UAV Plane Project

University Of Adelaide School of Mechanical Engineering 2008

Honours Project 637: Design and Build of a Pulsejet UAV

Ryan Anderson

1132309

Nicholas Lukacs

1133184

Mitchell O’Callaghan

1131620

Karn Schumacher

1133398

Michael Sipols

1133364

Terry Walladge

1133113 i

ii

Sir Ross and Sir Keith Smith Fund Acknowledgement and Disclaimer “Research undertaken for this report has been assisted with a grant from the Sir Ross and Sir Keith Smith Fund (Smith Fund) (www.smithfund.org.au). The support is acknowledged and greatly appreciated. The Smith Fund by providing funding for this project does not verify the accuracy of any findings or any representations contained in it. Any reliance on the findings in any written report or information provided to you should be based solely on your own assessment and conclusions. The Smith fund does not accept any responsibility or liability from any person, company or entity that may have relied on any written report or representations contained in this report if that person, company or entity suffers any loss (financial or otherwise) as a result”.

iii

iv

Executive Summary The pulsejet powered Unmanned Aerial Vehicle (UAV) was designed and manufactured by a group of six undergraduate engineering students from the School of Mechanical Engineering at the University of Adelaide. The students, studying a mix of mechanical and aerospace engineering, aimed to design and build a UAV powered by a valveless pulsejet engine, which was also developed throughout the year. The use of pulsejets in aviation history has been almost non-existent since the end of World War II. However, interest in pulsejet technology has increased in recent years, as they offer a cheap and viable alternative from turbojet and ducted fan engines. The design of the aircraft was based around the pulsejet engine and is ultimately intended for use as a high speed target drone or decoy aircraft.

The development of the valveless pulsejet engines followed of from work completed by Coombes et al in 2007, with the aim to produce an engine and fuel system capable for use in flight. A wide range of development was undertaken on three different engines throughout the year, with over 100 static tests performed by the students. Significant improvements were achieved in the areas of engine thrust, thrust specific fuel consumption, engine weight and engine fuelling; most notably achieving successful operation using liquid fuels.

The allowance for pulsejet engine installation meant that a conventional airframe design was not suitable. A classical approach was taken to determine the performance and stability of the airframe. This design incorporated low swept wings, dual vertical stabilizers and an elevated swept tail, to produce an airframe that is capable of pulsejet powered flight. The airframe was manufactured by the students under the supervision and assistance of the Mechanical Engineering Workshop staff, and was constructed primarily from composite materials.

Successful flight of the aircraft was achieved on a ducted fan as it was seen as a more conventional power source, which has similar operational characteristics to the v

pulsejet engines. The flight tests showed that the airframe was stable, controllable and maneuverable. A cruise speed of 150km/hr was achieved during a four minute flight. The aircraft performed all handling requirements during the test flight.

The project goals set by the students at the beginning of the project reflected the ambitious nature of the project. The extension goals were particularly ambitious and related primarily towards the performance of the aircraft and engine. While some goals were not completely achieved, most were well within the performance capabilities of the aircraft.

vi

Acknowledgements The group would like to acknowledge the many people who have helped make this project possible. We would especially like to thank our supervisor Dr. Maziar Arjomandi, whose guidance, support and technical knowledge has been invaluable in ensuring the projects success.

The group would like to acknowledge the Sir Ross and Keith Smith Fund, whose generous contribution was vital for the success of the project. Without the fund’s passion for the development of Aerospace design and technology in South Australia, the project would not be possible.

The group would also like to thank the School Of Mechanical Engineering, ASC and Australian Aerospace for their generous contributions to the project.

The authors would also like to thank and acknowledge all of the individuals who have spent countless hours with the group throughout the year. In particular, a special thanks to Bill Finch, from the Mechanical Engineering Workshop, whose technical knowledge and dedication were invaluable. The personal contribution of James Irvine, from Irvine Aeropulse, for his in-kind sponsorship, guidance and assistance in the development and operation of pulsejet engines was greatly appreciated. Finally, we would like to thank John Modistach, for both his time and effort spent assisting us with aircraft manufacture, as well as for passing on a wealth of knowledge, which assisted us in the manufacture of the airframe.

vii

viii

Disclaimer This statement confirms that the work presented in entirely our own, unless identified otherwise. The work presented was completed as part of the requirements for the Degree of Bachelor of Engineering (Aerospace and Mechanical respectively) at the University of Adelaide. This document describes the work carried out by the students, as recorded in individual project workbooks throughout 2008. The students acknowledge the penalties for plagiarism, fabrication and unacknowledged syndication and declare that the work presented is free of these.

Ryan Anderson

Nicholas Lukacs

-------------------------------------------

----------------------------------------

Date:

Date:

Mitchell O’Callaghan

Karn Schumacher

-------------------------------------------

----------------------------------------

Date:

Date:

Michael Sipols

Terry Walladge

-------------------------------------------

----------------------------------------

Date:

Date:

ix

x

Contents Sir Ross and Sir Keith Smith Fund Acknowledgement and Disclaimer ............................ iii Executive Summary........................................................................................................... v Acknowledgements .........................................................................................................vii Disclaimer .........................................................................................................................ix Contents............................................................................................................................xi List of Figures .................................................................................................................xvii List of Tables ................................................................................................................. xxiii 1

2

Introduction .............................................................................................................. 1 1.1

Project definition................................................................................................ 2

1.2

Project Aims ....................................................................................................... 2

1.2.1

Pulsejet Development................................................................................. 2

1.2.2

Airframe Development ............................................................................... 3

1.3

Project Goals ...................................................................................................... 3

1.4

Extension Goals .................................................................................................. 3

1.5

Scope .................................................................................................................. 4

Feasibility Study ........................................................................................................ 5 2.1

What is a Pulsejet............................................................................................... 5

2.2

Advantages + Disadvantages.............................................................................. 6

2.3

Pulsejet Engines in Aviation History................................................................... 7

2.4

Market Research and Benchmarking................................................................. 7

2.4.1

V-1............................................................................................................... 8

2.4.2

ENICS Drones .............................................................................................. 9

2.4.3

AMT Pulsejet Hobby Aircraft .................................................................... 10

2.4.4

Comparison to turbine engine UAVs or Target Drones............................ 10

2.5

Mission Profile Specifications .......................................................................... 11

2.5.1

Mission Profile .......................................................................................... 11

2.5.2

System Requirements............................................................................... 11

2.5.3

Takeoff methods....................................................................................... 14

2.5.4

Landing Options ........................................................................................ 15 xi

2.6

2.6.1

Valveless Pulsejet - Thermodynamic Cycle ...............................................15

2.6.2

Review of Previous Work ..........................................................................17

2.6.3

Alternative Engine Designs........................................................................21

2.6.4

Exhaust Pipe Development .......................................................................24

2.6.5

Liquid Fuelling ...........................................................................................27

2.6.6

Thrust Augmentation ................................................................................29

2.7 3

Power plant Design...........................................................................................15

Feasibility Study Summary................................................................................30

Conceptual Design...................................................................................................31 3.1

Aircraft Conceptual Design Introduction..........................................................31

3.2

Selecting Preliminary Aircraft Concept.............................................................31

3.2.1

General Configuration...............................................................................31

3.2.2

Fuselage Configuration..............................................................................32

3.2.3

Engine Configuration.................................................................................32

3.2.4

Wing Configuration ...................................................................................32

3.2.5

Empennage Configuration ........................................................................32

3.2.6

Landing Gear Configuration ......................................................................33

3.2.7

Basic Wing Parameters..............................................................................33

3.3

Developing concept for selected configuration ...............................................34

3.3.1

Concept Sketches ......................................................................................35

3.3.2

Statistical Calculations...............................................................................35

3.4

Designing technical parameters for concept....................................................38

3.4.1

Weight Estimation.....................................................................................38

3.4.2

Matching Diagram.....................................................................................39

3.4.3

Aerofoil Selection ......................................................................................42

3.5

Finalisation of Preliminary Aircraft Concept ....................................................45

3.5.1

Variation of Pulsejet Position in Concept Development...........................45

3.5.2

Empennage Design....................................................................................47

3.6

Finalization of Preliminary Aircraft Concept ....................................................52

3.6.1 3.7

Preliminary Conceptual Fuselage Design ..................................................52

Practical Modifications to Final Concept..........................................................53 xii

3.8

Engine Design ................................................................................................... 54

3.8.1

Exhaust Design – Two Stroke Exhaust Similarities ................................... 54

3.8.2

Steady State Diffuser Design..................................................................... 57

3.9

FWE Bellmouth Development.......................................................................... 63

3.9.1

Starting Vortices........................................................................................ 63

3.9.2

Bellmouth Design...................................................................................... 64

3.9.3

Final Design............................................................................................... 67

3.9.4

Flight considerations................................................................................. 68

3.10

Flight Engine Development .......................................................................... 71

3.11

Liquid Fuel System Design ............................................................................ 76

3.11.1

4

Fuel Choice ............................................................................................ 77

3.12

Fuel Injector Design ...................................................................................... 78

3.13

Conceptual Design Summary........................................................................ 83

Detailed Design ....................................................................................................... 85 4.1

Fuselage Structure Design................................................................................ 85

4.1.1

Fuselage Structural Layout ....................................................................... 85

4.1.2

Fuselage Structure Selection .................................................................... 88

4.2

Wing Design ..................................................................................................... 89

4.3

Wing Structural Design..................................................................................... 89

4.3.1

Lifting force profile.................................................................................... 89

4.3.2

Spar Design ............................................................................................... 95

4.3.3

Torsion ...................................................................................................... 97

4.4

Wing Connection............................................................................................ 101

4.5

Control Surface Sizing .................................................................................... 103

4.5.1

Aileron Sizing........................................................................................... 103

4.5.2

Elevator Sizing......................................................................................... 104

4.5.3

Servo Motor Sizing.................................................................................. 105

4.6

Pulsejet Engine Mount ................................................................................... 106

4.6.1

Mounting Locations ................................................................................ 106

4.6.2

Thermal Isolation .................................................................................... 107

4.6.3

Vibration Isolation .................................................................................. 108

4.6.4

Vibration Isolation Method: ................................................................... 110 xiii

4.6.5

Engine Mount Materials..........................................................................112

4.6.6

Final Design .............................................................................................113

4.6.7

Engine Modal Analysis.............................................................................117

4.7

Pulsejet Launch System ..................................................................................118

4.7.1

Launch process........................................................................................118

4.7.2

Launch Stand Components .....................................................................119

4.8

Electrical and Electronic Components............................................................121

4.8.1

Pump and related components...............................................................121

4.8.2

Radio Controller ......................................................................................122

4.9

Ducted Fan......................................................................................................123

4.9.1

Purpose of fan .........................................................................................123

4.9.2

Selection of fan system ...........................................................................123

4.9.3

Modifications to the airframe for Ducted Fan Testing ...........................127

4.10

5

6

Final Stability Analysis.................................................................................129

4.10.1

Longitudinal Moment Analysis ............................................................130

4.10.2

Roll Stability Analysis ...........................................................................134

4.10.3

Ground Performance...........................................................................135

Airframe Manufacture ..........................................................................................137 5.1

Available Manufacturing Methods.................................................................137

5.2

Wing Construction..........................................................................................138

5.3

Empennage Construction ...............................................................................140

5.4

Fuselage Construction ....................................................................................141

5.5

Internal Fuselage Construction ......................................................................144

5.6

Internal Access................................................................................................145

5.7

Propulsion System ..........................................................................................146

5.7.1

Ducted fan ...............................................................................................146

5.7.2

Pulsejet ....................................................................................................147

5.8

Landing Gears and Wheels .............................................................................148

5.9

Control System Installation ............................................................................149

Testing ...................................................................................................................151 6.1

Engine Testing.................................................................................................151 xiv

6.1.1

Phase One Testing................................................................................... 152

6.1.2

Phase Two Testing .................................................................................. 155

6.1.3

Phase Three Testing................................................................................ 160

6.2

6.2.1

Wing Structural Testing .......................................................................... 168

6.2.2

Electrical Component Testing................................................................. 169

6.3

7

8

Aircraft Testing............................................................................................... 168

Aircraft Pre-flight Tests .................................................................................. 170

6.3.1

C.G. Test .................................................................................................. 170

6.3.2

Other pre-flight checks ........................................................................... 171

6.3.3

Location for flying ................................................................................... 172

6.3.4

Pilot ......................................................................................................... 173

6.3.5

Engine and Flight Tests ........................................................................... 173

6.4

Pulsejet Flight Test ......................................................................................... 179

6.5

Discussion of experimental results ................................................................ 180

Management......................................................................................................... 181 7.1

Time Management ......................................................................................... 182

7.2

Financial Management................................................................................... 184

7.3

Risk Management........................................................................................... 185

Conclusion and Future Work ................................................................................ 187 8.1

Review of project goals .................................................................................. 187

8.1.1

Extension Goals....................................................................................... 189

8.2

Project Concerns ............................................................................................ 190

8.3

Future Developments and Recommendations .............................................. 191

References .................................................................................................................... 195 Appendix A - Configuration Selection........................................................................... 199 Appendix B- Weight Calculation Method ..................................................................... 211 Appendix C – Matching Diagram .................................................................................. 221 Drag polar estimation ........................................................................................... 221 Initial estimate of drag polar ................................................................................ 222 Climb requirements .............................................................................................. 223 Stall Requirement ................................................................................................. 223 Takeoff Field Length Requirement ....................................................................... 224 xv

