859 design and build of a uav with morphing con gurationrahauav.com/library/unmanned...

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


4.2.6 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


4.5.7 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.2.3 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.3.3 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.


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).


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.


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)


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



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


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.



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.



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.



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).



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.



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)


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



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.


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


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.



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


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.



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


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.



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.


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.



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


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



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


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



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.


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.



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



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.



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.



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



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)



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)



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



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.



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.


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-



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


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.



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.


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



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


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.



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.


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



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.


3.11. CONCEPTUAL DESIGN SUMMARY 52


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


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



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.


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



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-


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



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


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



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


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◦



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.


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



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.


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.



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)


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



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.


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.



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


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)



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.


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.



Figure 4.21: Shear diagram for the fuselage and empennage

Figure 4.22: Bending moment diagram for the fuselage and empennage



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



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



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


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



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.


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



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)



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.



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



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



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


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



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


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



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



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.



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



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.



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.



(a) Fuselage shear diagram

(b) Fuselage bending moment diagram



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.


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



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



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.


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.



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),



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.



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



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



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.


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



telescoping should be theoretically possible the the optimal morphing configurations for

a typical reconaissance mission were discussed.


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


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



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


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.



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


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.


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.



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.


5.7. AIRCRAFT ASSEMBLY 116

Figure 5.4: Empennage assembly drawing


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



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.


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.



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.


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.



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.


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


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


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


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.



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.


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



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.



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



Figure 6.5: Attempt 1

Figure 6.6: Attempt 1 - GPS output



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



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



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



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.


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



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


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.


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-


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


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.



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.


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



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


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.


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


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.


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:


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.


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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


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◦


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.


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)


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.


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)


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.


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


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


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


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



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



• Attach 16” x 8” propeller to motor and ensure it is safely secured


positioned

• Release E-stop (up), motor now operational.


• Set power at 0 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




positioned

• Release E-stop (up), motor now operational


• 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



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.


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


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


Shoebox lid 1 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



(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


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.



• 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.


• 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


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.



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.



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



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.



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



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


G. Heavy model certification

184

185 APPENDIX G. HEAVY MODEL CERTIFICATION


H. Heavy model requirements

186

187 APPENDIX H. HEAVY MODEL REQUIREMENTS


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


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



• 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


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



• 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


I.2. RANGE CHECKS 194

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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


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



I.4.2 Procedure

1. **Ground Checks

2. **Electronics start-up/shutdown (morphing included)


4. Run down runway

• Slow taxi along runway/ground

• Move elevator and ensure aircraft doesn’t pitch up


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


8. Thrust motor off


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.


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



4. Taxi

5. Ground run

6. return to start on runway

7. Taxi

8. Take off

9. Climb to 2 metres



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.


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



I.6.5 Apparatus

Component No. reqd We Have To Borrow from




8.4V 5350 mAh Li-Po thrust motor batteries 2 Y

90A ESC 1 Y



Transmitter 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





Lead connectors 1 Y

Signal amplifier 1 N Electrical workshop


Various known weights 6 N Final year study room

String 1 ball N Woolworths


Rubber bands 20 Y

Spanner 1 Y

Screw drivers 1 Y

Allen keys 1 set Y




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.



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



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.







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)



I.7.4 Related/required tests

Propulsion system static thrust test

I.7.5 Apparatus

Component No. required We Have To Borrow from



18.5V, 5350 mAh Li-Po thrust motor batteries 2 Y

90A ESC 1 Y



Transmitter 1 Y


Receiver 1 Y

4.8V 1100 mAh receiver battery back 1 Y





Lead connectors 1 Y


Cable ties 4 Y

Spanner 1 Y

Screw drivers 1 Y

Allen keys 1 set Y



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.



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.



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.


