” nylon bolts and blind nuts. This allowed the main landing gear to
shear off during a heavy landing, preventing damage to both the main landing gear and
the aircraft. The nylon bolts are easily replaced if damaged or broken.
A photo of the main landing gear installation is included in Appendix E.
5.8 Electronics installation
The electronics installation involved the installation of electrical and electronics com-
ponents, such as the electronics for the propulsion system, morphing system and radio
control system.
5.8.1 Propulsion system installation
The propulsion system installation involved the installation of the thrust motor, ESC and
Li-Po batteries.
Motor
The thrust motor was bolted onto the nose former with bolts and blind nuts, to allow
for easy removal of the motor if required. Once the motor was bolted to the former, the
cowling was attached with four screws, and the propeller and spinner were bolted on using
the washer and nut supplied by the motor manufacturer.
ESC
The ESC was mounted onto a plywood shelf, which was subsequently mounted onto the
longerons with screws. The plywood shelf is easily removed when required. Due to the
limited space at the nose of the aircraft, the ESC shelf is often removed to install the
batteries.
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5.8. ELECTRONICS INSTALLATION 120
Batteries
The Li-Po batteries were cable tied onto a plywood shelf and slid into the battery bay.
The plywood shelf was secured using wing nuts, and the battery bay hatch was secured.
The batteries are easily removed for charging or replacement when required.
A photo of the propulsion system installation is included in Appendix E.
5.8.2 Morphing system installation
The morphing system installation involved the installation of the morphing motors, ESCs,
PCB, limit switches and battery.
Motors
The morphing motors were screwed onto the gearboxes and motor mounts. The shaft
couplers were bonded to the gearbox output shafts with epoxy resin to prevent the shaft
couplers coming loose under loading, and the threaded rods were attached to the shaft
couplers with grub screws. Loctite was used to prevent the grub screws in the shaft
couplers coming loose under loading.
ESCs
The ESCs for the wing morphing motors were mounted on the landing gear mounting
plate inside the fuselage. These were cabled tied in position to prevent movement during
transport or flight. Extension leads were required between the ESCs, the morphing re-
ceiver, and the morphing battery. The morphing Li-Po battery powering the ESCs was
cable tied to a shelf at the rear of the aircraft to prevent movement during transport or
flight.
PCB
The PCB was mounted on the landing gear mounting plate inside the fuselage, next to
the ESCs and the morphing motors. The PCB was placed in this location, as it was easily
accessible from the hatch directly above the landing gear mounting plate. The PCB was
cabled tied in position to prevent movement during transport or flight.
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121 CHAPTER 5. MANUFACTURING
Limit switches
Several limit switches were mounted in the aircraft to detect the limits of the extension
and retraction of the wings and tail. The limit switches for the wings were mounted
on the root ribs and tip ribs, while the limit switches for the tail were mounted on the
fuselage formers. The limit switches had to be mounted in such a way that the sliding
blocks would contact them during the morphing process. Once the optimal position of
each limit switch was determined from morphing tests, the limit switch mounting plates
were bonded in position with epoxy resin.
A photo of the morphing system installation is included in Appendix E.
5.8.3 Radio control system installation
The radio control system installation involved the installation of the servos and receivers.
Servos
A hot paint scraper heated in a Bunsen burner was used to cut the servo cavities in the
wings and tail, and a hot rod heated in a Bunsen burner was used to cut the holes for
the servo leads. Hardwood blocks for hatch mounting were bonded into the servo cavities
with epoxy resin, and 3mm thick plywood hatches were made to fit the cavities. The
aileron servos and ruddervator servos were screwed onto the servo hatches with screws.
These hatches were then screwed into the wings and tail. The servo arms were linked to
a horn via a pushrod. The horns were screwed into the control surface. A clevis was used
so that adjustments to the length of the pushrods can occur if desired. The nose gear
servo was screwed directly onto the nose gear former with screws, and linked to the nose
gear via a pushrod.
Receivers
Both the morphing receiver and the flight receiver were cable tied to shelves on opposite
sides of the fuselage internal structure at the rear of the aircraft. This was to allow free
movement of the tail block within the fuselage during the morphing process. The flight
receiver required its own Ni-MH battery pack, which was also cable tied to a shelf at the
rear of the fuselage. All antennas for the receivers were exited the fuselage through holes
to minimise RF interference.
A photo of the radio system installation is included in Appendix E.
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5.9. PAINTING AND FINISHING 122
5.9 Painting and finishing
The aircraft was painted and finished for improved flight performance and aesthetics. The
methods for painting and finishing the aircraft are outlined below.
5.9.1 Two-pack paint
Most aircraft components were painted with two-pack paint, as it is lightweight, provides
a high quality surface finish and usually only requires one coat. Other types of paint are
heavy and provide an inferior surface finish.
5.9.2 Solartrim
To protect the outboard wings from being damaged by the rollers during the morphing
process, it was decided to use Solartrim to cover the outboard wings. Solartrim is a self-
adhesive covering that can easily be applied and replaced if required. It is a cheap, time
effective, lightweight method of covering smaller components, and provides a high quality
surface finish. Solartrim was also used for all blue trimming on the aircraft, as it was
possible to achieve a variety of complex shapes in a minimal amount of time.
5.10 The completed Morpheus UAV
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
6. TestingThe tests were divided into two main sections: component testing and flight testing.
Component tests were performed to ensure that all mechanisms and components were
flight worthy, prior to the aircraft being flown. Flight tests were performed to achieve
project goals, and to demonstrate that the Morpheus UAV could successfully achieve
morphing flight. Each of the tests and their results are briefly outlined in the following
sections. Complete test procedures, results and the associated safety documents can be
viewed in Appendices I and I.
6.1 Component tests
Component tests were conducted to test the major systems and critical components of
the Morpheus UAV to minimise failures during flight. Wherever possible, systems and
assemblies such at the propulsion system, morphing mechanisms and wings were tested
as a whole, rather that as a series of smaller tests on individual components. Several
unofficial tests were also conducted as part of the manufacturing and assembly processes,
to ensure that the components were working as required, prior to installation in the
airframe.
6.1.1 Propulsion - Static Thrust Test
A static thrust analysis was performed on the selected propulsion system to investigate it’s
performance and verify that the components were capable of providing sufficient thrust.
The tests were conducted in the Turbine Propulsion Laboratory at the University of
Adelaide on a custom test rig capable of measuring the thrust using a load cell. A
photograph of the test rig can be seen in Figure 6.1. Tests were performed using Li-Po
batteries as a power source and conducted for two different propellers (a 16” x 8 ” and a
16” x 12”) to determine the optimal propeller sizing.
The test results indicated that the propulsion system, when using a 16” x 8” propeller
was capable of producing 8.5 kg of thrust, corresponding to a power output of 1398W.
When using a 16” x 12” propeller, a similar thrust was produced, but a higher power
output of 1644W was required to achieve this. The results from both tests was consistent
with the theoretical results, and can be viewed in Figure 6.2.
For a 7.5kg aircraft, the maximum power required for flight was calculated to be 1095W.
123
6.1. COMPONENT TESTS 124
Figure 6.1: Static thrust set-up
Figure 6.2: Thrust vs. power for the 16”x 8” and 16”x 12” propellers, and the corre-sponding theoretical curve.
Thus, the selected propulsions system is capable of providing adequate thrust using either
propeller. The 16” x 8” propeller was selected for flight as it produced the required thrust
for flight, but at a reduced power consumption than the other propeller. This allows the
endurance of the aircraft to be maximised.
6.1.2 Morphing Mechanism
The morphing mechanism was tested at multiple stages during the design, assembly and
flight testing phases of the project. The primary morphing test was undertaken to de-
termine the required motor power. Preliminary calculations were inaccurate as increased
loads due to misalignment, caused the initial morphing motor to burn out during testing.
The test was modified to utilise a larger motor and a variable power pack to determine
the motor power required. To perform the mechanism test, the mechanism, including
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125 CHAPTER 6. TESTING
motor, threaded rod, spars, and outboard wing, was clamped to a table, and the motor
power slowly increased until the wing demonstrated a constant rate of movement along
the rails.
To determine that the wing was capable of morphing under load, the wing block was
incrementally loaded with sandbags to 3.8G, and the morphing test re-conducted. This
demonstrated that the mechanism could operate under the required manoeuvring loads.
The voltage and current were recorded for each load condition, allowing the power required
by the morphing motor to be determined. These morphing tests concluded that the
maximum power required was 33.25W. The complete results from this test can be viewed
in Appendix F.
Once the final motor, ESC and battery were been selected, a similar test was performed
with the wing under load. The speed of the morphing was maintained at 8 seconds per
300mm of translation. The mechanism was successfully was tested up to the 3.8G (6.95kg)
loading. This morphing mechanism test can be viewed in Figure 6.3..
Figure 6.3: Second morphing test with 3.8G (6.95kg) loading
Several challenges were encountered when performing the morphing mechanism tests. The
primary issue which occurred was misalignment of the alignment of the mechanism rails,
sliding block and motor, which resulted in the mechanism to jamming. This highlighted
the importance of alignment during the manufacturing and assembly of the wing. To
ensure this was not a problem for the completed wing, the mechanism was tested after
the final wing assembly. To further reduce the effects of both alignment and friction, the
rails and threaded rod were lubricated with synthetic grease.
Another issue encountered during the mechanism tests was that the large starting torque
of the motor often resulted in the shaft coupler to detach itself from the threaded rod.
To prevent this from occurring in the final wing assembly, the shaft coupler was bonded
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6.1. COMPONENT TESTS 126
to the threaded rod.
The final issue identified during these tests was significant vibration in the spars, motor
and threaded rod. Investigation into this vibration indicated that the majority of the
vibration experienced during the initial testing was amplified by the excess length (almost
double the final size) of the spars and threaded rod. Foam packing was used to shorten the
effective length of the spars, and to damp the vibrations from the motor and threaded rod.
Final tests performed on the fully assembled morphing mechanism did not demonstrate
sufficient vibrations to warrant additional damping.
6.1.3 Wing - Structural Test
To ensure the wings and wing tongues were capable of carrying the required flight loads,
a structural load test was performed on the assembled wing unit. To achieve the desired
load distribution, the wing was divided into ten sections (six on the inboard wing and
four on the outboard wing), and a piecewise load distribution was produced. The wing
was incrementally tested up to 3G, which was sufficient loading to simulate normal flight
and heavy landing. The distributions can be seen in Table 6.1.
Table 6.1: Piecewise wing load distribution up to 3G total loadDistance from root (m) 75% load (g) 100% load (g)
0.042 450 6700.126 480 7100.210 500 7300.294 520 7500.378 520 7400.462 500 7100.546 650 8900.630 570 7800.714 430 5900.784 240 350
To increase the accuracy of the test, the rollers (not yet assembled) were simulated using
shims placed between the inboard wing tip and outboard wing. The load consisted of
sandbags of various weights placed on the underside of the inverted wings to simulate an
upwards lift force, a method commonly used in industry. This test is shown in Figure 6.4.
For each loading scenario, the outboard wing tip deflection was measured. The wings were
also inspected for any damage sustained during the test. This data is presented in Table
6.2. The measured deflections were lower than the theoretically calculated deflection.
This was expected as the theoretical wing deflection assumed that the spars carried the
entire loading, and ignored the effects of the skin and foam.
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127 CHAPTER 6. TESTING
Figure 6.4: Wing structural test
Table 6.2: Wing deflection under loadLoad Case Port Wing (mm) Starboard Wing (mm) Theoretical (mm)
60% 44 45 47.680% 58 56 63.4
6.1.4 Assembled Electronics, Morphing and Control Systems
Prior to deeming the aircraft flight worthy, the electronics, morphing and control systems
were tested to ensure that they were operating as required. The majority of the electronics
were tested simply by operating the morphing and control systems. The thrust and
morphing switches were also tested individually to ensure that there were no operational
errors. During these tests, the control surface deflection was measured, and the results
can be viewed in Section 6.2.4.
During these tests, several issues were identified. One immediately obvious issue was a
grinding noise heard in the nose gear servo during operation. Upon inspection of the
servo, it was discovered that dust from the manufacturing process was preventing the
servo from operating sufficiently, requiring the servo to be replaced. A second problem
encountered was chatter in the ruddervator servos when no control inputs were given.
The chatter was stopped by removing the mechanical v-tail mixer and utilising electronic
mixing on the transmitter instead.
6.2 Flight testing
The flight tests and associated ground tests were carried out over four days at three
different airfields. The procedures for each test can be found in Appendix I. In addition
to the flight tests, pre-flight were required prior to ch flight to ensure that the aircraft
as flightworthy. These pre-flight checks can be viewed in Appendix I. The initial three
flight tests were conducted in the presence of a heavy model inspector, as the Morpheus
aircraft required heavy model certification to fly.
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6.2. FLIGHT TESTING 128
6.2.1 Heavy model certification
As the Morpheus UAV had a dry weight exceeding 7kg (total of 8.2kg) when the flight
batteries and morphing mechanism were installed, the UAV needed to be certified as a
heavy model by a MASA approved Heavy Model Inspector. Certification required that
the inspector check the aircraft for structural integrity, and check that it is capable of
stable flight. The inspector was also required to be present for the pre-flight checks.
Heavy model certification of the Morpheus UAV was achieved at the Constellation Model
Flying Club, and a three-year licence was issued. This is included in Appendix H.
6.2.2 Balance & stability
Prior to each test flight the longitudinal stability of the Morpheus UAV and the lateral
balance was determined to ensure stable flight. For each test flight the aircraft was found
to be laterally balanced to within ±5mm of the aircraft centreline. The static margin for
the first test flight was 16.5% and 10% for the second test flight.
The reduction in static margin was the result of repairs from the first crash and a request
from the pilot to reduce the weight on the nose gear. The third test flight was also con-
ducted with a 10% static margin. The longitudinal stability for each flight configuration
from the morphing flight is given in table 6.3.
Wing configuration Tail configuration Static marginExtended Extended 10.3Retracted Extended 20.4Retracted Retracted 15.9Extended Retracted 10.8
Table 6.3: Static margin for each configuration obtained during the morphing test
6.2.3 Ground test - range checks
Prior to flight tests, a range check was performed to determine the maximum range
between the transmitter and the receiver which maintained full control of the control
surfaces and morphing mechanisms. These tests were performed both with and without
the thrust motor in operation, and with the transmitter antenna both retracted and
extended. The morphing mechanism range test as only conducted when in-flight morphing
was to form part of test. The results of these tests can be viewed in Appendix F
The range check conducted prior to the first test indicated minimal range. The range
achieved was however still sufficient to be approved by the heavy model inspector. When
this rage check was conduced prior to the second attempted flight, this range was deemed
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
129 CHAPTER 6. TESTING
insufficient. The size the whip antennae and interference between e receiver and ESC
were deemed to be the cause. To solve this problem, the morphing mechanism receiver
was exchanged for the flight receiver, as the Spektrum system operates at a much higher
frequency than the ESC signals. Interference between the ESC and the 3.6 MHz radio
control system now utilised for the morphing system was still a possibility, however, this
was deemed an acceptable risk, as a failure of the morphing system during flight would
not directly cause the aircraft to crash. This seemed to resolve all range issues, as no
further problems were identified during the third and fourth test flights.
6.2.4 Ground handling tests
The ground handling tests consisted of a series of three tests to ensure successful flight.
These included control surface tests, a taxi test and ground run tests.
Control surface tests
Prior to each flight, the control surfaces and nose gear were checked to ensure adequate
movement for flight control, takeoff and landing. The deflections were measured and
changed if required. The results of these tests can be seen in Table 6.4.
Table 6.4: Ground handling - Control surfacesComponent Deflection - low rates (mm) Deflection - high rates (mm)Left elevator 24, -19 26, -21
Right elevator 24, -16 29, -21Port aileron 20, -18 29, -28
Starboard aileron 19,-20 28-31Left rudder 23, -16
Right rudder 16, -24Starboard flap -19 -19
Port flap -17 -17Original nose gear 46 (Port) 38 (Stbd)
New nose gear 40 (Port) 39 (Stbd)
Taxi test
The second ground handling test was a taxi test, to ensure that the thrust motor control
surfaces and nose gear were operating as expected, and that there was adequate nose gear
authority. This was achieved through a series of ground manoeuvres. On the initial flight
test, it was noted that the nose gear movement was too great, and this was subsequently
limited o prevent incidents on takeoff and landing. These results can be viewed in Table
6.4.
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6.2. FLIGHT TESTING 130
Ground run test
Several ground runs were performed prior to each flight to allow the pilot to become
acquainted with the UAV, and to ensure the rudder and nose gear were sufficient for
tracking on the runway. The test also assessed the longitudinal stability and the elevator
pitching response of the UAV. In these tests, the UAV was positioned at the start of the
runway and run at half throttle down the field whilst the ruddervators were operated to
determine the response of the aircraft. Further runs were made with increasing speed
until takeoff speed was approached. The results of these tests demonstrated sufficient
rudder and nose gear authority, and elevator. In all cases, the pilots commented that the
aircraft had excellent ground handling qualities.
6.2.5 Stability Test
This test involved takeoff, a short, straight-line flight just above the ground, and a landing.
This task proved to be very difficult for both pilots, and placed the aircraft in increased
danger. This test was therefore abandoned after the first two flight attempts.
Attempt 1
The first flight test of the Morpheus UAV was plagued by propulsion system problems,
resulting in a significant crash after the completion of only half a circuit. The crash oc-
curred when the aircraft lost all thrust, hence resulting in stall, and the subsequent crash.
Prior to this occurring, the pilot also experienced fluctuations in the thrust, independent
of any pilot input. The decision w therefore made to land the aircraft, however before
this could occur, the UAV lost all thrust, resulting in the crash.
Significant damage was sustained by the UAV, requiring the front half of the fuselage to
be remanufactured. This resulted in an increase in a 367g increase in aircraft weight due
to the repairs. The opportunity was also taken to improve access to the batteries. The
modifications to the nose of the UAV also resulted in a removable nose section, allowing
easier maintenance and access to the nose electronics. Minor repairs were also required
to the wings, and tail block. Component damage pictures can be seen in Appendix F.
The aircraft took 12 days to be completely repaired and fully tested ready for the second
test flight. A photo of the crashed aircraft is included in Figure 6.5, and one of the GPS
outputs showing altitude and ground speed over time is included in Figure 6.6
investigation and analysis The cause of the loss in thrust and thrust fluctuations
was determined to be a result of RF interference between the ESC and the 36 MHz radio
control used for the control surfaces. This resulted in the thrust surging between full
throttle and idle as the radio signal dropped in and out of range. Due to the fail safe
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
131 CHAPTER 6. TESTING
Figure 6.5: Attempt 1
Figure 6.6: Attempt 1 - GPS output
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6.2. FLIGHT TESTING 132
settings of the receiver, when the aircraft looses contact with the transmitter, it returns
to idle. As a result, the aircraft lost flight speed, and tip stalled due to a large bank
angle, and spun into the ground. The problem has since been overcome by using a 2.4
GHz radio control system for the control surfaces, which does not interfere with the ESC.
Initial analysis of the crash, lead the group to believe that the error was possibly the
result of a timing problem between the ESC and the thrust motor. Investigations into the
reliability of the propulsion system included wind tunnel testing of the entire propulsion
system to simulate an airspeed of up to 90kph. The test procedures for this test can be
viewed in Appendix F. During the wind tunnel tests, a spark was seen emanating from the
motor. It was found that the motor had overheated and melted the cables, drawing them
into the motor casing. The most likely reason for the motor overheating was a timing
error between the ESC and the motor, again indicating an error in the ESC. As a result
of tis, the decision was made to conduct the following test utilising the propulsion system
(motor, ESC and propeller) from a previous UAV project.
Attempt 2
The second attempt at the stability test resulted in a very short flight. However, just
after takeoff, a loud screeching sound was heard from the motor, and the UAV lost all
power. The pilot managed to flare the UAV containing the damage to the landing gear,
and smoke was observed from the motor. As the damage was contained toe landing gear,
this was quickly fixed in 2 days. A photo of the aircraft after attempt two is included in
Figure 6.7, and one of the GPS outputs showing altitude and ground speed over time is
included in Figure 6.8
Figure 6.7: Attempt 2
Investigation and analysis
Tests on the propulsion system with a new motor, determined that the second crash
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
133 CHAPTER 6. TESTING
Figure 6.8: Attempt 2 - GPS output
was caused by inappropriate settings on the ESC. The ESC default settings were not
adequate for the aircraft set-up, and as such, and the ESC was incorrectly detecting a
reduced number of battery cells (8 instead of 10). This caused the ESC to provide a 70
amp current, which is much higher than the motor’s rated continuous current of 40A. The
motor subsequently burnt out just after takeoff.
This problem was overcome by using a new ESC which automatically detects the correct
number of cells, and limiting the throttle settings on the transmitter so that the motor
does not experience more than 40A current draw at full throttle. The original propeller
was also utilised, as another possibility for the crash is that the aircraft was ’over propped’.
The propulsion system was thoroughly tested with static thrust tests, which maintained
maximum thrust for approximately 7 minutes until the batteries approached the safe
discharge voltage. The test procedure for these tests can be viewed in Appendix F.
6.2.6 Airworthiness test
The airworthiness test was performed to show that the UAV was capable of stable flight
and a conventional takeoff and landing. This was achieved on the third flight attempt.
The airworthiness test involved takeoff, climb, trim circuits, pitch and roll response, a
short cruise and flutter test, loiter, landing approaches and a landing. The flight was
performed with a 156g GPS as payload which, along with the 367g of weight added in
the crash repairs, gave us a total payload weight of 523g. During the flight, the UAV was
trimmed, the longitudinal and lateral stability of the aircraft was evaluated, and the UAV
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6.2. FLIGHT TESTING 134
endurance was determined.
The overall test was a success, and the pilot commented that the aircraft had excellent
stability, excellent control authority, and was a pleasure to fly. Photos of the test, including
takeoff, flight and landing are included in 6.9.
Figure 6.9: Airworthiness test flight images
(a) Take off (b) Flight (c) Landing
6.2.7 Morphing test
The Morphing Test was completed on the fourth flight attempt. This test aimed to ensure
a 50% change in wing span and a change in tail position was possible during flight. In
addition to this, it was the first opportunity to gauge the UAV’s stability in the retracted
configuration. The test involved takeoff, and trim followed by wing and tail morphing
before landing. Wing morphing included half retracting, and fully retracting the wings
bore returning to the extended configuration. Tail morphing involved retracting the tail
to its 2/3 position and fully retracted position. The wings were then morphed whilst
the tail was retracted, to demonstrate the UAV’s fully retracted configuration. The UAV
wings and tail were then fully extended before landing.
To ensure that the UAV was morphing as expected, the morphing was performed on the
straightest, closest part of the circuit to allow for visual confirmation. This also allowed
the morphing to be performed without banking during the transition. Due to the large
number of morphing circuits required, specific speed parameters such as cruise and loiter
were unable to be obtained. The full flight consisted of 15 circuits over a time of 7minutes
and 31 seconds and covered 14 km distance during the flight. Ground tests indicated
that to fully extend or retract the wings took approximately nine seconds, however the
starboard wing was slightly faster and would take seven seconds to fully retract. The tail
was able to extend or retract in seven seconds.
In all configurations the UAV performed well, was stable and had good control authority.
The pilot observed that the UAV performed differently in the various configurations.
When the wings were fully retracted greater roll control was achieved and when fully
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
135 CHAPTER 6. TESTING
extended greater lift was experienced. It is expected that morphing the tail did have an
effect of the UAV performance, however wind conditions on the day, and the inability to
test roll rates (as the aircraft was uncertified for these manoeuvres) meant that the effects
were hard to determine. Once the UAV was in the most retracted configuration, two
circuits of the field were performed, one at the morphing altitude and one at about 20m
above ground level; in both cases the roll and pitch response was found to be satisfactory
and the aircraft handled very well. Unfortunately, due to the limited battery life, cruise
and loiter velocities were unable to be attained in the retracted configuration.
The overall test was a success, demonstrating that the UAV could successfully morph the
wing span and tail position during flight. Photos of the test, including takeoff, flight and
landing are included in 6.10.
Figure 6.10: Morphing test flight images
(a) Flight (b) Half retracted (c) Fully retracted
6.2.8 Endurance test
This test aimed to measure the endurance of the UAV with 500g payload in the extended
configuration. The UAV was to be flown at loiter speed for 3 circuits, land, and with the
UAV secure continue to run the motor on the ground at the loiter throttle setting until
the ESC cut off the battery power.
Unfortunately due to weather conditions this test was not able to be performed before
October 30th. Theoretical estimates indicate that a loiter time of 36 minutes was achiev-
able. Static ground testing at full thrust indicated that the batteries used in the Morpheus
project provide approximately 7 minutes endurance. During loiter at 1.5Vstall the throttle
setting is significantly lower and the motor is under less load. It is possible that the aim
of 30 minutes endurance could have been met. Had this test been initially unsuccessful,
replacement batteries could have been sourced to provide increased flight time, and the
test re-run.
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6.3. EVALUATION OF AIRFRAME AND FLIGHT PERFORMANCE 136
6.2.9 Flight parameters
The performance parameters test aimed to measure the takeoff distance, maximum speed
(100% throttle), cruise speed (85% throttle) and loiter speed (65% throttle) in both the
fully extended and the retracted wings - extended tail configuration. Having the tail
extended ensured adequate handling during takeoff and landing. A comparison of these
parameters would then have been utilised to determine the effect of morphing the UAV.
Unfortunately this test required calm weather conditions which did not occur in the
available flight windows. This test was therefore not able to be performed before October
30th. An analysis of some performance parameters in different UAV configurations is
included in the discussion of the Morphing Test in Section 6.2.7
6.3 Evaluation of airframe and flight performance
Throughout the flight testing phase the aircraft and airframe performance was evaluated in
all flight/morphing configurations. An onboard GPS was used to record speed, position,
heading and altitude to assist with analysis. Bad wind conditions during the second
successful flight resulted in inconsistent and often erratic data. The GPS used was the
Garmin Vista HCx which contained a barometric altimeter, however due to its installation
in the fuselage, the results were unpredictable. Due to the weather conditions much of
the analysis was performed through observations made by the pilot and test coordinator
during the flight.
6.3.1 Flight Performance
Analysis of the flight testing data revealed that the Loiter speed of the UAV in extended
configuration was 97kph. This is significantly higher than expected. The main cause for
this is probably pilot reluctance to drop the speed too significantly and put the aircraft
at risk. In addition to this, a tail wind was present on the most constant phase of the
run, it is assumed that over a longer period of loiter the average speed would be much
lower. The Maximum speed of the extended configuration was measured to be 130kph,
taken at the maximum throttle setting and taking into account variations due to wind
speed. This is approximately 88
The Morphing Test was unable to provide any accurate numerical data whilst in-flight
due to the high wind speed of between 13 and 22kph and very high wind gusts of up
to 30kph which gave skewed and erroneous velocity data. Due to the small size of the
Morpheus UAV and the fact that the average speed was about 120kph the high wind
speeds and gusts and flight path altitude changes had more effect of the UAV velocity
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
137 CHAPTER 6. TESTING
than the morphing. As such the variations in speed between configurations is small (less
than the variations due to the wind) and no conclusive evidence could been determined.
To obtain more accurate results, a pitot tube should be installed to more accurately assess
the airspeed.
A possible source of error in these results is that as the batteries are depleted, the motor
power is reduced, thus the thrust and speed of the UAV is also reduced. This would skew
the accuracy of the results, depending on the time at which these measurements were
taken. Had time and weather conditions permitted the UAV would have been tested for
a full flight in each configuration, to provide a more reasonable comparison.
Takeoff distance and speed were also measured for the extended configuration. The data
revealed an average takeoff distance of 92.5m (99m and 88m on the first and second
successful flight respectively), and average takeoff speed of 70kph (73kph and 67kph on
the first and second successful flight respectively). The theoretically calculated takeoff
distance was determined to be 63.2m, significantly lower than the measured difference.
The variation in takeoff speeds and distances may be due to errors in the GPS (up to 5m
variance) but may also be due to the pilots increased confidence on the second day.
6.3.2 Stability and Controllability
The success of both flights demonstrated that a stable and controllable platform was
designed, manufactured and flight tested. The same pilot was used for both successful
flight tests, and reported that the aircraft was stable and controllable under all tested
flight conditions, speeds and morphed configurations. The pilot commented that the
aircraft required minimal initial trimming, and that once trimmed, the aircraft remained
in trim. During the investigation of the cruise and loiter speeds, no instabilities occurred,
and the aircraft remained controllable at all times. During the morphing tests, the aircraft
was flown in four different morphed configurations, and all configurations were stable and
controllable at all times during the flight.
According to the pilot, noticeable changes were observed in lateral stability when the
wings were morphed, with the retracted configuration being more manoeuvrable, and no
longitudinal stability affects were observed. No affects on the aircraft were observed when
the tail was morphed. With the wings and tail retracted, increased roll control authority
was observed, but no change in pitch control authority was observed.
6.3.3 Morphing Mechanism Performance
During static tests, the morphing mechanisms functioned as designed, successfully trans-
lating the wings and tail by the appropriate stroke, even under a full design loading of
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6.3. EVALUATION OF AIRFRAME AND FLIGHT PERFORMANCE 138
3.8G. The morphing mechanisms also performed flawlessly when they were utilised dur-
ing flight. The Morpheus UAV successfully morphed into four different configurations
within the same flight. The performance of the morphing mechanisms, both statically
and in-flight, exceeded all expectations.
6.3.4 RC System Performance
From the ground and flight tests, as well as advice from aeromodellers, it was determined
that 36MHz radio control systems interfere with the frequency that the ESC uses to oper-
ate. This can cause the ESC to send incorrect signals to the motor, causing fluctuations
in the throttle settings. Once the 36MHz radio control used for the control surfaces was
replaced with the 2.4 GHz radio control system, the problem was eliminated. As such, it
is strongly recommended that only 2.4 GHz radio control systems be used around large
electric motors and their associated ESCs and electronics.
Further research revealed that certain 2.4 GHz receivers are only suitable for park flying
model aircraft, which are of a small scale and meant to be flown close to the pilot. Only
with the advice of aeromodellers was it noticed that the 2.4 GHz receiver that was initially
being used was inappropriate for larger model aircraft. Hence, an alternative 2.4 GHz
receiver was sourced that was suitable for larger model aircraft. This resulted in two
successful flight tests, with no radio control issues detected.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
7. ManagementThe effective management of the project was an important contributor to the success-
ful completion of all project requirements as the management strategies ensured that all
project goals were met on time, on budget, and to a high standard. The project manage-
ment included the definition of the management structure and positions, risk management,
and resource management.