Cruise Requirement...............................................................................................224 Adjusting to take-off values ..................................................................................225 Appendix D – Sensitivity Analysis..................................................................................227 Appendix E – Engine Mounting Calculations.................................................................233 Appendix F – Liquid Fuels..............................................................................................235 Appendix G – Component Weight Breakdown .............................................................239 Appendix H – Test Log Books ........................................................................................241 Appendix I – Fuselage Stress Analysis ...........................................................................279 Appendix J- Gantt Charts...............................................................................................281 Appendix K- Risk Register..............................................................................................285 Appendix L- Meeting Minutes.......................................................................................287 Appendix M- Drawings..................................................................................................367

xvi

List of Figures Figure 1 - Valved and Valveless Pulsejet Designs ............................................................. 5 Figure 2 - Comparison of Engine Costs ............................................................................. 6 Figure 3-View of the V-1 ................................................................................................... 8 Figure 4-E-95 Ramp Launch .............................................................................................. 9 Figure 5- Flight Profile..................................................................................................... 11 Figure 6 - Statistical Trends of Target Drone UAVs ........................................................ 13 Figure 7 - Ideal Lenoir Cycle............................................................................................ 16 Figure 8-Pobezhimov modified Lenoir cycle................................................................... 17 Figure 9 - Valveless Pulsejet Engine (Carolina State University) .................................... 18 Figure 10 - Focus Wave Energy (FWE) Pulsejet Engine .................................................. 20 Figure 11-Chinese Valveless Pulsejet Engine.................................................................. 21 Figure 12-Lockwood Prototype ...................................................................................... 22 Figure 13-Escopette Valveless Engine ............................................................................ 23 Figure 14 - Interaction of Escopette Pressure Waves .................................................... 24 Figure 15 – A Focused Wave (FWE) Pulsejet engine ...................................................... 25 Figure 16 – A Lockwood-Hiller style Pulsejet engine,..................................................... 26 Figure 17 – A Focused Wave engine variation, the “FWE VIII - Lady Anne Boleyn”. ..... 27 Figure 18- optimised thrust augmenter as used on a valved pulsejet ........................... 29 Figure 19-Configuration Concept Sketches .................................................................... 35 Figure 20: Statistical Thrust Loadings of Jet UAVs.......................................................... 36 Figure 21: Wing Loading Versus Weight of Jet UAVs ..................................................... 37 Figure 22: Statistical Concept ......................................................................................... 38 Figure 23: Matching Diagram......................................................................................... 40 Figure 24: Example of Early Design................................................................................. 46 Figure 25: Second Phase Design Example ...................................................................... 46 Figure 26: Final Engine Position...................................................................................... 47 Figure 27: Centre of Gravity Excursion Diagram............................................................. 49 Figure 28: Longitudinal X-Plot......................................................................................... 50 Figure 29: Lateral Stability X-Plot.................................................................................... 52 Figure 30 -Conceptual Fuselage Design.......................................................................... 53 xvii

Figure 31: Modifications to final aircraft concept...........................................................53 Figure 32- Advancements in Two Stroke Exhaust Design ...............................................55 Figure 33 – The effect of expansion angle on wave behaviour. .....................................56 Figure 34: Loss Coefficient for a Conical Diffuser ...........................................................58 Figure 35 – The UFLOW1D model used to investigate expansion angles.......................59 Figure 36 – Combustion chamber pressure extremes for different expansion angles. .60 Figure 37 - Statistical data showing exhaust expansion angles from similar engine designs.............................................................................................................................61 Figure 38 - statistical data showing a trend between combustion chamber diameter and expansion diameter..................................................................................................61 Figure 39 - The final expansion design............................................................................63 Figure 40 - PIV images of vortex interaction...................................................................64 Figure 41 - Bellmouth designs considered (Blair, Cahoon 2006) ....................................65 Figure 42 - Performance of bellmouth designs...............................................................65 Figure 43 - The data obtained in 2007 using UFLOW1D (blue) and textbook recommendations (red) ..................................................................................................66 Figure 44 - The adjustable bellmouth design..................................................................68 Figure 45- Three intake geometries ................................................................................69 Figure 46- Domain Layout ...............................................................................................69 Figure 47- Effect of intake geometry on mass flow rate.................................................70 Figure 48 -Static Pressure Contours of Aerodynamic Flare at 80m/s.............................70 Figure 49-Static Pressure Contours on Standard Flare at 80m/s....................................71 Figure 50- Statistical trend of Chinese and FWE engines ...............................................72 Figure 51-Variation of material properties of 310 stainless steel with temperature.....73 Figure 52-Operating pressure of the Escopette pulsejet................................................74 Figure 53- full engine mesh.............................................................................................75 Figure 54-pressure loading input for flexible dynamic solver.........................................75 Figure 55-Stress Results on Combustion Chamber End Cap...........................................76 Figure 56- Final Results of the Axi Symmetric Model .....................................................76 Figure 57- 12 hole swirl injector......................................................................................79 Figure 58- 6 hole opposed spray injector........................................................................79 xviii

Figure 59-BETE PJ Cone Spray Injector ........................................................................... 80 Figure 60-5mm stainless steel injectors ......................................................................... 81 Figure 61: Conceptual Design Three View...................................................................... 83 Figure 62: Overview of Fuselage Structural Layout....................................................... 85 Figure 63: Fuselage Internal Reinforcing Structure ........................................................ 86 Figure 64 - Schrenk's Approximation.............................................................................. 90 Figure 65 -Lifting Force Distribution ............................................................................... 91 Figure 66- wing shear distribution.................................................................................. 92 Figure 67 - wing bending force distribution ................................................................... 92 Figure 68 - Corrected Cl Distribution .............................................................................. 93 Figure 69 - Lift Distribution at 88m/s.............................................................................. 94 Figure 70 - Maximum spar thickness from root to tip of the wing ................................ 96 Figure 71 - Position of Centre of Pressure with AOA...................................................... 98 Figure 72 - Wing Torque at Takeoff (70km/hr)............................................................... 99 Figure 73 - Wing Torque at Climb Speed (150km/hr)..................................................... 99 Figure 74 - Wing Torque at Cruise Speed (300km/hr).................................................. 100 Figure 75- Wing Connection System............................................................................. 101 Figure 77 - The mounting extension on the front of the Chinese engine .................... 107 Figure 78 - Force transmissibility as a function of frequency ratio and damping ratio109 Figure 79 - Yield stress relative to room temperature as a function of temperature for 301,302,304,321,347 annealed stainless steels ........................................................... 113 Figure 80 - The final engine mount design. The modification made to the front of the engine is shown in green. ............................................................................................. 114 Figure 81 - Thermal analysis results of the engine mount. .......................................... 114 Figure 82 - Stress distribution within the initial design under an 80N load................. 115 Figure 83 - Stress distribution within the design under dynamic loading of 40N +- 20N ...................................................................................................................................... 116 Figure 84 - 208Hz vibration mode of the engine, mounted at ends ............................ 117 Figure 85 - Release tab attached to the intake of the engine, and release tab on the launch stand.................................................................................................................. 119 Figure 86 - Launch stand for pulsejet flight .................................................................. 120 Figure 87: Flight Works Fuel Pump .............................................................................. 121 xix

Figure 88 - Schubeler Ducted Fan .................................................................................123 Figure 89 - Lehner electric motors ................................................................................125 Figure 90 - ZIPPY-R battery pack ...................................................................................126 Figure 91 - Ducted fan mounted in the airframe ..........................................................127 Figure 92 - Ducted Fan Mounting Tabs .........................................................................128 Figure 93 - The cover design for the ducted fan ...........................................................129 Figure 94 - The cover installed on the plane.................................................................129 Figure 95: Centre of Gravity and Aerodynamic Centre Excursion Diagram..................131 Figure 96:Cm-Cl Graph (Power On)...............................................................................132 Figure 97: Cm-Cl Graph (Power On)..............................................................................132 Figure 98: Cm-Cl Graph (Pulsejet) .................................................................................133 Figure 99: Roll Stability Contributions...........................................................................135 Figure 100 - Rib Installation in Wings............................................................................139 Figure 101 - Wing structure schematic .........................................................................140 Figure 102 - Servo Installation.......................................................................................140 Figure 103 - a) Horizontal tail joined as a single piece, b) horizontatal tail after glassing, c) installation of vertical tail onto fuselage...................................................................141 Figure 104 - Fuselage plug.............................................................................................142 Figure 105 - Gel coat being applied to plugs in preparation for creating the moulds..143 Figure 106 - Fuselage.....................................................................................................144 Figure 107 - Location of bulkheads (blue) and longerons (red)....................................145 Figure 108 - The aircraft showing the both access panels a) removed and b) attached .......................................................................................................................................146 Figure 109 – Schubeler ducted fan (Schubeler Jets, 2008)...........................................147 Figure 110 - Front pulsejet engine mount ....................................................................148 Figure 111 - Front landing gear steering system...........................................................149 Figure 112 - Test System Layout ...................................................................................151 Figure 113 – Reducing thrust during extended operation............................................153 Figure 114 – Effect of fuel injection position on engine performance .........................153 Figure 115 – Effect of exhaust and intake length on engine performance ..................154

xx

Figure 116 - The adjustable FWE engine with expanding tail section and 100mm extension....................................................................................................................... 156 Figure 117 - Affect of injector position on engine thrust ............................................. 157 Figure 118 - Thrust Results ........................................................................................... 158 Figure 119 - Visible damage to ceramic coating........................................................... 159 Figure 120 - liquid fuel injectors placed mid way along the intake tube ..................... 161 Figure 121 - Opposed injector configuration................................................................ 163 Figure 122 - Performance of the Chinese engine with different injector placements. 163 Figure 123 - engine performance on various fuels....................................................... 164 Figure 124 - Performance of the Chinese engine for various lengths.......................... 166 Figure 125 - Aircraft testing flow chart......................................................................... 168 Figure 126 – Load zones for wing structural testing .................................................... 168 Figure 127 - Experimental Wing Deflection.................................................................. 169 Figure 128 - C.G. Test Setup.......................................................................................... 170 Figure 129 - C.G. Test Photo ......................................................................................... 171 Figure 130 - Ground Roll Test at Gawler Airfield.......................................................... 174 Figure 131 - Plotted flight path from GPS logger.......................................................... 176 Figure 132-Compact Gantt Chart.................................................................................. 183 Figure 133-Cost Breakdown.......................................................................................... 185 Figure 134: Mock graphic of selected configuration ................................................... 209 Figure 135- Graph of WE/WO Vs WO........................................................................... 212 Figure 136- Graph of WE/WO Vs WO........................................................................... 217 Figure 137- Graph of WE/WO Vs WO for Consistent Data........................................... 219 Figure 138: First Estimate of Drag Polar ...................................................................... 222 Figure 139: Sensitivity to fuel consumption ................................................................ 227 Figure 140: Sensitivity to Engine Weight ..................................................................... 228 Figure 141: Sensitivity of Cruise Speed to W/S ............................................................ 228 Figure 142: Sensitivity of Takeoff Distance to W/S ...................................................... 229 Figure 143: Sensitivity of Climb Rate to W/S................................................................ 229 Figure 144 : Sensitivity of Stall Speed to W/S............................................................... 230 Figure 145: Sensitivity of Cruise Speed to T/W ............................................................ 230 Figure 146 : Sensitivity of Takeoff Distance to T/W ..................................................... 231 xxi

Figure 147 : Sensitivity of Climb Rate to T/W ...............................................................231 Figure 148- Engine during Test......................................................................................245 Figure 149- Thrust Vs Time for Test 6 ...........................................................................247 Figure 150- Thrust Vs Time for Test 7 ...........................................................................248 Figure 151- Thrust Vs Time for Test 8 ...........................................................................248 Figure 152- Thrust Vs Time for Test 9 ...........................................................................249 Figure 153- Thrust Vs Time for Test 10 .........................................................................250 Figure 154- Thrust Vs Time for Test 11 .........................................................................250 Figure 155- Thrust Vs Time for Test 14 .........................................................................252 Figure 156- Thrust Vs Time for Test 15 .........................................................................252 Figure 157- Thrust Vs Time with injector 32mm from intake mouth ...........................259 Figure 158- Time Vs Thrust for FWE with expanding Exhaust ......................................268 Figure 159 - Engine performance on liquid fuels ..........................................................278

xxii

List of Tables Table 1- Specifications of the V-1 ..................................................................................... 8 Table 2- Specifications of ENICS E95 Target Decoy .......................................................... 9 Table 3- Specifications of AMT Pulsejet aircraft............................................................. 10 Table 4- Lockwood Performance Data ........................................................................... 23 Table 5 : Requirements and Input Data of Matching Diagram....................................... 40 Table 6 – Characteristics of Suitable Aircraft.................................................................. 41 Table 7: Characteristics of Possible Aircraft ................................................................... 41 Table 8- Initial Aerofoil Analysis ..................................................................................... 43 Table 9: NACA 4 Digit Aerofoil Analysis .......................................................................... 43 Table 10- Suitable Tail Aerofoils ..................................................................................... 44 Table 11........................................................................................................................... 82 Table 12: Fuselage Stress Analysis Results ..................................................................... 88 Table 13 - Aileron Dimensions ...................................................................................... 104 Table 14- Servo Requirements...................................................................................... 105 Table 15 - Spring stiffness and deflection under a 40N thrust load, for various frequecy ratios. ............................................................................................................................ 110 Table 16: Material Selection for Engine Mount............................................................ 112 Table 17 - Ducted Fan Parameters ............................................................................... 124 Table 18 - Parameters of Lehner 1950 Electric Motor ................................................. 125 Table 19 - Expanding Exhaust Test Results................................................................... 156 Table 20 - Maximum control surface/servo motor deflection ..................................... 170 Table 21 - Flight data from GPS logger ......................................................................... 176 Table 22- Hours Worked By Group Members .............................................................. 184 Table 23: General Configuration Decision Matrix ........................................................ 201 Table 24: Fuselage Configuration Decision Matrix ....................................................... 202 Table 25: Engine Configuration Decision Matrix .......................................................... 203 Table 26: Wing Configuration Decision Matrix............................................................. 204 Table 27: Wing Height Decision Matrix ........................................................................ 205 Table 28: Wing Sweep Decision Matrix ........................................................................ 206 Table 29: Empennage Configuration Decision Matrix .................................................. 207 xxiii

Table 30: Landing Gear Type Decision Matrix...............................................................208 Table 31: Landing Gear Arrangement Decision Matrix.................................................208 Table 32- Weight Data for Piston UAVs ........................................................................211 Table 33- Table of First Iterations .................................................................................216 Table 34- UAV Data .......................................................................................................216 Table 35- Consistent UAV Weight Data ........................................................................218 Table 36- Iteration Results ............................................................................................220 Table 37- Fuel Flash Point Data.....................................................................................235 Table 38: Fuel Energy Density Data...............................................................................236 Table 39: Fuel Optimal AFR Data...................................................................................236 Table 40- Fuel Flammability Limit Data.........................................................................237 Table 41- Latent Heat of Vaporisation Data..................................................................238 Table 42 : Component Weight Breakdown...................................................................239

xxiv

1 Introduction The purpose of this project was to design and manufacture a Valveless Pulsejet Powered Unmanned Aerial Vehicle (UAV), suitable for use as a Target Drone or Decoy UAV. The project aimed to develop an understanding of valveless pulsejet engines, as well as developing a prototype engine, with the aim of showing that they are a cheap and viable alternative form of propulsion.