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)


4. Taxi

5. Ground run


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



18. landing approaches with flaps (if required/possible)

19. land

20. taxi


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)


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


1. Bring wings fully in and complete one circuit

• Trim aircraft






1. Fully extend both wings and complete one circuit

• Trim aircraft




1. Retract tail one third the way in and complete one circuit

• Trim aircraft




1. Retract tail two thirds the way in and complete one circuit

• Trim aircraft




1. Retract tail all the way in and complete one circuit

• Trim aircraft




1. Fully retract both wings and complete two circuits

• Trim aircraft



• Maximum throttle (maximum speed)

• Cruise speed (same throttle as before)



I.9. F3 - MORPHING MECHANISM TEST 214

1. Fully extend both wings and complete one circuit

• Trim aircraft




1. Fully extend tail and complete one circuit

• Trim aircraft



• 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



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,


3:41:87 Start tail 2/3 in 11:18:25 Good stability and control,


4:02:41 Start tail full in 11:18:45 Good stability and control,


4:29:84 Start wings in 11:19:13 Good stability and control,


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,


6:06:93 Start tail out 11:20:50 Good stability and control,


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


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)


4. Taxi

5. Ground run


7. Taxi

8. Take off


10. trim (approx)



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



17. taxi



20. place in secure position

21. Thrust motor on

22. set throttle at loiter position run until ESC cuts battery power.


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



4. Taxi

5. Ground run


7. Taxi

8. Takeoff (retracted wings, extended tail configuration)- (measure TO distance)


10. trim (approx)




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



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



4. Taxi

5. Ground run




7. Taxi

8. Take off (extended configuration)


10. trim (approx)



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


23. Control switches off, place in secure position

24. Thrust motor on

25. set throttle at 65



I.12. F6 - DIFFERENTIAL SPAN ROLL CONTROL TEST 220

I.12.3 Results

Unable to complete due to weather conditions and damage to UAV.


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.


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


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.



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.


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



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


228

Ris

k R

efer

ence



their effectiveness

(A) –Adequate

(M) – Moderate


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



Ris

k R

efer

ence



their effectiveness

(A) –Adequate

(M) – Moderate


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



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



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


230

Ris

k R

efer

ence



their effectiveness

(A) –Adequate

(M) – Moderate


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



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)


The effectiveness of this strategy will be determined by the ease of repairs should they be required.


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)


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.


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


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)


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� 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



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


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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:



� 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



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



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



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



� 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.


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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



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


250

� BOM � Presentation �

7. 6:00PM- Close:



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


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:



� 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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3. 6:05PM- Technical task: a. Send this to Maziar to review.

4. 6:10PM- Close



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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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.



• 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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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



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


Carlee Stacey


260

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



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


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



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


Carlee 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.



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


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



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


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


268

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



� 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


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



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


272

Meeting 10 - 10/3/2009



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


Tasks to perform completed by everyone Kevin Chan Crystal Forrester Ian Lomas Simon Mitchell Carlee Stacey get forms to everyone to sign



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.


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 ______.



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


276

Meeting 11 - 18/3/2009



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


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



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


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.



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



Meeting 12 - 25/3/2009



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


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



� 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


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


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.



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



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



Meeting 14 - 8/4/2009

Meeting 14.1 Wednesday 8th April 2009


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


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:



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.



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


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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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



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


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.



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


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Meeting 16 - 22/4/2009

Meeting 16.1 Wednesday 22nd April 2009


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


Tasks to perform completed by everyone Kevin Chan Crystal Forrester Ian Lomas Simon Mitchell Carlee Stacey Contact AEA re. sponsorship



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.


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



Meeting 18 - 6/5/2009

Meeting 14.1 Wednesday 6th May 2009


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


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


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



� 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


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


L.

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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

[email protected]

Crystal ForresterCrystal ForresterCrystal ForresterCrystal Forrester 0403 430 916

[email protected]

Ian LomasIan LomasIan LomasIan Lomas 0410 132 319

[email protected]

Simon MitchellSimon MitchellSimon MitchellSimon Mitchell 0423 982 431

[email protected]

Carlee StaceyCarlee StaceyCarlee StaceyCarlee Stacey 0400 714 400

[email protected]

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


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


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.


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.


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


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