7.1 Management structure
The management structure was designed to ensure that all aspects of the project were
addressed and given sufficient consideration, whilst maintaining a flexible and informal
team environment. The management structure, shown in 7.1 consisted of six coordinators
and officers, each responsible for specific aspects of the project, as well as a Logistics Co-
ordinator and a Technical Coordinator who were responsible for the overall coordination
of the project. The titles of ‘coordinator’ and ‘officer’ were selected specifically to promote
unity and equality within the group, whilst indicating the type of responsibilities required.
The title of ‘coordinator’ was selected to reflect that these roles within the group were
to coordinate a particular aspect of the project, and not to take sole responsibility for it.
The title of ‘officer’ implied that the majority of such work was the sole responsibility of
the officer. Each role involved promoting the interests of the particular aspect by ensuring
that all related requirements were met and the interests of this aspect were considered
during all group decisions. This promoted flexibility within the group as it allowed all
group members to be involved in all aspects of the project to varying degrees. This was
important in a small group, as it allowed the team to work more efficiency on time critical
tasks by re-distributing labour as required. A detailed description of each coordinator
and officer role is included below.
7.1.1 Technical coordinator
The Technical Coordinator was responsible for ensuring that the final design of the aircraft
met all requirements outlined in the technical task. The Technical Coordinator was aware
of and responsible for coordinating and reviewing all technical aspects of the project. This
included coordinating all calculations, concept design and detail design, as well as ensuring
that all components of the final design could be integrated.
139
7.1. MANAGEMENT STRUCTURE 140
Figure 7.1: Management structure
7.1.2 Logistics coordinator
The Logistics Coordinator was responsible for coordinating all logistic aspects of the
project. This included arranging all group meetings, producing agendas and minutes;
maintaining the Gantt chart, identifying critical tasks and ensuring that all deliverables
were submitted before the deadlines; maintaining the project finances; ensuring that
tasks were equally distributed amongst the group and that resources were distributed
appropriately between different aspects of the project.
7.1.3 CAD officer
The CAD Officer was responsible for the generation and management of all CAD drawings.
No part or drawing was created without involvement from the CAD officer. This was
necessary to assist with the generation of the CAD assembly model, and to ensure that
manufacturing drawings represented the latest design. The responsibilities of this role
included managing all CAD files, identifying spatial problems associated with the design,
and ensuring that all parts could be assembled within the CAD environment. The CAD
Officer worked closely with the Technical Coordinator (to ensure that all parts for the
aircraft were cohesive) and the Manufacturing Coordinator (to produce the manufacturing
drawings).
7.1.4 Manufacturing coordinator
The Manufacturing Coordinator was responsible for arranging the construction of the air-
craft. This involved liaising with the workshop and any external manufacturers, as well
as ensuring that the final design was manufacturable. The Manufacturing Coordinator
worked closely with the CAD officer (during the production of the manufacturing draw-
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
141 CHAPTER 7. MANAGEMENT
ings), the Technical Coordinator (to ensure that the final design was manufacturable),
and the Procurements and Assemblies Coordinator (organising the requirements for as-
sembling the aircraft).
7.1.5 Procurements and assemblies coordinator
The Procurements and Assemblies Coordinator was responsible for the logistics aspect of
assembling the aircraft. This role included procuring all non-manufactured items of the
aircraft as well as determining the schedule for the procurement of parts and assembly
of the aircraft. The procurements and assemblies officer worked closely with the logis-
tics coordinator (scheduling and financial decisions), and particularly closely with the
manufacturing coordinator (organising the assemblyscheduale).
7.1.6 Quality assurance officer
The Quality Assurance Officer was responsible for ensuring that the manufacturing of
the aircraft was of a sufficiently high standard. This involved assessing the dimensional
accuracy of all components and assemblies, along with ensuring correct operation and
appropriate surface finishes. The Quality Assurance Officer worked closely with the Tech-
nical Coordinator and the CAD officer (to determine critical dimensions and operation
which require measurements), and with the Manufacturing Coordinator (to ensure all
items were assessed at appropriate times during the manufacturing phase).
7.1.7 Test coordinator
The Test Coordinator was responsible for coordinating all ground tests (including com-
ponent, structural and electronic tests) and flight tests throughout the project. The Test
Coordinator was responsible for making any spontaneous decisions that may be required
during theses tests. The Test Coordinator worked closely with the Technical Coordinator
(to determine which components required testing to validate the electrical, structural and
mechanical expectations of the aircraft), the Logistics Coordinator and the Procurements
and Assemblies Coordinator (to determine the test schedule), and with the Safety Officer
(to ensure that all safety requirements were satisfied).
7.1.8 Safety officer
The Safety Officer was responsible for ensuring that all safety documentation and re-
quirements were satisfied. During tests, the safety officer was required to provide the final
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
7.2. RISK MANAGEMENT 142
approval prior to beginning. The Safety Officer worked very closely with the Test Coor-
dinator and the Manufacturing Coordinator to ensure that all tests and manufacturing
were conducted in a safe manner.
7.2 Risk management
A risk management plan was developed a the beginning of the project to identify, analyse,
and develop strategies to reduce risks which could impact upon the project outcomes.
The risk management plan is included in Appendix J. The strategies developed in the
risk management plan included the use of a design review, a quality assurance officer,
and the manufacturing of spare components. These strategies were deemed necessary in
addition to the resource management strategies discussed in more detail in section 7.3 to
reduce project risks to an acceptable level.
7.3 Resource Management
Resource management addresses the management of the schedule, labour, and finances
of the project. The management of the resources was ultimately the responsibility of
the logistics coordinator, however input was required from all group members, with all
significant decisions made during group meetings.
7.3.1 Project meetings
Project meetings were held on a regular basis throughout the project as a means of report-
ing progress, discussing problems which had arisen, and for the allocation of resources.
There were three main types of meetings, each serving a specific purpose. Supervisor
meetings allowed an opportunity for the project supervisor to stay up to date with the
project progress, whilst providing a formal environment for the project group to gain
guidance and advice. Internal Allocation meetings generally followed supervisor meetings,
and were utilised to discuss the resource allocationsProgress meetings were generally held
when a large number of smaller, critical tasks were being conducted. These provided an
opportunity for the re-distribution of tasks to ensure all tasks were completed on time.
Meeting minutes were generally produced for all meetings, and the minutes from super-
visor meetings are included in Appendix K. These meetings were essential to the success
of the project as they allowed for open communication between all group members.
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143 CHAPTER 7. MANAGEMENT
7.3.2 Scheduling
There were three levels of scheduling within the project, global, secondary and tertiary.
Each level of scheduling was the responsibility of different group members, and served a
different purpose. Global scheduling was primarily concerned with meeting the deadlines
for the project deliverables and the completion of the different phases of the project. The
Logistics Coordinator was primarily responsible for this aspect of the time management.
The secondary level of time management was the responsibility of individual coordinators
and officers, and was concerned with the management of tasks associated with the specific
aspect of the project which the coordinator or officer was in charge. The third level of
time management was the distribution of tasks and setting of weekly deadlines. This was
the responsibility of the group as a whole and was determined during allocation meetings.
To ensure that the project deliverables were met, and to monitor the project progress and
compare this with the project schedule, milestones and deadlines were set by the group,
and discussed during group meetings. Major milestones were selected to provide a flexible
internal deadline prior to the actual deadline of the deliverables, ensuring that they were
completed on time. Deadlines were set by individual coordinators and the group during
meetings as a means to meeting milestones. Gantt charts (included in Appendix L were
also found to be an effective tool for the logistics coordinator to manage the overall project
schedule.
7.3.3 Labour
The labour distribution was determined during weekly allocation meetings, at which all
upcoming and ongoing tasks were discussed. The tasks were distributed based upon the
following considerations:
• Coordinator jurisdiction.
• Previous involvement/experience of each group member in the task, or in similar
tasks.
• The priority of the task
• The amount of work involved in completing the task
• The availability of each group member
Consideration of each of these points allowed group members to maintain a realistic
workload, whilst ensuring that tasks were distributed in a manner which promoted the
most timely and efficient completion of all tasks.
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7.3. RESOURCE MANAGEMENT 144
The hours contributed to the project by each group member were recorded, allowing a
labour cost to be calculated. Over the course of the project, the group spent a total of
8960 hours. Assuming a labour cost calculated at $26.00 per hour, this equates o a total
cost of $232,965. The distribution of these hours is shown graphically below in Figure
7.2. Graph (a) is a representation of the labour distribution between tasks throughout
the project. Graph (b) is a representation of the hours contributed by each individual
group member. Numerical data supporting these graphs can be found in Appendix M.
(a) Labour distribution between tasks throughout the project
(b) Hours contributed by individual group members
Figure 7.2: Labour distribution between tasks and members
7.3.4 Finances
The University of Adelaide provides each project with 40 workshop hours per person,
and $200 per person for approved expenses. Due to the nature and scope of the project,
this was insufficient and external sponsorship from industry was sought. This resulted in
a total budget of $6000, in addition to in-kind support and Workshop hours. Prior to
approaching potential sponsors, it was deemed necessary to brand the project to ensure
it was memorable. As a result of this, the project was informally named ‘The Morpheus
Project’ in December 2008. Approaching potential sponsors was achieved though appli-
cation forms, letters, e-mails, and phone calls. An example of the letter sent to interested
sponsors, and the brochure distributed during presentations to potential sponsors is shown
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
145 CHAPTER 7. MANAGEMENT
in Appendix N. When approaching potential interstate sponsors and large companies, the
business plan (Included in Appendix O) was also provided.
Many companies were approached, sixteen of which were sufficiently interested in the
project to warrant further discussions. The group was successful in obtaining sponsorship
both of an in-kind and financial nature from the following companies:
• Aeronautical Engineers Australia - Financial sponsorship to the value of $4000
• Australian Aerospace - arranged for and covered the cost of the CNC of the fuselage
plug, in addition to in-kind support
• Babcock Integrated Technology Australia - Financial sponsorship to the value of
$1000
To manage the project finances, a project budget was developed, to divide the spending
into several categories. This was updated periodically to ensure that the optimal budget
was maintained as more accurate cost estimates became available. This predicted budget
provided a method of ensuring that all aspects of the project were considered when dis-
tributing project funds. A graphical representation of the Actual spending distribution
is shown graphically in figure 7.3. It can be seen from this chart that the majority of the
project budget was spent on the airframe and mechanisms, and general manufacturing.
Figure 7.3: Usage of project funds
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
8. ConclusionDespite many setbacks and unexpected challenges, the majority of the primary goals
have been successfully achieved and significant was made toward the completion of the
extended goals. Possible future work has been identified, which could be undertaken
to improve upon the Morpheus UAV and further contribute to the understanding of
morphing technology.
8.1 Project Achievements
A series of primary and extended goals were defined at the commencement of the project
as a means of measuring success. Each of these goals and the level of completion is now
discussed individually.
8.1.1 Primary Project Goals
1. The UAV shall have a normal takeoff and landing method. The Morpheus
UAV demonstrated two successful takeoffs and landings from a grass landing strip.
The aircraft used a conventional tricycle landing gear and demonstrated an average
takeoff distance of 94m.
2. The UAV shall be capable of having a loiter time of at least 30 minutes.
At the time of writing this goal has been met theoretically but has not been tested
due to setbacks resulting from the first crash. Theoretical estimates indicate that
a loiter time of 36 minutes was achieveable, assuming that other flight operations
consumed 40% of the battery capacity. Static ground test at full thrust indicate
that the batteries used in the Morpheus project provide approximately 7 minutes
endurance. During loiter at 1.5Vstall the throttle setting is significantly lower and
the motor is under less load. It is possible that this goal may have been achievable.
Otherwise, replacement batteries could have been sourced to provide increased flight
time.
3. The UAV shall be capable of cruising within line of sight. This goal was
achieved for both the extended and retracted configurations of the aircraft. The
theoretical maximum speed was 147 km/h for extended configuration and 165.7
km/h for the retracted configuration. The actual recorded speeds were inconclusive
due to high winds and gusts.
146
147 CHAPTER 8. CONCLUSION
4. The UAV shall be capable of carrying a 500g payload. Repairs made to the
aircraft following the first flight test, in addition to the weight of the GPS, exceeded
the 500g payload weight by 356g. This was deemed sufficient payload to meet this
goal, as the repairs were not part of the original design, and the increased weight
was not associated with rectifying the cause of the failed flights. In the absence of
time constraints the fuselage could have been rebuilt and 500g of payload added
separately to the aircraft. By achieving successful takeoff, flight and landing the
repaired aircraft has demonstrated that the addition of a 500g payload to the initial
design would not prevent sustained flight.
5. The UAV shall morph the wing to achieve a wing span increase of at
least 50% of the original wing span during flight. The Morpheus UAV was
able to successfully demonstrate a 60% increase in wing span during flight, without
any loss of control. The pilot noted a change in the lateral stability of the aircraft,
resulting in a more manoeuvrable aircraft. Speed variations however could not be
determined due to weather conditions on the day. Ground testing of the mechanism
has shown that the aircraft is capable of morphing under 2.3G loading.
6. The UAV shall change the tail position to control the longitudinal stabil-
ity during flight. The Morpheus UAV also demonstrated the capability of moving
the tail location by 400mm during flight. The theoretically determined effect of this
was a 4.1% change in the static margin.
In addition to this, the aircraft demonstrated stable flight in four different configura-
tions, consisting of a combination wing and tail morphing, and transitions between these
configurations.
8.1.2 Extended project goals
1. To measure the performance of the aircraft in different configurations
during flight. This goal was partially achieved, with data being obtained on the
performance of the UAV in it’s the extended configuration. The data obtained
during the morphing flight, however, was inconclusive due to the effects of high
winds and gusts. Theoretically, it was expected that the aircraft would have the
flight characteristics as outlined in table 8.1.
2. To theoretically optimise the morphing parameters for a predetermined
mission. For a typical reconnaissance mission, the optimal wing and empennage
configurations were determined or, in the case where analysis beyond the scope of
the project was required optimal configurations were recommended. The phases
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
8.1. PROJECT ACHIEVEMENTS 148
Table 8.1: Morpheus UAV PerformanceParameter Retracted wings Extended wingsStall speed [km/h] 64.39 48.6Takeoff distance [m] 63.2 35.2Maximum speed [km/h] 165.7 147Endurance [minutes] 22 36Rate of climb [m/s] 12.1 13.3
of flight considered are listed with there optimal configuration in table 8.2. These
results represent the completion of this goal within the scope of the project.
Mission phase Wing configuration Tail configurationTakeoff Extended ExtendedClimb Extended ExtendedCruise Retracted ExtendedLoiter Extended RetractedDash Retracted RetractedLanding Extended Extended
Table 8.2: Optimal configurations for a reconaissance mission
3. To achieve roll control through differential span morphing. The net roll
moment produced by the full extension of one wing, with the other retracted, was
calculated to equivalent to a 6.1 degree aileron deflection. This result indicated that
roll control through differential morphing is theoretically feasible for the Morpheus
UAV, but is dependent on the response rate required by the pilot. To determine if
the response rate was sufficient, further flight testing would be required.
8.1.3 Additional achievements
A method of applying the matching diagram sizing method to morphing aircraft was
developed during the conceptual design phase of the project. This method enables a
morphing aircraft to be designed to meet the requirements of multiple met areas whilst
also including the limitations of the morphing method employed.
The Morpheus UAV was successfully certified by a qualified heavy model inspector at
the completion of the first flight test. This required the Morpheus UAV to meet several
MASA requirements to demonstrate that it was a flight worthy aircraft. Previous projects
at the University have not achieved heavy model certification.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
149 CHAPTER 8. CONCLUSION
8.2 Issues and setbacks
Throughout the project many issues arose, which needed to be resolved in order to com-
plete the project. The two main issue of concern were inaccuracies in estimates due to
insufficient experience and problems with the aircraft electronics.
The inexperience of the group in project planning, design and manufacturing resulted
in numerous underestimates in the schedule. Schedule problems began when the design
phase of the aircraft took significantly longer than expected due to the numerous itera-
tions required to ensure that the aircraft and morphing mechanisms could be sucessfully
integrated. The CAD phase of the project also took a significant amount of time due to
the inexperience the group with the schools chosen software, Pro-E. The manufacturing
phase was also more time consuming than expected due to complexities and high toler-
ances of the morphing mechanisms. This required many components to be glued ‘in-situ’,
thus preventing further work on the entire aircraft component for 12-24 hours at a time.
This resulted in a significant delays to the schedule. Despite the delays, sufficient time
to conduct all test flights was available, however the aircraft suffered as severe crashed
during the first test flight. This resulted in a two week delay. This prevented the group
from completing all the required test flights.
Underestimation was also a problem when estimating the weight of the aircraft. Due to
inexperience and unforseen additional components inaccurate estimates for each aircraft
sub-system were made during design, resulting in an aircraft which was greater than
the original design weight, and 1 kg greater that the maximum weight specified in the
technical task. As a result it was necessary to have the aircraft certified as a heavy model
aircraft.
The onboard electronics of the aircraft were responsible for the first two failed test flights.
The skillset of the group did not includeprevious experience with practical electronics. The
group was forced to rely on advice from various aero-modellers, which was often conflicting
and made troubleshooting potential issues difficult. The most likely cause for the first
failed test fight was interference problems between the receiver and transmitter due to
the use of an antenna which was too small. This was used on the aircraft upon advice
from an external third party and was not picked up by the group due to inexperience.
8.3 Future work and recommendations
Although the Morpheus project endeavoured to cover all aspects of the aircraft design,
there were invariably many areas which fell outside the project scope or were not com-
pleted due to time constraints. Possible future work on the Morpheus aircraft includes:
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
8.4. PROJECT SUMMARY 150
1. Conduct further flight tests to more extensively and conclusively determine the
effects of morphing on the Morpheus UAV flight performance. It is recommended
that more accurate and appropriate test equipment is used to perform this analysis.
2. The morphing mechanisms at present are operated by a co-pilot, who must judge
the rates and relative positions of the wings and tail. Although limit switches
are installed, these only ensure that the extended and retracted configurations are
repeatable. A future undergraduate mechatronics project could develop a more
user friendly interface between the co-pilot and the aircraft to allow greater control
during morphing or could automate the morphing process entirely.
3. There are many examples of over design on the Morpheus UAV, particularly in the
outboard wings, tail and fuselage internal structure. By modifying the design of
these components, the aircraft weight could be significantly reduced. It is expected
that had a built up structure been utilised, particularly for the wings, the aircraft
could have been significantly lighter.
4. The actual aerodynamic effects during the morphing process could be fully investi-
gated. This would require the use of a wind tunnel, and possibly CFD (computa-
tional fluid dynamics) analysis.
5. The step between the inboard and outboard wing could also be investigated and
optimised to reduce drag. This would possibly require the use of a wind tunnel or
CFD analysis.
8.4 Project summary
The 2009 Morpheus UAV project has successfully designed, built and flight tested a mor-
phing UAV. The UAV has been demonstrated to be capable of achieving a 60% wing span
increase and a 400 mm change in tail position under flight loads. The theoretical effects
of morphing have been investigated, but further flight tests are required determine the
actual morphing outcomes. The majority of primary goals have been achieved and the
remainder have been demonstrated as being theoretically achievable. Significant progress
has been made towards the completion of flight related extended goals. The 2009 Mor-
pheus UAV project has resulted in a flightworthy, stable aircraft which can be further
developed in subsequent years by future projects.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
Reference List
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Vehicles, J. U. A. & Targets 2002, Jane’s Information Group, <http://www.janes.com/>.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
A. Electronics subsystems
specification and design
A.1 Specifications of electronic components
A.1.1 Battery specifications
Batteries Thrust Morphing Receiver
Model Flight Power Evo-lite Flight Power Evo-lite JR Propo
Type Lithium Polymer Lithium Polymer Nickel Metal Hydride
Capacity (mAh) 5350 1700 1100
Voltage (V) 14.8 11.1 4.8
Weight (g) 585 143 134
Constant Current (A) 90 59 N/A
Burst Current (A) 149 N/A N/A
Discharge Rate 17C 35C N/A
A.1.2 Radio control specifications
Radio Control Main RC Morphing RC
Brand JR Propo Spektrum
Model X2610 DX7
Frequency 36 MHz 2.4 GHz
Modulation SPCM DSM2
No. of Channels 6 7
A.1.3 Motor and ESC specifications
Motors Thrust motor Morphing motor
Brand Dualsky Ultrafly
Model XM5060CA10 EZRun-25A-SL
RPM/Volt 305 5400
Maximum Efficiency Current (A) 40 N/A
Maximum Burst Current (A) 60 21
Power (W) 1650 160
153
A.2. WIRING DIAGRAM - THRUST SUBSYSTEM 154
ESCs Thrust ESC Morphing ESC
Brand Dualsky Hobbywing
Model DSXC9036HV EZRun-25A-SL
Number of cells 2 to 12 2 to 3
Volts (V) 7.4 to 44.4 5.6 to 12.4
Maximum Output (A) 90 25
Maximum Burst Ouput (A) 120 90
A.2 Wiring diagram - Thrust subsystem
A.3 Wiring diagram - Control surfaces subsystem
A.4 Wiring diagram - Morphing subsystem
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
B. Landing gear positioningThe main and nose landing gear were positioned to meet weight distribution, tipback angle
and rollover angle requirements. Landing gear were also positioned to utilise existing
mounting points on the fuselage internal structure to reduce weight. It was assumed
that the Morpheus UAV would takeoff in the extended tail configuration as this would
provide the greatest pitch control authority. Hence the landing gear was positioned for the
extended tail configuration only. The nose gear was positioned on the second former to
utilise existing structure. This gave a nose gear position of 90 mm from the aircraft nose.
Preliminary calculations showed that this would require a main landing gear position
between the wing spar formers. The main gear was then positioned using an iterative
process to meet the three requirements. Iteration was also required with the aircraft
centre of gravity as the position of the main gear had a significant effect on the aircraft
centre of gravity.
The main landing gear is required to carry between 80-90% of the aircraft weight (Brandt
et al. 2004). Insufficient weight on the main gear will either prevent rotation or require
excessive elevator during takeoff. Excessive elevator may result in an uncontrollable pitch
up moment immediately after takeoff. Excessive weight on the main gear will lead to poor
ground handling qualities as the nose gear will have insufficient traction. The percentage
of weight on the main gear was calculated using equation B.1 for the foremost and aftmost
flight centre of gravity of 0.475 m and 0.478 m from the nose.
%Wmg =xcg − xnose gear
xmain gear − xnose gear(B.1)
The tip-back criterion requires the angle between the main gear pivot point and the centre
of gravity to be greater than the smaller of the stall angle or the angle between the aftmost
point of the aircraft and the pivot point. Preliminary calculations indicated that the stall
angle of 15◦ would be the limiting criteria. The tip-back criterion ensures that the centre
of gravity remains forward of the pivot point as the aircraft rotates during takeoff and
that the aircraft does not rotate backwards and sit on its tail. The tip-back angle was
calculated using the values given in table B.1 and equation B.2.
θtip−back = arctan
(xmain gear − xcgpivot height
)(B.2)
The rollover criterion ensures that the aircraft does not rollover during ground turns. The
rollover angle must not exceed 63◦ (Brandt et al. 2004). The rollover angle was calculated
155
156
Table B.1: Parameters used for the tip-back angle calculationCGz [m] 0.04Wheel diameter (d)[m] 0.089Main gear height (h)[m] 0.195Pivot height (CGz+d/2+h) 0.2795
using a half-wheel base of 0.2475, which included half the wheel thickness. Equations B.3
to B.5 were used to calculate the rollover angle.
θnose−main = arctan
(half wheel base
xcg − xnose gear
)(B.3)
d = (xmain gear − xnose gear)sin(θnose−main) (B.4)
θrollover = arctan
(pivot height
d
)(B.5)
The weight distribution, tip-back angle and rollover angle were calculated for a variety
of positions. A main gear position of 0.565 m from the nose was determined to meet all
requirements. The weight distribution, tip-back angle and rollover angle for this main
gear position are given in table B.2
Table B.2: Landing gear positioning criteriaCriteria Foremost flight CG Aftmost flight CGMain gear % weight 81.1 81.7Tip-back angle 17.9◦ 17.3◦
Rollover angle 51.9◦ 51.9◦
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
C. Fuselage load calculationThe fuselage experiences inertia loads from all components and aerodynamic loads from
the wings and empannage. Aerodynamic loads generated by the fuselage itself has been
ignored to simplify the analysis. Fuselage loads were calculated for the three following
critical cases:
• Pull-up maneuver at maximum speed
• Full aileron roll
• Static thrust
A pull-up maneuver at maximum speed generates the greatest shear and bending loads
within the fuselage. A full aileron roll produces the maximum torsion throughout the
fuselage. The retracted wing, extended tail configuration was the critical case as this
resulted in the highest maximum speed and the greatest moment arm for empennage
loads. Static thrust, in which the empennage is held, produces the greatest axial load in
the fuselage.
C.1 Empennage aerodynamic loads
The maximum control surface deflection likely to be experienced is δ = 30◦ (Raymer 2006),
which will generate the maximum force during a pull-up. Downwash from the wings will
also affect the maximum force generated by the ruddervator deflection. This affect can be
determined from the effective angle of attack given by equation C.1. The effective angle
of attack was calculated to be −0.147◦ which gives an effective lift coefficient of -0.01.
αh = (α + iw)(1− δe
δα) + (ih − iw) (C.1)
The lift inciment due to a full ruddervator deflection may be calculated using equationC.2
(Raymer 2006). Kf is the correction factor for large deflections whilst ( δClδδf
) is the section
lift increment values for these parameters may be found in tables in Raymer (2006). The
tapered tail and unswept hinge line results in ruddervators which are not a constant chord
percentage. The chord percentage at the tail mean aerodynamic chord (28.5%) was used
as an approximation, giving values of Kf = 0.62 and ( δClδδf
) = 4.4. The total tail lift
increment, including downwash affects, was calcualted to be CLt = 1.1.
157
C.2. INERTIAL LOADS 158
∂CL∂δf
= 0.85× 0.9Kf (∂Cl∂δf
)SflappedSref
cos(ΛHL) (C.2)
The maximum downforce and sideforce generated by the tail are given in equations C.3
and C.4, where γ = 35◦ is the dihedral angle. The tail was calculated to be capable
of producing a total downwards force of 128 N during a pull-up maneuver and a total
sideforce of 88.7 N. Calculation of the total sideforce did not include the downwash effects
as the net effect on the sideforce would be zero.
LH = qStCLtcos(γ) (C.3)
LV = qStCLtsin(γ) (C.4)
C.2 Inertial loads
Inertia loads of the fuselage are due to the weight force of major components. The major
systems, weights and locations from the aircraft nose are given in table C.1. The centre
of gravity for this configuration is at 0.475 m and the total aircraft weight is 8 kg.
Table C.1: Aircraft weight breakdown summaryComponent Component CG [m] Mass [kg]Motor 0.0315 0.377ESC 0.107 0.125Batteries 0.223 1.295Fuselage structure 0.3743 2.043Payload 0.425 0.5Wings 0.539 2.41Main gear 0.565 0.345Morphing and receiver batteries 0.807 0.226Tail 1.253 0.654
C.3 Wing aerodynamic loads
The aerodynamic loads on the wings are due to lift and the wing pitching moment. The
lift generated by the wings is given by equation C.5, where n is the load factor of 3.8
and LH is the tail downforce. The total wing lift was calculated to be 425.1 N. The
wing pitching moment is calculated from the moment coefficient in equation C.6 and was
determined to be -14 Nm. This load and moment are applied to the shear and bending
moment diagrams.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
159 APPENDIX C. FUSELAGE LOAD CALCULATION
Lw = ngm+ LH (C.5)
M = qScCm (C.6)
C.4 Shear and bending moment diagrams
The shear and bending moment diagrams for the fuselage and boom are given in figures
C.1 and C.2. These diagrams are used in the stress analysis of the fuselage and empennage.
Figure C.1: Fuselage and tail boom shear diagram
C.5 Full aileron roll torsion
The torsion due to a full aileron roll was calculated by determining the difference in lift
generated by each wing at maximum speed. The lift-increment resulting from aileron
deflection was calculated using equation C.7. The maximum aileron deflection expected
for the Morpheus UAV was δ = 25◦. The inboard wing ailerons were not designed to
be a constant percentage of the chord and hence the aileron was analysed as a series
of strips. The aileron chord percentage for the strips varied between 14.7% and 16.7%,
giving values for the section lift increment of between 3.2 and 3.4. For a δ = 25◦ deflection
the correction factor Kf ≈ 0.78 for each of the strips.
∂CL∂δf
= 0.85× 0.9Kf (∂Cl∂δf
)SflappedSref
cos(ΛHL) (C.7)
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
C.6. STATIC THRUST 160
Figure C.2: Bending moment diagram for the fuselage and empennage boom
Summing the lift increments due to each of the aileron strips gave a total lift increment
of 0.237 which was applied to the nominal lift coefficient of 0.0697. The resulting lift
difference between the two wings was 269.6 N. The couple moment produced by the
differential aileron deflection was assumed to act at the mean aerodynamic chord of the
inboard wings. The net moment due to aileron deflection was determined to be 67.7 Nm.
C.6 Static thrust
The maximum static thrust found during testing was 82.4 N.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
D. Theoretical performance
calculationsThe theoretical performance of the Morpheus UAV was calculated to determine the the-
oretical effects of wing morphing. This analysis did not consider the effects of tail mor-
phing on performance. The parameters used to calculate the theoretical performance of
the Morpheus UAV are listen in table D.1.
Table D.1: Morpheus UAV parametersParameter Retracted wings Extended wingsWing planform area [m2] 0.44275 0.55315Wing span [m] 1.15 1.84Weight [kg] 8 8Aspect ratio 3.0 6.1Power [W] 1650 1650VNE [km/h] 225 147CL,max 0.90475 1.269*Oswald’s efficiency factor (e) 0.971 0.865
*Calculated as the area weighted average of the inboard and outboard wings.