A feasibility study was initially conducted in order to develop project goals and define a realistic scope. This stage included an extensive study of all valveless pulsejet engines, developed by academics and enthusiasts, in order understand the working characteristics of the engines, and to better understand how to optimise and improve the operating characteristics of these engines. A study of target drones and decoy aircraft, powered by both pulsejet and turbo jet engines was undertaken to help develop the fundamentals of the aircraft, as well as to identify some of the key issues which needed to be addressed in the following design stages.

The engine design was initially a continuation and modification of a Focused Wave Energy (FWE) valveless pulsejet engine developed by Coombes et al in 2007 at the University of Adelaide.

The aircraft design was produced initially using a combination of statistical and numerical analysis, in consultation with aircraft design literature. The aircraft design was then progressively refined in an iterative manner.

This project has involved a significant testing section, with over 100 static engine tests, conducted, and two successful aircraft flights. The data obtained from this project has helped to better develop the understandings of valveless pulsejet operational characteristics, particularly with liquid fuels.

This report shows the development steps which were utilised to ensure the project was completed on time; on budget and that all goals were achieved. This project has 1

Chapter 1 Introduction shown through proof of concept that valveless pulsejet engines are a viable form of propulsion for short range target drones and decoy UAVs.

1.1 Project definition The project continues on from the work conducted by Coombes et.al in 2007. This project expands from that work to concern the development of a valveless pulsejet powered UAV.

The preceding data and research from the detailed feasibility study and bench-marking has been synthesised to produce this project definition, which outlines the aims and objectives of the project. This project definition is categorised into the Pulsejet and Airframe development areas.

1.2 Project Aims The project aims to show that valveless pulsejets are a viable alternative engine for short range and low cost UAV aircraft. To ensure the project is completed the following must be achieved.

1.2.1 Pulsejet Development •

Continued development of a valveless pulsejet, with the aim of increasing the overall thrust of the engine and a reduction in engine weight based on 2007 results.

•

Research and development of a liquid fuel delivery system for pulsejet engines.

•

Develop a fuel system that is of suitably low weight for flight.

•

Successfully test and record key performance criteria of the pulsejet on both gas and liquid fuel mixes.

•

Development of alternative engine designs which may be suitable for future development.

•

Completion of all tasks within the allocated budget.

2

Section 1.3 Project Goals 1.2.2 Airframe Development •

Successful design of a low weight air-frame, based on the parameters of an estimated 3kg of thrust and of plausible 300km/hr max speed.

•

Successful manufacture of the airframe using composite materials within the specified weight.

•

Successful flight of the UAV.

•

Completion of all tasks within the allocated budget.

1.3 Project Goals The project success was based on the completion of the following goals: •

To modify, build and manufacture a valveless pulsejet, with the aim of producing 3kg of thrust, with an engine weight of 1.5kg or less. This goal will be quantified by the output received from the thrust measurement stand constructed during the 2007 Project.

•

Develop a liquid fuelled system for a pulsejet engine and integrate a flight weight version into the UAV design.

•

Based on the desired pulsejet specifications; design, develop and build a lightweight UAV capable of sustaining flight for 10 minutes with thrust supplied by a valveless pulsejet engine.

•

Achieve a cruise speed of over 200km/h. As measured by onboard GPS or a similar system.

•

Achieve a flight time of 10 minutes

•

Gain a better engineering perspective on the workings of pulsejets, with the aim of developing different engine design alternatives.

1.4 Extension Goals Completion of extension goals will show above expected outcomes from the project: •

Achieve 3.5kg of thrust from a valveless pulsejet engine. 3

Chapter 1 Introduction •

Achieve a cruise speed of 250 km/h or above.

•

Increase flight time of the proposed liquid fuelled pulsejet UAV to over 15 minutes.

•

Manufacture an alternative engine design for future development.

1.5 Scope The project scope is limited to the successful completion of the goals specified above. The project aimed to design an airframe capable of supporting a pulsejet engine. While this involved some optimisation of the airframe structure, the project only intended to develop a proof of concept aircraft. The project is aimed to develop an aircraft capable of flight, the scope of the project is limited to the initial development and manufacturing stages, further alteration and optimisation after successful flight was minimal.

The engine and fuel delivery systems were designed to be capable of producing 3kg of static thrust for an estimated 10 minute flight time. The scope of this project was therefore limited to the development of these systems to a level which will allow the aircraft to sustain flight for the desired time period. Continued development of the systems to optimise flight of the aircraft is not anticipated unless project goals are not fulfilled.

4

2 Feasibility Study The aim of this feasibility study was to construct realistic project goals, understand the challenges and risks involved with the project and formulate a logical and progressive development plan for the remainder of the project. This was completed through a literature review and market survey of both the airframe and pulsejet components of the project.

2.1 What is a Pulsejet A pulsejet engine is a form of combustion engine with few to no moving parts. The engine comes in two forms, valved or valveless (Figure 1). Both engines have similar layouts, consisting of an intake, combustion chamber and exhaust.

Figure 1 - Valved and Valveless Pulsejet Designs (Pulse-jet.com 2008)

The main difference between the two engines is in the use of a valve to direct the flow out of the exhaust tube. This valve was the main source of problems in early pulsejets.

As can be seen in Figure 1, the valve in the valved engine is positioned inside the combustion chamber. The combination of extreme heat and violent closing movement of the valve meant that valved pulsejets often experienced lifetimes only lasting several minutes before the vales fatigued.

5

Chapter 2 Feasibility Study The ‘valveless’ pulsejet uses an aerodynamic valve, created by the differences in length between the intake and exhaust, in order to sustain operation. This means the engine has no internal parts, and thus is significantly more reliable once and effective engine layout has been created. It is for this reason that valveless pulsejet engines have been investigated in this project.

2.2 Advantages + Disadvantages The main advantage of valveless pulsejet engines is in their extreme low cost, as shown in Figure 2. This is due to the engines simple design, and use of low cost and readily available materials and manufacturing methods. This makes them an excellent power plant for low cost target drones and decoy UAV (Tao 2006).

Figure 2 - Comparison of Engine Costs

Pulsejet engines however suffer mostly from their poor thermodynamic efficiency (outlined in section 2.6.1), which means the specific fuel consumption of the engines is significantly greater than that of common turbojet or turbofan aircraft.

The other disadvantage of pulsejet engines is the extreme levels of noise and vibration they emit. This factor rules out the use of pulsejets in anything other than military applications.

However the most interesting and exciting area of pulsejet engines is in the combustion mechanism. Pulsating combustion is self compressing, so that the air fuel 6

Section 2.3 Pulsejet Engines in Aviation History mixture does not burn steadily, but in bursts. This makes pulsejet engines an excellent research engine, as many of the fundamental theories have been investigated on pulsejets, before the construction of a large scale Pulse Detonation Engines (Wilson, Dougherty 2002).

2.3 Pulsejet Engines in Aviation History The pulsejet engine first found application in aircraft in 1891. Pulsejet engines have been used throughout aviation history in several applications, including unmanned military vehicles, early missile development, and vertical takeoff and landing (VTOL) research, however much of the recent research has been undertaken by model aircraft enthusiasts.

Sometime after the invention of the Pulsejet the Pulsejet powered German V-1 Missile was produced. This missile is the pulsejet powered aerial vehicle produced in the largest quantities with approximately 30,000 units manufactured. The V-1 missile utilised a valved pulsejet engine and during tests of the V-1 significant failures occurred, even though the aircraft only flew for less than 20 minutes (Goeble 2003).

In modern times much development in pulsejet engines has come from model aircraft hobbyists, due to its low cost and comparable ease of manufacture.

The Pulsejet Engine has been of interest to commercial manufacturers throughout several brief periods in history. Pulsejets have been used commercially and for the military as propulsion devices for target drones.

2.4 Market Research and Benchmarking The aim of this section was to gain an understanding into the capabilities and aircraft configuration styles of pulsejet powered aircraft. Due to the lack of such aircraft, the study was extended to both hobby aircraft and jet powered target drones and decoy aircracft. 7

Chapter 2 Feasibility Study 2.4.1 V-1 The V-1 was the first pulsejet powered aircraft, used by the German Air force during World War II as a low cost and high quantity missile. It was the first mass-produced guided missile and first jet powered aircraft. Specifications of the V-1 can be seen in Table 1. The design of the V-1 is shown in Figure 3.

Figure 3-View of the V-1 (Naughton 2001) Table 1- Specifications of the V-1 (Combined from: Werrel 1985, Goebel 2003, Naughton 2001)

Engine Thrust (kg) Take-off weight (kg) Speed (kph) Span (m)

Argus valved pulsejet ‘109-014’ 272 2150 630 5.3

While the size and weight of the aircraft is significantly larger than the anticipated UAV weight, it useful for analysis as it is one of few aircraft which has been powered by a pulsejet engine.

8

Section 2.4 Market Research and Benchmarking

2.4.2 ENICS Drones ENICS is a Russian company which provides pulsejet powered decoy aircraft for military training. ENICS produces fully manufactured drones and engines in three different configurations. Full details of their E95 target decoy can be seen in Table 2. Table 2- Specifications of ENICS E95 Target Decoy (Enics 2006a)

Engine Engine weight (kg) Thrust (kg) SFC (kg/kg/hr) Take-off weight (kg) Span (m) Speed (kph) Range (km) Endurance (min) Launch

Enics M44D pulsejet 900 mm length, 75mm diameter 0.9 20 6.61 70 2.4 400 70 30 Ramp, pneumatic

Figure 4-E-95 Ramp Launch (Enics 2006b)

The aircraft is larger than the estimated project design, however its use as a target drone and use of a valveless pulsejet engine make it an excellent aircraft for analysis. 9

Chapter 2 Feasibility Study 2.4.3 AMT Pulsejet Hobby Aircraft Pulsejets are moderately popular as propulsion systems for jet model aircraft. Pulsejets are attractive to many pilots as they are low cost and offer good thrust to weight ratios. In most cases commercially available ‘valved’ engines are used. This ‘AMT Pulsejet’ is a custom built delta wing aircraft with a modified valved pulsejet producing 8.7kg of thrust. The specifications of this aircraft can be seen in Table 3. This aircraft is useful for analysis as it is close to the expected weight of the aircraft and its use of a pulsejet allows analysis of expected fuel consumption during flight. Table 3- Specifications of AMT Pulsejet aircraft (AMT 1998)

Engine Engine length (mm) Engine diameter (mm) Thrust (kg) Take-off weight (kg) Empty weight (kg) Speed (kph) Span (m) Fuel Consumption

Custom valved pulsejet 880 90 8.7 7.5 5.9 390 1.12 500 mL/min [50% Kerosene, 40% Gasoline, 10% Propylene Oxide]

2.4.4 Comparison to turbine engine UAVs or Target Drones Turbine engines UAVs similar in size to the project aircraft have a large advantage in terms of thrust to weight comparison to pulsejet engines. This is due to the compactness of the engines, as well as comparably lower fuel consumption figures. However the main disadvantage of these types of engines is the cost of the engine for the similar amount of thrust, as shown in Section 2.2. For the statistical design of the aircraft, turbine powered UAVs will be included in the analysis, due to the low number of pulsejet powered aircraft, specifically of comparable size to the anticipated design size of the aircraft.

10

Section 2.5 Mission Profile Specifications 2.5 Mission Profile Specifications 2.5.1 Mission Profile Based on the analysis of the aircraft in Section 2.4 , the mission profile of the aircraft was developed. As the aircraft was aimed to be developed as a ‘proof of concept’ aircraft, it was decided that the mission profile would be kept simple. The profile can be seen in Figure 5.

Figure 5- Flight Profile

Further details are specified for some sections of this profile: •

Start up and warm up - with pulsejets this is especially critical, as the engines must be stable before launch. As a result, the fuel consumption during this period will be significantly higher than for other engine types.

•

Loiter – The goal of the flight is top remain airborne for 10-15 minutes with no set range goal therefore the flight will take place within the line of sight of the pilot.

2.5.2 System Requirements The system requirements define the abilities of the aircraft and its components. These values are determined from know requirements, calculations and the market research performed on similar aircraft.

11

Chapter 2 Feasibility Study Cruise Speed Requirements In section 2.4, different aircraft that utilized pulsejet engines or jet engines for power were analysed. A suitable requirement for the cruise speed can be decided based on that data and other calculations. A realistic cruise speed requirement was determined based on numerous things.

Direct Bench Marking The direct bench marking here refers to other pulsejet aircraft of similar size. A small pulsejet aircraft presented earlier that was similar was the AMT. This had a top speed of 390 kph, but also had a thrust to weight ratio greater than what we are aiming for.

Collated statistics of other Jet UAVs Of the aircraft identified in the research and benchmarking section it can be seen that most jet powered target drone aircraft have a cruise speed of approximately 400km/hr. However these aircraft have high thrust loadings and also high wing loadings which reduce wing area and thus drag/weight. These characteristics are allowed by the use of rocket-assisted launch and/or ramp launches and multiple or more powerful engines. The following graph shows the thrust loading for a variety of aircraft. It can be seen that the mean thrust loading is approximately 0.3. This was the basis for all further aircraft development.

12

Section 2.5 Mission Profile Specifications

Figure 6 - Statistical Trends of Target Drone UAVs

Estimated available thrust The amount of thrust currently available from the engine that is to be used for this aircraft is 2.3kg (Coombes et.al 2007). This is described in more detail in Section 2.6.2

Estimated possible speed Using calculations that estimate the drag based on the estimated drag in conjunction with thrust and weight, it was possible to estimate the possible top. For thrust around 2.3 kg, and with a thrust loading of 0.3, from the mean of the collated data, top speeds of 200-250 kph are possible.

Plausible Cruise Speed Requirement Based on the above considerations, a realistic cruise speed requirement was deemed to be 200 kph.

Control and Electronic Requirements There were two separate control functions for the aircraft, control of the flight and control of the engine.

Control of Flight

13

Chapter 2 Feasibility Study In selecting a control mechanism for the flight of the UAV it was determined that simplicity was of importance, due to this a common remote controlled system was deemed appropriate over other systems.