D.1 Wing and power loading
The wing loadings were calculated to be 18.1 kg/m2 and 14.46 kg/m2 for the retracted
and extended wing configurations respectively. The power loading was 4.85 kg/kW.
D.2 Stall speed
The stall speeds of the aircraft was calculated using equation D.1. The retracted stall
speed was 64.39 km/h whilst the extended stall seed was 48.6km/h.
Vs =
√2W
ρSCL,max(D.1)
161
D.3. TAKEOFF DISTANCE 162
D.3 Takeoff distance
The takeoff ground run distance of the aircraft was calculated using the FAR23 method
and equations D.2 to D.4. The retracted and extended takeoff groundrun distances are
63.2 m and 35.2 m respectively.
CL,TO =CL,max1.21
(D.2)
TOP23 =
(WS
) (WP
)CL,TO
(D.3)
STOG = 4.9TOP23 + 0.009TOP232 (D.4)
D.4 Drag polar
The zero lift drag coefficient for the Morpheus UAV was estimated using the wetted area
ratio and equivalent skin friction coefficient. An equivalent skin friction coefficient of
0.0055, similar to a single engine light aircraft, was used (Raymer 2006). The wetted area
ratio was calculated by assuming a cylindrical fuselage and that the surface area of the
wings and tail were equal to double their planform areas. The effects of the landing gere
were accounted for with an increase in drag coefficient of 0.025 (Roskam 1989).
The zero lift drag coefficients were 0.042 and 0.041 for the extended and retracted config-
urations respectively.
D.5 Maximum speed
The maximum speed of the aircraft was calculated by solving equation D.5 for V . This
equation was derived by equating thrust to drag, where thrust is given by T = P/V . An
altitude of 121.92m (400 ft) and a propeller efficiency of 0.7 were assumed. The maximum
speeds of the retracted and extended configurations were determined to be 165.7 km/h and
156.9 km/h respectively. The maximum speed attainable in the extended configuration
exceeds the structural velocity never exceeded limitation for this configuration. Hence
the maximum speed in the extended configuration was reduced to VNE = 147km/h.
Pηp =ρcrV
3SCD0
2+
2W 2
ρcrV SπAe(D.5)
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
163 APPENDIX D. THEORETICAL PERFORMANCE CALCULATIONS
D.6 Endurance
Endurance was calculated for each wing configuration assuming that the aircraft loiters
at V = 1.5Vstall and the total battery capacity available is 5000 mAh. It was also assumed
that 60% of the total battery capacity was available for loiter with other flight operations
consuming the other 40%. This assumption allows a comparison of the endurance of each
configuration in a realistic manner. Equation D.6 was used to calculate the endurance for
each configuration, where Jltr is the energy available for loiter and η is the total propulsion
efficiency of 0.63 (propeller efficiency of 0.7 and motor efficiency of 0.9).
E = (ηJltr)
(1
2ρ(1.5Vstall)
3SCD0 +2W 2
ρ(1.5Vstall)SπAe
)−1
(D.6)
The endurance of the Morpheus UAV was calculated to be 22 minutes and 36 minutes for
the retracted and extended configurations respectively.
D.7 Rate of climb
The rate of climb for each configuration was calculated using the FAR23 rate of climb
parameter and rate of climb sizing method. Equation D.7 describes the rate of climb
parameter, where CL)3/2/CD is given by equation D.8. The rate of climb can then be
calculated from equation D.9 and converted into SI units. The calculated rates of climb
were 12.1 m/s and 13.3 m/s for the retracted and extended configurations respectively.
RCP =
(ηp
(W/P )
)−(
(W/S)1/2
19((CL)3/2/CD)
)(D.7)(
(CL)3/2
CD
)=
1.345(Ae)3/4
(CD0)(1/4)(D.8)
RC = 33000RCP (D.9)
D.8 Performance summary
A summary of the Morpheus UAV’s performance is given in table D.2.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
D.8. PERFORMANCE SUMMARY 164
Table D.2: Morpheus UAV PerformanceParameter Retracted wings Extended wingsStall speed [km/h] 64.39 48.6Takeoff distance [m] 63.2 35.2Maximum speed [km/h] 165.7 147Endurance [minutes] 22 36Rate of climb [m/s] 12.1 13.3
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
E. Manufacturing photosComponent photos:
Fuselage internal structure Wing tip reinforcement Tail root reinforcement
Fuselage Plug
Fuselage Skin
Process photos:
Manual hot-wire cutting Wing wet layup
165
166
Installation photos:
Fuselage bonding to skin Sliding block Inboard Wing
Empennage Nose gear Main landing gear
Propulsion system Morphing electronics Radio control system
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
F. Component test procedures
F.1 Propulsion System Static Thrust Test
F.1.1 Aim
• To test the propulsion motor and two propellers at different power settings.
• To ensure that the propulsion motor and propellers can provide the required 900W
of thrust for a 6kg aircraft (1095W for a 7.5 Kg aircraft).
• To select the best propeller for flight.
F.1.2 Intended results
• Thrust vs. power curves for both propellers.
• Maximum thrust and thrust to power ratio for each propeller.
F.1.3 SOP required
Yes
F.1.4 Related/required tests
None
167
F.1. PROPULSION SYSTEM STATIC THRUST TEST 168
F.1.5 Apparatus
Component No. required We Have From
Propulsion motor 1 Y
16” x 8” propeller 1 Y
16”x 12” propeller 1 Y
8.4V 5350 mAh Li-Po thrust mo-
tor batteries
2 Y
90A ESC 1 Y
Main test rig 1 N Holden Labs
Motor test stand 1 N Holden Labs
Transmitter 1 Y
Transmitter battery pack 1 Y
Receiver 1 Y
4.8V 1100 mAh receiver battery
back
1 Y
Load Cell 1 N Electrical workshop
Ammeter 1 N Electrical workshop
Voltmeter 1 N Electrical workshop
150A relay 1 N Electrical workshop
E-stop 1 N Electrical workshop
Connection wires 1 N Electrical workshop
Lead connectors 1 Y
Signal amplifier 1 N Electrical workshop
Motor mounting bolts 4 Y
Various known weights 6 N Final year study room
String 1 ball N Woolworths
Masking tape 1 roll Y
Rubber bands 20 Y
Spanner 1 Y
Screw drivers 1 Y
Allen keys 1 set Y
Scales 1 N Electrical workshop
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
169 APPENDIX F. COMPONENT TEST PROCEDURES
F.1.6 Diagram
F.1.7 Method
Connections
1. Create circuit using wires and relay
Battery bharging
1. Charge Li-Po batteries
2. Charge transmitter battery pack
3. Charge receiver battery pack
Load cell and data logger calibration
1. Connect load cell to rig
2. Connect voltmeter to load cell
3. Connect ammeter to load cell
4. Measure voltage and amps via voltmeter and ammeter with no loading
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
F.1. PROPULSION SYSTEM STATIC THRUST TEST 170
5. Attach weights to load cell rig via string and measure voltages and amps via volt-
meter and ammeter
6. Create calibration curve for voltmeter
7. Create calibration curve for ammeter
8. Compare voltmeter results to data logger results
9. Compare ammeter results to data logger results
10. Calibrate load data logger
11. Calibrate data logger
Test Procedure
• Mount motor on test stand and ensure it is secured safely
• Connect ESC, batteries and safety circuit to motor
• Set throttle to 0%
• Turn on transmitter
• Connect receiver to ESC
• Connect load cell
• Connect voltmeter and ammeter to load cell
• Calibrate data logger and load cell
• Connect thrust Li-Po batteries to ESC while remaining vigilant of propeller, ensure
E-stop is activated (down)
• Vacate immediate area around motor test stand and ensure that everyone is safely
positioned
• Release E-stop (up), motor is now operational
• Vary power between 0 and 100 W to test motor response
– If motor responds, then continue
– If motor does not respond, then check all connections and try again
• Activate E-stop
• Disconnect thrust batteries
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
171 APPENDIX F. COMPONENT TEST PROCEDURES
• Attach 16” x 8” propeller to motor and ensure it is safely secured
• Vacate immediate area around motor test stand and ensure that everyone is safely
positioned
• Release E-stop (up), motor now operational.
• Vary power between 0 and 100 W to test motor response
• Set power at 0 W and hold for 20 seconds
• Set power at 50 W and hold for 20 seconds
• Set power at 100 W and hold for 20 seconds
• Set power at 200 W and hold for 20 seconds
• Set power at 300 W and hold for 20 seconds
• Set power at 400 W and hold for 20 seconds
• Set power at 600 W and hold for 20 seconds
• Set power at 800 W and hold for 20 seconds
• Set power at 1000 W and hold for 20 seconds
• Set power at 1200 W and hold for 20 seconds
• Set power at 1400 W and hold for 20 seconds
• Set power at 1500 W and hold for 20 seconds
• Set power at 1600 W and hold for 20 seconds
• Repeat steps 16 through 40 two more times
• With throttle set to 0%, activate E-stop (down)
• Disconnect Li-Po batteries from ESC
• Turn off transmitter
• Detach 16” x 8” propeller from motor
• Attach 16” x 12” propeller from motor
• Turn on transmitter
• Set throttle to 0%
• Connect Li-Po batteries to ESC
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
F.1. PROPULSION SYSTEM STATIC THRUST TEST 172
• Vacate immediate area around motor test stand and ensure that everyone is safely
positioned
• Release E-stop (up), motor now operational
• Vary power between 0 and 100 W to test motor response
• Repeat steps 19 through 32
• With throttle set to 0%, activate E-stop (down)
• Disconnect Li-Po batteries from ESC
• Turn off transmitter
• Disconnect data logger
• Disconnect load cell
• Disconnect receiver from ESC
• Disconnect voltmeter and ammeter
• Disconnect ESC and safety circuit from motor
• Detach 16” x 12” propeller from motor shaft
• Detach motor from test stand
• Tidy test area
• Save the computer results to a portable storage device
• Return all borrowed equipment
• Graph results and compare to theoretical data
F.1.8 To do
• Ensure all wires, leads and components have the correct connections
• Borrow all components from Electrical workshop
• Buy string
• Setup rig in propulsion Lab
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
173 APPENDIX F. COMPONENT TEST PROCEDURES
F.1.9 Results
The maximum power required for takeoff was calculated to be approximately 900W.
Hence, the motor using either propeller will be able to provide adequate thrust for flight.
The motor and 16 inch by 8 inch propeller can produce the required thrust at a reduced
power, and as such, was selected as the propeller to be used on the aircraft.
F.2 Mechanism motor test
F.2.1 Aim
To determine the power and current requirements for the morphing motors to power the
telescoping mechanism.
F.2.2 Intended results
1. The power, and current requirements for the morphing motors
2. Verification that the selected motor works
F.2.3 Project phase
Once outboard wings are fully assembled
F.2.4 SOP required
No SOP required. Moving parts consist of threaded rod and small motor. The speeds
and forces involved are therefore restricted, and can be stopped at any time.
F.2.5 Other/related tests required
NA
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
F.2. MECHANISM MOTOR TEST 174
F.2.6 Apparatus
Component QTY We
Have
Needs manufac-
turing
To Borrow from
Carbon wing spars 2 Y (sanded)
Outboard wing + connected
wing block
1 Y
Clamps (small to medium
sized)
6 Y/N Workshop
Spare wooden ribs with
holes positioned
2 Y
Wood block (20-50mm
thick)
2 N Workshop
Aluminium threaded rod 1 Y Y - Machined to
fit motor con-
nector
Motor Bracket -wood 1 N Y
Motor 1 Y
Motor Battery 1 Y - con-
nections
Sandbags Up to
5kg
N Y - smaller in-
crements
iSoar
Scales 1 N Electrical workshop
Shoebox lid 1 Y
Masking tape 1 roll Y
Lubrication for the
threaded rod (silicon
spray)
1
tube
N workshop
Large table 1 N workshop
Shaft coupler
F.2.7 Method
1. Set up the test rig
(a) Inspect the wing for any visible cracks and defects and note these before be-
ginning the test
(b) Cover the spars and threaded rod with silicon spray for lubrication
(c) Thread the port-side spars through the runner tubes on the port-side wing
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
175 APPENDIX F. COMPONENT TEST PROCEDURES
(d) Thread the spars through the holes in the rib and clamp these to the thick
pieces of wood. Clamp the wood securely to the table so that the spars are
about 40mm above the table
(e) Attach secure the threaded rod to the motor via the shaft coupler
(f) Place the threaded rod through the root rib and thread it into the wing such
that the wing is located approximately at of the length of the free threaded
rod
(g) Screw the motor onto the motor support and clamp this to the root rib
2. Attach the power pack to the motor and run the wing to within 100mm of the end
of the rails
3. Switch the power leads and extend the wing to the end of the rig. Switch the power
leads and retract the wing to the root end of the rig
4. Repeat steps 2 and 3 and expand or reposition the ribs and rib holes until the
block/wing moves freely.
5. Once running smoothly repeat steps 2 and 3 twice and record the voltage and current
used each time try to maintain similar speed each time.
6. Secure the shoebox lid to the wing-block using masking tape (longest side parallel
with ribs)
7. Place 1.8 kg of sandbags on the shoebox and secure with tape, being careful to not
spill sand on the lubricated rails repeat steps 2 and 3 twice and record the voltages
and current used each time.
8. Repeat step 7 until the design load of 6.95kg is reached, reapplying lubrication as
required.
F.2.8 To Do
• Borrow all required equipment and lubricant
• Manufacture connections for the motor battery
F.2.9 Results
Motor 1
Run 1 Run 2 Run 3 Run 4
Load (G) Load (%) Load (kg) Volts Amps Volts Amps Volts Amps Volts Amps
0 0 0 2 5 . 5 . 5 1.5 4
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
F.2. MECHANISM MOTOR TEST 176
Motor 2
Run 1 Run 2 Run 3 Run 4
Load (G) Load (%) Load (kg) Volts Amps Volts Amps Volts Amps Volts Amps
0 0 0 1.5 5 2.5 4 3 4
1 2.6 1.83 2 5 2 6
2 5.3 3.66 2 7 2.5 8.5 2.5 9
2.3 60 4.17 3 7 4 8 3.5 6 2.5 9.5
3 79 5.48 3.5 9.5 3.5 9 3 9
3.8 100 6.95 NA
Discussion
• The first motor was not powerful enough and burnt out at the 1G load case. This
was an unknown motor that was found.
• The second motor was much bigger and provided adequate power however we were
not able to test the 3.8G case as the current draw was too high for the power supply
pack and needed to be geared down.
• The Maximum power drawn was 9.5A*3.5V = 33.25Watts
Problems
• Alignment of rails and mechanism - small misalignments caused mechanism to jam
• Vibration and whipping of aluminium rod - will be less in reality as test piece was
of double length
• Could decrease vibration by lining the tube to damp vibrations and prevent whip-
ping.
• Vibration of motor - very large motor so will not be as bad in the final mechanism.
Was damped using foam blocks
• Vibration of rails - will be reduced with smaller motor and will be damped by the
foam in the wings. The spars are also double their final length so vibrations will be
reduced.
• Lubrication - the carbon on carbon mechanism would easily seize up and so a Silicon
based lubricant was used however under heavy loadings this was not sufficient to
prevent binding. A synthetic grease was tried and this was found to be adequate.
• Bending of spars - due to the spars being double the required length. When sup-
ported at the correct length, only negligible bending was observed.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
177 APPENDIX F. COMPONENT TEST PROCEDURES
• Shaft coupler slipping - due to the loads experienced in the mechanism the shaft
coupler often slipped off the threaded rod. This was solved by epoxying the threaded
rod into the shaft coupler.
F.3 Wing Structural Tests
F.3.1 Aim
To ensure the complete unit of outboard and inboard wings can take the design load of
3G when supported only by the two wing tongues.
F.3.2 Intended results
• Deflection at Load data/graphs
• Determine whether able to support the design Load of 3G
F.3.3 SOP required
No
F.3.4 Related/required tests:
NA
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
F.3. WING STRUCTURAL TESTS 178
F.3.5 Apparatus
Component No. required We Have Needs manufac-
turing
To Borrow from
Tables* 4 Y Workshop
Sandbags 68 (see below) N Make ourselves Sand from workshop
Ruler 2 Y
White Paper 1 Y
Texta 2 (red, black) Y
Video Camera 1 Y
Freezer/sandwich bags 136 N Woolworths
Wing tongue units (includ-
ing brackets)
2 Y
160x200x9mm piece of ply
wood
1 Y
Fully assembled Out-
board/inboard wing units
2 Y
Balsa shims 8x20x1 mm 8 N Make ourselves
Masking tape 1m approx Y
Small/medium clamps 4 Y
*two tables must be same width and have a narrow ridge below the tabletop
F.3.6 Method
Create paper ruler
1. Create two large rulers by drawing lines along the paper at 5mm increments from
zero to 140mm, with zero at the top of the paper. For ease of reading make the 10’s
black and the 5’s red.
2. Attach a ruler to one side of each paper ruler.
Sandbags
37. Fill the freezer-bags with sand until the required weight is reached, see table below,
and then tie securely. Ensure that the sand bag is fairly flat to enable a greater
weight distribution.
38. Place the freezer-bags in a second bag and tie securely to prevent sand seepage.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
179 APPENDIX F. COMPONENT TEST PROCEDURES
Test Wings - both wing units simultaneously - perform experiment in a quiet area.
39. Inspect the wings for any visible cracks and defects and note these before beginning
the test.
40. tip tables onto their long sides with the legs facing away from each other and top
surfaces together.
41. Position the brackets on the wing tongues 120 mm apart (as in fuselage) then bolt
the brackets in position onto the piece of wood. Clamp this to the tipped tables via
the ridges.
42. Position the shims under the root ribs in the 4 roller positions on the top surfaces
of the wings to simulate the support provided by the rollers.
43. Assemble the wings in the wing tongues ensuring they are upside down such that
the loading simulates an “upwards” lift force.
44. Position the paper ruler behind the wing tips such that the tip is aligned with the
zero position.
45. Mark the loading positions onto the wings at the required positions via masking
tape, 6 per inboard wing and 4 per outboard wing. (see below for positions).
46. Position a video camera so it can record the deflection of the wing against the ruler
during the experiment. (Start recording)
47. Place the sandbags on the wings at the 60% loading condition (see below) ensuring
the loads are places on each wing simultaneously to prevent tipping of wings or
uneven loading. Also ensure someone is supporting the inboard and outboard wings
in the unloaded position whilst loading to prevent longer excessive times.
48. Once the bags are positioned slowly lover the wings and remove hand supports,
leave for 30 seconds and measure the tip deflection. Listen for cracks and if heard,
stop experiment immediately and unload wing.
49. Raise wings and support them at the unloaded position whiles removing the sand-
bags.
50. Inspect the wing for any new visible/audible cracks, if any are found stop the ex-
periment.
51. Test the wings by tapping along the wing and listening for any change in sound
indicating failure or cracks, if any defects are found stop the experiment.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
F.3. WING STRUCTURAL TESTS 180
52. Repeat steps 13, 17 with 80% load conditions at 30 seconds.
53. Repeat steps 13, 17 with 90% load conditions at 5 seconds.
54. Repeat steps 13, 17 with 100% load conditions at 5 seconds
55. Remove sandbags from the wing.
56. Check wing for any cracks/defects that were not present at the beginning of the
test.
Note
• If any cracking is heard stop experiment immediately and find failure point. The
cracking indicates de-bonding. Fix with epoxy resin (as a filler) and re-enforce
section if required.
• If delamination occurs stop experiment immediately and talk to Maziar.
• Once wing is fixed and re-enforced test can be resumed.
F.3.7 Loading conditions
Sandbag loads at: 75% of total load (2.3G)
Distance from root rib [m] Load [g] Rounded [g] COMBINATIONS [g]
0.042 449 450 50 400
0.126 480 480 80 400
0.21 500 500 500
0.294 519 520 20 500
0.378 519 520 20 500
0.462 502 500 500
0.546 649 650 50 600
0.63 569 570 70 500
0.714 427 430 30 400
0.784 242 240 40 200
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
181 APPENDIX F. COMPONENT TEST PROCEDURES
Sandbag loads at: 100% of total load (3G)
Distance from root rib [m] Load [g] Rounded [g] COMBINATIONS [g]
0.042 671 670 70 600
0.126 706 710 10 700
0.21 726 730 30 400 300
0.294 745 750 50 400 300
0.378 738 740 40 200 500
0.462 708 710 10 200 500
0.546 890 890 70+20 400 400
0.63 783 780 80 500 200
0.714 594 590 70+20 500
0.784 347 350 50 300
Number of sand bags required
Weight (g) number required
10 6
20 4
30 2
40 2
50 4
60 2
70 6
80 2
200 6
300 8
400 8
500 10
600 4
700 4
F.3.8 Results
Carried out by Crystal, Kevin, Ian and Simon in S225 room in Engineering south 25/09/09.
• Simon and Ian load and unload wings =¿ place load at same points at same time
• Kevin and Crystal support wings =¿once loaded, lower wings to natural defelction
(no support) =¿ read deflection =¿ raise and support wings for unloading.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
F.3. WING STRUCTURAL TESTS 182
Load Case Port Wing Deflection (mm) Starboard Wing Deflection (mm)
75% 44 45
100% 58 56
75% Load
• Very light creaking when loading in port wing (occurs on all load cases)
• No cracks heard or seen
• No change in wings with tap test
• Successful test
100% Load
• Light creaking when loading in port wing
• No cracks seen or heard
• No change in wings with tap test
• Successful test
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
183 APPENDIX F. COMPONENT TEST PROCEDURES
F.3.9 Assembly electronics, morphing and control test checklist
COMPONENT GROUP COMPONENT RESULTS PASSED
Servos (x5) left elevator L24, -19; H26, -21 Y
(deflection U/D via Remote) right elevator L24, -16, H29, -21 Y
Port aileron L20, -18; H29, -28 Y
Stbd aileron L19,-20; H+28-31 Y
left rudder 23, -16 Y
right rudder 16, -24 Y
Stbd flaps -19 Y
Port flaps -17 Y
nose gear Port46; stbd38 Y **
New nose gear Port40; stbd39 Y
Mixer Original mixer Chatter in elevator N ***
transmitter mixer No chatter Y
ESC (x4) Thrust operational Y
(working) left morphing operational Y
right morphing operational Y
tail morphing operational Y
Motor (x4) Thrust operational Y
Morphing Left operational Y
Morphing Right operational Y
Morphing tail operational Y
Receiver (x2) main unit operational Y
morphing unit operational Y
Batteries (x2) Thrust operational Y
Morphing operational Y
Switches thrust on/off operational Y
Switches Morphing on/off operational Y
Switches Morphing Port Wing on/off operational Y
Switches Morphing Starboard wing on/off operational Y
Switches Morphing tail on/off operational Y
Wires (test with associated components) operational Y
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
G. Heavy model certification
184
185 APPENDIX G. HEAVY MODEL CERTIFICATION
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
H. Heavy model requirements
186
187 APPENDIX H. HEAVY MODEL REQUIREMENTS
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I. Flight test procedures
I.1 Pre-flight ground checks
I.1.1 Things to Check
• Hatches - Cowl, ESC, batteries, morphing mechanism, landing, fairing
• Four screws tightened
• Wing tongues - four screws
• Batteries to be longitudinally secure
• Batteries secured
• Payload secured
• Wing nylon bolts X4
• Receivers are secure
• Path to be clear for sliding block
• Transmitters, receivers, batteries - fully charged and working
I.1.2 Actual Tests
• Control Surfaces (also Receiver + switches)
• Ailerons - rough deflection + range + direction
• Flaps - rough deflection + range
• Ruddervators (elevators + rudder) - rough deflection + range
• Nosegear - rotation + range
• Morphing (also receiver, motor, receiver batteries, morphing batteries + switches)
• L-wing → in/out → close range + max range (or cont surf range if smaller)
• R-wing → in/out → close range + max range (or cont surf range if smaller)
• Tail → in/out → close range + max range (or cont surf range if smaller)
188
189 APPENDIX I. FLIGHT TEST PROCEDURES
• Limit switches
• Emergency stop
• Kill Switch
• Thrust Motor: (also main receiver, main batteries, connections)
• Works
• Control surfaces and morphing works at max range with thrust motor on
• SM:
• Check is correct
• Tipback test:
• Ensure correctly balanced
I.1.3 Electronics start-up procedure
(if not morphing remove all morphing references)
• Turn off phones, wireless internet, and bluetooth devices
• Remove antennae shroud
• Finalise payload arrangements - secure to plates
• Single cable-tie for each battery on mounting plate
• Longitudinal cable ties on batteries
• Install thrust batteries (i.e. physically put in) - 4 swivel clamps attaches
• Battery hatch on (4 screws)
• Secure batteries longitudinally via ESC hatch
• Install and secure morphing + receiver Batteries (1 cable tie each) sticky-tape re-
ceiver battery plug to prevent disconnection- check receiver secure and tail block
free movement area
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.1. PRE-FLIGHT GROUND CHECKS 190
I.1.4 Inboard wing installation
• Push onto carbon tubes, feed servo, limit switch and motor cables into fuselage
• Mount nylon wing bolts via battery hatch and fairing.
• Attach motor leads as marked (3 on each side - 6 total)
• Connect limit switch channels to PCB as colour marked - (2) on each wing
• Attach servos (1 on each wing) to main receiver (big one) → A = aileron channel,
X=auxiliary 1
• Control surface check now possible
I.1.5 Outboard wing Installation
• Clean and lube rails and threaded rod
• Place threaded rod in Aluminium insert
• Reach through morphing mechanism hatch - turn motor clockwise so thread has
caught
• Repeat last two steps for other wing
• Follow morphing procedure to position wings all the way
• Attach limit switch able to extension lead
• Put tip rib on
• Put balsa front on (ignore for reality)
I.1.6 Tail
• Remove fairing
• Clean and lube rails and threaded rod
• Line up threaded rod
• Connect servos to the receiver as marked
• 10 turns
• Motor on - position tail part way
• Insert rails into block and on fuselage former
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
191 APPENDIX I. FLIGHT TEST PROCEDURES
• Replace fairing
• Screw tight
• Motor on - position tail fully
I.1.7 Ready to fly
• Connect receiver battery and morphing LiPo check receiver is secure
• Secure fairing
• Place frequency key in the club board to ensure no interference from another trans-
mitter.
• Check switches on transmitters
• Main receiver = all low rates and flaps off
• Morphing receiver all low rate and kill switch off
• Switch order: (if doing range checks do 2on, 3on test 3off, 2off, 1on, 2on, 3on)
1. Both transmitters on (don’t need morphing on if not morphing)
2. Connect Thrust LiPos
3. Control surface switch
4. Morphing switch
5. Morph to most extended position wing and tail (check equal wing extension)
6. Activate kill switch = check nothing should happen
7. Replace ESC hatch
• FLY
I.1.8 Landing
• ESC Hatch off
• Switch order
1. Morphing switch
2. Control surface switch
3. Disconnect thrust LiPo
4. Turn off transmitters
• → SAFE
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.1. PRE-FLIGHT GROUND CHECKS 192
I.1.9 End of consecutive flights
• Remove fairing
• Disconnect morphing LiPo
• Replace fairing
I.1.10 End of flight/day
• Remove and charge all batteries
I.1.11 To Change thrust batteries
• ESC hatch off
• Cut longitudinal support cable tie for batteries
• Remove battery hatch
• Remove battery plate
• Repeat start up procedure
I.1.12 Trouble shooting Ground Checks
• Electrical connections (don’t need to check if range checks work):
• Batteries
• Main → attached + properly connected
• Morphing → attached + properly connected
• ESC - X4
• Main X1 → Connected to batteries + receiver; secured to internals
• Morphing X1 → connected to receiver, battery, PCB, motor + secured to internals
• Receivers - X2
• Main → Secured to internals; attached to
• Morphing → Secured to internals; attached to
• Motors
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
193 APPENDIX I. FLIGHT TEST PROCEDURES
• Main → connected to batteries + ESC + Prop; Secured to former
• Morphing → connected to battery + ESC + PCB + prop + secured to wing
• Switches - X
I.1.13 Top 10 trouble shooting
• Morphing not work
– Killswitch position
– Check PCB connections
– All 3 soft switches
– LiPo charged
– LED’s of three morphing ESCs
• Control surfaces (nose gear)
– Control surface switch on
– Receiver battery connected to receiver lead
– Check receiver battery charge
– Check receiver connections - 7 inputs
• Thrust motor not working
– Check cables connected - thrust LiPo connection
– Control surface switch on
– Receiver battery connected to receiver lead
– Check receiver battery charge
– Check receiver connections - 7 inputs
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.2. RANGE CHECKS 194
I.2
Ra
ng
eC
hec
ks
Ran
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stre
sult
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em
orphin
gsy
stem
san
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esy
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s
Con
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Surf
ace
Ran
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gM
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ism
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ayan
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,
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d
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,
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d
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trol
Set
up
Mot
oroff
,
Ante
nna
retr
acte
d
Mot
oron
,
Ante
nna
retr
acte
d
6/10
/09
Hol
dfa
st36
MH
z-
whip
aeri
alSp
ectr
um
NA
NA
18/1
0/09
Bar
ossa
36M
Hz
-w
hip
aeri
alSp
ectr
um
NA
NA
18/1
0/09
Bar
ossa
36M
Hz
-w
ire
aeri
alSp
ectr
um
NA
NA
18/1
0/09
Bar
ossa
Sp
ectr
um
-sa
tellit
eae
rial
36M
Hz
NA
NA
24/1
0/09
Con
stel
lati
onSp
ectr
um
-sa
tellit
eae
rial
>70
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Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
195 APPENDIX I. FLIGHT TEST PROCEDURES
I.3 PF1 - Ground handling - Taxi test
I.3.1 Aim
• Ensure thrust motor is operational
• Obtain taxi throttle setting
• Ensure control surfaces + nose gear operates as expected with thrust motor opera-
tional
• Ensure adequate nose gear authority
• Check ground handling of aircraft
• Nose over
• Tail over
• Tip back
• Ensure aircraft capable of traversing over small bumps and rocks.
I.3.2 Procedure
1. **Ground Checks
2. **Electronics start-up/shutdown (morphing included)
3. Position aircraft on runway
4. Slow taxi along runway/ground - loop when get to end
• Move all control surfaces
• Figure 8s Left
• Figure 8s Right
• Spiral Left → find min turn radius
• Spiral Right → find min turn radius
5. Bring aircraft to starting/end position
6. Thrust motor off
7. **Electronics start-up/shutdown
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.4. PF2 - GROUND RUN 196
I.3.3 Results
First flight day
• Throttle setting = approximately 12
• Nose gear is a bit flimsy
• On sharp turns the nose gear digs into the ground, tipping the UAV.
• The UAV handled very well and was able to successfully complete tight turns at
slow speeds.
• The nose gear movement was reduced to prevent very tight turns.
Second flight Day
• Performance after the crash repairs was very good and no modifications to the nose
gear movement were required.
• Handling was excellent and the UAV easily completed the spirals and figure eight
manoeuvres.
• Pilot commented that the UAV had excellent ground handling qualities.
Third and fourth flight Days
• Nose handling was excellent
• No adjustment required
I.4 PF2 - Ground Run
I.4.1 Aim
• Ensure adequate nose gear authority
• Ensure rudder and nose gear adequate for tracking on runway
• Determine longitudinal stability
• Determine the elevator/pitching response
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
197 APPENDIX I. FLIGHT TEST PROCEDURES
I.4.2 Procedure
1. **Ground Checks
2. **Electronics start-up/shutdown (morphing included)
3. Position aircraft on runway
4. Run down runway
• Slow taxi along runway/ground
• Move elevator and ensure aircraft doesn’t pitch up
5. Bring aircraft to starting/end position
6. run down runway
• Move elevator and ensure aircraft doesn’t pitch up
• Repeat whilst increasing speed until reach just below TO speed or until feel
like aircraft wants to takeoff
7. Bring aircraft to starting/end position
8. Thrust motor off
9. **Electronics start-up/shutdown
10. Quality Assurance check
I.4.3 Results
Flight day 1
• Elevators during the first/slow pass did not cause a pitching moment
• Elevators during a speed just below TO speed did not cause a pitching moment
• At a TO speed without elevators the nose gear caught in the rough during a slight
turn, however the plane sustained no damage. Therefore speed should be reduced
before turning.
• Good handling and tracking at speeds near TO speed.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.5. F1 - STABILITY TEST 198
Flight day 2
• Elevators during the first/slow pass did not cause a pitching moment
• Elevators during a speed just below TO speed did not cause a pitching moment
• Good handling and tracking at speeds near TO speed.
Flight day 3 and 4
• Since no major modifications had been made to the UAV the full ground run test
was not required. Instead, a single run was made to allow the pilot to become used
to the Morpheus UAV and to ensure adequate nose gear authority.
• On both days this test was successful and the UAV demonstrated excellent ground
handling qualities.
I.5 F1 - Stability test
I.5.1 Aim
• Prove conventional takeoff and landing
• Prove aircraft is capable of short cruise
• Trim aircraft and demonstrate straight and level flight
I.5.2 Procedure
1. **Ground Checks
2. **Electronics start-up/shutdown
3. Position aircraft on runway
4. Taxi
5. Ground run
6. return to start on runway
7. Taxi
8. Take off
9. Climb to 2 metres
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
199 APPENDIX I. FLIGHT TEST PROCEDURES
10. Cruise for five seconds
11. Trim aircraft
(a) Test pitch response
(b) Test roll response
12. When straight and level flight has been achieved for 5 seconds, land
13. Taxi
14. Thrust motor off
15. Control switches off
16. Full check of aircraft and components
I.5.3 Results
Flight day 2
Manoeuvre Pilot comments
Taxi Throttle response was good
Takeoff Screeching sound just after takeoff, accompanied by loss of power
Land Aircraft flared to reduce impact forces
Main landing gear sheared as designed
Smoke emitted from motor
Weather conditions
• Light winds
• Light gusts
I.6 Propulsion System Static Motor Test
I.6.1 Aim
• To test the propulsion motor and two propellers at different power settings.
• To ensure that the propulsion motor and propellers can provide the required 900W
of thrust for a 6kg aircraft (1095W for a 7.5 Kg aircraft).
• To select the best propeller for flight.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.6. PROPULSION SYSTEM STATIC MOTOR TEST 200
I.6.2 Intended results
• Thrust vs. power curves for both propellers.
• Maximum thrust and thrust to power ratio for each propeller.
I.6.3 SOP required
Yes
I.6.4 Related/required tests
None
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
201 APPENDIX I. FLIGHT TEST PROCEDURES
I.6.5 Apparatus
Component No. reqd We Have To Borrow from
Propulsion motor 1 Y
16” x 8” propeller 1 Y
16” x 12” propeller 1 Y
8.4V 5350 mAh Li-Po thrust motor batteries 2 Y
90A ESC 1 Y
Main test rig 1 N Holden Labs
Motor test stand 1 N Holden Labs
Transmitter 1 Y
Transmitter battery pack 1 Y
Receiver 1 Y
4.8V 1100 mAh receiver battery back 1 Y
Load Cell 1 N Electrical workshop
Ammeter 1 N Electrical workshop
Voltmeter 1 N Electrical workshop
150A relay 1 N Electrical workshop
E-stop 1 N Electrical workshop
Connection wires 1 N Electrical workshop
Lead connectors 1 Y
Signal amplifier 1 N Electrical workshop
Motor mounting bolts 4 Y
Various known weights 6 N Final year study room
String 1 ball N Woolworths
Masking tape 1 roll Y
Rubber bands 20 Y
Spanner 1 Y
Screw drivers 1 Y
Allen keys 1 set Y
Scales 1 N Electrical workshop
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.6. PROPULSION SYSTEM STATIC MOTOR TEST 202
I.6.6 Diagram
I.6.7 Method
Notes from the Safe Operating Procedure
• All personnel must be familiar with the emergency stop procedure of this test.
• Personnel must not, under any circumstances be in the immediate vicinity of the
propeller if the propeller is rotating.
• Personnel must not remain in the plane of the propeller if the propeller is rotating.
• Personnel must ensure that long hair is secure and that their clothing and accessories
cannot be trapped in the propeller.
• Gloves must be worn when connecting and disconnecting the Lithium Polymer Bat-
teries,
• Personel are to make themselves aware of any tripping or electrical hazards before
the start of this test.
Wind tunnel
1. Attach motor to the Metal rig stand
2. Secure all components to be tested to a single rig, and ensure that the rig fits within
the open cross section of the wind tunnel. Ensure that the rig weighs at least 30 kgs
by adding heavy, secure ballast to a safe location. Qualitatively ensure that the rig
can resist a 20kg force in the forward and backward direction without a tendency
to slip or rotate.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
203 APPENDIX I. FLIGHT TEST PROCEDURES
3. Connect motor to ESC
4. Connect ESC to receiver (connected to receiver battery) and safety relay
5. Connect safety relay to E-stop connected to a 12 V power supply, and connect the
coaxial voltage terminals to a multimeter to measure voltage.
6. Connect current clamp to surround the positive wire between the ESC and the
safety relay, and connect a second multimeter to measure current.
7. Cover ESC with a small, secure container with restricted ventilation to simulate
flight environment
8. Position wireless, digital thermometer in container, next to the ESC.
9. Connect safety relay to batteries and test safety relay and E-stop functionality.
10. Remove current experiment in the wind tunnel
11. Position table in the wind tunnel such that the propeller is facing the oncoming flow
12. Using the conventional procedure, start the motor and esc. Set the ESC throttle
range as given in the manufacturer’s instructions. Bring the propeller to the slowest
possible rotation. Ensure that the propeller is rotating in the correct direction and
ensure the propeller is rotating true.
13. Test the emergency stop functionality and wait for the propeller to completely stop.
14. Disconnect the batteries and recheck the security of the components, especially the
propeller and batteries. Check the security of the rig.
15. Start the airflow at 80kph
16. Test speed of the flow via handheld speed device
17. Using the conventional procedure, start motor on idle throttle via transmitter.
18. Slowly increase throttle to 100% hold for 5 minutes
(a) Monitor voltage and Amps and record values every 15 seconds unless significant
and lasting drop observed (record min value and duration)
(b) Note any unusual noises, drops in propeller/motor speed, or problems in trans-
mitter (record Volts and Amps at this point)
(c) Monitor temperature of ESC - record every 15 seconds
19. Repeat steps 11 to 14 at 90kph
20. Repeat steps 11 to 14 at 100kph
21. Repeat steps 11 to 16 at 75% throttle
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.6. PROPULSION SYSTEM STATIC MOTOR TEST 204
Conventional procedure for starting the propeller
1. Ensure all electrical connections are complete and secure, excepting the connection
of the thrust LiPos.
2. Ensure that all batteries (LiPos, receiver batteries, transmitter batteries) are charged.
3. Ensure the security of the thrust rig, and ensure a working zone free of trip hazards,
loose objects, and other environmental hazards.
4. Ensure the emergency stop has been activated and is in the locked position.
5. Using gloves, connect the LiPo batteries.
6. Move the throttle to the zero position and turn on transmitter.
7. Verify that the beeps from the ESC are as expected
8. Check operation by applying a small amount of throttle, and verify the direction of
prop rotation.
Using old ESC
1. Start motor on idle throttle via transmitter.
2. Increase throttle to 100% hold for 5 minutes
(a) Monitor voltage and Amps and record values every 15 seconds unless significant
and lasting drop observed (record min value and duration)
(b) Note any unusual noises, drops in propeller/motor speed, or problems in trans-
mitter (record Volts and Amps at this point)
(c) Monitor temperature of ESC - record every 15 seconds
3. Decrease throttle to 75% hold for 3 minutes
Using New ESC
1. Set up rig as before, but using the new ESC
2. Start motor on idle throttle via transmitter.
3. Increase throttle to 25% hold for 5 minutes
4. Increase throttle to 50% hold for 5 minutes
5. Increase throttle to 75% hold for 5 minutes
6. Increase throttle to 100% hold for 5 minutes
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
205 APPENDIX I. FLIGHT TEST PROCEDURES
I.6.8 Results
This test was performed three times in the Holden Labs at the University of Adelaide.
The first trial utilised the ESC, receiver and batteries that were installed in the aircraft
when it crashed, however a new motor of the same type as before was purchased. In this
test, the motor was very responsive to any throttle input and was working correctly at
full throttle and a headwind (provided by the windtunnel). The propulsion system was
operating smoothly until the wind speed reached 60kph and a spark was observed between
the motor and the ESC. Upon inspection of the equipment the ESC-motor cables were
very tight against the metal rig and a puncture was seen in the cables. It was assumed
that the spark was caused by the wire discharging to the rig. The cables and rig were
then isolated and the experiment was re-performed. Again a spark was seen in the region
between the ESC and motor along with a loss in thrust, however this time the spark
occurred at a lower windspeed of 25kph with full throttle. When the components were
inspected it was found that the motor had overheated and drawn the ESC leads into
itself causing them to become taught against the rig. Burn marks inside the motor also
indicated that the motor was the cause of the sparks. This could have been caused by
either a timing error between the ESC and the motor or by the motor receiving too much
power.
Upon discussions with Chris French, a member of the 2007 Fuel Cell UAV group, it
was found that the Fuel Cell system was acceptable and this was utilised in our design.
Repeating the test with the new components caused no sparking or loss of power.
I.7 Propulsion System Static Motor Test
I.7.1 Aim
• To test To determine the appropriate ESC and throttle settings required for suc-
cessful propulsion system performance in flight
I.7.2 Intended results
• Appropriate ESC settings for successful propulsion system performance in flight
• Maximum throttle setting required to prevent the motor drawing more than 40 amps
I.7.3 SOP required
Yes (covered by propulsion system static thrust test SOP)
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.7. PROPULSION SYSTEM STATIC MOTOR TEST 206
I.7.4 Related/required tests
Propulsion system static thrust test
I.7.5 Apparatus
Component No. required We Have To Borrow from
Propulsion motor 1 Y
16” x 8” propeller 1 Y
18.5V, 5350 mAh Li-Po thrust motor batteries 2 Y
90A ESC 1 Y
Main test rig 1 N Holden Labs
Motor test stand 1 N Holden Labs
Transmitter 1 Y
Transmitter battery pack 1 Y
Receiver 1 Y
4.8V 1100 mAh receiver battery back 1 Y
Voltmeter 2 N Electrical workshop
150A relay 1 N Electrical workshop
E-stop 1 N Electrical workshop
Connection wires 1 N Electrical workshop
Lead connectors 1 Y
Motor mounting bolts 4 Y
Cable ties 4 Y
Spanner 1 Y
Screw drivers 1 Y
Allen keys 1 set Y
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
207 APPENDIX I. FLIGHT TEST PROCEDURES
I.7.6 Diagram
I.7.7 Method:
Connections
1. Create circuit using wires and relay.
Battery charging
1. Charge Li-Po batteries.
2. Charge transmitter battery pack.
3. Charge receiver battery pack.
Load cell and data logger calibration
1. Connect voltmeters
2. Measure voltage and amps via voltmeter
Test Procedure
1. Mount motor on test stand and ensure it is secured safely.
2. Connect ESC, batteries and safety circuit to motor.
3. Set throttle to 0%.
4. Turn on transmitter.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.7. PROPULSION SYSTEM STATIC MOTOR TEST 208
5. Connect receiver to ESC.
6. Connect voltmeter and ammeter
7. Connect Li-Po batteries to ESC while remaining vigilant of propeller.
8. Vacate immediate area around motor test stand and ensure that everyone is safely
positioned.
9. Vary power between 0 and 100 W to test motor response.
• If motor responds, then continue
• If motor does not respond, then check all connections and try again.
10. Switch circuit off.
11. Attach 16” x 8” propeller to motor and ensure it is safely secured.
12. Vacate immediate area around motor test stand and ensure that everyone is safely
positioned.
13. Vary power between 0 and 100 W to test motor response.
14. Increase throttle gradually to full throttle.
15. Record current draw
• If current draw on full throttle is 40A, end test
• If current draw on full throttle is not 40A, then continue
16. Throttle back to 0%
17. Adjust transmitter settings to set new throttle limit
18. Repeat steps 14 and 15 until current draw on full throttle is 40A.
19. Throttle back to 0%.
Pack-up:
1. Disconnect Li-Po batteries.
2. Disconnect receiver from ESC.
3. Disconnect voltmeter and ammeter.
4. Disconnect ESC from motor.
5. Detach 16” x 8” propeller from motor shaft.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
209 APPENDIX I. FLIGHT TEST PROCEDURES
6. Detach motor from test stand.
7. Tidy test area.
8. Save the computer results to a portable storage device.
9. Return all borrowed equipment.
I.7.8 TO DO:
• Ensure all wires, leads and components have the correct connections
• Borrow all components from Electrical workshop
• Setup rig in propulsion Lab
I.7.9 Results
The aim of the test was to determine the appropriate ESC and throttle settings required
for successful propulsion system performance in flight. Unlike previous thrust tests, this
test did not aim to determine the thrust of the motor. Rather, the throttle settings and
ESC settings were the main parameters of interest. Analysis of the propulsion system after
the accident revealed that the motor was drawing too much current from the batteries.
The motor is rated to 40A continuous current draw and 60A burst current draw for
15 seconds. Hence, the aim of the test was to set the ESC to the most appropriate
settings and limit the throttle setting on the transmitter so that the motor only drew 40A
continuous current at all times.
Prior to the test, the ESC was reset to default settings, with the exception of the timing
setting, which was set to ’high’ as per advice from aeromodellers. Once the test com-
menced, it was determined that the motor was drawing more than 40A current at full
throttle, so the settings on the transmitter were adjusted until full throttle corresponded
to a 40A current draw. Then, the motor was run at full throttle until the batteries ap-
proached their safe discharge voltage of 15V each (3V per cell at 5 cells per battery).
The test was then stopped, and an endurance of 7 minutes was recorded. The motor ran
flawlessly throughout the entire test, with no signs of any issues. The test was repeated
with no changes to any throttle settings or ESC settings. An endurance of 7 minutes was
once again recorded, wit the motor running flawlessly throughout the entire test.
I.7.10 Conclusion
The propulsion system settings are suitable for flight.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.8. F2 - AIRWORTHINESS TEST 210
I.8 F2 - Airworthiness test
I.8.1 Aim
• Prove conventional TO an Landing
• Capable of short cruise
• Determine Loiter endurance
• Trim aircraft
I.8.2 Procedure
1. **Ground Checks
2. **Electronics start up/shutdown (morphing included)
3. Position aircraft on runway
4. Taxi
5. Ground run
6. return to start on runway
7. Taxi
8. Take off
9. climb - straight line to trim altitude
10. trim (approx 10 circuits)
(a) test pitch response
(b) test roll response
11. short cruise (note throttle setting)
12. flutter test - low pass
13. Climb to altitude
14. loiter velocity (2 circuits)
15. deploy flaps =¿ response
16. undeploy flaps
17. landing approaches without flaps
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
211 APPENDIX I. FLIGHT TEST PROCEDURES
18. landing approaches with flaps (if required/possible)
19. land
20. taxi
21. Thrust motor off
22. control switches off
23. Electronics start up/shutdown
24. Quality Assurance check
25. Full check of aircraft and components
I.8.3 Results
• Very good ground handling
• Minimal trim required - very good pitch and roll response, not too extreme.
• Short cruise occurred on full throttle
• Flutter test showed no apparent flutter in the control surfaces.
• Loiter velocity occurred slightly higher than planned at about 75
• Flaps not required
• Landed on third landing approach
• No problems with landing
Pilot said the UAV was stable, had a very good control response rate and was very good
to fly.
I.9 F3 - Morphing mechanism test
I.9.1 Aim
• Ensure aircraft is capable of 50
• Ensure aircraft is capable of morphing the tail
• Ensure aircraft is stable in the retracted configuration
• Measure performance parameters (max speed, cruise speed, loiter speed, endurance)
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.9. F3 - MORPHING MECHANISM TEST 212
I.9.2 Procedure
• **Ground Checks - morphing range check included
• **Electronics start-up/shutdown (morphing included)
• Practice extending and retracting the wings at the same rate
• Practice extending and retracting the tail
• Position aircraft on runway
• Taxi
• Ground run
• return to start on runway
• Taxi
• Take off
• Climb - straight line to trim altitude
• Trim (approx)
1. test pitch response
2. test roll response
• Cruise velocity
• When on straightest part of circuit:
1. Bring wings half in and complete one circuit
• Trim aircraft
1. Test roll response
2. Test pitch response
• When on straightest part of circuit:
1. Bring wings fully in and complete one circuit
• Trim aircraft
1. Test roll response
2. Test pitch response
• When on straightest part of circuit:
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
213 APPENDIX I. FLIGHT TEST PROCEDURES
1. Fully extend both wings and complete one circuit
• Trim aircraft
1. Test roll response
2. Test pitch response
• When on straightest part of circuit:
1. Retract tail one third the way in and complete one circuit
• Trim aircraft
1. Test roll response
2. Test pitch response
• When on straightest part of circuit:
1. Retract tail two thirds the way in and complete one circuit
• Trim aircraft
1. Test roll response
2. Test pitch response
• When on straightest part of circuit:
1. Retract tail all the way in and complete one circuit
• Trim aircraft
1. Test roll response
2. Test pitch response
• When on straightest part of circuit:
1. Fully retract both wings and complete two circuits
• Trim aircraft
1. Test roll response
2. Test pitch response
• Maximum throttle (maximum speed)
• Cruise speed (same throttle as before)
• When on straightest part of circuit:
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.9. F3 - MORPHING MECHANISM TEST 214
1. Fully extend both wings and complete one circuit
• Trim aircraft
1. Test roll response
2. Test pitch response
• When on straightest part of circuit:
1. Fully extend tail and complete one circuit
• Trim aircraft
1. Test roll response
2. Test pitch response
• Land
• Taxi
• Thrust motor off
• Control switches off
• Place in secure position
I.9.3 Results
Wing morphing rates:
Component Half ext. time [s] Fullext. time [s] Half ret. time [s] Full ret. time [s]
Starboard wing 4 9 4 7
Port wing 4 9 5 9
Tail morphing rates
Component1/3 ex-
tension
time (s)
2/3 ex-
tension
time (s)
Full ex-
tension
time (s)
1/3 re-
traction
time (s)
2/3 re-
traction
time (s)
Full re-
traction
time (s)
Tail 2 4 7 2 4 7
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
215 APPENDIX I. FLIGHT TEST PROCEDURES
Flight phases
Elapsed time (mins) Manoeuvre Real time Pilot comments
0:00:00 Takeoff 11:14:43 Plenty of power, good climb
rate, good trim
1:27:14 Start wings 1/2 in 11:16:10 Good stability and control,
no trimming required
1:52:01 Start wings full in 11:16:35 Good stability and control,
greater roll control author-
ity, no trimming required
2:26:08 Wings out 11:17:09 Good stability and control,
felt greater lift once wings
were extended, no trimming
required
3:07:45 Start tail 1/3 in 11:17:47 Good stability and control,
no trimming required
3:41:87 Start tail 2/3 in 11:18:25 Good stability and control,
no trimming required
4:02:41 Start tail full in 11:18:45 Good stability and control,
no trimming required
4:29:84 Start wings in 11:19:13 Good stability and control,
no trimming required
4:53:39 Start circuit retracted 11:19:36 Good stability and control,
greater roll control author-
ity, two circuits completed
(one high and one low), no
trimming required
5:41:28 Start wings out 11:20:24 Good stability and control,
no trimming required
6:06:93 Start tail out 11:20:50 Good stability and control,
no trimming required
6:45:18 Start land 11:21:29 Good stability and control,
good glide path and throttle
setting
7:31 Land 11:22:14 Good landing approach,
more elevator input needs
to be given to flare as
required
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.10. F4 - ENDURANCE TEST 216
I.9.4 Weather conditions
• Winds up to and including 20 km/h
• Gusts up to and including 30 km/h
I.10 F4 - Endurance test
I.10.1 Aim
Ensure capable of 30 minutes loiter with 500g payload
I.10.2 Procedure
1. **Ground Checks
2. **Electronics startup/shutdown (morphing included)
3. Position aircraft on runway
4. Taxi
5. Ground run
6. return to start on runway
7. Taxi
8. Take off
9. climb - straight line to trim altitude
10. trim (approx)
(a) test pitch response
(b) test roll response
11. Loiter velocity - slowly decrease until reach 65% throttle if possible (2 circuits)
12. deploy flaps → response
13. undeploy flaps
14. landing approaches without flaps
15. landing approaches with flaps (if required/possible)
16. land
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
217 APPENDIX I. FLIGHT TEST PROCEDURES
17. taxi
18. Thrust motor off
19. control switches off
20. place in secure position
21. Thrust motor on
22. set throttle at loiter position run until ESC cuts battery power.
23. Thrust motor off
I.11 F5 - Performance parameter tests
I.11.1 Ext Goal 1
A few short tests where the performance parameters are measured such as takeoff distance,
cruise speed & range, endurance, landing distance, dash speed, or turn rate. These tests
may be performed in conjunction with other tests previously mentioned. (Extended Goal
1) (TO speed, Max speed, loiter speed (65% throttle), range at max speed, endurance
I.11.2 Procedure
1. **Ground Checks - morphing range check included
2. **Electronics startup/shutdown (morphing included)
3. Position aircraft on runway
4. Taxi
5. Ground run
6. return to start on runway
7. Taxi
8. Takeoff (retracted wings, extended tail configuration)- (measure TO distance)
9. climb - straight line to trim altitude
10. trim (approx)
(a) test pitch response
(b) test roll response
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.12. F6 - DIFFERENTIAL SPAN ROLL CONTROL TEST 218
11. 1 circuit full throttle - (measure maximum speed)
12. 1 circuit cruise 85% throttle - (measure cruise speed and range)
13. 1 circuit loiter 65% throttle - (measure cruise speed and range)
14. On straightest part of circuit morph wings out fully.
15. 1 circuit full throttle - (measure maximum speed)
16. 1 circuit cruise 85% throttle - (measure cruise speed and range)
17. 1 circuit loiter 65% throttle - (measure cruise speed and range)
18. land
19. taxi
20. Thrust motor off
21. control switches off
I.11.3 Results
Due to aircraft damage and weather conditions this test was unable to be performed by
the 30th of Ocrober.
I.12 F6 - Differential span roll control test
I.12.1 Aim
To complete one circuit using only the morphing mechanism and associated change in
wingspan to complete one circuit without using ailerons.
I.12.2 Procedure
1. **Ground Checks - morphing range check included
2. **Electronics startup/shutdown (morphing included)
3. Position aircraft on runway
4. Taxi
5. Ground run
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
219 APPENDIX I. FLIGHT TEST PROCEDURES
6. return to start on runway
7. Taxi
8. Take off (extended configuration)
9. climb - straight line to trim altitude
10. trim (approx)
(a) test pitch response
(b) test roll response
11. cruise velocity
12. When on straightest part of circuit →
(a) Slowly bring Port wing 1/2 in
(b) See response of UAV
13. Trim
14. Turn loop - when on straightest part of circuit →
(a) Slowly bring Port wing 2/3 in
(b) See response of UAV
15. Trim
16. Repeat steps 15 and 16 listening to pilot to change wing extension until the correct
bank angle is achieved.
17. Extend Port wing out
18. One circuit with no aileron control
19. On straightest part of circuit - retract starboard wing to neutralise bank
20. Once neutral bank achieved - extend starboard wing to fully extended configuration
21. land and taxi
22. Thrust motor off
23. Control switches off, place in secure position
24. Thrust motor on
25. set throttle at 65
26. Thrust motor off
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
I.12. F6 - DIFFERENTIAL SPAN ROLL CONTROL TEST 220
I.12.3 Results
Unable to complete due to weather conditions and damage to UAV.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
J. Risk management Plan
Risk Management Plan for
859: Design and build of a UAV with morphing configuration
A University of Adelaide undergraduate project
Prepared by:
Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey
221
222
Introduction:
The purpose of this risk management plan is to identify and manage the risks associated with the
Adelaide University final year undergraduate project 859: Deign and build of a UAV with
morphing configuration, informally dubbed ‘The Morpheus Project’. This plan investigates first the
context of the project, followed by a detailed Risk identification table. Risk reduction strategies
have also been developed to reduce unacceptable risks to an acceptable level.
Context
Internal Influence
The internal factors of influence for The Morpheus Project have been analysed using the SWOT
method. This provides a framework to analyse the project’s structure, financial constraints,
obligations, and to determine their influence upon the project.
Strengths
All five members of the project team are very committed to producing an exceptional project to the
highest standard.
All five team members have worked well together in the past and are familiar with each others
strengths and weaknesses.
One group member has experience building and flying model aircraft.
Weaknesses
The group has no experience working on projects of this scale.
The group has very limited manufacturing experience.
Three group members are overloading to 125% subject load in the first semester.
Opportunities
The group has the opportunity to expand their knowledge in a variety of ways, including academic,
interpersonal, liaising with technicians, manufacturing methods.
The group has the opportunity to prove to themselves, their peers, engineering staff and sponsors
what they are capable of.
The group has the opportunity to obtain good academic results for their final year project.
The group has the opportunity to become aquatinted with and work with members in industry.
Threats
Should the project group not succeed, their honours grade will be affected
Should the UAV not be test flown in sufficient time for inclusion in the major deliverables of the
exhibition and the final report, this will significantly affect the success of the project.
Interpersonal issues would threaten the group and the outcome of the project if it is not quickly
resolved. Due to the size of the project group and the overlap of all tasks, it is important that the
project group can overcome any interpersonal issues.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
223 APPENDIX J. RISK MANAGEMENT PLAN
Should a group member become unavailable, or unable to contribute to the project for a significant
length of time, this will have a significant impact of the ability of the project group to complete the
project on time.
That the project is too ambitious to complete in the required timeframe.
External
The External factors of influence on the Morpheus Project have been analysed using the PERT
analysis method. This provides a framework to analyse external factors which could impact upon
the project.
Political
Potential political issues resulting from differences of opinion between workshop staff and the
project group, the department or the university, differences of opinion between the academic staff
and the project group, department, or the university.
Economic
Due to the economic crisis, it is possible that it will be more difficult than usual for the project
group to secure industry sponsorship.
Societal
It is important for the group to have a good working relationship with other project groups as this
allows for the exchanging of advice and ideas.
Technologic
The concept selected was deemed to have a technological level sufficiently low that the required
technology should be available to students.
It is possible that some components may be difficult to source due to the unique usage and size of
these components. It is also possible that interfacing components intended for different uses may
become an issue.
Stakeholders
The major stakeholders involved in the Morpheus Project are listed in the table below. Both internal
and external stakeholders are listed, along with the expectations of each stakeholder from their
association with the project, and the opportunities and vulnerabilities to the project from the
stakeholder. This information is summarised in Table 1.
Table 1: Stakeholders
Stakeholder Stakeholder
Expectations
Opportunities for the
project to be gained
by association
Project vulnerabilities
due to association with
the stakeholder
Internal
Group members
� To achieve a good grade for the project
� Successful completion of the project by
committed group
members
� Should one or more group members not
contribute sufficiently
to the project it is
possible that the project
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
224
will not be completed in
time.
Supervisor
� For students to complete the project
� For the project group to gain advice and
encouragement in
regard to the project
� Should the supervisor not provide sufficient
support and interest to
the project, it is
probable that the project
group will not succeed,
or will overlook aspects
of the design/project
process.
Workshop staff � To be provided with sufficient information
to allow for the
manufacture of required
components
� To have sufficient contact with the group
to ensure that the
manufactured
components are as
required.
� To provide manufacturing and
technical advice
� To provide quality components
�
� Should workshop staff not take an interest in
the project, it is possible
that this could result in
delays in the
manufacturing of the
project, potentially
affecting the ability of
the project to be
completed.
The school of
Mechanical
engineering
� That the school reputation is upheld
� Funding from the school (as provided to
all final year projects)
� The use of the school reputation,
� The use of the school’s staff and
contacts for advice
� That the project will not be taken seriously
by some suppliers due
to the idea that it is just
a ‘student project’
The University
of Adelaide
� That the university reputation is upheld
� The use of the University logo and
reputation
� That the project will not be taken seriously
by some suppliers due
to the idea that it is just
a ‘student project’
External
Sponsors
� To have their support recognised in all
deliverables
� To remain up to date on project progress
� Financing
� Advice
� Contacts within the industry
� Possible damage to the groups reputation
should the project not
succeed, or should the
sponsors be displeased
with the outcome.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
225 APPENDIX J. RISK MANAGEMENT PLAN
Suppliers
� To provide goods at a cost to the project.
� Possibly to gain further business from
the university or other
students by word of
mouth advertising.