Control of the Engine The thrust produced by the engine could be controlled by varying the fuel flow rate supplied to the engine by the fuel pump.

2.5.3 Takeoff methods There are numerous ways that a UAV system can be launched including trolley launched and fixed gear. These methods have significantly different characteristics and will be discussed.

Trolley Launched The idea of a UAV being launched from a trolley or with a detachable landing gear is that once the aircraft leaves the ground the trolley or gears detach from the aircraft. This can be done using the propulsive power of the aircrafts own propulsion system and or with a supplementary propulsion system such as rockets or sling shot. This takeoff method requires that the aircraft has an alternate landing method other than via landing gear. The advantage of this launch method is that it has no need for a landing gear which would decrease drag during flight. However the main disadvantage is that it requires an alternate landing method such as a parachute.

Fixed Gear For an aircraft taking off from an attached landing gear both fixed and retractable types of landing gear designs can be considered. The main advantage of a fixed gear is that the system is reliable and simple however it has a disadvantage of increased drag during flight.

14

Section 2.6 Power plant Design 2.5.4 Landing Options There are several alternative landing methods for UAVs without a conventional landing system. These options have been considered to determine the overall risk and feasibility of designing an aircraft without a conventional style landing gear. The four options considered are parachute, belly landing, net catch and air cushioned landing.

Parachute There are numerous advantages to landing an aircraft with a parachute. Recovery parachutes are commercially available at a relatively low cost and they produce minimal extra drag in comparison to a fixed landing gear system. The main disadvantages of a parachute recovery are the complexity and weight of the system and the high loads experienced when the parachute is first deployed.

Belly Landing The use of a belly landing for an aircraft has numerous benefits, primarily the minimal effect on drag, the slight effect on the weight of the aircraft and the low complexity of the system. For a belly landing the underside of the aircraft is reinforced to withstand the forces created by the impact of the aircraft with the ground, which is the main disadvantage of this system.

2.6 Power plant Design The power plant for the engine was defined by the initial project outline. This section outlines the initial research that was conducted by the group into the workings, research and challenges that exist in designing a valveless pulsejet engine.

2.6.1 Valveless Pulsejet - Thermodynamic Cycle Pulsating combustion is the main area of confusion for researchers attempting to successfully understand the operation of pulsejet engines. In the research conducted, it has been noticed that different authors associated the behaviour of the engines to 15

Chapter 2 Feasibility Study different phenomenon. The self sustaining, periodical nature of the combustion is generally associated with either wavy, acoustic or vortex nature (Pobezhimov 2006), however models using these analysis generally can only describe parts of the combustion process accurately. A thermodynamic approach can be used to explain the operating process of a pulsejet engine, and show the advantages that exist in pulsating combustion.

The operating cycle of a pulsejet engine can be described by

modifications to the Lenoir Cycle which can be seen in Figure 7.

Figure 7 - Ideal Lenoir Cycle (Pobezhimov 2006)

The operating cycle is described in thee steps: 1-2

Constant volume (isochoric) heat addition

2-3

isentropic expansion.

3-1

Constant pressure (isobaric) heat rejection - compression to the volume at the start of the cycle.

The main difference between the Lenoir cycle and a pulsejet cycle is that during heat addition the process is neither isochoric or isobaric, as there is a combination of pressure release, and heat release (McCalley 2006). This is because the engine operates from wave compression which is relatively weak; therefore combustion is not confined to the combustion chamber, but occurs down the length of the engine. A more realistic diagram can be seen in Figure 8.

16

Section 2.6 Power plant Design

Figure 8-Pobezhimov modified Lenoir cycle (Pobezhimov 2006)

2.6.2 Review of Previous Work From the research conducted, it was found that the development of pulsejet engines has been the recent study of several universities. The two of interest to this project were studies conducted by North Carolina State University and The University of Adelaide. The work conducted by these two bodies allowed for a better understanding on the fundamentals of pulsejet engine operation and optimisation.

North Carolina State University Within the past decade, numerous investigations have been conducted by North Carolina State University Masters students, under the direction of Dr. William L. Roberts into various areas of pulsejet engine development. Studies have included:

•

Experimental Investigations into Pulsejet Engines

•

Experimental Investigations Into The Operational Parameters of a 50 Centimetre Class Pulsejet Engine

•

Experimental Investigations in 15 Centimetre Class Pulsejet Engines

•

Experimental Investigations of 8 Centimetre Class Pulsejet Engines

•

Experimental Investigations of Liquid Fuelled Pulsejet Engines

•

Numerical Simulations of Pulsejet Engines

17

Chapter 2 Feasibility Study These investigations have aimed to better understand the operating characteristics of valved and valveless pulsejet engines, as well as attempting to develop small scale engines for use with small size UAV and MAV aircraft, as the efficiency of commonly used turbojets becomes lower as the size of the engine decreases (Tao 2006).

This section covers some of the key research conducted by these projects, with focus on fundamental operating theories and engine performance.

Work into the

development of liquid fuelled pulsejet engines, as conducted by McCalley in 2006, is covered in section 27.

Valveless Engine Studies Valveless engines studies conducted at Carolina State University have revolved around the analysis of straight exhaust valveless engines, known as Schubert jets, (Figure 9), with varying lengths and diameters of the intake and exhaust pipes. Schubert jets are known for their ease of manufacture, but low thrust and high specific fuel consumption.

Figure 9 - Valveless Pulsejet Engine (Carolina State University)

Experimental data was taken from the engines via a number of different mechanisms, including instantaneous pressure sensors, manometers, thermocouples and SPL meters.

18

Section 2.6 Power plant Design Experiments in varying the length of the valveless pulsejet engine showed a direct correlation between operating frequency and length. This frequency can be linked directly to the Helmholtz frequency for the intake pipe (Equation 1) and a 1/4 wave tube frequency for the exhaust (Equation 2).

Equation 1

Equation 2

f =

C 4L

It was found that both these frequencies act together to give the engine operating characteristics which are similar to that of a 1/6 wave tube. This was compared to tested data and was found to be accurate to within 5%. The equation is temperature dependant, which suggests that changes in area in the engine can cause a change in the operation frequency. Also, changes in fuel will alter the burn temperature and thus affect the engines operating characteristics. However it was found that if the intake and exhaust frequency are within 10% of each other, the engine will successfully operate.

Studies by Ordon in 2006 showed that this frequency characteristic is altered significantly by changes in geometry, as these cause reflections in the waves, which effect how the jet operates. It was found later by Kiker that the operating frequency of the pulsejet scales as the inverse of the inlet length and reducing the exhaust diameter of the pulsejet has very little effect on its operating frequency. With respect to combustion chamber pressures, Kiker found that pressure scaled inversely with exit diameter and directly to fuel flow rates. He also investigated the use of platinum coating in a 5cm pulsejet to act as a catalyst and increase chemical reaction time.

19

Chapter 2 Feasibility Study 2007 Study – University of Adelaide In 2007, a study by Coombes et.al Al 2007, was conducted at Adelaide University into the devolvement and testing of a valveless pulsejet engine and thrust measurement stand. The work aimed to create an engine capable of 3kg of thrust, with an engine weight of under 2kg, a stand capable of accurately measuring the engines thrust during tests and a software package to be used to predict pulsejet performance and allow the optimization of engines The groups work focused on the development of a Focus Wave Energy (FWE) Valveless pulsejet engine, as shown in Figure 10, which was originally developed by notable pulsejet engine developer, Larry Contril.

Figure 10 - Focus Wave Energy (FWE) Pulsejet Engine

Two engines were developed, the first based on statistical design, with adjustable lengths. This engine aimed to investigate the effect of the intake and exhaust lengths on engine performance, and a fixed length engine, developed based on findings from the engine prediction program. The work completed produced an optimum engine configuration which produced 2.392kg of static thrust with a total length of 1035mm. The second engine developed was not successful in achieving sustained combustion.

20

Section 2.6 Power plant Design The notable areas of interest are in the relationships which lead to the design of their statistical based engine, the testing procedure they utilized, the theory behind the development of the engine design software, and finally the problems and risks they encountered throughout the project.

2.6.3 Alternative Engine Designs Numerous different valveless engine designs have been developed, with the aim of improving the performance of the engines. In selecting a valveless engine for use on a UAV, thrust output, fuel consumption and aerodynamic performance must be considered. This section outlines some of the most successful pulsejet engines which have been developed, with the aim of identifying the most suitable engine for a flight weight aircraft.

Chinese Pulsejet Engine The Chinese Pulsejet engine was developed in the 1960s by CS manufacturing, a 2stroke motor designer from Shanghai. The engine is characterised by its expanding tail exhaust and cylindrical combustion chamber (Figure 11). CS manufactured two commercially available engines, which were designed to run on regular gasoline. In 1993 the designs for the engine became public, and it has since been developed by enthusiasts for use with propane fuel systems.

Figure 11-Chinese Valveless Pulsejet Engine (Beck 2008)

21

Chapter 2 Feasibility Study

The engine is streamline in design, with rearward facing exhaust and intake to ensure all thrust created acts in the same direction. No analytical research has been conducted into this specific design, however specific fuel consumptions of between 3kg/kg/hr and 6kg/kg/hr have been noted from enthusiasts. Thrust to weight ratios of between three and five have been achieved.

Lockwood Valveless Engine The Lockwood valveless engine has been the most successful valveless pulsejet developed in recorded history. The engine was investigated between during the 1960s as a form of propulsion for vertical takeoff and landing (VTOL) aircraft. The engine is a U-shape, with the exhaust bent around 180 degrees to direct both the intake and exhaust thrusts in the same direction. A table of the final engine performance claims can be seen in Table 4, however it should be noted that these values have never been achieved using the patented design, specific fuel consumptions closer to 5kg/kg/hr have been seen, with thrust results approximately 25% less than claimed. The aerodynamic performance of the engine is also poor, in comparison to the Chinese and FWE designs shown earlier.

Figure 12-Lockwood Prototype (Lockwood 1957)

22

Section 2.6 Power plant Design Table 4- Lockwood Performance Data (Lockwood 1957) Model HH 5.25-7 Valveless Engine Military max thrust (lbs) Maximum continuous (lbs) Minimum idle (lbs) Idle to mil. max time (secs) Fuel/thrust (lb/lb/hr) Dry weight (lbs)

300 280 30 0.1 0.85 30

Escopette The Escopette was developed by the French research agency SNECMA (Societe Nationale d'Etude et de Construction de Moteurs d'Aviation) in 1950. The engine was the first developed with a rearward facing intake, and with expanding sections in the exhaust.

Figure 13-Escopette Valveless Engine

The engines operating characteristics are different to a normal pulsejet, due to the unique exhaust design and separation between the curved intake and the main engine.

The split intake allows the engine to behave as if its length were variable – long during the exhaust phase of the cycle and short during the intake phase. During expansion, it treats the curved intake as a part of the effective length of the engine and uses it to turn the escaping gas around and increase thrust (Ogorelec 2004). During the intake cycle the curved section is not used. This reduces the effective length of the intake and lets the engine inhale more easily.

The tailpipe is a series of steps of increasing section. Each transition from a straight 23

Chapter 2 Feasibility Study section into a diffusing section represents a point from which the pressure waves travelling up and down the tube will reflect. Each of these waves reflects in an area of varying temperature, and therefore they all travel at different speeds. The interaction and timing of these waves are critical to the engines operation (Figure 13).

Figure 14 - Interaction of Escopette Pressure Waves (Belfast University 1983)

The unique design of the engine means that it inhales twice for each expansion cycle, with the aim of increasing the amount of cool air drawn into the exhaust section. This increases the mass of the air in the exhaust and thus allows energy from the combustion process to be converted more efficiently into thrust.

The original engine produced 108N of thrust, with a fuel consumption of 19.8kg/hr. The engine however was extremely long at over 2.6m.

2.6.4 Exhaust Pipe Development From the analysis of the pulsejet engines in Section 2.6.3, it can be seen that the performance of a pulsejet engine is reliant on the behavior of the dominant waves in the engines exhaust. Modifications of the engine exhaust characteristics can have a dramatic effect on the engine performance. Studies by Artt and Balair in 1983 found that altering the exhaust of a valved pulsejet engine could improve its performance by up to 25%. The following section investigates the various exhaust designs, 24

Section 2.6 Power plant Design characteristics and theories, in order to provide a knowledge base from which modifications to the existing engines can be made.

Straight Pipe Straight pipe exhausts are generally found on basic engines designed for first time builders. The most common engine design to use a straight pipe is the “Focused Wave Engine”, shown in Figure 15.

Figure 15 – A Focused Wave (FWE) Pulsejet engine (Beck, 2008)

The advantage of this type of exhaust is primarily ease of manufacture and cost reduction, as commercially available pipe can be used, without the hassle of forming conical sections. The section only operates on a single refraction wave returning from the end of the exhaust, significantly reducing the engines throttle range, and thrust output (Artt 1983).

Expansion Pipe This type is the most common exhaust found on designs that produce reasonable to high levels of thrust. Popular designs include the Lockwood-Hiller engine (Figure 16), as well as the “Chinese” engine.

25

Chapter 2 Feasibility Study

Figure 16 – A Lockwood-Hiller style Pulsejet engine, (Kontou 2007)

The most common justification of the expansion pipe is related to the drawing of cool air into the exhaust section after the combustion pulse. The larger entrance volume draws a larger volume of cool air into the exhaust. This dense air then acts as a ‘cold air piston’ which is accelerated by the combustion wave leading to more thrust.

The expansion may also be justified through energy and momentum laws. The thrust a jet engine creates is related to the momentum of the expanded gasses leaving the engine, and hence proportional to the product of mass and velocity. The energy of a particle however is proportional to the product of mass and the square of velocity. Hence, it can be deduced that for a given energy input, the thrust of the engine can be maximized by reducing the velocity of the gasses leaving the engine. By expanding the exhaust pipe, the flow velocity will reduce, resulting in a larger momentum of the gasses leaving the engine, and hence a thrust increase.

Advanced Pipes Some advanced pulsejet engines feature unique exhaust designs that help to improve the efficiency of the engine. An example of this is the SNECMA “Escopette” engine, 26

Section 2.6 Power plant Design which was explained in the previous section. Another unique exhaust design is found on the “FWE VIII - Lady Anne Boleyn” engine as shown in Figure 17. The exhaust on this engine has a diverging then converging section between the combustion chamber and the final expansion.