� To supply off the shelf components.
� There may be delays in receiving the goods if
other work is deemed
more important than a
small, once off student
project.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
226
Risk Identification A comprehensive list of risks associated with the Morpheus Project were identified and analysed in
Table 3 to determine the consequences and likelihood of each risk. The categories used to analyse
these risks are as follows:
� CO,SEQUE,CES:
� Catastrophic 5: death or large number of serious injuries, huge cost, >1 month delay,
prevent the achievement of a primary goal
� Major 4: serious injury or extensive injuries, major cost, > 2 week delay, impacts upon
the extent of the completion of a primary goal or prevents the achievement of an extended
goal.
� Moderate 3: medical treatment required, high cost, > 8 day delay, impacts upon the extent
of the completion of an extended goal
� Minor 2: first aid treatment required, some financial impact, > 4 day delay, no impact
upon the project goals
� Insignificant 1: No injuries, low financial impact, <1 day delay, no impact upon the
project goals
� LIKELIHOOD:
� Almost Certain 5: expected to occur in most circumstances or could be expected to occur
for most components
� Likely 4: will probably occur in most circumstances or will probably occur for most
components
� Possible 3: could possibly occur at some time, or could possibly occur for some
components
� Unlikely 2: could occur at some time or could occur for some components
� Rare 1: may occur only in exceptional circumstances, or may occur to only a few
components
By assessing each risk using these categories, Table 2 was used to determine the ‘value’ of the
existing risk level. From this, the acceptability of each risk could be analysed. Table 2: Risk assessment matrix
Likelihood
Consequences
Catastrophic
5
Major
4
Moderate
3
Minor
2
Insignificant
1
Almost certain: 5 10 9 8 7 6
Likely: 4 9 8 7 6 5
Possible: 3 8 7 6 5 4
Unlikely: 2 7 6 5 4 3
Rare: 1 6 5 4 3 2
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
227 APPENDIX J. RISK MANAGEMENT PLAN
Table 3: Risk identification and analysis
Ris
k R
efer
ence
The Risk Source and Curent Liklihood Level Impact and Curent Consequence Level
Current control Strategies and
their effectiveness
(A) –Adequate
(M) – Moderate
(I) – Indadequate Cu
rren
t R
isk
Lev
el
Acc
epta
bil
ity
(A/U
)
1 Requirement to
re-design and/or
re-manufacture
components or
assemblies
Design errors which are not discovered prior to
manufacturing
This will probably occur in most circumstances
due to a lack of communication about different
design aspects. It is expected though that some
components will not require any re-design or
remanufacture
Likely (4)
Could result in significant delays if a major
component or assembally problem is
discovered.(possibly > 1 month). This could also
result in significant cost to re-manufacture the
comnents
Catastrophic (5)
Should a minor problem (i.e. Remanufacture of a
single conmonent dueto a fault found early in the
manufacturing phase), this cold still result in >2
week delay if workshop is required to
remanufacture a part. This could also result in
major cost to re-manufacture the comnents
Major (2)
Weekly meetings with the project
supervisor.
With present control strategies, it is
almost certain that some components
will require re-design and re-
manufacture as it is not possible for
all aspects of the design to be
discussed in these meetings, or with
the project supervisor.
Inadequate.
9 U
2 Delays in
manufacturing
whilst waiting
for the
procurement of
off the shelf
components or
components to
be delivered
deliveries
Postage delays, components not in-stock,
suppliers not delivering components on-time,
components not being procured with sufficient
time to arrive before they are required.
This could be expected to occur for some items.
Possible (3)
Could result in minor delays in the scheduale.
Should such delays occur, it is unlikley to impact
significantly upon the scheduale as most
manufacturing tasks ru in parallell. Also, should
this become an issue, alternative suppliers, or
express potage can be used to reduce the delay.
Minor (2)
The appointment of a procurements
and assemblies officer to manage all
the scheduling and procurement of
the long lead time procurements, and
the procurement of critical
components and items not readily
available off the shelf.
Adequate
5 A
3 Test flight
delays due to the
weather
Bad weather resulting in the flight delays
This could possibly occur for some test flights
Possible (3)
Possibly some delays which may impact upon
completion of the extended goals.
Moderate (3)
Scheduling of a back-up test flight
for each test flight.
Adequate
6 A
4 Inability of the
group to work
Unresolved differences or infighting. This could
result from sending too much time in each others
should this occur for a short time, this would
result in some very minor delays as it is still
Each group member desires to do
well. It is therefore up to individuals
6
A
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
228
Ris
k R
efer
ence
The Risk Source and Curent Liklihood Level Impact and Curent Consequence Level
Current control Strategies and
their effectiveness
(A) –Adequate
(M) – Moderate
(I) – Indadequate Cu
rren
t R
isk
Lev
el
Acc
epta
bil
ity
(A/U
)
together company, high levels of stress, insufficient sleep
or other similar reasons
It is reasonable to expect that this will almost
certainly occur for a short period at least once
during the project.
for a short period of time:
Almost certain (5)
for a long period of time:
Rare (1)
possible for group members to work
independently
Insignificant (1)
Should this occur for an extended amount of time,
this could have Major consequences, although
should this occur, strategies could be devised for
independent working
Major (4)
to ensure that hey put aside any
differences to concentrate on the
project.
It is the responsibility of the
Logistics coordinator to coordinate
the group and ensure that such
situations are avoided if possible, and
managed appropriately if they should
arise.
Adequate
5
5 Incapacitation of
a group member
for a significant
amount of time,
or one group
member unable
to complete the
project
This could occur due to personal reasons, or
major injury or illness. This would only occur
under exceptional and unforeseen circumstances.
Rare (1)
This would have a significant effect on the ability
to complete the project, and could possibly affect
the completion of the primary goals.
Catastrophic (5)
It is not possible to control this risk
as it deals with unforeseen
circumstances
6 A
6 A required
manufacturing
method
becoming
unavailable
This could be caused by a workshop machine
breaking down, or a backup of work in the school
workshop.
Possible (3)
This could result in delays in manufacturing,
either waiting for the required method to become
available, or in delays during re-design. This
could also have a high cost impact to outsource
the component.
Moderate (3)
Inbuilt lag time in the project
schedule.
Inadequate
6 A
7 minor damage to
the aircraft
during test
flights
Mechanical failure, electronic failure,
aerodynamic problems, pilot error, acts of god.
Given the past history of Adelaide university
UAV projects, and the complex structure,
mechanisms, and flight requirements of the
aircraft, it is almost certain that the aircraft will
This would result in minor delays in the schedule.
Possibly require the test to be re-conducted
Minor (2)
Inbuilt lag time into the project
schedule
Adequate
4 A
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
229 APPENDIX J. RISK MANAGEMENT PLAN
Ris
k R
efer
ence
The Risk Source and Curent Liklihood Level Impact and Curent Consequence Level
Current control Strategies and
their effectiveness
(A) –Adequate
(M) – Moderate
(I) – Indadequate Cu
rren
t R
isk
Lev
el
Acc
epta
bil
ity
(A/U
)
sustain some damage.
Minor (2)
8 major damage to
the aircraft
during test
flights
Mechanical failure, electronic failure,
aerodynamic problems, pilot error, acts of god.
Given the past history of Adelaide University
UAV projects, and the complex structure,
mechanisms, and flight requirements of the
aircraft, it is possible that the aircraft will sustain
some major damage.
Possible (3)
This would either result in significant delays and
major costs to re-built, or could impact upon the
completion of the primary goals and/or extended
goals
Major (4)
significant lag time in the project
schedule
Inadequate
7 U
9 complete loss of
the aircraft
during test
flights
Mechanical failure, electronic failure,
aerodynamic problems, pilot error, acts of god.
Given the complex structure and mechanisms,
and flight requirements of the aircraft, it is
unlikely rather than rare that the aircraft will be
completely lost.
Unlikely (2)
This would either result in catastrophic delays
and major costs to re-build. This could impact
upon the completion of the primary goals and/or
extended goals.
Catastrophic (5)
significant lag time in the project
schedule
Inadequate
7 U
10 Aircraft is
overweight
>7kg
Underestimation of component weight during
design, increase of weight due to unaccounted
design changes or repairs.
Weight of paint, glue, bolts etc. exceeding
estimations.
Requirement for increased structure, or different
materials due to simplicity, availability, and the
outcome of structural tests and calculations.
possible (3)
This will affect the aircraft performance, and will
require the aircraft to be certified for flight.
Certification will have some impact upon the
project schedule.
Insignificant (1)
This could either affect the completion of the
endurance goal, or require the purchase of new
batteries. This could either affect the completion
of a primary goal, or have a major cost associated
with it.
Major (4)
Lag time in the schedule to allow for
certification.
A weight budget is to be maintained
by the technical coordinator to
control the aircraft weight during
design and manufacturing.
Adequate
4
7
A
A
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
230
Ris
k R
efer
ence
The Risk Source and Curent Liklihood Level Impact and Curent Consequence Level
Current control Strategies and
their effectiveness
(A) –Adequate
(M) – Moderate
(I) – Indadequate Cu
rren
t R
isk
Lev
el
Acc
epta
bil
ity
(A/U
)
11 Aircraft is
overweight
>>7kg, such that
the ability of the
UAV to fly is
affected
Underestimation of component weight during
design, increase of weight due to unaccounted
design changes or repairs, weight of paint, glue,
bolts etc. exceeding estimations. Requirement for
increased structure, or different materials due to
simplicity, availability, and the outcome of
structural tests and calculations.
Rare (1)
This will affect the aircraft performance, and will
require the aircraft to be certified for flight.
Certification will have some impact upon the
project schedule.
Insignificant (1)
This could either affect the ability of the aircraft
to fly. This would have a significant impact on
the primary and extended
Catastrophic (5)
Lag time in the schedule for
certification.
A weight budget is to be maintained
by the technical coordinator to
control the aircraft weight during
design and manufacturing.
Adequate
2
6
A
A
12 Serious injury to
a group member
Not following correct safety protocol during
dangerous manufacturing or testing operations
Rare (1)
This could result in serious injury, or even death.
Catastrophic (5)
A safety officer has been appointed
by the group to look after safety
protocol, Safe Operating Procedures
etc.
Adequate
6 A
13 Minor injury to a
group member
Lack of proper care during manufacturing or
testing
Almost Certain (5)
The impact of this be minor first aide only
Insignificant (1)
General common sense and advice
from workshop staff and the safety
officer should be adhered to. This is
up to individual group members.
Adequate
6 A
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
231 APPENDIX J. RISK MANAGEMENT PLAN
Risk Treatment, Monitoring and Reviewing
For risks which are identified as unacceptable, either treatment methods, monitoring or reviews of
the risk should be implemented to bring the risk to an acceptable level. Possible methods to reduce
each unacceptable risk are discussed in this section.
� Risk # 1: Requirement to re-design and/or re-manufacture components or assemblies
This risk is based on the re-design and re-manufacturing required due to errors made during the
design phase. The main risk is associated with the project schedule and budget should these errors
not be discovered until the components involved have been manufactured.
To reduce this risk to an acceptable level, a design review should be implemented prior to the
commencement of manufacturing to ensure that all major errors and most minor errors are
discovered prior to the beginning of manufacturing.
The Technical coordinator shall be responsible for the design review.
This risk mitigation strategy will reduce the severity of the consequences to a level of Minor (2)
The likelihood will remain the same at a level of Likely (4)
This provides a new risk level of 6, and the risk is deemed acceptable.
The effectiveness of this strategy will be determined by the number of design flaws found during
the design review process.
� Risk # 8: Major damage to the aircraft during flight tests
This risk is based on the possibility of major damage to the aircraft sustained during a flight test.
The main risk is associated with the project schedule and budget should major repairs be required.
This could in turn affect the ability of the group to achieve primary and/or secondary goals should
the damage occur before the goals are achieved.
To reduce this risk to an acceptable level, the aircraft should be built to be easily repaired should the
need arise. A quality assurance officer should also be utilised to ensure that the aircraft is
manufactured to the design dimensions and requirements to ensure the highest possibility of
success.
These risk mitigation strategies would be the responsibility of the manufacturing coordinator and
technical coordinator to ensure the ease of repair, and the quality assurance officer to ensure the
quality of the manufacturing.
This risk mitigation strategy will reduce the level of the consequences to a level of Minor (2)
The likelihood will remain the same at a level of Possible (4)
This provides a new risk level of 6, and the risk is deemed acceptable.
The effectiveness of this strategy will be determined by the ease of repairs should they be required.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
232
� Risk # 9:Complete loss of the aircraft during test flights
This risk is based on the possibility of complete loss of the aircraft during a flight test. The main
risk is associated with the possibility that the primary and/or extended goals of the project would
not be met.
To reduce this risk to an acceptable level, the aircraft should be built to be easily re-built should the
need arise. This involves ensuring that spare components are either readily available, or already
manufactured. A quality assurance officer should also be utilised to ensure that the aircraft is
manufactured to the design dimensions and requirements to ensure the highest possibility of
success.
These risk mitigation strategies would be the responsibility of the manufacturing coordinator to
ensure the ease of re-build, and the quality assurance officer to ensure the quality of the
manufacturing.
This risk mitigation strategy will reduce the level of the consequences to a level of Major(4)
The likelihood will remain the same at a level of Unlikely(2)
This provides a new risk level of 6, and the risk is deemed acceptable.
The effectiveness of this strategy will be determined by the time required to re-build should this be
required..
Conclusion There are many several main risks which may affect the ability of the group to successfully
complete the Morpheus project. This risk analysis has shown that the majority of risks are not
significantly high enough to cause major concern in regard to the project outcome. It is possible to
put in place risk reduction strategies to reduce the level of risk associated with the higher level risks.
From the risks considered in this report, it is deemed that if the risk management strategies listed are
implemented, then the risks associated with this project will be reduced to a suitable level.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
K. Meeting minutesThe folowing documents outlined in this appendix are the official meeting minutes taken
during group meeetings with the project supervisor.
233
234
Meeting 1 - 8/12/2008
Meeting 1.1 Monday 8th December 2008
17:30-21:30 Meeting was held in two parts. The first, with Dr Maziar Arjomandi in attendance. The second part was an internal meeting.
Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Rachel Harch, Ian Lomas, Simon Mitchell, Carlee Stacey
Summary: Project was chosen to be MORPHING AIRCRAFT
Next meeting: With Maziar: Monday 15th Dec, 5:00PM Adelaide Uni
(Also another meeting Monday the 22nd at 5:00PM) Internal meetings: Friday 12th Dec, 5:30PM;
Sunday 14th Dec, 1:30PM
Actions before next meeting: Internal Meeting:
� Project definition
� Research morphing aircraft. o Summarise information. 2-3 lines on each document. For important
document, save/ copy entire document if possible well as summarise. o Need to review have 30-40 documents o Post research on the Google group.
� Consider who you wish to nominate for the positions of Logistics manager and Technical manager.
� Get a logbook (folder) Meeting with Maziar:
� To have chosen topic, and defined the project
� elected logistics and technical managers
� reviewed 30-40 documents
� Prepare a presentation aimed at potential sponsors (approx 10-15 slides). o This will include a general project definition (scope, technical tasks
e.g. size, fly at), as well as a bit of a literature survey.
� Prepare a list of sponsors o Rank them. 2 lists, in-kind sponsorship and cash sponsorship
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
235 APPENDIX K. MEETING MINUTES
Meeting Minutes: PROJECT SELECTION: Ideas discussed (in meeting with Dr. Maziar): � Submarine UAV (POSSIBLE)
o Difficulties: water proofing (manufacturing challenges), testing (logistics), VERY EXPENSIVE (approx $60,000-70,000 needed in sponsorship)
o Positives: Interesting, very engineering focused outcome o Main learning would be of manufacturing techniques,
Admin/management, obtaining sponsorship
� Endurance rotor (REJECTED) o Maziar will not support a rotor aircraft project o Insurance will not support o Very expensive (blades cost $1200 each)
� Varying Anhedral/dihedral (INCORPERATED INTO MORPHING AIRCRAFT)
� Morphing Aircraft (POSSIBLE) o Varying Anhedral/dihedral is a possible part of this project o Could possibly change fuselage length, wing span, horizontal tail span,
possibly could extend the goals to include changing aerofoil shape o Main learning would be aerodynamics and aircraft control o Will achieve if we simply have the wings move during flight. o Estimated cost $10,000. Biggest cost will be the mould.
� Blended Wing (REJECTED) o Would be Similar to fuel cell OR o Low aspect is too simple OR o High aspect is harder, but is mainly a control problem
Potential projects identified: Submarine UAV OR Morphing Aircraft Maziar prefers the Morphing Aircraft idea. Ideas discussed (in internal meeting): � MORPHING AIRCRAFT was chosen. � Vote was 5:1
o FOR: Kevin Chan, , Rachel Harch, Ian Lomas, Simon Mitchell, Carlee Stacey
o AGAINST: Crystal Forrester o There were no strong objections raised in regard to the Morphing
aircraft project. LOGISTICS AND TECHNICAL MANAGER � The two managers need to get along.
� These positions require approx extra 50% more time than other group members.
� By the end of the year, each person will have a management job.
� Will be chosen at the net internal meeting (Friday 12th of Dec 08)
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
236
� Logistics Manager: o Meeting minutes o Keeps track of contacts such as sponsors, documents o Keeps records of phone calls o Later becomes the financial manager.
� Technical Manager: o Coordinates technical decisions o Makes decisions about technical issues o Would be good to have some manufacturing experience
PROJECT INFOMATION: Contract � There is a contract which we need to sign
� probably not on access Adelaide yet- to be done next year
� Has 2 parts. o A project definition (inc. expected achievements and extended goals.
These cannot be changed unless entire group and supervisor agree). o The second part is submitted at the end of the year, and is based on the
achievement of these goals. Log book � Gets marked.
� Best to use a folder (thick folder) so can just add loose sheets of paper.
� Put everything in there. Including minutes, sketches, calculations, phone calls/e-mails to contacts
� This is like a time record of everything which you do.
� Do NOT include hard copies of everything you have read
� It is OK for it to be messy, but must be useful and readable
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
237 APPENDIX K. MEETING MINUTES
Meeting 2 - 17/12/2008
Meeting 2.1 Wednesday 17th December 2008
17:00 – 19:00
Attendance: Maziar Arjomandi, Richard Jones, Kevin Chan, Crystal Forrester, Rachel Harch, Ian Lomas, Simon Mitchell, Carlee Stacey
Summary: Main discussion was focused on sponsorship Also discussed was preparing a technical presentation which will identify 3 possible configurations and project definitions.
Next meeting: With Maziar: Wednesday 8th Jan, 5:00PM Adelaide Uni
(TO BE CONFIRMED) Internal meetings: Monday 22nd Dec, 5:10PM;
Actions before next meeting: Internal Meeting:
� Vote for the project online so the project can be closed off.
� Skim read the Aircraft design notes Meeting with Maziar:
� Run through of the sponsor presentation o Rachel and Kevin to do o Approx 15 min
� Present sponsorship letter
� Sponsorship list, including contact numbers
� Technical presentation / project definition o Given by Kevin (Tech manager) o 20-25 min o Discuss 3 configuration designs o Need project definition for each of the 3 designs
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
238
Meeting Minutes: Sponsorship Restructure presentation � Project information goes first
o Do not have a who are we slide etc. just list names on the first slide
� Motivation o Benchmarking o More info about why morphing wing; why they are important o Very active in UAV and larger aircraft o Mention names of people who are looking into morphing wing aircraft.
Particularly relate it to big companies � DARPAR � Boeing etc.
� Definition o Project is from scratch o Design… o This could be technical parameters
� Who are we- emphasise the university o School of mech. eng. has history of such projects, successful… o Sell the uni
� What we offer them o Logo on deliverables o Copy of report o Invite to exhibition o Tax write off o Recruitment method
� DO NOT give a clear description of cost. Leave all cost to talking.
� Ask them at the start to ask questions throughout the presentation Approaching the sponsors
� Approach sponsors carefully
� Easiest sponsors are the ones who don’t care what happens to their money.
� We CANNOT give them IP
� Choose companies we approach carefully o Some companies will not let you approach other companies o Big companies want to know who else.
� Go online, look for example of presenting to a company
� Lots of companies select people at exhibition
� Purpose of getting sponsorship is for us to sell our project to people outside university
Suggest we approach:
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� Australian Aerospace (maybe $5000?)
� Tales
� ASC
� Nova
� BAE
� Aeronautical engineers Australia
� Qantas
� Model flight (not now, but later. They often give a discount)
� Eccenture Approach people whom we have contacts for first. Other sponsorship information discussed
� Prepare a list o Contact numbers o Tailored letters o Tailor our motivations to the company’s values
Technical presentation � To be given by Kevin (Tech manager)
� 20-25 min
� Discuss 3 configuration designs o Sketches o Explain configuration o Have technical backup; particularly identify technical challenges (i.e.
to morph tail, wings, and fuselage, weight would be an issue.
� Hand sketches supported by rough calculations o Weight o Wing area etc.
� Different types of Morphing
� Need to find 3 configurations
� Remember, usually your firs idea is your best!
� Generate Bill of Material (BOM). o This will be simple now. o This will eventually become a very large spreadsheet o Includes everything. i.e. how many actuators etc.
� Project definition for each configuration. UAV should be 5-7 kg. MUST be under 7 to fly it. Course notes, Rainer and Roskam should help. Morphological Analysis � Investigate, give score, rank
� Solution selection analysis
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Meeting 3 - 8/1/2009
Meeting 3.1 Thursday 8th January 2009
Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Rachel Harch, Ian Lomas, Carlee Stacey
Apologies: Simon Mitchell
Summary: � Content of agenda and minutes � No further actions required to vote for involvement in the morphing UAV
project, except that Kevin needs to contact the coordinator to be added to the list � Sponsorship (letter, presentation, approaching the sponsors
Next meeting: With Maziar: Wednesday 21st Jan, 5:00PM Adelaide Uni Internal meetings: Monday 12th Jan, 5:15PM
Tasks before next meeting: Internal Meeting: Meeting with Maziar: Morphological analysis 3 concept designs Investigate propulsion (propeller vs. ducted fan)
Summary of Actions: Tasks to perform completed by Kevin Chan Contact Ben Cazz RE. to be put on the project list ASAP Crystal Forrester
Rachel Harch Ian Lomas Carlee Stacey
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Meeting Minutes:
1. Agenda / minutes: a. Agenda:
� Include in the agenda who will discuss what � Agenda is to be very similar from week o week � Technical agenda, therefore will have more detail
b. minutes: � needs to include a table of actions � summery � time � Should cover ‘who, when, where, what, how’ for all decisions
made. 2. Voting for the project:
� Cannot vote since we have been selected and locked in. � Kevin needs to contact Ben Caz to have his name put on the list of
final year students. � No one else needs to do anything in regard to voting.
3. Sponsorship: a. Letter
� E-mail Maziar a copy of the letter for checking � Cannot use the university logo. See marketing website for relevant
policy. � Include in the letter the information that sponsors will receive a
copy of the report � Letter is OK. � Will usually contact the person to find the correct person to send
the letter/e-mail to. � E-mail is more likely to be what we send- faster etc. but cal them
first. b. Presentation
� Slides: 1. Template:
a. it is OK to use the logo in this instance b. change the ‘UAV Project 2009 to Morphing
UAV 2009 or similar… can put name when we have decided.
2. slide 1 a. Put supervisor below team members
3. slide 2 a. needs a schematic
4. slide 3 a. remove some of the technical information b. perhaps just include the application
5. slide 4 a. Background is a problem. The graduated
background with the white picture does not work.
6. slide 5 a. very texty, but good info.
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7. slide 6 a. say what we want to do b. talk about how out of about 40 projects each
year, these UAV projects have been recognised by industry as some of the top projects
c. write what the projects actually were (i.e. not iSOAR
8. slide 7, 8 no comments 9. slide 9
a. need more photos demonstrating Teamwork b. Crystal obtained a copy of these from Maziar
after the meeting. 10. talk more about the seminar-list of external exhibition
juges etc. 11. 12.
� � � Rachel: Remember: you don’t need to repeat the information on the
slide. Forget what is on the slide. You do not need to use the best words as already written on the slide.
� Carlee: careful of fidgeting when presenting. �
c. Other � Crystal will be the main point of contact as she is the only person
with unrestricted phone access over the next few weeks. � Start contacting companies ASAP � When you phone the potential sponsor,
1. we are students looking for… 2. find out who you need to contact 3. contact them 4. e-mail
Restructure presentation 1. Project information goes first
a. Do not have a who are we slide etc. just list names on the first slide 2. Motivation
b. Benchmarking c. More info about why morphing wing; why they are important d. Very active in UAV and larger aircraft e. Mention names of people who are looking into morphing wing aircraft.
Particularly relate it to big companies � DARPAR � Boeing etc.
3. Definition f. Project is from scratch g. Design… h. This could be technical parameters
4. Who are we- emphasise the university i. School of mech. eng. has history of such projects, successful… j. Sell the uni
5. What we offer them
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k. Logo @ … l. Copy of report m. Invite to exhibition n. Tax write off o. Recruitment method
6. DO NOT give a clear description of cost. Leave all cost to talking. 7. Ask them at the start to ask questions throughout the presentation
Approaching the sponsors
8. Approach sponsors carefully 9. Easiest sponsors are the ones who don’t care what happens to their money. 10. We CANNOT give them IP 11. Choose companies we approach carefully
p. Some companies will not let you approach other companies q. Big companies want to know who else.
12. Go online, look for example of presenting to a company 13. Lots of companies select people at exhibition 14. Purpose of getting sponsorship is for us to sell our project to people outside
university Suggest we approach:
15. Australian Aerospace (maybe $5000?) 16. Tales 17. ASC 18. Nova 19. BAE 20. Aeronautical engineers Australia 21. Qantas 22. Model flight (not now, but later. They often give a discount) 23. Eccenture
Approach people whom we have contacts for first. Other sponsorship information discussed
24. Prepare a list r. Contact numbers s. Tailored letters t. Tailor our motivations to the company’s values
Technical presentation
25. Given by Kevin (Tech manager) 26. 20-25 min 27. Discuss 3 configuration designs
u. Sketches v. Explain configuration w. Have technical backup; particularly identify technical challenges
(i.e. to morph tail, wings, and fuselage, weight would be an issue. 28. Hand sketches supported by rough calculations
x. Weight y. Wing area etc.
29. Different types of Morphing 30. Need to find 3 configurations
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31. Remember, usually your first idea is your best! 32. Generate Bill of Material (BOM).
z. This will be simple now. aa. This will eventually become a very large spreadsheet bb. Includes everything. i.e. how many actuators etc.
33. Project definition for each configuration. UAV should be 5-7 kg. MUST be under 7 to fly it. Course notes, Rainer and Roskam should help. Morphological Analysis
34. Investigate, give score, rank 35. Solution selection analysis
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Meeting 4 - 21/1/2009
Meeting 4.1 Wednesday 21st January 2009
5:05-6:00PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Rachel Harch, Ian Lomas, Simon Mitchell, Carlee Stacey
Summary: Item 1: Recommendations regarding possible configurations were made based
on the calculations and the 4 checklists. the checklists need to be combined
Concept design and feasibility study should be done by February Item 2: BOM needs a lot of work Item 3: propeller vs. ducted investigation needs more research tractor propeller recommended Item 4: We need to start contacting companies The presentation is to be given at the next meeting Item 5: we also need to prepare the technical task
Next meeting: With Maziar: Wednesday 28th Jan, 5:00PM Adelaide Uni Internal meetings: 24th Jan, 11:00-5:00PM (as much time as you want to be there), Holdfast model aero club
Summary of Tasks Combine the checklists Complete the BOM (reformat and finish) Prepare lots of Concept sketches
1. work in pairs 2. each pair to present their top three 3. minimum of 9 analysed sketches to discuss at the next meeting
Give the sponsorship presentation at the next meeting. Technical specifications
Summary of Actions:
Tasks to perform completed by Kevin Chan ASAP Crystal Forrester
Rachel Harch Ian Lomas Simon Mitchell
Carlee Stacey Contact the school office to ask for access to the project room (S237)
ASAP
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Meeting Minutes:
2. 5:05PM – Morphing Selection: a. Calculations:
� Summarised by Kevin: 1. Wing Area >= Sweep 2. Sweep is still significant 3. Tail area/moment was not used in any calculations
� Resulting recommendations: 1. change area 2. if we can, change sweep 3. tail morphing should be selected using the other
parameters 4. tail is used only to counteract area/sweep
� Maziar’s comments 1. for small scale morphing, should probably look at area 2. sweep
a. impacts roll stability b. efficient for sub and super-sonic c. could look at a system for this d. changing the horizontal area requires similar
effort to changing the moment arm e. changing sweep of a rectangular wing is not
good. b. Innovation Checklist:
� Summarised By Rachel 1. Looked at the components 2. Tail arm seems more innovative 3. Need CG management when not using fuel- this seems
good for this 4. Folding wings are more innovative than telescopic 5. Sweep is not innovative
� Recommendations: 1. tail arm 2. folding wings
� Maziar’s comments 1. we need to compare the effect of a telescopic tail and an
increasing area tail on the drag coefficient 2. decreasing the tail might decrease the drag 3. telescopic tail effects vertical and horizontal surface
effectiveness c. Stability Checklist
� Summarised by Rachel 1. this could have been either very complicated or very
simplified 2. went with the simplified version 3. Tail arm and area are similar 4. sweep gives more stable rolling
� No other comments made d. Controllability checklist
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247 APPENDIX K. MEETING MINUTES
� Summarised by Carlee 1. overlaps a lot with Rachel and Ians checklists 2. has not yet been reviewed
� Recommendations 1. telescopic, tail moment arm adjustment, with or
without sweep depending on the complexity desired � Maziar’s comments
1. need to merge this with stability and manufacturing checklists
e. Manufacturing checklist � Summarised by Ian
1. includes manufacturing, tooling, labour and testing 2. fits in with the BOM
� Recommendations: 1. Tail arm and area are about the same 2. sweep is best wing change 3. telescopic is next best for wing change 4. folding is least preferred
f. Discussion about selection: � Next task is to select � We need to combine the checklists � Next week need to be 80% clear about
1. what the aircraft configuration will be 2. what the aircraft does 3. what it will look like 4. We will finish this by the end of February
� By the end of February, 1. concept design will be done 2. feasibility study will be done
� we will need to limit ourselves on manufacturing, but not just yet
� this week we need to make sketches. 1. produce as many sketches as possible 2. suggested that we work in 3 teams of 2 3. make sure all ideas actually work
a. i.e. shapes which can actually telescope-cylinder, not curve!