Figure 17 – A Focused Wave engine variation, the “FWE VIII - Lady Anne Boleyn”. (Cottrill 2006)

2.6.5 Liquid Fuelling There are a number of critical design aspects when it comes to using liquid fuel to power a valveless pulsejet engine. These include the choice between direct injection and carburetion for supplying fuel to the combustion zone, as well as the advantages and disadvantages of different fuels depending on the desired use of the engine. These aspects of liquid fuelling are discussed in more detail below.

Carburetion or Direct Injector An atomisation, or carburetion system, is potentially the simplest system available for liquid fuelling. The principals are similar to those for carburettor in an internal combustion engine. By Incorporating a venturi into the intake the pressure within the pipe at the injector point is reduced, causing fuel to be drawn from the fuel tank.

While this system is simple in design it is very sensitive to pressure head, and therefore fuel tank placement. This is compounded in flight by the aircraft performing manoeuvres. The other disadvantage to a carburetion system for pulsejet applications is that the engine cannot be throttled. In a normal carburetion system the engine is 27

Chapter 2 Feasibility Study throttled by increasing and decreasing the air flow rate, and thus the draw of the venturi, in a pulsejet, the intake flow cannot be varied, as it will affect the operating characteristics of the engine.

A direct injection fuel system has the added complication of a fuel pump, which is used to provide a constant pressurised fuel supply to the engine. A pump requires its own power supply and requires the operator to control the fuel flow rate into the engine. Due to the constant fuel supply from the pump, slightly more fuel is wasted during the combustion cycle of a direct injection system when the exhaust gasses are pushed out of the intake pipe. The system however is not affected by fuel tank placement or aircraft angle and has been shown to be able to be throttled (McCalley 2004).

Fuels There are a wide range of fuel properties that must be considered when determining the type of fuel to be used to power a valveless pulsejet engine, such as flash point, energy density, air-fuel ratio and latent heat of evaporation. Unlike many engines, a pulsejet engine can run on a wide range of liquid fuels, without the need for major redesign (Simpson, 2006). However, as valveless pulsejet engines are heavily dependent on the acoustic lengths of the engine to obtain resonance, temperature changes within the engine will directly impact the operation of the engine. Therefore, different fuels, with different combustion temperatures, will directly impact the engines operation. Successful operation of a valveless Schubert jet on kerosene and unleaded petrol was achieved by McCalley in 2006.

Injector Design The studies by McCalley in 2006 were directed towards the operation of a valveless pulsejet engine on heavy fuels, such as kerosene. The investigation suggested that injector design and fuel atomisation were vital in ensuring the engines operation. It was also noted that the engine must be brought up to operating temperature using propane first, as the compression of the pulsejet was not enough to cause combustion at room temperature. 28

Section 2.6 Power plant Design

The study developed numerous pinhole injector designs, varying from single hole injectors, to 6 hole swirl patterns. Results suggested that a high pressure injector was required, as low restriction injectors (multiple hole injectors) were not capable of producing atomisation. These injectors were also found to lead to boiling of the fuel within the fuel injector head, starving the engine of fuel. Successful operation of the jet on liquid fuel was achieved, however it was noted that variations in the injector position were vital in achieving self sustained combustion.

2.6.6 Thrust Augmentation Thrust augmenters or ejectors, are a simple addition which can be made to pulsejet engines. The augmenter is positioned behind the exhaust section of the pulsejet and is used to capture the starting vortices which form with each cycle of the engine (Paxton 2006). The augmenter uses these starting vortices to draw in and mix cold air with the hot exhaust gasses adding extra mass to the exhaust flow. Studies have shown that exhaust thrust augmentation can increase engine thrust by over 100%, with designs which are much simpler and more compact that conventional steady state ejectors (Kailasanathan 2007). However, as it can be seen in Figure 18, the addition of an optimised ejector to a pulsejet engine would increase the drag of the engine significantly in flight.

Figure 18- optimised thrust augmenter as used on a valved pulsejet

29

Chapter 2 Feasibility Study 2.7 Feasibility Study Summary The above data from the detailed feasibility study and bench-marking has been synthesised to produce the project definition, which outlines the aims and objectives of the project, as well as develop the direction for further work.

30

3 Conceptual Design The conceptual design stage aimed to develop a feasible and realistic concept aircraft and engine which met all the criteria outlined within the feasibility study and project definition. The aircraft design process entailed an initial statistical overview of existing aircraft and configuration options before sizing of the aircraft to meet the requirements outlined in the project definition. The engine development involved an initial testing of the engines developed by Coombs et al. (2007) and the design of modifications to the second prototype engine developed by the group. In order to manage the risks associated with this project a second ‘Chinese’ valveless engine was investigated and designed from statistical data and professional opinion.

3.1 Aircraft Conceptual Design Introduction The first step was to determine a configuration then using concept sketches and statistical considerations to evolve towards a concept based on the configuration.

3.2 Selecting Preliminary Aircraft Concept To develop a concept, a general configuration first had to be selected. The initial configuration of the aircraft was determined through use of a decision matrix method. A summary is presented here while the complete breakdown is presented in Appendix A and the review of pulsejet and target drone aircraft conducted in the feasibility study has been taken into consideration when determining the rankings

3.2.1 General Configuration The general configuration is the overall configuration of the aircraft, entailing options such as monoplanes, biplanes. In this area, the available layouts were rated in the areas of pulsejet suitability, complexity, weight, stability and drag. The highest rated option was a conventional monoplane due to high ratings in all areas, particularly pulsejet suitability. 31

Chapter 3 Conceptual Design 3.2.2 Fuselage Configuration The fuselage configuration entails the design of the fuselage, with several variations on a conventional tubular fuselage available. The fuselage options were rated in the areas of cost, complexity and weight, considering that no payload would be carried. After consideration, the highest rated fuselage was a conventional fuselage due to its strength in all areas.

3.2.3 Engine Configuration The engine configuration includes both the number of engines and the mounting method used as a pulsejet was specified in the project brief. These options were rated in the areas of cost, reliability, weight, heat and performance. A single engine was optimal for all areas and so was highest rated. However, in this area final design thrust required may be the determining factor if two engines are required.

3.2.4 Wing Configuration The wing configuration included shape, height and sweep areas with common options considered for all. In the area of a linearly tapered wing was preferable due to a balance of all factors. The best wing height was a low wing due primarily to minimal landing gear weight. A non-swept wing was determined as the preferred wing but aft sweep may be required for stability purposes depending on future calculations.

3.2.5 Empennage Configuration The empennage configuration options included a conventional tail, T-tail, H-tail and Vtail. These options were rated in the areas of suitability to a pulsejet, complexity, weight and size. The H-tail was the preferred option based on these factors, primarily due to the suitability of this design to a pulsejet.

32

Section 3.2 Selecting Preliminary Aircraft Concept 3.2.6 Landing Gear Configuration Landing gear configuration encompasses both the ability to be retracted and wheel arrangement. In the area of the ability to be retracted, fixed, retractable and removable landing gears were considered as well as a design with no landing gear requiring a launcher. These options were rated based on drag, complexity, weight and reliability with a fixed landing gear the best option for this aircraft. Wheel arrangements considered were tail-dragger, bicycle and tricycle designs. The areas of consideration for these designs were drag, stability and weight with a tricycle landing gear being the selected design.

3.2.7 Basic Wing Parameters For later conceptual calculations, some basic wing parameters were required to be determined.

Aspect Ratio Aspect ratio is usually determined by a trade study between structural weight and aerodynamic efficiency, however the scope of this project does not allow for such a complex optimisation therefore the aspect ratio of this aircraft was be selected based on statistics. A medium aspect ratio of 6.5 will be selected here, because it is statistically consistent with other similar aircraft and provides a balance between drag and weight.

Taper Ratio The taper ratio of an aircraft can increase its aerodynamic performance by more closely approximating an ideal elliptical wing, reducing induced drag. For the determination of taper ratio, Raymer (1992) determines that a taper ratio of 0.45 is ideal for a non-swept wing and as the wing used for this aircraft is slightly swept, a slightly smaller taper ratio such as 0.4 should be used.

33

Chapter 3 Conceptual Design Wing Sweep The wing sweep of the aircraft is normally zero for an aircraft which does not approach a sonic Mach number. However, sweep is also used to adjust the aerodynamic centre of the aircraft and allow aircraft. As this aircraft will have a heavy engine located at the rear, stability will be a concern and so the wing will be swept so that the trailing edge is straight with the taper ratio and the front edge swept.

Wing Twist Wing twist is used to prevent the onset of stall at the wingtips and enable the pilot to retain control of the aircraft during stall. The non-use of twist would also simplify manufacture, although this is not the overriding requirement.

Wing Incidence Angle The wing incidence angle is set to balance between trim drag in flight and drag at takeoff. However, generally the wing incidence angle is set to that which will minimise cruise drag. This will be decided once an aerofoil has been selected for this aircraft.

Dihedral Angle A dihedral angle can be applied to the wings to provide additional roll stability. However, the use of an H-tail to provide two surfaces and rearward sweep on the wings reduces the requirement for dihedral angle. As a result, a dihedral angle is not required on this aircraft.

3.3 Developing concept for selected configuration Basic concept based on the selected configuration were developed using sketches and statistics.

34

Section 3.3 Developing concept for selected configuration 3.3.1 Concept Sketches Some preliminary concept sketches were prepared, shown in Figure 19.

These

outlined various options for positioning the pulsejet engine on the aircraft. The top concept has the pulsejet mounted underneath and an aerofoil shaped fuselage. While the third concept is a sketch whose main feature is an open fuselage allowing suitable cooling airflow past the pulsejet.

Figure 19-Configuration Concept Sketches

3.3.2 Statistical Calculations An extensive collection of all the available pulsejet and jet UAV statistical data was compiled from which trends were noted in some important statistics.

Statistical Estimates of Thrust and Wing Loadings Thrust loading is the thrust to weight ratio of the aircraft. The majority of thrust loadings of the aircraft surveyed were between 0.3 and 0.6, providing a reasonable minimum thrust for flight. All aircraft before this loading used an alternate launch 35

Chapter 3 Conceptual Design method such as release from aircraft or rocket assisted. The median thrust loading of jet UAVs is approximately 0.4 and this is approximately the required thrust loading for the project vehicle. A graphical representation of statistical thrust loadings is below in Figure 20.

Thrust Loading of UAVs 1.4

1.2

1

T/W

0.8

0.6

0.4

0.2

0

Figure 20: Statistical Thrust Loadings of Jet UAVs

The range of wing loadings of UAVs was also analyzed, showing correlation with weight of the aircraft. Below in Figure 21 is a graph of the wing loading versus the weight of statistical UAVs, showing this relationship. This is as reasonably expected as a larger aircraft would allow a longer runway length for takeoff and landing, hence allowing a greater wing loading

36

Section 3.3 Developing concept for selected configuration

Wing Loading - Weight 700

600

Wing Loading

500 W/S = 83.872Ln(W) - 182.62 400

300

200

100

0 0

100

200

300

400

500

600

700

800

Weight (kg)

Figure 21: Wing Loading Versus Weight of Jet UAVs

Statistical Estimates of Length and Span From statistics of weight and length, for the estimated size of the aircraft of approximately 8 kg, it was possible to estimate the main spatial parameters length and span from trend-lines found in plots of length and span versus length. Via this method, the length was statistically estimated at approximately 1.4 m while span was statistically estimated at approximately 1.6 m.

Statistical Concept Based on the statistical review, a statistical concept was produced this is shown in Figure 22. It has a span of 1.6m, length of 1.4 m, and conventional layout with H-tail.

37

Chapter 3 Conceptual Design

Figure 22: Statistical Concept

3.4 Designing technical parameters for concept Following the development of a basic concept, the technical parameters of the concept to meet the performance requirements in the project goals were determined. The weight is being determined, and the size of the aircraft to meet climb, cruise and requirements was determined.

3.4.1 Weight Estimation The weight estimations for the aircraft were completed based on statistical information obtained from Munson (2001) for similar sized piston powered UAVs. Piston powered UAVs were used as a basis due to expected size similarity to the pulsejet UAV and ease of data acquisition. The additional weight of a pulsejet and additional fuel was taken into account as a payload for a piston powered aircraft. Weight estimation calculations aimed to determine: •

We – empty weight of the aircraft

•

Wf – weight of the fuel the aircraft can carry

•

Wo –the takeoff weight of the aircraft.

38

Section 3.4 Designing technical parameters for concept The calculations are based on equating Equation 3 and Equation 4 below to compare values required by the mission and statistical values from past UAVs. Equation 3:

WO = WWOF WO + WWOE WO Equation 4:

WE C = AWO K S WO The values for A, c and Ks were determined from a statistical review of piston powered UAVs. An iterative procedure was then utilised in the Microsoft Excel spreadsheet program to determine Wo. From the converged procedure weight values were obtained. The empty weight of the aircraft was 6.1kg, fuel weight was 1.9kg and takeoff weight was 8kg. A full outline of the weight calculation estimation procedure can be seen in Appendix B.

3.4.2 Matching Diagram A matching diagram is a plot of take-off thrust loading versus take-off wing loading. It is used to find two or three possible ‘limit case’ combinations of wing loading and thrust loading to evaluate and determine an optimal aircraft. For this aircraft the matching diagram contains a stall requirement, a cruise requirement, a single climb requirement and a takeoff length requirement.

Details of the matching diagram

calculation are contained in Appendix C – Matching Diagram

Summary of requirements and assumptions The requirements used to produce the matching diagram are summarized in Table 5.

39

Chapter 3 Conceptual Design

Table 5 : Requirements and Input Data of Matching Diagram

Parameter

Requirement

Stall Speed

60km/hr

Cruise Speed

260km/hr

Climb rate

8.33% at 78km/hr

Takeoff distance

100m

Input Data Takeoff Weight

8kg

Fuel Weight

2kg

Specific Fuel Consumption

7.2kg/kg/hr

Clmax

1.2

Aspect Ratio

6.5

Matching Diagram Plot All of the sizing requirements that were found were plotted onto the matching diagram seen in Figure 23.