� Each pair should identify their top 3 � At the internal meeting we should rate the sketches. � Innovative ideas are good � We need to catch people’s attention on 3 occasions, at the
presentation, the exhibition, and in the report 1. Maziar’s example: cylinder with triangle wings and tail.
No ailerons, this way in one configuration you could have a delta wing aircraft
� For our project, we will need to make lots of aerodynamics calculations, but the first priority is stability and control.
1. We wll need 1-2 students who focus strongly on this. 2. Suggest we look at the text book by Nelson. This is
simple, and Roskam is old and scary text book.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
248
3. Nelson is the text book for aircraft design taught by Con Doolan in the 2nd semester.
3. 5:31PM- BOM: a. Pulse jet report is online b. Testing
� Best wind tunnel Adelaide uni has is the one in the Holden labs � To small � Forget about measuring the lift and drag or any other wind
tunnel testing � Can test by putting the aircraft on the roof of a car, but this is a
project on its own. Talk to Brad Gibson-he tried to do it, but did not have time
c. Number of hours � Was based on ISOAR and tailless and a bit more added � Maziar suggested that we cannot use ISOAR, fuel cell and
pulse jet are more appropriate for us to look at. d. What Maziar suggests we need to do:
� 3 types of components 1. structure
a. wing assembly i. raw materials, labour, tooling, testing
(structural testing included) b. fuselage assembly c. etc.
e. ten we can say how much f. The main point of the BOM at the moment is that we need to
understand the components involved.
4. 5:42PM- Propulsion: a. Ducted vs. propeller
� Summarised by Rachel and Ian 1. PowerPoint presentation prepared 2. Ducted fan recommended
� Suggested by Maziar that more investigation is required into: 1. battery weight (missed from the analysis) 2. ducted should need more batteries 3. should graph Power/weight(thrust load) vs. aircraft
weight 4. consider the rotational speed (safety) ducter~30-40,000
rpm, propeller ~8-10,000 rpm 5. ducted is less availible 6. look into engine controller price. This will make a big
difference 7. for low speed we should not need ducted. 8. we should choose the propeller based on weight 9. prop should be ½ ducted weightings 10. pilot is harder to get for a ducted propulsion system
b. Pusher vs. tractor � Summarised by Simon
1. pusher could cause take off issues as propeller can hit the ground
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249 APPENDIX K. MEETING MINUTES
2. tractor means prop has clean airflow, but fuselage does not
� recommends that tractor is better � no objections were raised
5. 5:52PM- Sponsorship a. Reported on by Carlee b. Letter
� Simon has completed � Still being reviewed � Maziar does not wish to see the letter again
c. Presentation � Rachel has completed � Still being reviewed � Wewill look at this at next weeks meeting.
1. usually 3 members go to a meeting (minimum of 2) 2. 1st person responsible or the general discussion, opens,
introduces, covers finance 3. 2nd person covers the technical stuff 4. 3rd person is an internal auditor- takes not of the
peformance etc. for the purpose of feedback. 5. 1st part if the slides is the technicl suff 6. 2nd hal is about us, and the money etc. 7. should be a 12minute presentation 8. should be presented next week
d. Phone prompts � Ian has completed � Reviewed by Carlee, but not yet by Kevin
e. Contacts � No contacts have yet been contacted. � We need to do that this week.
6. 5:56PM - Next weeks tasks/other business/Close a. Not previously discussed this meeting:
� What to morph 1. see other prototypes to gain an idea of why 2. area is usually changed for altitude 3. tail because they have to for stability 4. sweep for manoeuvrability and speed
� Prepare technical specifications list 1. the project definition will be taken directly from this 2. it should deal with specific numbers 3. we can prepare this for a larger project, but say that as
students we are creating a prototype as the why we are only doing … much.
4. Life impact needs to be included in the definition 5. clearly define the limitations
b. this weeks tasks: � Prepare lots of Concept sketches
1. each pair to present their top three 2. minimum of 9 analysed sketches to discuss at the next
meeting � Technical specifications
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
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� BOM � Presentation �
7. 6:00PM- Close:
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251 APPENDIX K. MEETING MINUTES
Meeting 5 - 28/1/2009
Meeting 5.1 Wednesday 28th January 2009
5:05-6:00PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Rachel Harch, Ian Lomas, Simon Mitchell, Carlee Stacey
Summary: Item 1: the presentation needs to be fixed up
sponsorship contacts need to be made. Item 2: Calculations need to be performed to determine the feasibility of the
concepts. A concept needs to be selected
Item 3: Technical task to be sent to Maziar for checking
Next meeting: With Maziar: Wednesday 28th Jan, 5:00PM Adelaide Uni Internal meetings: 24th Jan, 11:00-5:00PM (as much time as you want to be there), Holdfast model aero club
Summary of Tasks Fix sponsorship presentation Finalise sponsorship presentation Perform calculations etc. to determine which is the best concept Send Technical Specifications to Maziar Complete the BOM (reformat and finish)
Summary of Actions:
Tasks to perform completed by everyone Calculations to determine the feasibility of the
concepts
Kevin Chan ASAP Crystal Forrester
Send Technical Specifications to Maziar
Rachel Harch Ian Lomas Simon Mitchell
Carlee Stacey Contact the school office to ask for access to the project room (S237)
ASAP
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Meeting Minutes:
1. 5:05PM – Presentation: a. Feedback on the presenters:
� Crystal: 1. presentation was a bit boring 2. don’t use notes 3. had good eye contact
� Simon: 1. good intonation 2. no fidgeting
� General 1. generally the presentation was higher than average.
b. Slides and setup: � Slide by slide breakdown (number corresponds to the slide)
1. OK 2. needs more info added. Should put focus on morphing
structures. We cannot make an aircraft, so we will make a UAV
3. replace the picture of the P3 with predator or Global hawk, combine speed and manoeuvrability with high altitude
4. the first row is the previous generation of morphing aircraft. These A/C concentrated on sweep. The second row should be the second generation of aircraft. The dates need fixing, the reference dates and names can easily be confused with the dates and names of the aircraft.
5. Boring. Need to remove the first and last point, and just say the 1st point. Need to add in 2-4 pictures of nature as this is the focus of this slide (birds), concentrate on the benefits.
6. the reference for ISOAR is incorrect. The school has run the exhibition for many years for projects. Need to talk more about the school. Conc on the schooll, not the prizes. (aside note, EMCS courses are all open except aerospace, aero is the largest program in the school)
7. Should be more complete after todays meeting. Need to add in more slides to talk about what we want to achieve. We could include the sketches we have done so far.
8. same as or 7 9. talk more about the procedure. i.e. what we need to do
this year (design, build, test, present, report). Discuss how we receive support from the school for our project, however that for a project of the size of ours, we require external support
c. General:
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253 APPENDIX K. MEETING MINUTES
� We need to have a general cost breakdown etc. before we go to talk to them, however we don’t present it, but need to have it to discuss in case it comes up.
� If they ask how much we want, we need to know what our answer is.
� We need to finalise the strategy during our next internal meeting.
� We should be asking for $15-$20,000, if we get $10,000 it should do.
2. 6:35PM Concept: a. Telescopic:
� Rachel and Ian’s concept b. Sliding Sheaths:
� Crystal and Simon’s Idea c. Blackbird:
� Kevin and Carlee’s Idea � Fuselage could be shaped, but we still need roomfor the
payload d. Next week:
� Suggest that we look at what each concept can give us. � Consider:
1. can it generate sufficient lift 2. can it remain longitudinally stable (the blackbird has the
A/C of a rocket. 3. what else can we change/ other things (i.e. blackbird
idea can also be cannon launched, sliding has many different configurations)
� List what we get out of it. (subjective is OK for this, this is one ranking for consideration)
� Determine which ones can satisfy the requirement for flight (we can easily say that the telescopic idea will satisfy this criteria). We need to consider
1. lift 2. longitudinal stability 3. we will need to do some calculations, but not matching
diagrams. a. Simple moment calculation (consider forward
most configuration of the CG. This must balance)
b. Consider the maximum and minimum position of anything.
4. lateral balance a. look at the area b. lok at other aircraft
5. consider for high-speed, low speed, and takeoff configurations. (i.e. for 150km/hr cruise)
6. we can in some considerations use morphing for control � Kevin has been asked to also look into the biplane idea. Biplane
effectively increases the aspect ratio of the wing. � Next week we must be able to say which is the best aircraft
(and it could be a combination of the different ideas.
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254
3. 6:05PM- Technical task: a. Send this to Maziar to review.
4. 6:10PM- Close
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255 APPENDIX K. MEETING MINUTES
Meeting 6 - 5/2/2009
Meeting 6.1 Thursday 5th February 2009-02-08
5:10-7:00 PM Chairperson: Crystal Forrester Attendance: Kevin Chan, Crystal Forrester, Ian Lomas Apologies: Carlee Stacey, Simon Mitchell Summary: Meeting covered discussion on sponsorship strategy and progress, and calculations and feasibility of the three concept designs. Next Meeting: With Maziar: Wednesday 10th February 5:00-6:00 pm Allocation meeting: Wednesday 10th February 6:00 – 7:00pm Internal Meetings: Friday 7th February 5:15 – 7:00pm Summary of Tasks:
1. Concept feasibility a. Calculations b. Research if required c. Written paper of rejected concept.
2. Sponsorship a. Letters (due by Saturday night) b. One company presentation meeting by Wednesday 10th February
Summary of Actions:
Tasks to perform completed by everyone Further calculations on concepts Wednesday Sponsorship letters with contacts to Carlee Saturday night sponsorship letters to companies and arrange
presentation ASAP
Kevin Chan Crystal Forrester
Ian Lomas Send sponsorship presentation to Maziar ASAP Simon Mitchell
Carlee Stacey
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256
Meeting Minutes: 1. 5:10 Meeting Started
2. Sponsorship:
a. Presentation: • TO BE EMAILED TO MAZIAR FOR REVIEW • Discussed changes that had been made:
• Made changes as discussed last meeting • Inserted 3 slides for the current concepts • Changed the aircraft comparison slide to a Global Hawk and changed the data to
a discussion on the different attributes
b. Sponsorship Strategy: • NEED AT LEAST ONE PRESENTATION TO BE MADE BY NEXT
WENDESDAY • ASC started - suggested that they won’t sponsor the project
- Maziar said to look into their $2000 sponsorship from last year - Need to show them that they need to support MORPHEUS if they
want good engineers. • Crystal is not working at the moment and can be a second speaker for a presentation • Ian finishes work at 2:30 on Fridays so can help with a presentation • As soon as the company contact is found, give them the letter and arrange a meeting
ASAP.
3. Concept Calculations and Feasibility: a. Kevin and Carlee’s Concept (Rocket cruise concept):
• Moving the tail with the wing doesn’t work – tail moves back as wing moves forward, therefore use solid fuselage with T-tail that slides over the top
• In swept back configuration at 90kph cruise the CL=0.4 at 3 degree angle (flat plate) for W = 8kg
• Beyond 50 degrees sweep the UAV becomes unstable (via scissor graph) • “Scissor Graph” = plot AC and CG position with Sweep => traditionally used to
calculate the area of the horizontal tail • Don’t know what happens between 80->90 degrees sweep • BENEFITS
• High speed configuration, launch from torpedo or tall launcher, short takeoff distance, high manoeuvrability and high ceiling.
• MAZIAR • This design fail for now. • Suggested using a hinge and rotation mechanism for the wings (i.e. move root
forward as sweep backwards) – similar to the Bell X-5 – modern aircraft not use this as they have a high wing load (e.g. F1-11) and don’t want the wing inside the fuselage
• Use table of weights in the Aircraft Design II lecture notes – last lecture. • AC not at 25% MAC. • Should assume AC = 30% fuselage and 25% wing. • Has Ian’s span changes but also has maximum change in AR.
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257 APPENDIX K. MEETING MINUTES
• TO DO • Calculate roll rate - refer to flight control textbook by Arthur Nelson • Need to fix the wings and try calculations again, look at fuselage which is
mainly flat plate, look at Clark airfoil (1 surface) • Calculate/ prove all “Benefits”
b. Ian’s Concept (Standard Configuration Area and Tail Arm morphing concept): • Will definitely flu – has plenty of lift • Minimum cruise is 90kph • Stall Speed: small area configuration Vst = 50 kph
Large area Configuration Vst = 20 kph • Control Surface area is acceptable although there may be a problem with actuation • Possibly will not need to move the tail – check double span = double distance • BENEFITS
• Shorter takeoff and landing distance, higher ceiling • MAZIAR
• Need better calculations for position of tail - Do equation of moment around the CG.
• Look at Aspect Ratio => higher when expanded therefore lower induced drag. • Calculate/ prove all “Benefits” • Look at the frequency response of linear actuators and see if they can be used as
ailerons. i.e. use movement for roll control. • Calculate how much area change is required for control (see Roskam II and
Aircraft Design notes)
c. Crystal and Simon’s Concept (Delta-wing sweep and tail morphing concept): • Approximately half the wing consists of the sliding plate sweep mechanism. • Mechanism area assumed to be a flat plate with ½ CL of airfoil (rest of the wing) • Calculated for both square and triangular wings as unsure how the triangular section
will affect the results. • For CL= 1.2 at take off the takeoff speed needs to be 60kph • For cruise speed of 90kph the CL = 0.9 for triangle wing and 0.72 for square wing.
This is very high. • MAZIAR
• CL,cr is very high • The entire wing section containing the sweep mechanism (for the entire chord
length) produces minimal lift as you can’t have an airfoil shape. • It is not possible to do the delta wing configuration without wind tunnel testing
as there is a double airfoil shape due to the wing then tail sections. Problems also arise due to very small spacing between wing and tail – cannot calculate this properly.
• Fix calculations with new area data • Recalculate the tail position (balance around CG) • This concept not work -> wrap up the concept with a 2->3 page report on why it
doesn’t work.
4. Technical Task:
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258
a. To be discussed at next meeting.
5. Other discussions: a. Team member change (Rachel change projects)
• Shouldn’t have lost a day – having only 5 people will not greatly affect the project • Technical Coordinator and Logistics Coordinator were both re-elected
Logistics: Carlee Stacey Technical: Kevin Chan • We need to be more open with each other and let each other know of any problems
that are arising. b. Technical Coordinator
• Should work 20% more than other members – this will not give any higher marks at the end of the year.
• Need to be very tolerable of other’s opinions • Should expect to be arguing with team members more than other people
6. 6:10 Meeting Close 7. Debrief and arrange allocation meeting 6:10 – 7:00
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
259 APPENDIX K. MEETING MINUTES
Meeting 7 - 11/12/2009
Meeting 7.1 Wednesday 11th February 2009
5:05-6:00PM Attendance: Maziar Arjomandi, Kevin Chan (arrived at 5:40), Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey(arrived at 5:10)
Summary: Item 1: We need to contact more potential sponsors. Item 2: Calculations need to be performed to determine the feasibility of the
concepts. A concept needs to be selected
Item 3: Technical task to be sent to Maziar for checking
Next meeting: With Maziar: Wednesday 18th Feb, 5:00PM Adelaide Uni Internal meetings: Monday 16th Feb, 11:00-5:00PM Adelaide Uni
Summary of Tasks Find more sponsors Fix up Calculations done this week. Finalise calculations to consider ruling out the rocket/plane idea Consider forward sweeping Determine the configuration for each phase of flight for the telescopic idea (inc. AR, area, matching diagrams)
Summary of Actions:
Tasks to perform completed by everyone Research potential Sponsors Kevin Chan Crystal Forrester
Rachel Harch Ian Lomas Simon Mitchell
Carlee Stacey
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Meeting Minutes:
1. 5:___PM – Sponsorship: a. Australian Aerospace:
� Ian has given Tony Bernardo the letter � They seem more interested in in-kind support � Want to talk to us in March. � Want a cost breakdown to see what their money would be
going to � We do not really have much use for in-kind support. This can
generally be gained directly from the as we need it. b. QANTAS:
� We should see the uni website/ the Adelaidian for contacts in QANTAS (in regard to the deal Adelaide uni did with QANTAS last year
� c. ASC
� Letter submitted, but very doubtful that we will get any money. d. Eccenture
� Crystal will start this weekend e. Thales
� The ‘Big Boss’ is coming to SA on Friday, Simon will talk to him then
f. TRY MORE COMPANIES � At least 20-30 (try small companies as well) � Avionics � Electronics:
2. 5:15PM Concept: a. Telescopic:
� Research � Calculations
1. generally confirmed the research, a. Finding the proper AR is usually an optimisation
problem i. see the lecture notes
2. range calculation disagreed a. for electric motor, constant fuel weight b. simplifies the equation c. changes stepwise e.g.
i. 1 cell = 1 hour ii. 2 cell= 2 hour… iii. As the range increases, L/D changes
(due to weight change of more batteries required)
3. Polar drag a. Simplified calculations may not cover
everything. 4. Takeoff and Landing
a. Takeoff requires Increased area for inc. lift b. Landing requires inc lift & inc drag
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261 APPENDIX K. MEETING MINUTES
c. For landing, AR is not good because of the wing load.
5. Cruise a. Low area for inc. manoeuvrability and speed.
� Need to do a sensitivity analysis � Need to look at A/C control � No ailerons/flaps gives no coefficient of lift change which can
affect the takeoff and landing b. Sliding Sheaths:
� Is still being written up (finalising the rejection of the idea) c. Rocket plane:
� It works in a Canard configuration. 1. Canard configuration is unstable 2. Maziar has not seen a stable Aircraft with canard
configuration 3. AC needs to be trimmable for all phases of flight. AC
and CG, elevators 4. If we look at canard configuration, it must be tested in a
wind tunnel. 5. Look at the
a. S-37 Berkut ( spelling of this name may not be correct!) (this has forward sweep with a canard
b. Variez c. Velocity
� Works if we load up the nose 1. this is not practical as it is difficult to nose up (i.e. for
takeoff) � Look at morphing the wing forward � Difference between rockets and aircraft is that a rocket used
electronics for stability, aircraft does not d. e. Next week:
� Put 80% effort into the telescopic idea. � Need to look into what else we can do with the telescopic idea. � Look at what other ideas can be used in combination with the
telescopic idea. Use the telescopic idea as a basis, and build on it. � � Determine the configuration for each phase of flight
1. wing area 2. aspect ratio 3. matching diagrams 4. (we should have a strange matching diagrams since our
aircraft shape is changing) 5. look at roll rate 6. need to have a good sketch (would be good to have it in
CAD). We need a 3-View and Isometric drawing � Document the Rocket design � Look into forward sweeping for rocket design
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262
3. 6:05PM- Technical task: � Need to look up the standard requirements � Need to be specific � Look at CASA 101
1. This has the required safety factors. � look at the other projects � This is what is given to the customer � Write what will help deliver the product we are after � The specifications should not be too limiting � i.e. ‘the aircraft should provide power for the payload and the
platform’, OR ‘there should be 2 isolated power sources, one each for the payload and the platform’
� limit the temperature 1. above 40˚C you need special electronics (inc. cost) 2. above 60˚C cannot use composites
� Simon is to be the ‘bad guy’ at the meeting 1. Question everything
� Consider do we want the a/c to be built in modules for airfield repairs?
� Number of hours of flight before maintenance is required? � All info should come from the technical task � Find pilots standards (re. weight) � Reference where everything came from � Payload weight
1. Try a camera(inc. battery) 0.5 kg – 0.3kg total weight for this.
4. 6:00PM- Close
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
263 APPENDIX K. MEETING MINUTES
Meeting 8 - 18/2/2009
Meeting 8.1 Wednesday 18th February 2009
5:05-6:10PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey,
Summary: Item 1: Sponsorship. We need to contact more sponsors Item 2: more calculations need to be performed as well as matching diagrams,
and a 5-view sketch Item 3: Technical task is to be progressed
Next meeting: With Maziar: Wednesday 4th March, 5:00PM Adelaide Uni Internal meetings: Wednesday 25th Feb, 5:00-6:00PM Adelaide Uni
Summary of Tasks Find more sponsors Calculations. Feasibility study Technical task
Summary of Actions: For a comprehensive outline of tasks, see meeting 8.2 (allocation)
Tasks to perform completed by everyone Kevin Chan Crystal Forrester
Rachel Harch Ian Lomas Simon Mitchell
Carlee Stacey
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
264
Meeting Minutes:
1. 5:05 PM – Sponsorship: a. Thales
� Simon is feeling less confident about getting sponsorship from Thales
� Thales apparently prefer reimbursement rather than just giving out money.
1. this is fine as long as we have an official letter sent to Maziar stating this. i.e. ‘Thales will sponsor the Adelaide uni final year project morphing UAV group up to a cost of $____’
b. NOVA Aerospace � Crystal has contacted Nickelov.
c. Eccenture � Crystal is talking with HR to find out who to talk to.
d. BAE � Simon has e-mailed Ian Touey. He should be back on the 23rd
Febuaryb, and Simon will talk to him in person not long after that at the AIAA meeting.
e. Babcock � Ian has contacted them � They seem interested
f. QANTAS: � Ian � Need to find another contact. We cannot get in touch with the
contact we currently have. g. Australian Aerospace:
� Ian � No progress to note
h. ASC � No progress.
i. TRY MORE COMPANIES � NEED to contact Aeronautical Engineers Australia. � NEED to look for even more companies again.
2. 5:15PM Concept: a. We have been calculating the static margin incorrectly.
� We need to have a clear understanding of the static margin, what it is, and how to calculate it.
� Also the meaning of neutral point and aerodynamic centre b. Sweeping:
� Need to re-look into it once we are calculating the static margin correctly.
� Should look at using a canard and a tail. We need to consider the angles of attack and the elevators (consider the change required for change in trim
� We need to finalise this concept. � For old static margin calculations;
1. Have found a way to make it work.
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265 APPENDIX K. MEETING MINUTES
2. Possible with tail 2/3 way down body, and wing root at 850 mm.
3. other possibility to consider is using a tail and a canard � look at the Clark aerofoil for the fuselage
c. Telescopic: � Determined this week:
1. the aircraft can take off, fly around etc. 2. we need numbers
a. get these from the TS and the stat analysis � We want a sensitivity analysis.
1. i.e. by changing this, we get double, half, etc of the …. (altitude, lift, drag, stall speed etc.) compared with the other design
2. we need to produce comparative matching diagrams. a. show on same diagram what we get in
i. configuration 1, ii. configuration 2 iii. a normal aircraft
3. the actual matching diagrams will be done later. We are still looking at the feasibility of the aircraft at the moment. We will do actual matching diagrams later.
4. We just want to know what happens if we change the span by X amount
a. see aero 1 notes (flight profile, altitude vs. Velocity)
b. Velocity vs altitude (with constant load factor (n) line and corner speed
5. We need the 2 configurations to give 2 different areas, which exceed the area given by the ‘normal’ aircraft.
a. Need to consider: i. the performance parameters
ii. the load factor iii. weight (morphing is heavier than
normal) iv. Compare with a non-morphing aircraft v. Show what we can get from this aircraft
� Non-conventional roll control. 1. rotating wing:
a. usually connected by a spar b. not applicable to a large scale aircraft
i. it should be scalable c. look into this perhaps for trim only
i. not sure if this is used for large scale aircraft or not.
d. This would result in a single point of contact. Single points of contact are very expensive on a large scale aircraft.
� Telescopic control 1. deflecting the aileron changes the CL, rolling the
aircraft
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266
2. need to determine for a normal roll rate how much the change in CL is, and then determine how much change in area (and hence length) is required,
3. Determine the actuation speed required for similar control to a normal aircraft.
� We need to construct a 5-view drawing 1. this should be done with a CAD package (we will need
to select a program to use) a. when choosing compare the draft program in
Catia and pro-E (uni no longer has a solid edge licence)
b. we need to establish a good base c. talk to ex-students to figure out which is better d. CEASAR was very successful in their drawings
(pro-E) e. Not much difference in programs, depends on
our level of expertise. f. One of us will have to focus on the drawings g. Fuel cell also had good drawings (solid edge)
� We need to consider the technology involved in manufacturing the wings and fuselage
3. technical task a. needs to be progressed
4. 6:10PM- Close
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
267 APPENDIX K. MEETING MINUTES
Meeting 9 - 4/3/2009
Meeting 9.1 Wednesday 4th March 2009
3:10 - 4:05PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey,
Summary: Item 1: Sponsorship. Updates Item 2: Technical Task Item 3: Other Business Consider conferences Look at paperwork requirements Gant Chat Item 4: FEA/CFD To be further discussed at the next meeting Item 5: Concepts FLYING BODY IDEA WAS REJECTED TELESCOPING WINGS IDEA WAS SELECTED
Next meeting: With Maziar: Wednesday 11th March, 5:00PM Adelaide Uni Internal meetings: Monday 9th March, 10:00-11:30AM Adelaide Uni
Summary of Tasks Continue with sponsorship Sensitivity Analysis Matching diagrams 5-view Technical task Gant chart Paperwork for project (see my uni) SELECT A DRAWING MANAGER
Summary of Actions: For a comprehensive outline of tasks, see meeting 8.2 (allocation)
Tasks to perform completed by everyone Kevin Chan Finalise flying body concept Crystal Forrester
Ian Lomas Simon Mitchell
Carlee Stacey Gant Chart Find forms on MyUni which require completion
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Meeting Minutes:
1. 3:10PM – Sponsorship: At our next internal meeting, we need to further discuss the good and the bad of the presentation. Especially the bad so we can improve. a. Babcock
� Presented to them on Monday � They are not an aerospace company, so they will not sponsor us
a lot, but the indication was that they would sponsor us some money � General indication was that it was a good presentation � They would like to have seen a brochure � We should prepare a brochure and send it to them.
b. BAE � Have been given a contact in Melbourne � Simon chasing this up
c. Australian Aerospace: � Trying to arrange a time
d. Aeronautical Engineers Australia � Waiting for a call. � The person we need to speak to is away until the end of next
week. � Maziar has suggested that we should contacting Mat Mulner
(crystal knows who he is), and ask for a time to go and talk to them. e. Thales
� Have sent Simon an e-mail saying that they have not forgotten us, but that they require more time.
� Simon is feeling less confident about getting sponsorship from Thales. He is receiving the e-mails they are sending between themselves to discuss possibly sponsoring us (since they are ‘replying to all’. It does not look promising.
f. Boeing � We have been reaching dead ends.
g. NOVA Aerospace � No Progress.
h. QANTAS: � No progress
i. ASC � REJECTED.
j. Eccenture � REJECTED.
k. TRY MORE COMPANIES � NEED to contact Aeronautical Engineers Australia. � NEED to look for even more companies again.
2. 3:20PM – Technical Task a. Revision progress
� We have made significant changes. � This is still to be sent to Simon � Once Simon has reviewed, and the draft has been further
modified, this is to be sent to Maziar
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269 APPENDIX K. MEETING MINUTES
� This needs to be sent to Maziar before the weekend if we wish it to be reviewed before the next meeting.
� b. Included regulations
� We have used CASA 101. Not many regulations actually apply to aircraft of the size we intend to build.
� Talk to Todd Sandercock (other UAV project). He is a pilot and should know what documentation applies.
� FAR23 61kt max stall speed �
3. 3:25 - Concept: a. Statistical Analysis
� Stall speed � T/O distance � Look at models and try to define a mission
1. We want two objectives since we are morphing 2. Use this to obtain numbers from the statistical analysis
� Can look at previous years UAV’s to give us an idea. b.
4. 3:30 - Other Business a. Conferences
� AAEE conference is in Adelaide this year (therefore cheap for us)
1. Students often present papers 2. They want academics involved in the papers though 3. This is NOT part of project, but an external thing we
can do. (‘ this does count somehow, but it does not contribute to our marks’)
4. The best combination for writing paper is 3 people (1 is supervisor), but could have 2 or 4 (we would probably have to work in teams)
5. Conference is about education 6. Topics should be relevant to project, but we should
actually have to do some extra work. 7. Paper is easy to write, and there is lots of time 8. Maziar believe that our group on average can find the
time to do this! 9. This could help us out in regard to job applications 10. Possible ideas include things such as:
a. The importance of the decision making process in the engineering environment
b. Teaching the next generation communication c. Importance of communication in project based
learning d. Statistical analysis of something is always good.
(Ben and Brad presented this type of info. They wrote a simple questionnaire and got everyone to answer it)
e. What information do we like most (and what do we want to know about (possible questionnaire
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
270
about which university values are most important to students
f. Last year they focused on Quality Control g. We have lots of time to consider this. h. Check out the AAEE website
� AIAA student Conference 1. There is not enough quality at this conference 2. If we wish to, we can consider submitting something to
this conference. 3. We need to find out the dates. 4. Simon receives e-mails about this. As of yet, nothing
has been organised. b.
5. 3:40 - concepts a. We need to choose a concept to go on with. We need to progress. Also we
need to know what we are doing so we can come up with a project for our CFD and FEA courses.
� Although we will do CFD and FEA, we will not rely on these numbers. They will only really give us pretty pictures. They will be used only to further support our Hand Calculations
b. Flying Body: � Stable between 0-80degree sweep. 80% confident it will remain
stable for the rest � Method of calculations and assumptions was discussed
1. Model body as wings and wings as strakes 2. We can achieve stability without moving the tail. 3. Issue: strake<<<<less dominant that the wing. This is
not really true in our design. � CONCERNS
1. Manufacturing 2. All calculations required are unknown. There is not a lot
of information out there in regard to equations for flying bodies
3. Without the use of a wind tunnel and CFD we are not sure if it will actually fly
c. IDEA SELECTED: � We will go with the TELESCOPIC WINGS idea � Flying Body Idea has been REJECTED
1. The idea is interesting 2. Usually the stability of tailless aircraft is a function of
shape, and an ‘s’ shaped wing is required. This is the same for a wingless aircraft.
3. It can be calculated, but what we determine will be different to the actual shape that we build. This therefore requires wind tunnel testing in order to find out the effects of the shape that we actually built.