Figure 23: Matching Diagram

40

Section 3.4 Designing technical parameters for concept The matching diagram combining all of the requirements (climb, cruise, take-off and stall) is plotted above. Only one intersection point (designated A) fits both cruise and stall requirements. However, if the stall speed was extended another point (designated B) could be considered at the intersection of the cruise and takeoff requirements. The data for the intersection points can be seen in Table 6. Table 6 – Characteristics of Suitable Aircraft

Point

T/W

(W/S)TO [kg/m2]

A

0.375

20.81

B

0.323

29.31

Discussion of Suitable Points: A & B Two possible aircraft can be found from the above diagram. Table 7 below summarises the performance of Aircraft A and Aircraft B. Table 7: Characteristics of Possible Aircraft

T/W Vs Vcr sto Sref T W/S (kg/m2) (kg/kg) (km/hr) (km/hr) (m) (m2) (kg) Aircraft A 20.8 0.38 60 260 57.8 0.38 3.04 Aircraft B 29.3 0.32 71.2 260 100 0.27 2.56 From the above data it can be seen that Aircraft A has superior performance in both stall speed and takeoff distance. Aircraft B would have superior efficiency in cruise due to a lower range of operational speeds. The result of this improved efficiency is a lower thrust loading requirement. Aircraft B also requires a smaller wing area which would reduce the susceptibility of the aircraft to turbulence. After a consideration of the above characteristics, Aircraft A has been selected due to the lower risk of failure in the design. The lower stall speed is significant to reduce the chance of failure on landing. A shorter runway length will also allow more freedom in possible launch locations. For a proof of concept project such as this a safe design is 41

Chapter 3 Conceptual Design more critical than optimal efficiency as efficiency can easily improved in future designs. A sensitivity analysis was performed for Aircraft A, with the results available in Appendix D – Sensitivity Analysis

3.4.3 Aerofoil Selection To enable the tail to be sized correctly, aerofoil characteristics for both wing and tail were required. Different aerofoils were required for the wings and tail due to varying requirements.

An aerofoil for the wing was required to meet various parameters to be suitable for use on the aircraft and provide performance as expected. The aerofoil was also required to perform adequately across the range of Reynolds numbers which would be encountered during operation. The selected wing aerofoil was the NACA 4412.

The tail aerofoils were selected to provide suitable symmetric performance in the variable conditions at which they will be operating. The NACA 0012 aerofoil was selected for the horizontal and vertical tails via a similar method to the wing aerofoil.

Wing Aerofoil Selection An aerofoil for the wing was required to meet various parameters to be suitable for use on the aircraft and provide performance as expected in the initial sizing. The aerofoil needed a three dimensional lift coefficient equal to or greater than 1.2, minimal drag across the expected range of operation (from a Cl of 1.2 to 0.1 at maximum speed level cruise).

Initially, various aerofoils with characteristics indicative of their families were tested in ‘Javafoil’ analysis to determine which were worthy of closer examination. The Clark, SD, NACA 4, 5 and 6 digit series aerofoils were analysed with the results of this testing presented in Table 8. This initial analysis also operated as a test of the accuracy provided by ‘Javafoil’ by comparison with two dimensional wind tunnel test data from Abbott, von Doenhoff and Stivers’ 1945 report. 42

Section 3.4 Designing technical parameters for concept Table 8- Initial Aerofoil Analysis Profile

Clmax

α(Clmax)

Cm(Clmax)

(L/D)max

CDo

α(CDo)

NACA 4412

1.22

12

-0.12

23.6

0.011

-4

Clark Y

1.12

11

-0.10

22.9

0.012

-2

SD7034

1.12

12

-0.08

23.8

0.011

-3

NACA 23015

1.13

15

-0.04

20.4

0.010

-1

NACA 63-615

1.21

11

-0.15

22.0

0.011

-4

Experimental

Data

NACA 4412

1.28

11

-0.10

N/A

0.010

-4

NACA 23012

1.24

10

-0.02

N/A

0.010

-2

NACA 63-615

1.38

10

-0.12

N/A

0.011

-4

The three dimensional approximations used by ‘Javafoil’ produced data that approximates that determined from NACA wind tunnel testing and so were suitable for use. The test reports by Abbott, von Doenhoff and Stivers (1945) give lift coefficients of between 5% and 14% more than those from ‘Javafoil’ and drag coefficients within 8% as shown in Table 8.

The initial testing shows that the NACA 4 digit series is the best for this application. As a result, the family chosen for closer analysis is the NACA 4 digit series which has superior aerodynamic efficiency to other families tested.

Various NACA four digit aerofoils were analysed in ‘Javafoil’ three dimensional analyses to determine the optimal camber and thickness. These tests were performed at the Reynolds numbers of 2.9 x 105 and 1.3 x 106 as before. The worst value for each variable parameter between the tested Reynolds numbers is recorded in Table 9. Table 9: NACA 4 Digit Aerofoil Analysis Profile NACA 2412 NACA 4409 NACA 4412 NACA 4415 NACA 4418 NACA 6409

Clmax 1.07 1.04 1.22 1.37 1.5 1.24

α(Clmax) 12 9 12 15 15 10

Cm(Clmax) -0.06 -0.11 -0.12 -0.13 -0.14 -0.16

(L/D)max 23.2 25.1 23.6 21.9 20.6 22.6

CDo 0.009 0.009 0.011 0.012 0.013 0.014

α(CDo) -2 -4 -4 -4 -4 -3

43

Chapter 3 Conceptual Design From this data it can be seen that the NACA 4412 is the best available aerofoil. A smaller thickness such as in the NACA 4409 aerofoil will reduce the Clmax below 1.2 while a higher thickness will reduce the available (L/D)max value. Reducing the camber reduces the maximum lift coefficient and the maximum L/D so a less cambered wing is inferior in both key areas despite decreasing the moment coefficient generated. An excessively cambered wing would result in poor cruise performance for an aircraft with such a speed range as this aircraft and exceptional wing moment. As a result, the NACA 4412 aerofoil profile was chosen as the best option to meet the requirements outlined above.

Empennage Aerofoil Selection The tail aerofoils were selected to provide suitable symmetric performance in the variable conditions at which they will be operating. The requirements for the horizontal and vertical tails are similar, with a reasonable aerodynamic performance and slow stalling key requirements. For these aerofoils, as symmetry was required and performance requirements were not particularly rigorous only the NACA 4 digit aerofoil series was considered as these aerofoils are in common use on many aircraft tails due to their gradual stall characteristics. The results of ‘Javafoil’ testing of three reasonable thickness aerofoils are presented in Table 10. This shows the worst value from the two tested Reynolds numbers and the difference between the upper and lower values. Table 10- Suitable Tail Aerofoils Profile

Clmax

α(Clmax)

Cm(Clmax)

(L/D)max

CDo

α(CDo)

NACA 0009

0.54

9

-0.01

17.1

0.007

0

NACA 0012

0.67

12

-0.01

17.5

0.009

0

NACA 0015

0.83

15

-0.02

16.2

0.010

0

This initial testing shows that the NACA 0012 is the best for this application. This aerofoil provides the best aerodynamic efficiency and an angle of attack versus coefficient of lift graph shows that the stall of this aerofoil is suitably gradual for use. 44

Section 3.5 Finalisation of Preliminary Aircraft Concept The angle of attack of stall is around that of the wing and the delayed stall will enable this tail to stall after the wing. The NACA 0012 has better aerodynamic efficiency and gradual stall than other options and so has been selected. Although this testing is for the aspect ratio of the horizontal tail, it is expected that the aerofoil properties would not be too dissimilar for the vertical aerofoil and so the vertical tail will also be fitted with a NACA 0012 aerofoil.

3.5 Finalisation of Preliminary Aircraft Concept With the configuration and basic parameters determined, the aircraft concept was able to be finalised.

3.5.1 Variation of Pulsejet Position in Concept Development The main variable in sketches was the engine position, and various engine positions were tried in a preliminary CAD model to work out a final concept considering longitudinal stability, thrust line, practicality and overall plane length. This analysis can be characterised into three phases with differing pulsejet positions, attempting to solve the inherent challenges of pulsejet engine airframe design. The first phase utilised an engine mounted external to the airframe, directly above the fuselage. The second phase used a twin boom layout, enabling the engine to be mounted inline with the airframe. The engine was mounted inside a fuselage cavity in the final phase of development.

The initial designs focussed on a usable airframe, with pulsejet mounting considered a less important aspect of design. As such, the pulsejet was initially mounted above the fuselage, enabling sufficient air flow for cooling but with additional drag and an uncertain mounting method. Eventually, when these designs were analysed the drag provided by the airframe was exceptionally high due to the pulsejet mounting location.. An example of an early design with engine mounted above the airframe is shown below in Figure 24.

45

Chapter 3 Conceptual Design

Figure 24: Example of Early Design

The inline engine phase was characterised by mounting the pulsejet inline with the fuselage, reducing profile drag and moment produced by the engine thrust. Mounting the engine inline with the fuselage allowed more efficient flight but the exceptional length of the pulsejet engine lead to stability and size issues. An example of an aircraft with this engine position is shown below in Figure 25.

Figure 25: Second Phase Design Example

The final phase of design combined the previous two by mounting the engine out of the oncoming airflow in a fuselage cut out to reduce drag without compromising stability. This phase developed from designs with engine mounted above a cut down fuselage section in the second phase, by moving the engine forward so that it would be almost over the centre of gravity of the aircraft.

However, this engine position was not without drawbacks. Heat generated by the pulsejet was a critical issue and without cooling air passing over most of the pulsejet surface the mountings and connections required additional heat resistance through the use of coatings. The direct mounting of the engine to the fuselage meant that extra vibration isolation was required in the design of the pulsejet mountings. Considering all areas however, designs such as shown in Figure 26 were superior to those that came before them and have been continued into detailed design. 46

Section 3.5 Finalisation of Preliminary Aircraft Concept

Figure 26: Final Engine Position

3.5.2 Empennage Design To determine the required horizontal and vertical tails, longitudinal and lateral stability of the aircraft must be considered. This will enable the exact size of these tails to be determined for a reasonable stability margin. Stability will be considered in more detail once other details of the aircraft have been determined.

Longitudinal Stability Longitudinal stability is based on the margin between aerodynamic centre and centre of gravity. When the centre of gravity is ahead of the aerodynamic centre, an aircraft which pitches upwards will have a natural moment downwards as the lift increases due to an increased angle of attack, returning it to the original attitude. An unstable aircraft would continue to pitch upwards until stall. However, in all aircraft and particularly this aircraft with a large fuel fraction the centre of gravity will move depending on the load configuration. As a result, the movement of the centre of gravity must first be determined to find the critical most aft centre of gravity. The aerodynamic centre of the aircraft can then be computed for varying horizontal tail size to find the tail size for a defined stability margin.

47

Chapter 3 Conceptual Design

Centre of Gravity Envelope An envelope of potential centre of gravity positions was determined using a spreadsheet computation. The centre of gravity is determined through Equation 5 by taking into account the centre of gravity of all (n) components of the aircraft and their relative masses. Equation 5: n

∑xM i =1 n

i

∑M i =1

i

= xcg

i

With no payload or crew for this UAV, four loading options were considered, based on variable depletion of the fuel bags used. As the weight of various structural components was still uncertain, initially statistical assumptions were made with the use of equations. This method used empirical equations based on many aircraft properties to estimate the weight of aircraft components. Once the components had been manufactured, the measured weights and centres of gravity of these were used. A component weight breakdown is presented in Appendix G – Component Weight Breakdown A together with applied lever arms from the front of the aircraft used to determine the position of the centre of gravity. When the above variable configuration options are considered the centre of gravity excursion diagram presented in Figure 27 can be determined. This diagram shows how the location of the centre of gravity varies with the loading of the aircraft as well as the locations of the wing and whole aircraft stick fixed aerodynamic centres. The aerodynamic centre of the aircraft (also known as the neutral point) was determined from Equation 6 (Raymer 1992) and shows the stability margin which is determined by later tail sizing.

48

Section 3.5 Finalisation of Preliminary Aircraft Concept Equation 6:

F pα ∂α p Sh ∂α h C l αh X ach + Xp Sw ∂α qS w ∂α S ∂α h F pa C lα + η h h C lαh + Sw ∂α qS w

C lα X acw − C mαfus + η h X np =

Figure 27: Centre of Gravity Excursion Diagram

Horizontal Tail Sizing The horizontal tail size required was determined through use of an X-Plot which is a graph of neutral point and centre of gravity versus horizontal tail size. A required minimum stability margin of approximately ten percent of the mean aerodynamic chord is required for an aircraft with inherent static stability (Roskam 1985b). As at this stage the location of the stick-free aerodynamic centre was unknown, a slightly greater margin was used to allow for the stick-free point being further forward than the stickfixed position calculated. The most aft centre of gravity (with no tail) was determined from the above excursion diagram. The neutral point of the aircraft was determined by use of Equation 6 (Raymer 1992) for each tail size tested. The X-Plot determined could then be used to determine the required horizontal tail size by applying the required stability margin as shown in Figure 27. During iterations of this procedure, it was found 49

Chapter 3 Conceptual Design that the downwash from the wings was having a detrimental effect on the performance of the tail. This was then corrected by utilising the modified T-tail layout outlined earlier. Sweep was incorporated on both the vertical and horizontal tails to maximise the moment arm between the wing and the tail by moving the aerodynamic centre of the tail rearward. This allowed a significantly smaller horizontal tail area and hence a more efficient aircraft. The figure shown below is for the final centre of gravity position, showing a greater stability margin than required as the theoretical centre of gravity was significantly further backwards than the This method gave a horizontal tail size of 0.145m2 for the aircraft. Longitudinal X-plot 120.0%

100.0%

Location (% MAC)

80.0%

Centre of Gravity 60.0%

Static Margin Neutral Point

40.0%

20.0%

0.0% 0

0.05

0.1

0.15

0.2

0.25

0.3

S(m2)

Figure 28: Longitudinal X-Plot

Lateral Stability The required vertical tail size was determined through use of a directional X-Plot of vertical tail size versus lateral stability margin. To enable this sizing to be completed, the coefficient of directional stability (Cnβ) for the aircraft was required. This was influenced by the required stability margin and three area specific coefficients Cnβf, Cnβi and Cnβp which were considered separately. Cnβp is only relevant to aircraft with propeller engines and so was zero for this design. For a low winged aircraft, Cnβi is 0.024 (Torenbeek, 1982) while Cnβf is dependent on fuselage geometry and can be 50

Section 3.5 Finalisation of Preliminary Aircraft Concept calculated from Equation 7 (Torenbeek 1982) (correcting for specific geometry using Equation 8). Equation 7

C nβf

S fs l f  h f 1  = −κ β Sb  h f 2

 b f 2   b 1  f

   

1

3

Equation 8

κ β = 0 .3

l cg lf

+ 0.75

h f max lf

− 0.105

Equation 9

S v l v C nβ − (C nβf + C nβp + C nβi ) = Sb C yvα (Vv / V ) 2 Equation 9 (Torenbeek 1982) can then be used to determine the stability margin for a given vertical tail size. This can be repeated for varying vertical tail sizes to determine the required size for a certain stability margin. The stability margin required can be estimated via consultation with various sources, Roskam (1985b) recommends 0.001 deg-1 while Torenbeek (1982) recommends a range from 0.04 to 0.1 rad-1 so a required margin of 0.06 rad-1 has been chosen to fit within these values. The lateral X-plot can then be drawn as in Figure 29 and used to determine required vertical tail area. From this a vertical tail size of 0.05 m2 was determined. However, as the vertical tail size also contributes to the effectiveness of the horizontal tail by raising it upwards from the downwash of the main wing the vertical tail size was somewhat larger than required to achieve the stability margin set. The uncertain effects of locating the pulsejet between the vertical tails altering the flow profile also required a greater margin of error. As a result, the combined area of both vertical tails was set at 0.11m2.

51

Chapter 3 Conceptual Design Lateral X-Plot 0.35

0.3

0.25

Stability margin (rad-1)

0.2

0.15

0.1

0.05

0 0

0.02

0.04

0.06

0.08

0.1

0.12

-0.05

-0.1 2

Svertical tail(m )

Figure 29: Lateral Stability X-Plot

3.6 Finalization of Preliminary Aircraft Concept The engine data development (such as selection of Chinese engine and engine length) was now able to be incorporated into the aircraft concept.

3.6.1 Preliminary Conceptual Fuselage Design With the main conceptual parameters determined such as engine length and based on some of the sketches developed earlier, a conceptual design for the fuselage was able to be produced.

The overall shape of the fuselage is roughly cylindrical, with a cut out section for the pulsejet and an aerofoil shaped nose. This circular shape enables a simpler manufacturing process for both the fuselage and bulkheads and easier, more accurate structural analysis. This circular shape also allows the use of a rotated aerofoil cone as the nose of the fuselage, providing minimal drag and hence better aircraft performance.

52

Section 3.7 Practical Modifications to Final Concept The conceptual fuselage diameter selection is a balance between providing minimal centre of gravity movement and minimizing drag. A smaller diameter fuselage would reduce the cross sectional area and as a result, the drag of the fuselage. The Chinese pulsejet used required a cavity of 80mm to be left in the top of the fuselage to provide space and clearance for the engine’s exhaust (diameter 74mm) without allowing too much of the pulsejet to protrude from the fuselage. This protrusion would significantly increase the drag of the aircraft as the front of the engine is simply a cylinder with no streamlining. The final design can be seen in Figure 30.

Figure 30 -Conceptual Fuselage Design

3.7 Practical Modifications to Final Concept (a)

(b)

Figure 31: Modifications to final aircraft concept

With all the conceptual parameters of the aircraft the first concept showing all the elements of the design is shown in Figure 31 (a). However some modifications had to be made to this concept and this is shown in Figure 31 (b). A fillet has been added around the root of the wing to reduce interference drag and to reduce the number of 53

Chapter 3 Conceptual Design sharp edges needed to be cut using CNC milling. The linearly tapered engine cut-away has been changed to a flat engine cutaway as the tapered cutaway would be difficult and expensive to manufacture due to complex moulds being required. The tail has been changed to a modified T-tail design to, this allows for elevated positioning of the horizontal tail to limit impacts of main wing downwash on the tail. Sweep has been incorporated to give the tail later stall characteristics than the wing, while the swept vertical allows for further rearward position of the horizontal tail to improve longitudinal stability. Finally, the diameter of the fuselage has been increased slightly from 150 mm to 160 mm.

3.8 Engine Design Initial research performed in the feasibility study section suggested the use of an expanding tail section would improve engine performance. The following section investigates the design aspects related to the expansion design. The similarity between pulsejet expansions and the exhaust systems used on two-stroke internal combustion engines was noted and used as a partial basis of the expansion theory. This theory was combined with steady state diffuser theory, fluid dynamics software and statistical research.

The effect of intake and exhaust flares was investigated based on the research conducted by Coombes et al in 2007, flares for both the intake and exhaust were developed for subsequent testing based on a literature review and 1-D fluid dynamics software.

3.8.1 Exhaust Design – Two Stroke Exhaust Similarities The evolution of pulsejet engine exhausts appears to carry some similarities to the evolution of 2-stroke engine exhaust design. The use of this observation as a guide into the workings of the advanced exhaust designs may provide a means of understanding the complex behaviour that is present in the systems, as well as provide the possibility to produce an exhaust that dramatically improves the performance of the current engine. 54

Section 3.8 Engine Design 2-Stroke exhaust design evolution: Two stroke engines typically have one of three exhaust designs (Figure 32). Early engines used straight pipes, which evolved into ‘megaphone’ expansions, and finally the ‘expansion chamber’ style pipes that are seen currently. (Jennings, 1987)

Figure 32- Advancements in Two Stroke Exhaust Design

The Initial straight pipe exhaust systems on 2-stroke engines could be tuned by simply changing the length of the pipe based on the desired speed range. (Jennings, 1987) Thi allowed for significant performance gains.

It was then discovered that a “Megaphone” produced a stronger low pressure region behind the pressure wave. This helps the exhaust extract all the burnt gasses from the combustion chamber, allowing more fresh air and fuel into the cylinder. It proved possible to build an expansion that worked so well that it would draw unburnt fuel into the start of the exhaust pipe. (Jennings, 1987)

The addition of a convergent section to the end of the expansion results in a pressure wave being sent back towards the exhaust port. It is found that if this wave was timed correctly, raw mixture drawn into the start of the exhaust could be pushed back into the combustion chamber just prior to the port closing. This effectively supercharges 55

Chapter 3 Conceptual Design the engine, forcing more air into the combustion chamber than it could normally hold. The extent of this effect is very significant, often producing volumetric efficiencies of around 140% on modern engines. (Jennings, 1987)

Effects of chamber characteristic lengths: The diffuser angle is responsible for the intensity and duration of the returning low pressure wave. A higher angle results in a stronger wave with a shorter duration. This relationship is seen in Figure 33.

Figure 33 – The effect of expansion angle on wave behaviour. (Jennings, 1987)

On a 2-stroke engine, it is claimed that an 8 degree included angle provides the best energy recovery. The use of multi-cone sections is common, as it reduces the angle required between cone steps, and hence allows greater angles before flow separation occurs. This optimal angle is a compromise between peak power and the power range, and hence the optimal angle for a pulsejet, which has a constant operating frequency, will possibly be larger than commonly found on 2-stroke engines. (Jennings, 1987)

As with the diffuser angle, the convergent cone angle determines the intensity and duration of the returning wave. This angle is typically about twice the diffuser angle. This angle tends to effect the tuning location of an engine dramatically, with subtle 56

Section 3.8 Engine Design changes in location producing noticeable translation in the tuned frequency range. The small exit pipe is the main dimension that determines the pressure inside the exhaust chamber. A higher pressure in the chamber results in stronger pressure waves, and as such can be used to improve the performance of the engine. (Jennings, 1987)

A study using 2-stroke exhaust theory with valved pulsejet engines was conducted in 1983 by Arrt and Blair through Belfast University. The experiment showed that the megaphone design had the effect of improving the specific fuel consumption of the engine by 44%, whereas the expansion chamber style engines increased the engine fuel consumption and reduced thrust.

3.8.2 Steady State Diffuser Design Background research suggested that the use of an expanding exhaust section results in an increase thrust output from the engine, for a number of reasons. Firstly, the expanded section results in a more efficient use of available energy within the engine. A valveless pulsejet engine does not lack energy, as can be deduced by the high temperature exhaust gasses that exit the engine (Coombes et al, 2007). Instead, the major challenge for a valveless pulsejet engine is being able to convert the available heat energy into momentum and thrust (Equation 10 and Equation 11) Equation 10

KE =

1 2 mv 2

Equation 11

p = mv

The precise sizing of the design of an expanding section is, as with most pulsejet related design, not documented. However, approximate sizing’s can be estimated from a range of different sources including statistical reviews, information provided by enthusiasts and some fluid dynamics theory. 57

Chapter 3 Conceptual Design

The design of the expanding tail pipe is effectively a conical diffuser during the combustion cycle and a contractor during the intake cycle. It is therefore expected to behave as such, which will make the expansion angle critical. Figure 34 below shows results for the behaviour of a typical diffuser. Due to the complex nature of flow through a diffuser, the figure can only be use as a guideline, with Individual flows varying with characteristics of the design such as area ratio and Reynolds number (Munson, Young, Okiishi, 2006).

Figure 34: Loss Coefficient for a Conical Diffuser (Munson, Young, Okiishi, 2006)

Figure 34 shows that for small expansion angles, the diffuser is excessively long, with large losses a result of wall shear stress. For moderate to large angles, losses are a result of flow separation, which effectively reduces the diameter of the expanded section. Commonly, an optimum diffusion angle is about 8° with angles similar to this (+/- 1°) used widely in pulsejets, including the Messerschmidt pulse-ram jet, the Escopette and the Laird Chinese. Half angles of 4° or 5° are very common, with optimum performance angles for individual engines generally requiring extensive testing.

Figure 34 shows that it is very difficult to efficiently decelerate a fluid. Acceleration of a fluid in a contractor on the other hand is more efficient and less complicated, with loss 58

Section 3.8 Engine Design coefficients only ranging from 0.02 at 30° to 0.07 at 60° (Munson, Young, Okiishi, 2006). Therefore, as the expanding section acts as both a diffuser and a contractor, it was decided that the section should be designed to minimise losses for expansion, which is during the combustion stage of the engines cycle.

3.8.2.1 Exhaust Analysis In order to produce the optimal expansion section, a 1 dimensional analysis of the engine with various characteristic lengths was performed using the ‘UFLOW1D’ software package. Uflow is a one dimensional compressible flow program for analysing flow in pipes.

The basic geometry that was used for the investigation was based on the engine developed by Coombes et al in 2007. These characteristics include the combustion chamber, as well as the diameter of the middle sections. The lengths of the remaining sections, as well as the expansion angle, were then investigated using the software.

The performance of particular engine geometries was gauged by the ability of the engine to produce a large pressure swing in the combustion chamber. This criterion was chosen as it was considered to be a key factor in determining the power the engine can produce. The basic engine was therefore modelled using varying expansion angle sections, in order to find the optimal divergence angle. The model used in the process is shown in Figure 35. The results obtained from the investigation are shown in Figure 36.

Figure 35 – The UFLOW1D model used to investigate expansion angles.

59

Chapter 3 Conceptual Design

Figure 36 – Combustion chamber pressure extremes for different expansion angles.

The ‘UFLOW1D’ analysis shows that the combustion chamber pressure swings are significantly increased by the presence of the divergent section. Results suggest that while the best cone angle is around 10 degrees, most of the expansion effect is seen by around 6 degrees.

A statistical review of some similar engines is presented in Figure 37, which suggests that the expansion angle should be around 8 degrees. The reasoning for the use of 8 degrees was not known, but was assumed to be as a result of the onset of flow separation. The engine with the greatest expansion angle in the statistical data was, however, the most efficient engine that was considered, which suggests that the use of a 10 degree expansion is reasonable, and perhaps optimal in some circumstances.

60

Section 3.8 Engine Design

Figure 37 - Statistical data showing exhaust expansion angles from similar engine designs

The design expansion angle was 10 degrees, chosen mainly due to the length constraints on the exhaust pipe discussed in the expansion location section.

3.8.2.2 Expansion Diameter From statistical research, it was found that the expansion typically expands to approximatly the diameter of the combusition chamber, as shown in Figure 38. It was decided to comply with the trend, as this would remove a variable that would be diffucult to alter. Hence, the expansion diameter was chosen to be the same as the combustion chamber.

Figure 38 - statistical data showing a trend between combustion chamber diameter and expansion diameter.

61

Chapter 3 Conceptual Design 3.8.2.3 Expansion location To investigate the effect of moving the expansion closer towards the combustion chamber, the ‘UFLOW1D’ model was altered. It was found that the location of the expantion along the exhaust length did not make an appreciable change to the pressure swing in the combustion chamber. However, this finding does not compliment the general trend of engines seen. A common trend seen in engines is that the the exhaust sections are approximatly equal lengths, resulting in the expansion being located in the middle of the exhaust section.

In order to achive the required expansion angle and diameter, it was required that the expansion section be 285mm long. Using this length, and assuming equal length sections, the complete exhaust would be 855mm long. This was seen as a potential problem, as the optimal length of the exhaust with a straight pipe was found to be approximately 600mm, significantly shorter than the new design. The length of the pipe is also a dominant reason for the use of a 10 degree expansion angle, as it reduced the length of the expansion section. It was decided that, in order to shorten the length of the pipe, the end section should be shortened.

As a reuslt of this, the expansion was designed with a 150mm extension at its end, rather than the 285mm extension that was originally intended.

3.8.2.4 Adjustability: In order allow for investigation into the behaviour of the engine, it was required that the exhaust section be made adjustable, in a similar fashion to the 2007 development. In order to aid manufacture, and to make testing of the full range of lengths possible, it was decided that only one characteristic length would be adjustable. The most suitable length to adjust was the centre section, as it will modify the length of the engine easily, and sections of centre pipe were already available. The shortest exhaust length sections from straight pipe testing can be used for this, however some shorter lengths still required to be manufactured. The final design is shown in Figure 39. 62

Section 3.9 FWE Bellmouth Development

Figure 39 - The final expansion design

3.9 FWE Bellmouth Development Based on the research conducted by Coombes et al, it was noted that the omission of flares from the engine design could possibly effect the ability of the engine to achieve a sustained thrust. To determine the effect of the flares on engine performance, suitable intake and exhaust flares were developed for a FWE engine.