4. One small mistake could result in catastrophic failure 5. The idea is therefore not feasible
d. Telescoping �
6. 3:50 - Other Business
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271 APPENDIX K. MEETING MINUTES
a. We need to produce a Gant Chart � 1.5 months before we start manufacturing � Beginning of the mid-semester break we need to start
manufacturing the UAV � Mid semester break- finalise all the design and manufacturing � Mid-year break – completely complete the product � End of mid-year break we need to start testing � Nothing but writing to be done after the mid-semester break (all
testing to be completed) b. Look at MyUni for the various forms which need filling out
� SPPA (student participation ___ agreement) (3 copies, very important)
� Project definition c.
7. 3:55 - Progressing the concept (Telescoping) a. Roll methods:
� We can get differential roll � Double wing area and get equivalent of 12.5˚ aileron deflection � Calculations are wrong.
b. Before the next meeting e need to: � Complete the matching diagram for both configurations � Aerofoil selection � Propulsion selection (including the propeller) � Sort out the roll rates � Lots of sensitivity analysis
1. Determine the sensitivity of EVERYTHING. Look into everything in the A/C design notes, and then more.
2. We what to know the sensitivity of all aspects toward morphing.
� 5-view drawing IN A CAD PACKAGE c. WE NEED A DRAWING MANAGER
� They should establish a good filing system now, to save time later
� Once this is done, 1-2 people will sit and help the drawing manager to make the components
d. BOM e. Start looking into the detailed design of components
8. 4:03 - CFD/FEA a. We will not trust this for the project. b. Everything is to be doubled up with hand calculations c. Design of the spar d. Cannot use FEA on any composite structure. It is too hard, and we will not
achieve usable results. e. Could look at adding bush resin to the spar ( method used to bolt things to
composites) f. We will discuss this more at the next meeting. g. Abstract for FEA project is due next Friday.
9. 4:05 - Close
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
272
Meeting 10 - 10/3/2009
Meeting 10.1 Wednesday 10th March 2009
3:00-4:15PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey,
Summary: Item 1: Sponsorship. We need to contact more sponsors Item 2: more calculations need to be performed as well as matching diagrams,
and a 5-view sketch Item 3: Technical task is to be progressed
Next meeting: With Maziar: Wednesday 18th March, 3:00PM Adelaide Uni Internal meetings: Wednesday 10th March, 4:15-6:00PM Adelaide Uni
Tuesday 17th March, 10:00-11:00AM Adelaide Uni
Summary of Tasks � Sponsorship � Matching diagrams � Sensitivity analysis � Performance calculations � Drawings (using CAD package) 5- view requested, expected to complete a 4-
view due to time constraints arising from the Avalon trip � Technical task (to be completed) � The required forms are to be signed � Gantt chart draft to be completed � Table of values to be generated (as used in existing calculations) (low priority) � Decision matrix used in the propulsion selection to be e-mailed to Maziar � Generation of ideas for morphing mechanisms
Summary of Actions: For a comprehensive outline of tasks, see meeting 10.2 (allocation)
Tasks to perform completed by everyone Kevin Chan Crystal Forrester Ian Lomas Simon Mitchell Carlee Stacey get forms to everyone to sign
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
273 APPENDIX K. MEETING MINUTES
Meeting Minutes:
1. 3:00 PM – Sponsorship: The only thing we cannot advertise at the exhibition is alcohol. a. Avalon:
� Usually big companies prefer everything o go through the university
� Concentrate on smaller companies � Sponsorship is a tax deduction for local companies only.
b. Aeronautical engineers Australia � Got Matt Maloney’s phone number
c. Boeing � No further progress
d. BAE � Want a business plan � Usually this is how much a project will earn. We are students
and therefore this is not so relevant to us � Prepare and send to them instead a grant application. � BAE does not have much money at the moment. � Leave this for a couple of weeks, see ho we go, and if we still
need money, we can ten try and chase them. e. Red Bull
� Tell them how many people and how many times their logo will be shown
� Mention that Hungary Jacks sponsored pulsejet last year � In 2006 iSOAR spoke to red bull. They were given cans of red
bull f. Australian Aerospace:
� No response yet from Tony Bernardo g. s
� Simon is feeling less confident about getting sponsorship from Thales
� Thales apparently prefers reimbursement rather than just giving out money.
1. This is fine as long as we have an official letter sent to Maziar stating this. i.e. ‘Thales will sponsor the Adelaide uni final year project morphing UAV group up to a cost of $____’
h. NOVA Aerospace � Rejection.
i. Babcock � Ian will follow up
j. QANTAS: � No Response.
k. Virgin � 2-3 weeks we should know.
l. Tiger � We should contact
m. AIAA � We should contact.
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274
2. Forms to be filled out: a. These must be completed and signed before the next meeting
� SSPP � 1-5 form
1. Fill out all 5’s. If we think something is not a five, bring it up at the next meeting.
� Contract 1. 2 parts 2. these are important for marking 3. The moderator uses this to determine their grade. 4. Maziar and the moderator then have to sit down and
agree to a mark. 5. The moderator knows nothing about our project except
for parts a and b. 6. be careful in pat a to ensure that we can show that we
have achieved these goals when we have to complete part B
7. Give numbers, but ensure they are achievable. 8. in the project specifications
a. do not talk in respect to budget with actual numbers
b. include a small gantt chart (timetable of deliverables
9. technical specifications are under the goals 3. 3:20 progress on the concept
a. Statistical analysis � Run into a problem with the turn rate.
1. Check FAR 23, JAR 23 for lateral and roll controllability etc.
2. Far 23 is a very tight standards, FAR25 is more general. 3. we can assume we are building an aircraft similar to a
far23 aircraft 4. Also check Roskam 6 (half of this book is basically an
explanation of FAR 23. � Look at FAR 23 for the takeoff length
b. Matching Diagram � We have been getting weird numbers � MUST COMPLETE the matching diagrams � Want to then get the flight envelope � Need 2 people to work on these � Code the matching diagram, and then change the numbers
c. Sensitivity Analysis � Ian got numbers
1. takeoff 2. not sure what to do with these numbers 3. it is a way of checking numbers 4. should look at he sensitivity of takeoff etc. to the energy 5. look at the sensitivity of both configurations 6. Should tell us what ____ should be if we want ______.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
275 APPENDIX K. MEETING MINUTES
7. just get derivative d. 5-view drawings
� Should have been done. e. Calculations
� Wing span is not a good parameter to use. � Instead should consider the % increase/decrease for all the
values. � We need to get a table of our values, and then determine what
happens to all of them. � CD is wrong. � Roll rate is strange, inc rate with inc span (moment), but also
inc. stability � Need to look at the turn rate � We now have the minimum speed from the ceiling calculations � These calculations are still very elementary. � Need to consider tail parameters. This is generally to d o with
drag equations. 1. see Raimer
� We still have a very elementary analysis. f. Differential telescoping for roll control
� We need to double 1 wing to gain an equivalent change to a 25% deflection of the ailerons.
� Go to Java foil and check the calculations that way.
4. Progress to date – why have things not been done � If Carlee could not find anyone else to do these drawings,
Kevin should do them. � We should not say we are going to do something if we cannot
get it done � We should find a way to get things done. � Drawings and detail design should be done in parallel. � If we need a value, phone Kevin, and he should be able to give
you a value, or make it up. � We need to see a result. � We need to use an engineering, not research approach. We
should go ahead even with best guess numbers instead of waiting for the exact numbers. As an engineer, we need to use more of a trial an error method
� We need to look more at the quantity rather than quality �
5. Other Tasks a. We need to look into morphing methods. (tail and wing)
� E.g. what about if we have a hole in the wing, i.e. extend out the tip and don’t fill in the space
b. We need to be able to justify what we are doing, which method we are choosing.
c. Think innovatively 6. 4:15PM- Close
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
276
Meeting 11 - 18/3/2009
Meeting 11.1 Wednesday 18th March 2009
3:00-4:10PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey,
Summary: Item 1: Sponsorship Item 2: Item 3:
Next meeting: With Maziar: Wednesday 25th March, 3:00PM 2nd floor meeting room Internal meetings: Wednesday18th March, 5:00-6:00PM 4th year project room Thursday 19th March, 4:00-4:45PM 4th year project room Tuesday24th March, 9:0 – 11:00 Rumours Café
Summary of Tasks � To Be done TODAY
o Sign SPPA forms (X3 each) o Fix contract and e-mail to Maziar
� Fix the contract. It is due on Friday (2 copies, one to office, one to Maziar) � Continue with sponsorship tasks � Sponsorship presentation to Aeronautical Engineers Australia (Crystal, Ian,
Carlee), 1:00 PM Monday the 22rd of March � Talk to Red Bull re. sponsorship 11:00 AM Tuesday the 23rd March (Rumours
café) � Continue with the drawings � We should have lots of different concepts by next week � Look in the use of the existing plugs for the aircraft fuselage � Elect a test manager
Summary of Actions: For a comprehensive outline of tasks, see meeting 11.2 (allocation)
Tasks to perform completed by everyone Sign SPPA forms
Fix and sign contract Elect a test manager
Kevin Chan Crystal Forrester
CAD models
Ian Lomas Simon Mitchell
Carlee Stacey
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
277 APPENDIX K. MEETING MINUTES
Meeting Minutes: 1. 3:05 PM – Sponsorship:
The reimbursement is online on my uni. If we spend any money, we need to fill this out in order to be reimbursed.
� Generally we will not be reimbursed for printing costs, petrol, meals for potential sponsors unless there is sufficient money left over in the budget at the end of the project.
� We should fill out these forms for immediate reimbursement for larger, more expensive items, and sites brought directly for the project
� We should keep a list of items other than for this, and reimbursement will be determined depending on the budget at the end of the project.
� It is the policy of the head of school that dinner is not something which we will get reimbursed for. If we do get money out of the company, then we will probably be able to claim the money back.
b. Babcock � We have been promised $1000 � Ian has e-mailed Rae Taylor re. sending an invoice �
c. Aeronautical Engineers Australia � We have a meeting on Monday at 1:00PM
d. Avalon: � Australian Aviation
1. we are very hopeful of sponsorship � Mincham
1. maybe in-kind only 2. they could be useful if we need to make a plug for the
fuselage. These can be very expensive ($5000-$6000) � DMO via Lockheed
1. we should follow this up 2. unlikely to get anywhere. DMO is a big company 3. good to get the name of the university and our project
out there. � American representatives
1. were very interested in our idea 2. we should follow these up a well. US$ are great with the
exchange rate at t he moment! � CAE
1. representative at Avalon Seemed very interested and gave us the contact details for ______.
2. Crystal, Ina and Simon took the CAE representative out to dinner on Monday the 16th o march to Café Piatto’s
3. seems promising as a potential sponsor. 4. They missed out on the opportunity to sponsor a
Melbourne university project, and were too late for the Adelaide uni careers fair
5. they are very interested in being more involved with students
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
278
6. Maziar commented that taking a potential sponsor out to dinner was good idea.
7. They would like a business plan �
2. Forms to be filled out: a. SPPA
� We need 3 copies of this � We cannot just photocopy the forms since we cannot photocopy
a signature b. Expectations form
� Submitted c. Contract
� Return to later in the meeting 3. Decision Matrix
� Was sent to Maziar � We only have a matrix for the Propulsion selection � Pusher tractor matrix- will be written up after feedback is
received on the propulsion selection matrix � Decision on combustion vs. electric
1. we nee a decision matrix for the preliminary report 2. in the final report, this will just be a paragraph
4. Drawing � Pictures of the 3D model were shown � According to Maziar, it is ugly � Crystal is still doing the drawings. � The aircraft does not have an aerofoil for the wings yet � Crystal has only completed about 1/3 of the tutorials for Pro-E � This is a first impression only � Maziar would like to have seen a more advanced drawing � We need to present a 5-view with lateral and longitudinal
cutaways next week � We need to push harder with the drawings to get them done � Maziar was confused by the tail shown on the model, it does
not really demonstrate the morphing capabilities � The aircraft in the model was done on the concept which we ha
been assuming � The fuselage will need a plug in order to be constucted. These
are very expensive ($5000-$6000). The school currently has 2 plugs, one for iSOAR, and 1 for fuel cell. We could possibly use one of these plugs, or part of one of these lugs.
� For next week, Maziar would like to see: 1. Drawing showing the details inside the aircraft 2. these should e a laterally and longitudinally cut view 3. the drawings should sow mechanisms and the landing
gear 4. Crystal needs to start working with someone on the
drawings 5. 2-3 people should spend the whole week n drawings.
We should be able to show lots of different concepts. a. These can be done by hand if w are still learning
the CAD software.
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279 APPENDIX K. MEETING MINUTES
b. Maziar would like to se big progress at the next meeting.
5. 3:40 Contract a. The revised version needs to be sent to Maziar tonight b. Section 1
� This is an introduction. We do not need to include n aim c. Section 2
� This should be a table of deliverables � We should not mention the budget. Our project is not about
getting the sponsorship � This should be a list of the deliverables as outlined by the
school (deliverables and the dates d. Section 3
� This is what should be done by the end of the project. � These should be more specific
1. can carry a 0g payload 2. can loiter for 30 minutes 3. can cruise in line of sight 4. can takeoff and land normally 5. can morph in the sky with a 180% span increase
e. Section 4 � Extension goals � To measure the performance parameters in different
configuration whilst in the sky � To theoretically optimise the morphing configurations
6. Gantt Chart a. Generally on the right track
� Need to remove the university breaks 1. These were only included to assist in planning the Gantt
chart � The Gantt chart has main parts, Technical and admin
1. Technical includes tests, CAD, design, concept phases etc.
2. the technical components need similar breakdown a. i.e. testing, design and manufacturing should all
include a sub section ‘wing design/test/manufacturing’
b. this week, we need to go back and fix the Gantt chart c. The Gantt chart is very important in internal meetings. It should remind us
about what we should be doing, and prevent us from focusing on one task for too long.
7. Matching diagrams and sensitivity analysis etc. a. Matching diagrams
� We can make our aircraft smaller than 7kgs. This would make the aircraft easier to build we could look at a 5kg aircraft, and ake it for 4.5kg
� By next week we need to have a firm value for the weight, wing area, aircraft length and power.
� We need to know all the dimensions so we can give them to crystal for drawing.
b. Sensitivity analysis
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� The sensitivity analysis is working better � 1km→30km range � 1kg→3kg � Since we are using batteries, this should result in a stepwise
function � This code is still being developed � Still need Takeoff range and endurance � We should ignore the sensitivity analysis this week.
8. Morphing concepts/Drawings � We need to prepare lots of configuration sketches and
drawings. � Be imaginative � Look at how to morph, and at the mechanisms � Crystal’s main focus is on learning the software and including
the internal components. 9. Tasks
a. Main tasks this week are: � The matching diagrams, to obtain final decisions on the
numbers previously mentioned � Lots of concepts sketches � 50% of time →drawing � 30% of time →matching diagrams � 20% of time →everything else � Kevin and crystal and Ian to consult with carlee re. Gantt chart
to determine what needs to be done by the end of the mid semester holidays in order to get the manufacturing drawings done.
� Test manager needs to be assigned this week. 1. need to start determining how many test and when 2. this can be anyone but Kevin and Ian
a. not Kevin since the test and technical manager should argue
b. not Ian since manufacturing and testing roles should occur simultaneously
3. one other manager position still to be determined. (Safety officer)
� The Gantt chart needs to be discussed and presented next week. 10. Other
� We can use transmitters and a few actuators from previous years projects.
� If an item is not part of the aircraft, then we can use it � The BOM will be looked at again next week.
11. 4:05PM- Close
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
281 APPENDIX K. MEETING MINUTES
Meeting 12 - 25/3/2009
Meeting 12.1 Wednesday 25th March 2009
3:05-4:10PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey,
Summary: Item 1: Sponsorship Item 2: Matching Diagrams Item 3: CPM/Gantt Chart Item 4: Manufacturing Item 5: Drawing Item 6: Aerofoil Selection Item 7: Concepts
Next meeting: With Maziar: Wednesday 1st April, 3:00-4:00PM 2nd floor meeting room Internal meetings: Wednesday 25th March, 4:00-5:00PM 2nd floor meeting room Tuesday 31st March, 10:00-11:0AM 4th year project room Tuesday24th March, 9:0 – 11:00 Rumours Café
Summary of Tasks � Continue with sponsorship � Review and continue with the matching diagrams � Talk with the workshop in regard to the fuselage plug, moulds etc. � Sketches of the aircraft to determine the configuration � Select an aerofoil � Determine which concept we are going to go with � Choose a motor and a propeller � Determine a test procedure for the motor
Summary of Actions: For a comprehensive outline of tasks, see meeting 11.2 (allocation)
Tasks to perform completed by everyone Kevin Chan Crystal Forrester test plan for the motor Ian Lomas Simon Mitchell Carlee Stacey find out about CPM
determine if the Gantt chart needs to be handed in.
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Meeting Minutes: 1. 3:05 PM – Sponsorship:
a. AEA � Met with them on Monday
b. Babcock � $1000 – should receive at end of the month
c. Magazine � Still following up
d. CAE � Met last week � Wants a business plan
2. Matching Diagrams: � Done
1. Sized for 6kg and 7kg 2. Includes battery weight when the statistics are given 3. 6kg⇒1.1kW engine, wingspan → 0.6m2 4. ⇒We=4.5kg, no batteries 5. 10 cells for batteries ⇒7.5kW
� Needs to be reviewed � Big issues, results are questionable
1. the numbers for cruise are to big � We need to get the aircraft to work for Takeoff, land, stall are 3
most important to get it to the sky � The data from the matching diagrams needs to be presented in a
better way. � We have now got the span etc. � Put the matching diagram away � Check the calculation performance parameters and compare. � V-n diagram (since in extended; can withstand less loading) � H-V diagram (altitude vs. velocity (should have two different
profiles)) � Also looked at conventional aircraft to compare
1. hold of on this at the moment. 2. look at the weight this week. 3. want to look at the structure more 4. check CD0 for home built aircraft
3. CPM/Gantt chart a. Find out what the critical path method is. Look into it.
� Carlee needs to know what is critical at the moment and make sure these things get done
� At meetings Carlee needs to raise which tasks are most important (i.e. which tasks are critical to the completion of the project)
� Find out how the Gantt chart needs to be handed up. If it is to the supervisor, Maziar says this is not necessary.
4. 4:40 Manufacturing a. Ian needs to talk to Billy (workshop) re. plugs, moulds etc.
� We could temporarily modify the mould 1. ask how
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283 APPENDIX K. MEETING MINUTES
� is there another method than using a mold � need to investigate other ways
1. molds are tie consuming and we cannot make them ourselves
5. Drawing: a. Aesthetics
� We need to make the fuselage look prettier. � We now have dimensions therefore we need to determine
possible layouts of the aircraft. 1. we should sit together and sketch the aircraft
a. where is the landing gear? b. Show everything on the aircraft c. Look at internal structures d. Need a cut off view of the tail boom etc.
6. Aerofoil selection � Need to select an aerofoil. � Go for a thicker chord (~16%)
1. This is easier to manufacture � Look at 3 curves
1. Clα 2. Cl Cd 3. Cmα
7. Mechanism/Wing Concepts: a. Gear and pinion
� Not enough space in the wing b. Pulley
� Not very reliable c. Rotating screw
� Will require guides for the outer wing section � We should follow this up.
d. External sheath � for the external sheath concept, the ailerons limit the design. � We need to look into different alternatives fro roll control
1. possibly movement of the wing tips (rotating them up and down)
2. possibly use slats instead of ailerons � We need to re-look at differential span roll control
e. Taper idea � Good idea too
f. Considerations in our decision matrix: � Manufacturability � assembly
8. Other Items: a. Ordering a motor and propeller
� A 1.5kW motor should be sufficient (base this on 8kg) b. Crystal was elected to be the test manager.
� We wish to test the motor in 2 weeks time. 9. 4:05 PM - Close
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Meeting 13 - 1/4/2009
Meeting 13.1 Wednesday 1st April 2009
3:10-4:10PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey,
Summary: Item 1: project description – write a 1 page project description Item 2: Sponsorship – no need to worry. Continue looking. Item 3: Existing part for use – can get batteries and actuators Item 4: design development - need to reconsider. Go back and look at other
concepts. - single boom V- tail was selected for the tail
- forget re-using the iSOAR plug, just design what we want, and go from there - we need 2 fuselage layouts – one for each concept - NEXT WEEK we will be only discussing the sketches
Item 5: Safety officer role /testing Item 6: Close
Next meeting: With Maziar: Wednesday 8th April, 3:00-4:00PM 2nd floor meeting room Internal meetings: Thursday 9th April, 4:00-5:00PM EM, level 3 Sunday 12th April, 10:00-we are finished, Crystal’s house Tuesday24th March, TBD
Summary of Tasks � 1 page project description to be posted on Maziars website. No immediate hurry. � Sponsorship � Consider another idea. If this is completed by Monday, talk it over with Maziar. � Purchase and test the motor
Summary of Actions: For a comprehensive outline of tasks, see meeting 11.2 (allocation)
Tasks to perform completed by everyone Kevin Chan Crystal Forrester test plan for the motor
find out what needs to happen to the test rig to use it/get it working.
Ian Lomas Simon Mitchell Talk to Ian McNair re safety requirements Carlee Stacey Define a purchasing process.
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285 APPENDIX K. MEETING MINUTES
Meeting Minutes: 1. 3:05 PM – Project Description
a. Concern was raised by the workshop staff that we did not have a description of our project online
� Maziar says not to worry about this. It is not important � Maziar has his own website where he posts details of his
projects. � We need to send him a 1 page description of our project to be
posted on his website. � When we have finished our project, this will then be replaced
with the outcome � This is of no immediate concern. But something we should do
when we get time! � The target audience is the general public. � Just in word format
2. 3:15 PM - Sponsorship: a. The group expressed concern regarding the lack of sponsorship we seem
to be going to get. � Maziar said not to worry about this. It is not the most important
of part of the project. � We can keep trying to get sponsorship until about August, but
we will not spend as much time on it. b. CAE
� Rejected our application � Their reason was that they could not afford the money this year. � Maziar said that this is ok. The school will accept a promise of
money to be paid next year. c. AIAA
� Simon looking into has been given a new contact. � Not very hopeful
d. QANTAS, BAE – still being looked into e. Magazine – crystal still looking into f. DMO
� Has crystal’s business card. They will phone her when they have more time.
g. AEA – Carlee is going to call them to follow up. h. Aus Aero – meeting with them on Monday i. DSTO –
� They have been kinder this year. They have given money to one of the other groups.
� In ’07 and ’08 they were very keen on looing into morphing technology.
� We should phone them first to see if it is worth sending the e-mail.
3. Existing parts which we can use: a. We will not need to buy batteries as we can use these from previous
projects b. We do not need to buy actuators since we already have these left over from
previous projects. c. We will need a motor, speed controller, propeller, raw materials
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d. PROCESS FOR PURCHASING PARTS � Carlee is responsible for the budget � Carlee is to define a process fir purchasing. � The process for purchasing parts is that it needs to go through
Kevin (technical manager) to approve, and Carlee (financial approval).
� Maziar should receive all purchase orders through Carlee. 4. 3:25 – Design development:
a. Final Wing Design � 3 different concepts looked at. The chosen concept was uses
internal telescoping without taper. � External sliding sheath (not chosen)
1. tape ration >1 . therefore wing loading is greater on the tip
2. we can have a solid internal section 3. This concept maintains roll control better than the other
concepts as the ailerons move out with the wing. This also increases wing loading.
� Tapered design 1. taper ratio < 1 but not by much ( basically negligible). 2. the effects of taper are negligible at reducing the drag
since the ratio is so small. To reduce the taper ration, the area of the extended section will also be reduced. Although our goal is to increase the span, the main point of morphing the wing is to change the wing area. Therefore there is no point in extending a very narrow wing section.
3. this will be more expensive to manufacture as each rib is a different shape and needs to be loaded into the CNC machine separately. Also, if we require spares, this will be a problem.
4. the tapered idea was ruled out as the taper gained is not worth the other problems associated with this design.
� Sketch of the internal telescoping: 1. Sketch not to scale 2. 300mm chord, composite structures, built up structure 3. this allows for 15mm of structure 4. Maziar made the comment that we should have had a
sketch like this 1 month ago. 5. Maziar does not like this concept
a. It will be difficult to make the inner section b. The ribs cannot take any torsion loads c. The inboard section needs to be very stiff d. The idea is too complex e. We cannot have 700mm of wing section without
support, the external section must sit on the internal section.
f. The concept is limited since we cannot apply a taper ration.
g. The ailerons are inboard => not in a good position
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287 APPENDIX K. MEETING MINUTES
h. Maziar expects we will have manufacturing problems due to the tolerances we require.
i. The aileron actuators being in the fuselage is not good.
� External sheath design has similar problems 1. Foam wings might be better. 2. The 5-10 cm solid tip would solve the problem of the
aileron actuators. For the external sheath. 3. Maziar has suggested that we do not bother
investigating this further � Maziar suggested that we go back and look at our old concepts.
1. in particular, the folding wing concept a. in this design, the pieces are independent, and
we would not have problems with the aileron b. We could also manufacture using a foam core.
2. If we have a concept by Monday, drop by Maziar’s office and discuss it with him. (9:00 Monday would be good!)
3. The concept needs to be developed to the point that the current concept has been developed.
b. Final Tail Design � Considered first Boom tail , twin boom tail, fuselage mounted
tail. 1. A single boom tail was selected to reduce weight and
drag. a. Maziar wished to know if we had considered
interference drag and ?form? drag 2. Maziar agreed that a single boom tail would be OK.
� The next choice was which type of tail. Many we considered. 1. decision matrix resulted in a conventional tail being
selected. 2. inverted V – required longer landing gear 3. V-tail, rudders are an issue
a. Don’t need a rudder, therefore the V-tail will work. If we wanted a rudder, we could just use a mixer to get the same effect.
b. This will change the decision matrix outcome from a conventional tail to a V-tail
c. A V-tail is ‘sexier’ than a conventional tail. 4. The decision made was to go with a single boom, V-tail
aircraft. c. Fuselage design
� We will need 2, one for our current idea, and one for our new concept.
� d. NEXT WEEK we will be only discussing the sketches.
5. Safety Officer Role/ Testing a. Main rolls:
� To prepare the safety documents 1. risk assessments 2. SOP
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� Once we have all the safety documents signed, then the school assumes the responsibility for safety
b. When we run a test, there are 3 main people involved; the test officer, safety officer, and manufacturing (whoever was involved in the design/selection of the thing being tested)
� The safety and test officer both need to prepare a test checklist. � Simon as the safety officer gives the final OK to go ahead
c. The school safety officer is Ian Macnair. � His office is in the workshop near Richards � Simon should go and talk to him to find out what needs to be
done. d. We should be testing the motor next week.
� Buy motor, get reimbursed � The school has the batteries, so purchase a motor which goes
with the batteries. 1. batteries we have are normal Li-Po batteries. 2. the batteries are currently in the electronics workshop
a. to see them , go and talk to Phil or Sylvio b. guess is 14.0 Volts
� Crystal need to come up with a test for the motor 1. thrust 2. look for the test stand.
a. It needs to be fixed. It is missing a part. 6. 4:10 PM - Close
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
289 APPENDIX K. MEETING MINUTES
Meeting 14 - 8/4/2009
Meeting 14.1 Wednesday 8th April 2009
3:00-4:10PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey,
Summary: Item 1: Propulsion System Item 2: Sponsorship Item 3: BOM development Item 4: Preliminary report Item 5: Drawings Item 7: Calculations Item 8: Design Item 9: Tasks after this week Item 10: Tasks for this week. Item 11: Close
Next meeting: With Maziar: Wednesday 15th April, 3:00-4:00PM 2nd floor meeting room Internal meetings: Tuesday 14th April, 10:00-5:00PM Study room
* technical meetings will occur over Easter. To be organised depending on availability. KEVIN to organise.
Summary of Tasks � Purchase and test the propulsion system � Sponsorship � Consider another idea. If this is completed by Monday, talk it over with Maziar. � Purchase and test the motor
Summary of Actions: For a comprehensive outline of tasks, see meeting 11.2 (allocation)
Tasks to perform completed by everyone Kevin Chan Crystal Forrester test plan for the motor
find out what needs to happen to the test rig to use it/get it working.
Ian Lomas Simon Mitchell Talk to Ian McNair re safety requirements Carlee Stacey
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Meeting Minutes: 1. 3:15 PM – propulsion system:
a. We need to purchase the Propulsion system. � Motor � Speed controller � At least 2 propellers � (one climb and one cruise). Plastic propellers are cheap. � MOUNTING-
1. sometimes the engines come with a different mounting depending on weather they re pusher or puller
2. we do not need to necessarily buy this now, but we do need to get the dimension, weight etc.
� When we get the motor, try to also get a graph of the thrust curve for comparison.
2. Sponsorship. a. DSTO
� We should be more active in following up the DSTO b. CAE
� Simon to chase this up c. We should no consider companies not related to aerospace engineering.
� We could even put an environmental twist onto our project for the purpose of sponsorship applications
3. BOM development: a. Later we will add a reference to a drawing number b. Everything with a drawing number needs to be in here
4. Preliminary Report: � Need to start looking into the structure of the report � Draft report is more of a detailed plan, not the whole report � Usually the 0 days between the draft and the actual report being
due are spent entirely on the prelim report. � The actual report is approximately 120 pages
b. Detailed Plan � Our detailed plan should be approximately 30-35 pages. � Should include the chapters, sub chapters, and one bullet point
sentence describing each paragraph. We should also include the number of pages required
� We should start writing up the pieces of the report related to what we are working on.
c. The report should not be written in book format. It should include only what is relevant to the project. i.e. it is not telling a story.
d. At the end of the holidays, we should start with a shorter version again. We should have basically a content list if the chapters to be sent to Maziar
e. If we send in the draft early, we will get it back early. � It usually takes 2-4 days to mark.
5. Drawings: a. We NEED to get these done b. Maziar can give us an example if we would like
� Co-axial had < 200 drawing sheets last year 6. Calculations:
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291 APPENDIX K. MEETING MINUTES
a. Structural calculations need to be done. b. ANSYS is very popular
� This gives lots of pretty pictures. � We do not have the understanding to do this properly though. � We will rely on hand calculations � We could be disadvantaged if we do not use ANSYS. These
pictures always look good. c. Structural Calculations:
� We will need to find some books on hand calculations of composites.
� Michael Niu 1. Has 3 books. (structural calculations, composite
materials and manufacturing, and _____) 2. Maziar thinks we need the green book.
� Megson 1. Aircraft structural calculations for aerospace
engineering ( or something like this) 2. this book has 5-6 examples for composites 3. we need to look at the load calculations 4. composite structural calculations
� Lift distribution is the first thing that we need to look at. 1. Need to determine the shear, bending moment and
torsion due to lift. 2. then calculate the stress on the structures 3. consider all loads 4. can start with Raymer and Roskam 5. we need a V-n diagram for the aircraft
� Carry out the structural calculations in the following order: 1. V-n diagram 2. Load distribution on the wing (shear, bending and
torsion) 3. Simple structural calculations on everything.
a. We can do local calculations and calculation on load bearing structures
� We do not need to consider: 1. vibration 2. dynamic loading 3. fatigue
� Need to consider accessibility of the mechanism � Consider that the load bearing sections are separated from the
skin � We need to look into materials. Wood or aluminium. � We cannot use composites as we cannot cut composites.
7. Design: a. We need final decisions on the mechanism at the next meeting. b. We need to bring a structural design
8. Next tasks- week after this one. a. Next week we will chose the aerofoil
� This is to do with cruise speed for a large aircraft � For a small aircraft we look for laminar flow, then look at the
other properties
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� Usually during to the manufacturing method, the aerofoil is not the same as is selected.
1. Because of this, we cannot select an aerofoil which has a strongly defined curve on the bottom.
2. We need to select a simple aerofoil. 3. Look at the book ‘theory of wing section’ by Abbott
� Kevin has previously looked into aerofoils 1. Selected for the previous design a S4233 Sigel aerofoil 2. Selected it for its CL, thickness and pitch. 3. Maziar asked what it’s post stall characteristics were
� When looking at aerofoils, Maziar has suggested that we will need to find 3, and then choose.
1. We should be careful, as some aerofoils are more efficient with a flap. This is not good for us.
2. We need to consider post stall characteristics. a. This is very important. b. We do not want a stall curve with a fast drop off.
This makes it more difficult to recover the aircraft.
c. We need to look at the Cmax and the Cα d. We need to look at the sensitivity of the aerofoil
characteristics to the angle of the trailing edge. (We cannot achieve an exact trailing edge angle).
e. We can then select our aerofoil f. We will need 2 aerofoils. One for the main
section and one for the extending section. b. Look into manufacturing
9. FOR THIS WEEK: a. We need to have a VERY detailed drawing
� It can be either CAD or hand drawn � Simon is to be the ‘bad guy’ and criticise the design.
1. We should be able to answer all his questions. � We need to consider possible options for the design. � Start considering the structure.
1. Loads usually take the shortest path to transfer their loads.
2. We need to be careful when considering moving and sliding components.
a. There should always be at least 2 points in contact
b. It must sit at 2 points 3. The sections should be load bearing for their entire
length. 4. The aileron needs to be attached to a hard section. 5. The structure required should dictate the wing shape,
not the aerofoil. 6. The installation of the aerofoil will also help to dictate
the shape of the wing. 7. We need to use the load bearing points to transfer the
loads.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
293 APPENDIX K. MEETING MINUTES
Meeting 15 - 15/4/2009
Meeting 15.1 Wednesday 15th April 2009
3:00-4:00PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell
Summary: Item 1: Sponsorship Item 3: Motor Testing Item 4: Drawings and structure Item 5: Tail Design Item 6: Close
Next meeting: With Maziar: Wednesday 22nd April, 3:00-4:00PM 2nd floor meeting room
Summary of Tasks
•••• SM and CS to complete test procedure and safety protocols for this test •••• SM to define a procurement procedure to be adhered to for each procurement •••• Todd might know of the whereabouts of a receiver for the motor testing. IL to
contact Todd. •••• We need to do a full aircraft sketch to ensure that there are no problems •••• MA would like us to work out how many bearings will be needed and what the
forces through them will be •••• We need to calculate if cutting a hole in the inboard wing foam in order to add in
an auxiliary rib is worthwhile. •••• We need to cut the foam by next week to practise with the hot-wire. We don’t
need to use the correct template; we can use any one we find. SM to work out safety for this, and IL to work out procedure.
•••• Design and draw the tail blocks properly and present them to Maz
Summary of Actions: For a comprehensive outline of tasks, see meeting 11.2 (allocation)
Tasks to perform completed by Everyone o We need to do a full aircraft sketch to
ensure that there are no problems o We need to cut the foam by next week to
practise with the hot-wire. We don’t need to use the correct template; we can use any one we find. SM to work out safety for this, and IL to work out procedure.
Anyone o MA would like us to work out how many bearings will be needed and what the forces through them will be
o We need to calculate if cutting a hole in the inboard wing foam in order to add in an
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294
auxiliary rib is worthwhile. o Design and draw the tail blocks properly
and present them to Maz Kevin Chan Crystal Forrester Ian Lomas o Todd might know of the whereabouts of a
receiver for the motor testing. IL to contact Todd.
Simon Mitchell o SM and CS to complete test procedure and safety protocols for this test
Carlee Stacey o SM and CS to complete test procedure and safety protocols for this test
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
295 APPENDIX K. MEETING MINUTES
Meeting Minutes: 1. Sponsorship
a. SM – No luck with CAE, they are “in no position to commit funds for the next financial year”
b. CS – Unable to contact AEA, our contact took a long Easter holiday and was out of the office
c. IL – Tony from AA wants our materials list and drawings, as well as a schedule. They seem to have given us a verbal promise for in-kind support, with nothing written down.
d. CF – Crystal left a message with DSTO, still unable to contact. 2. Motor testing
a. SM - Purchase order for motor, ESC, and two propellers has been submitted to the value of $538.00 excluding GST. Mech Eng office recommends a 2-3 day turnaround on the purchase order being dealt with.
b. IL – Has computed a thrust curve, MA showed no real objections (IL used McCormack text book)
� Maz suggests using online tools to help with the thrust calculations but the report needs to have the proper calculation procedure. Propeller selection will be an “important chapter in the report”
c. SM and CS to complete test procedure and safety protocols for this test
d. SM to define a procurement procedure to be adhered to for each procurement
e. MA – suggests we probably don’t need a folding prop on our final aircraft, but this is a problem best left to deal with later.
f. CF – to contact electronic workshop to organise the load cell and the data logger. The thrust to amperage relationship is what we are trying to obtain from this test.
g. MA – suggests we should use the receiver from the co-axial project (originally used on the airship). Todd might know of its whereabouts. IL to contact Todd.
3. Drawings and structure a. Drawings of the wing layout, tail layout, mechanism, and rib and spar
design were shown to MA b. Kev described the roller concepts (ball bearings etc) to MA c. We need to do a full aircraft sketch to ensure that there are no
problems d. MA suggests that the ball bearings could be very expensive but they have
good alignment characteristics e. MA advises, “Attaching Al to Al is easy, but attaching something else to
Al is quite hard”. f. SM provided graphs of load distribution due to bending, MA showed no
major objections. MA doesn’t recommend using a tapered spar due to machining costs.
g. MA recommends getting the wing design and structure correct now, as weight issues will result if these issues aren’t dealt with.
h. MA would like us to work out how many bearings will be needed and what the forces through them will be
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
296
i. Maz suggests that we only need supports at the tip and a little bit in from the tip, and not load bearing rollers further in.
j. We may need to embed a hard strip on the surface (or just under) of the outboard wing section to contact with the rollers. This may not be able to be made out of composites due to the tolerance required (and composites are mainly hand made by students).
k. He suggests using Aluminium or plywood. Plywood is better but it can’t be used on the surface due to poor surface finish. If bearings are being used, we can’t use plywood and we should use Aluminium, but if we aren’t using bearings (for example a rubber wheel), we can embed the plywood strip under a composite surface.
l. In normal construction, the spars don’t normally have a free end in the foam; they have an end cap on each end. We can use the ply discussed about as spars but we must join the ply on the top and the ply on the bottom together. This is normally done by cutting foam out in the middle, joining the two, and replacing the foam. We need to calculate if this is worthwhile.
m. For the inboard spar design, we don’t need as many ribs as we currently have if we use a monocoque structure. This will also cut back on the spars required and provide us with more room for the aerofoil. Spars can run above and below the wing section (like giant stringers) instead of right through the middle.
n. Maz wants to cut the entire structure from foam (in two or three pieces which are then glued together) and add a few ribs to the foam section (3 ribs?)
o. MA – we can manufacture by creating the foam sections, cutting in grooves for the spars, adding ribs at the joins of the foam sections, glueing the foam sections together, and maybe cutting away foam to add another rib if necessary.
p. MA – composites can be applied directly to the foam and don’t need a medium in between.
q. MA – suggests attaching the rollers directly to the ribs or spars r. We need to cut the foam by next week to practise with the hot-wire.
We don’t need to use the correct template; we can use any one we find. SM to work out safety for this, and IL to work out procedure. We don’t need to purchase the foam, “it’s cheap”, and we need to ask Bill in the workshop.
s. MA – “Cutting something [with a hot-wire] that is 100mm across is easy, 300mm is much harder.”
t. MA wants us to stop all calculations and report writing and focus on the 1:1 drawing
u. When designing the rollers, we need to ensure that the rollers we’d like are available and need to make a physical check that they are appropriate and as expected.
v. MA suggests that we are two weeks behind schedule 4. Tail design
a. Maziar is happy with using the three carbon rods in a triangle arrangement for the tail sliding mechanism, however he is a little concerned about how the carbon rods are to be joined.
b. MA – “Triangle is more draggy than cylinder, so thank about this” c. KC can justify the joining of the triangle well and MA is happy with this.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
297 APPENDIX K. MEETING MINUTES
d. MA suggests using an ellipsoid shape instead of a triangle. e. MA – We may be able to get away with using carbon strips instead of
carbon tubes, and they are easier to work with. f. MA thinks we need a little more support internally for the tail, but he
thinks the concept is good. g. MA requests that we design and draw these tail blocks properly and
present them to him h. MA also says that we should leave the tail for now and finish the other
drawings first i. We may need to taper the end of the mechanism off to reduce drag.
5. 4:00 PM - Close
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
298
Meeting 16 - 22/4/2009
Meeting 16.1 Wednesday 22nd April 2009
3:00-4:05PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey,
Summary: Item 1: Aircraft design –1-1 scale drawing –some suggestions made for improvement –some issues with the design pointed out Item 2: Sponsorship –Continue looking and following up.. Item 3: Aerofoil Selection –NACA 2416 was selected for the inboard wing
–NACA 4412 was selected for the outboard wing
Item 4: Propulsion Test –we can use the Jet propulsion Lab Item 5: Close
Next meeting: With Maziar: Wednesday 29th April, 3:00-4:00PM 2nd floor meeting room Internal meetings: Monday 27th April, 10:00AM-LATE, FYP study room
Summary of Tasks � Fix the design, as recommended. � Sponsorship � Purchase and test the motor
Summary of Actions: For a comprehensive outline of tasks, see meeting 11.2 (allocation)
Tasks to perform completed by everyone Kevin Chan Crystal Forrester Ian Lomas Simon Mitchell Carlee Stacey Contact AEA re. sponsorship
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
299 APPENDIX K. MEETING MINUTES
Meeting Minutes: 1. 3:05 PM – Aircraft Design – 1-1 scale drawing.
a. General discussion in regard to possible improvements and potential problems with the deign.
� Attachment of the wings to the fuselage 1. We could possibly use a tongue in the fuselage, instead
of into the wing 2. this way the tongue is part of the wing and is inserted
into the fuselage where it is attached with screws. 3. we could alternatively have a metal/wooden c-section
along the length of the fuselage which sticks into the rib.
4. we only require 2 points of connection between the fuselage and the wing. This is all most light planes have, so it should be fine for us If we use anymore, ten all the load is taken by only 2 anyway (result of tolerances etc.)
� Tail attachment. 1. the tail structural members need to be connected to the
boom. 2. possibly look at PVC pipe connections.
� Maziar says to go an make it! �
b. Screw Thread: � Before we look into alternatives etc. we need to design the
thread (i.e. determine what we would like it to be) , and then look into alternatives.
� Design and calculate the pitch, height of thread etc. �
2. Sponsorship a. Magazine
� Rejected b. DSTO
� Crystal Talked to Simon Henbest on the phone � They are not sure about their financial situation � They have been sent the business plan � It did not sound very promising � Maziar says that Simon is quite high up in the organisation, and
new to the position. c. AEA
� Carlee has still been unable o contact Mick Kaesler. � Maziar has suggested that if we have not been able to get an
answer by next week that we should phon their financial manager. Maziar has her phone number from fuel cell last year.
d. Australian Aerospace. � Ian has been in contact with Tony. � Tony has talked to Holbright engineering on our behalf and
Holbright has indicated that they will see what they can do to help if we contact them
1. Holbright only do machining.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
300
2. might be worth finding out what they do. It is possible we might be able to get them to make us a plug for the fuselage.
3. 3. Aerofoil Selection
a. Inboard wing � Seilig S8037 (16% thickness) OR NACA 2416 (very similar)
1. this has a greater lift coefficient, and less pitch moment 2. Both are used in model aircraft, therefore both are
laminar � NACA2416 was selected
1. easier to make, very similar to S8037, can use a similar aerofoil on the inside.
b. Outboard Wing � NACA 4412 and SG6042 were considered � NACA 4412 was selected � Since the aileron is inboard, we do not need to be greatly
concerned with tip stall 4. Propulsion Test
a. Yes, we can use the Jet propulsion lab. 5. 4:05 PM - Close
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
301 APPENDIX K. MEETING MINUTES
Meeting 18 - 6/5/2009
Meeting 14.1 Wednesday 6th May 2009
3:00-4:10PM Attendance: Maziar Arjomandi, Kevin Chan, Crystal Forrester, Ian Lomas, Simon Mitchell, Carlee Stacey,
Summary: Item 1: Sponsorship Item 2: Testing Item 3: Procurements Item 4: Design Item 5: CAD Item 6: Structural Calculations Item 7: Report Item 8: Close
Next meeting: With Maziar: Wednesday 13th May, 3:00-4:00PM 2nd floor meeting room
Internal meetings: To be arranged as required
Summary of Tasks � Report � Finish CAD � Get fuselage plug manufactured � Get motor test completed
Summary of Actions: For a comprehensive outline of tasks, see meeting 18.2 (allocation)
Tasks to perform completed by everyone Kevin Chan Crystal Forrester Ian Lomas find out how, and when we can get the plug
manufactured for the fuselage arrange a meeting with Airspeed
Simon Mitchell Carlee Stacey
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
302
Meeting Minutes: 1. 3:00 PM – Sponsorship
a. DSTO � We have heard from them � Need to follow this up � CF to phone around 9:00 AM tomorrow
b. AEA � We need to get E-mail confirmation of sponsorship. We need to
send this through to Maziar so we can get the school to send an invoice to AEA.
c. Redbull � We can get cans of Redbull, no money.
d. Airspeed � We need to go and talk to them once we have the
manufacturing drawings � We need to make the plug � Put them in the prelim report � By this time next meeting, MA would like us to have met with
airspeed. 1. We need to take the drawings of the fuselage 2. We should also talk with Bill to find out an idea of the
time required to get the plug made. 3. At the next meeting, Ian should present to us about how
the fuselage will be manufactured. 2. Testing
a. We have got the motor b. Electrical workshop has set up circuit c. We are still waiting on one part of the circuit. We are waiting for a student
to return the part. d. We wont be using a computer, but rather will be reading from a multimeter e. Expect this test to have been done by next week f. We need to make some corrections to improve the safety g. When the other group wishes to use the test stand next week, they will be
directed to us. We should not return the stand to the workshop, . we should go with the group to the workshop and swap .
� T h. The person to speak to in regard to the testing Beau… i. We need to look into the threaded rod test for the mechanism.
� We are considering using the thrust rig � Get a rough idea from calculations first � Test using the real rod eventually.
3. 3:15 - Procurements a. Parts from model flight:
� These should have come in � Model flight have not yet been in voiced � We need to go and see Wendy or Yvette � Next time, we need to record the number on the in voice, and
then we can check with the office if it has been payed. b. We needed to buy sone safety equipment for the test
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
303 APPENDIX K. MEETING MINUTES
� Next time we need to check with Maziar to find out if we already have the parts availible. From previous projects.
c. We are looking a getting the tail boom in from New Zealand. d. We now have a procurements proceedure e. We need to get the threaded rod lathed to be coupled to the motor for the
test. f. Need a drawing to do this.
4. Design: a. Landing gear:
� Position and angle 1. KC calculated to be 21° 2. MA says this does not have to be 21°. It should be about
16°. 21° is too much as the maximum angle of attack is ___°.
3. MA says that tip back should be close to the angle of attack. We should start with the angle of attack. If the angle of attack is 12°, then we should make the tip back angle 14 °or 15°.
4. we do not need to consider the landing gear just yet. b. Wing block design
� We have changed the step down section so that it is no longer there.
c. Fuselage � We have another 1-1 Drawing � Nose is very forward � Batteries are actually smaller � To find the CG
1. include the Aircraft actual size. Also include the moment of inertia
a. this is not so good for gusts. 2. need to consider the CG envelope.
� first need to find the minimum tail arm, then play with the layout. Then increase the nose if necessary.
� Try to change the tail to move the motor forward toward the nose.
� We need to make the fuselage as compressed as possible � Can make holes in the frame with no problems (i.e. O shaped
frames) and just increase the thickness of the ply. 8mm ply is just as good as steel.
� We may have a problem with attaching � We need to make the aircraft shorter � Find CG in the 2 configurations, play with the aft CG to find
the stability margin, then get the static margin. � Move the tail forward to reduce the stability � Something strange is happening with the CG �
d. The tail deflection must be less than 6mm. 5. Drawings:
a. Ian needs to prioritise the manufacturing drawings 1. Ribs
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
304
a. Need templates for cutting foam 2. wing
a. need ribs, b. template c. spar d. rollers e.
6. Report a. We start the report with the literature review and the technical task b. We need to start with the introduction c. Significance section:
� Why morphing � This it the chance to sell the project
d. Aims and project specification � Tech task should finish this section. Put in as is, but with some
modifications. e. Concept vs detail design:
� Stability 1. envelope is concept 2. phase 2 which we do not do is detail
� everything for the first sketch is concept design � structural calculation is detail. � Analysis of design is detail design
7. 4:00 PM - Close
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
L.
Gantt
Chart
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follow
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305
M. LabourTable M.1 shows the labour contributions toward the project by each group member over
the course of the project.
Table M.1: Labour contributions by each group member
KC [hrs] CF [hrs] IL [hrs] SM [hrs] CS [hrs] Total [hrs]November 0.00 8.50 6.50 8.00 10.50 33.50December 127.65 130.25 35.75 25.25 32.00 350.90January 114.90 83.75 74.00 58.25 78.30 409.20February 92.00 61.50 68.00 53.50 58.32 333.32
March 137.50 111.48 110.00 80.75 104.55 544.28April 146.50 123.85 114.50 137.00 115.27 637.12May 217.00 209.00 172.25 186.25 145.25 929.75June 73.00 121.73 62.00 90.00 51.66 398.39July 204.00 149.59 205.25 229.50 209.81 998.15
August 187.50 195.50 230.25 209.75 287.71 1110.71September 342.00 304.03 294.25 275.50 270.17 1485.95October 390.00 360.08 410.00 328.00 355.49 1843.57
TOTAL 2032.05 1850.77 1776.25 1673.75 1708.53 9041.35
306
N. Documents used in obtaining
sponsorshipThe first document included in this appendix is an example of a letter sent to all poten-
tial sponsors, requesting a face-to-face meeting. The second document is a copy of the
brochure produced by the group for distribution during meetings with potential sponsors.
Mr. Mick Kaesler Carlee Stacey Assistant Engineering Manager-Adelaide MORPHEUS Project Team Aeronautical Engineers Australia School of Mechanical Engineering 8 Douglas Drive The University of Adelaide Mawson Lakes, S.A. 5095 Adelaide, S.A. 5005 Phone: 0400 714 400 Email: [email protected]
18 March 2009 Dear Mr Mick Kaesler, I am writing on behalf of the University of Adelaide MORPHEUS final year Aerospace Engineering project team to give you some information regarding our project prior to our scheduled meeting this Monday the 23rd of March. The primary purpose of this meeting is to present to you the possibility of Aeronautical Engineers Australia sponsoring the MORPHEUS project. As part of the final year of Aerospace Engineering, it is a requirement for students to complete a major engineering project. These projects allow the students to gain practical experience in all aspects of the engineering process, from concept generation through to manufacturing and testing. The MORPHEUS project involves the design and build of an Unmanned Aerial Vehicle (UAV) with a morphing configuration. We are currently in the final stages of the concept selection phase of this project, and are beginning the detailed design phase. The selected concept involves increasing the wing area by morphing the wing span, as well as morphing the tail by changing its position to maintain a balanced aircraft in all configurations. The final design will result in a multi-mission platform which reduces the need for performance compromise during different flight phases. Aeronautical Engineers Australia is an Australian company heavily involved in the aircraft industry here in Australia. As such, we would like to present you with the opportunity to sponsor our project. We are very enthusiastic to gain your company’s support, as Aeronautical Engineers Australia has an excellent reputation within industry for supporting engineering. As a sponsor, Aeronautical Engineers Australia would receive invitations to the project seminars and exhibition, be recognised in all deliverable tasks including the final report, project seminars and the project exhibition, and your logo will be displayed on our UAV. This will give Aeronautical Engineers Australia the opportunity to assist in the education of future engineers, whilst gaining exposure to students, academics and the wider engineering community. Please contact us with any questions that you may have. We look forward to discussing the MORPHEUS project with you in person on Monday. Yours sincerely, Carlee Stacey On behalf of Kevin Chan, Crystal Forrester, Ian Lomas and Simon Mitchell
307
308
Kevin ChanKevin ChanKevin ChanKevin Chan 0416 339 183
Crystal ForresterCrystal ForresterCrystal ForresterCrystal Forrester 0403 430 916
Ian LomasIan LomasIan LomasIan Lomas 0410 132 319
Simon MitchellSimon MitchellSimon MitchellSimon Mitchell 0423 982 431
Carlee StaceyCarlee StaceyCarlee StaceyCarlee Stacey 0400 714 400
Supervisor:Supervisor:Supervisor:Supervisor:
Dr. Maziar Arjomandi
“The Design and build
of an Unmanned Aerial
Vehicle with morphing
capabilities”
The University of The University of The University of The University of
AdelaideAdelaideAdelaideAdelaide
School of Mechani-School of Mechani-School of Mechani-School of Mechani-
cal Engineeringcal Engineeringcal Engineeringcal Engineering Due to the large scale of this project, fund-
ing is required for successful completion.
As a Sponsor your company will receive:
• Company logos on all deliverables
• Recognition at all public events
• Invitations to project related events:
• Project Seminar
• Project Exhibition
• Final project report
• The Opportunity to invest in future en-
gineers
• The opportunity to invest in future UAV
and Morphing technologies
iSOAR UAV — 2007 Project
2008 wind turbine project exhibition
Hy-Five fuel-cell
UAV —2008 pro-
ject
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
309 APPENDIX N. DOCUMENTS USED IN OBTAINING SPONSORSHIP
Definition:Definition:Definition:Definition:
To Design, build and test an Unmanned Aerial Vehicle with the aim of determining its effec-tiveness in multiple flight phases.
Goals:Goals:Goals:Goals:
1. Determine preferred morphing configu-rations for multiple phases of flight
2. Design, build and test multiple morph-ing mechanisms and integrate these into a custom designed remote con-trolled UAV
3. Achieve stable and sustained flight in at least one configuration.
4.
Extended Goals:Extended Goals:Extended Goals:Extended Goals:
1. Flight test the UAV in different configu-rations.
2. Achieve stable and sustained flight whilst morphing between configura-
tions.
Concept 1:Concept 1:Concept 1:Concept 1:
A variable sweep, wing area and
tail position UAV capable of
morphing between a flying
body/ rocket configuration for
high speed to a standard high lift
configuration.
Concept 2:Concept 2:Concept 2:Concept 2:
A variable (telescoping) wing
area and tail position UAV
which allows high altitude and
low stall speed in one configura-
tion and high speed and long
range in the second.
After an early start in November the team
has completed a detailed literature review,
statistical analysis, technical specifications,
concept generation and propulsion sys-
tem selection. We are currently in the con-
cept selection and development phase.
Ian explaining how
a model aircraft is
designed and flies
at our research
flight trial day.
Experience in:Experience in:Experience in:Experience in:
• The entire engineering process from concept generation through to manufacturing and testing
• Project management
• Financial management
• Systems engineering
• Practical engineering
• Teamwork
A high performance aircraft that can op-
erate efficiently in multiple flight regimes
by changing its external shape.
Benefits:Benefits:Benefits:Benefits:
• Increased fuel efficiency
• Reduced noise
• Improvement of aerodynamic
properties
• Milti-mission capabilities with one
aircraft
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
O. Business plan
Morpheus Final Year ProjectBusiness Plan
Kevin Chan
Crystal Forrester
Ian Lomas
Simon Mitchell
Carlee Stacey
The University of Adelaide - School of Mechanical Engineering
310
311 APPENDIX O. BUSINESS PLAN
Background Information
Who we are:
As part of the final year of the Aerospace Engineering degree at The University of Adelaide, it is a requirement
for students to complete a major engineering project. These projects allow students to gain real experience in
all aspects of engineering process from concept generation, through to manufacturing and testing. The Uni-
versity of Adelaide has a strong reputation for excellence in final year projects. Projects from The University
of Adelaide frequently receive national news coverage, or release professional academic papers on their topics.
The MORPHEUS project involves the design and build of an Unmanned Aerial Vehicle (UAV) with a
morphing configuration. Detailed design is well underway and the team is in the initial stages of component
testing. The UAV consists of a two-part telescoping wing (allowing variable wing area) and a boom-tail with
a telescoping mechanism to stabilise the aircraft. The final design will result in a multi-mission platform
which reduces the need for performance compromise during different flight phases.
What is a morphing aircraft?
A morphing aircraft is a high-performance aircraft that can operate efficiently in multiple flight regimes
by changing its external shape. The morphing is normally achieved by using smart materials or dynamic
structures. Such aircraft morph by changing the wing or tail location, area or sweep. The location of the
wings or centre of gravity, or the dihedral are examples of possible characteristics to morph. The focus of
this project is to flight test these morphing mechanisms on an unmanned aerial vehicle.
An example of morphing aircraft can be seen by comparing the F/A-18 Hornet and the Global Hawk UAV.
The two aircraft are completely different; the Hornet is a fighter aircraft designed for speed and maneuver-
ability whereas the Global Hawk is a reconnaissance and surveillance aircraft. It is designed to have high
endurance and high stability, but is not too fast. These two aircraft could not trade roles. A Global Hawk is
not maneuverable enough to perform air-to-air combat, and an F/A-18 cannot fly slow enough for surveil-
lance, and would have to refuel numerous times to stay in the air for the required time. The solution to this
problem is to design a morphing aircraft to perform both roles. The Global Hawk has a larger wing span for
high altitude and endurance, and a conventional aerofoil profile for high lift at low speeds. The F/A-18 in
contrast has a smaller wing span and a diamond shaped aerofoil profile for high speed and maneuverability.
The perfect morphing aircraft in this circumstance would change its wing span and aerofoil shape to perform
both roles.
Through morphing aircraft geometry the aircraft can increase fuel efficiency, decrease emissions, reduce noise
and improve aerodynamic characteristics.
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
312
Project definition and goals:
The project has been formally defined as follows:
Definition:
To design, build and test an unmanned aerial vehicle with morphing configuration, primarily as a test bed
for morphing technology.
Primary goals:
1. The UAV shall have a normal takeoff and landing method.
2. The UAV shall be capable of having a loiter time of at least 30 minutes.
3. The UAV shall be capable of cruising within line of sight.
4. The UAV shall be capable of carrying a 500g payload.
5. The UAV shall morph the wing to achieve a wing span increase of at least 50% of the original wing
span during flight.
6. The UAV shall change the tail position to control the longitudinal stability during flight.
Extended goals:
1. To measure the performance of the UAV in different configurations during flight.
2. To theoretically optimise the morphing parameters for a predetermined mission.
3. To achieve roll control through differential span morphing .
4. To encourage continued undergraduate and postgraduate development of unmanned aerial vehicles.
5. National and/or international recognition for aeronautical research at the University of Adelaide.
6. To encourage high school students to study Aerospace Engineering at a university level.
The requirement of industrial sponsorship
Projects such as this rely on the support of industry to come to fruition. Traditionally, businesses sponsor
these projects and have their business advertised to high school and tertiary students, industry professionals,
academics and the public. The sponsor maintains a good relationship with the University, and has numerous
recruitment and exposure advantages.
The project team is required to facilitate the design and manufacture of the aircraft, as well as the pro-
curement of all components. The aircraft will be manufactured at the university workshop so that we can
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey
313 APPENDIX O. BUSINESS PLAN
gain a better understanding of the manufacturing process. The university supports the students in kind by
allowing us a limited number of hours in the workshop. Excess hours (which will be required for a project of
this nature) will involve extra costs at the expense of the group. Manufacturing is the most expensive aspect
of our project since in order to meet weight restrictions, composite materials will need to be used. The will
require the manufacture of moulds and plugs which of considerable expense. Other costs also include the
electronic components, motors, propeller and batteries.
The Morpheus final year project will require $20,000, to be spent as follows:
• Propulsion system: $3000
• Airframe and mechanisms: $14000
• Control systems: $2000
• Imaging system: $1000
Benefits of industry sponsorship
Should your company choose to sponsor our project, it will receive the following benefits:
• Company logos on all deliverables (aircraft, report, seminar presentations and exhibition);
• Recognition at all public events;
• Invitations to all project related events (Project seminar, Project exhibition);
• A copy of the final project report;
• The sponsorship is tax deductible;
• The opportunity to invest in future engineers;
• The opportunity to invest in future UAV and morphing technologies
More information:
If more information is required, feel free to email the group member who contacted you or phone one of us
directly:
Kevin Chan: 0416 339 183
Crystal Forrester: 0403 430 916
Ian Lomas: 0410 132 319
Simon Mitchell: 0423 982 431
Carlee Stacey: 0400 714 400
We look forward to hearing from you. Regards,
MORPHEUS Final Year Project - The University of Adelaide
Date: October 30, 2009 Chan, Forrester, Lomas, Mitchell, Stacey