It was proposed that the engine did not sustain due to the lack of an intake bellmouth, resulting in a reduced ability for the engine to draw air into the combustion chamber (Tao 2007). Based on the bellmouth research performed in 2007, it was also proposed that a flare on the exhaust of the engine would strengthen the intensity of the returning waves, and hence improve the chances of reaching sustained operation. This hypothesis can be confirmed through research conducted by Tao in 2007 into the effect of starting vortices on unsteady combustion devices.

3.9.1 Starting Vortices Starting vortices have been shown to be a significant contributor into the operation of pulsejet engines, due to their unsteady operational characteristics. Studies of vortex ring formation date back to 1900’s and this phenomenon has been extensively investigated theoretically, numerically, and experimentally. The link between to pulsejet engines was made by Gharib et al. (1998), and was shown to improve low pressure swing, allowing the engine to draw more fuel and air during the intake cycle. 63

Chapter 3 Conceptual Design The vortex structure, captured using particle image velocimetry (PIV), shows the interaction of vortices during the interchange between intake and exhaust cycles (Kailasanathan 2007).

Figure 40 - PIV images of vortex interaction

3.9.2 Bellmouth Design Bellmouth shape In order to optimize the effect of the bellmouth on the engine, the shape the design was considered. A shape that produced the good results and was relatively simple to manufacture was desired for the project. Due to the future use of the engine on a UAV, the overall size of the bellmouth was also considered, as increased diameter would result in a larger drag on the engine, reversing any positive effects seen by addition. The article ‘Best bell’ (Blair, G and Cahoon, M 2006) was used as a basis of the investigation of the performance of various bellmouth designs. Figure 41 shows the bellmouth designs that were considered for the second 2007 engine.

64

Section 3.9 FWE Bellmouth Development

Figure 41 - Bellmouth designs considered (Blair, Cahoon 2006)

The performance of the bellmouth sections were laboratory tested in comparison to a standard radius bellmouth, producing the data in Figure 42. It is estimated that the minimum pressure in the combustion chamber will be around 0.7 bar, resulting in a pressure ratio of up to 1.4 at the intake during the intake stage of operation.

Figure 42 - Performance of bellmouth designs

The article shows that in low pressure ratio conditions, the best results are seen with the elliptical and aerofoil shaped bellmouths. However, the improvement seen by the 65

Chapter 3 Conceptual Design use of a more complex shape is minimal, in the realm of 2% for the pressure ratio expected for the engine. In the interests of ease of manufacturing, it was decided that a simple radius bellmouth would be best suited to the intake of the engine.

Radius The 2007 project group conducted an investigation into the effect of bellmouth radius on the intensity of a returning pressure wave. The data was obtained using the one dimensional fluid dynamics software UFLOW1D. The results from the investigation are shown in Figure 43. These results obtained were confirmed with the statement “the minimum radius of the bellmouth lip is specified as being best between 0.25 and 0.3 times the diameter of the tube.” (R. Alan Wallis, 1983).

Figure 43 - The data obtained in 2007 using UFLOW1D (blue) and textbook recommendations (red)

Adjustability Engine 2 was designed by the 2007 project group using scaling of a significantly smaller engine. As a result the engines state of tune is questionable, and the ability to adjust the intake and exhaust lengths is desirable. Due to this, it was decided that an adjustable bell mouth design would be beneficial to the engine development. It was also noted that a continuously variably pipe length would provide better resolution 66

Section 3.9 FWE Bellmouth Development data than discrete length sections, and hence allow better understanding of the engine behaviour.

Manufacturing The ease of manufacturing of the bellmouth was considered in order to ensure that the final product was at the desired standard, and could be produced at a reasonable cost. Initially, the bellmouth was only be used for testing purposes, and hence was not required to be optimised for flight purposes. Discussion with workshop staff resulted in the conclusion that manufacturing a stainless steel bellmouth to the optimal radius would be very difficult, as the metal would not allow the amount of deformation required to be deformed into the desired shape. As such, it was decided that a machined curve would be much easier to produce, and would produce the best solution.

3.9.3 Final Design The final design was manufactured from a combination of pipe and machined billet. The machined curve was welded to the pipe section, which was sized to slide over engine pipe. This solution was particularly easy to manufacture, and provided a smooth curve to improve air flow. The model can be seen in Figure 44.

The greatest concern with this solution was the presence of a lip where the engine pipe ends, when the flow transfers to the inner surface of the bellmouth. While this situation was not ideal for flow, it was not of a major concern at the time of manufacture. This was justified by the presence of steps and uneven joints in the engine, due to the presence of welds and metal irregularities. It was decided that if the inner lip of the pipe were ground to smooth the lip, the transition would produce less interference to the flow than the joints that existed in the pipe, and hence it was not a significant concern for the performance of the engine.

67

Chapter 3 Conceptual Design

Figure 44 - The adjustable bellmouth design.

3.9.4 Flight considerations While the bellmouth manufactured for the second engine was not designed to be optimised for a flight condition, future engines will require consideration in this area. Concerns have been identified regarding the bellmouth and its impact on the plane aerodynamics. The greatest concern is that the frontal area of the bellmouth will create a significant amount of drag, which will not only reduce the performance of the UAV, but will create a low pressure region in the area that the intake draws from, potentially limiting the volume of air that the engine can intake. This would be detrimental to the performance of the engine, as thrust would dramatically reduce as the UAV velocity increases, limiting the aircrafts’ top speed.

In order to improve engine performance at high air speeds, a more streamlined bellmouth was designed, such that drag is reduced, and the pressure in the intake maximised.

To perform this analysis a Axi-Symmetric model of the intake was created using computational fluid dynamics (CFD) programs Gambit and Fluent. Three geometries were modelled as shown in Figure 45.

68

Section 3.9 FWE Bellmouth Development

Figure 45- Three intake geometries

The domain was setup with flow from left to right, over the intake, with a low pressure zone in the intake to simulate the intake cycle of the engine (Figure 46). Compressible flow was modelled as flow speeds reached over Mach 0.4 inside the intake. The height of the domain was set to 20 times the radius of the intake, whilst the length was 30 times longer than the intake pipe, to ensure the flow had stabilised before reaching the far field pressure outlet.

Figure 46- Domain Layout

Results were obtained at three different flow speeds, 10m/s, 40m/s and 80m/s. The results confirmed that the addition of a flare increases the mass flow rate into the 69

Chapter 3 Conceptual Design engine, however little difference in mass flow was shown wit the addition of a flare. The results can be seen in Figure 47. Mass Flow Rate Comparisons 0.16

Mass Flow Rate at inlet (kg/s)

0.14 0.12 0.1

No Flare

0.08

Standard flare Aerodynamic Flare

0.06 0.04 0.02 0 0

10

20

30

40

50

60

70

80

90

External Flow Velocity (m/s)

Figure 47- Effect of intake geometry on mass flow rate

The major gain found from the aerodynamic flare was a reduction in the drag coefficient by up to 50% at 80m/s. A comparison of the pressure field plots for the standard and aerodynamic flare at 80m/s can be seen in Figure 48 and Figure 49.

Figure 48 -Static Pressure Contours of Aerodynamic Flare at 80m/s

70

Section 3.10 Flight Engine Development

Figure 49-Static Pressure Contours on Standard Flare at 80m/s

3.10 Flight Engine Development From the analysis conducted in Section 2.6.3 it was determined that the Chinese Pulsejet Engine was the most suitable engine to develop for flight, most notably due to its documented success on liquid fuels. An extensive search produced numerous models developed with thrust outputs ranging from 2.6lbs to 50lbs. Overall, seven fully dimensioned engine models were discovered. This allowed a simple statistical analysis to be performed in order to determine the approximate size required for a 3kg engine.

The statistical analysis produced extremely accurate results with R2 values no lower than 0.97. This allowed trend lines to be used to predict all dimensions of the engine. This also provided information on the critical dimensions that must be considered when attempting to scale a pulsejet.

71

Chapter 3 Conceptual Design

Comparison of Thrust vs Engine Length 60 Thrust (Lbs)

50 40 Chinese Engine

30

FWE

20 10 0 0

20

40

60

80

100

Length (inches)

Figure 50- Statistical trend of Chinese and FWE engines

From the statistical analysis it was determined that an engine of similar length, but significantly smaller combustion chamber size, when compared to the current FWE design could be produced to achieve the required 3kg. This can be seen from the data collected in Figure 50.

From contact with Irvine Aeropulse, a small Brisbane based company, it was determined that the exhaust section of the statistical Chinese model could be better optimized to reduce engine length and improve performance. As mentioned earlier, the expansion for the exhaust can be optimized at around 5 degrees. The statistical engine had an expansion section with half angle of approximately 3.4 degrees, therefore by expanding this angle and maintaining the same area of the sections, in accordance with studies by Mcalley (2006). the final engine design reduced the length by 200mm and the phone diameter increased by 10mm, giving a final engine length of 850mm.

Material Selection and Stress Analysis To ensure the weight of the engine was kept to a minimum a finite element analysis (FEA) was conducted. Varying pressure and temperature loads were applied to a full engine model. Based on the results a separate axi-symmetric model of the combustion 72

Section 3.10 Flight Engine Development chamber end cap was created, as this was found to be the area of highest stress. The material chosen was 310 Stainless Steel, as it was the only stainless steel found which was capable of withstanding cyclic loading at 1150 degrees. Figure 51 shows the material properties of 310 stainless steel at high temperatures.

310 and 310s Material Properties 700

Strength (MPa)

600 500 310 Yield 400

310 Tensile

300

310s Yield 310s Tensile

200 100 0 0

200

400

600

800

1000

1200

Temperature (C)

Figure 51-Variation of material properties of 310 stainless steel with temperature

The analysis required coupled thermal and pressure loads, with load data based on the performance of the SNECMA Escopette (SNECMA 1951). From Figure 52 it can be seen that the pressure waves fluctuate in a sinusoidal manner between 50kPa to 150kPa (absolute) during each cycle. It can also be seen in Figure 52 that the pressure waves are constant in magnitude along the length of the engine. For this modelling case the engines operational frequency was assumed to be 200Hz.

73

Chapter 3 Conceptual Design

Figure 52-Operating pressure of the Escopette pulsejet

The model was created in pro engineer, and imported into ANSYS workbench where it was thinned and simplified for modelling. The analysis was conducted using shell elements with the aim of reducing the computational time of the coupled solver. A wall thickness of 0.5mm was initially investigated, as this was found to be a common value used in hobby pulsejet manufacture.

The meshing of the engine was created using element sizing and manually setting the divisions along the length. This ensures the mesh remains aligned and reduces element deformation. Elemental deformation leads to errors in both the temperature and stress gradients, causing inaccuracies in the model. The final mesh of the model is shown in Figure 53.

74

Section 3.10 Flight Engine Development

Figure 53- full engine mesh

The thermal and pressure loads were input into the flexible dynamic solver using Equation 12. The solver used, completes the thermal analysis first and determines thermal stresses on the model, before the pressure model is solved. Equation 12

P = 50000 * sin((2 * pi * 200 *180/pi) * time )

Figure 54-pressure loading input for flexible dynamic solver

The results of the full body model showed that the highest stresses existed around the combustion chamber end cap, as shown in Figure 54. Based on this analysis an axisymmetric model of the combustion chamber end cap was created to determine the strength of the cap with the inclusion of fillet welds.

75

Chapter 3 Conceptual Design

Figure 55-Stress Results on Combustion Chamber End Cap

The axi-symmetric model showed high stress regions around the base of the welds. The stress of 109MPa represents the yield stress of the material at 950 degress, which was determined to the operational temperature of the combustion chamber. Therefore the thickness of this section was increased to 1mm for manufacture, and the thickness of the TIG welds was also increased.

Figure 56- Final Results of the Axi Symmetric Model

3.11 Liquid Fuel System Design This section contains comparisons between a number of potential liquid fuels and the fuel characteristics that were identified as important for use in a valveless pulsejet engine.

76

Section 3.11 Liquid Fuel System Design 3.11.1 Fuel Choice The choice of fuel was based on several key chemical properties. These include the energy density, flash point, latent heat of vaporisation and flammability limits. Based on these properties it was determined that, petrol, kerosene and methanol were the most suitable fuels for testing. A detailed analysis of the fuels that were considered can be found in Appendix F – Liquid Fuels.

The chemical properties of petrol suggest that it was the most suitable fuel to use to power a valveless pulsejet engine. The high energy density means a relatively low fuel flow rate is required to produce a given thrust, thereby reducing the mass of fuel that would need to be carried onboard the aircraft. The low flash point of petrol (-46°C) means that it will ignite easily which should allow the engine to be started relatively easily. The main disadvantage of petrol is its small flammability range (1.3% – 6%), potentially limiting the throttleable range of the engine. Petrol also has a low latent heat of vaporisation, which will cause the engine to run hot. Too much heat may eventually affect the structural integrity of the engine, especially under static operating conditions.

A fuel such as methanol’s main advantages over a fuel such as petrol is its latent heat of evaporation and larger flammability range. As methanol expands, it absorbs a large amount of heat from its surroundings, resulting in an overall cooler burn temperature. The air being mixed with the methanol will therefore also be cooled, producing a higher density mixture in the combustion chamber. Therefore, a larger mass is able to be ejected from the chamber, thereby producing more thrust compared to lower density, higher temperature combustion mixes. The large flammability range would likely help to increase the throttle range of the engine. The trade-off for the cooler burn and large flammability range is a much lower energy density, being approximately half that of petrol. Therefore an engine running on methanol will require approximately twice the fuel compared to an engine running on petrol, to produce the same thrust. As a result, the weight of fuel needing to be carried to produce the same amount of thrust for the same time period will also increase by around a factor of 2. 77

Chapter 3 Conceptual Design For flight purposes, the fuel mass required will likely be the limiting factor for methanol.

Outcome The successful operation of a valveless pulsejet engine is highly based on acoustic resonance within the engine. Therefore, the combustion temperature of a given fuel will impact on the acoustic properties and hence the performance of the engine. To test the effect of the fuel properties on the engine a range of different fuel mixes were tested. From the information gathered, petrol will be used as the main fuel, due to its chemical properties and because it can be easily obtained. Petrol-methanol mixes were also tested, with as based on the fuel properties it reduced the fuel consumption of the engine. Fuels such as shellite (a.k.a naphtha) can also be mixed with petrol as based on its chemical properties it increases the frequency of the engine by raising the combustion temperature.

3.12 Fuel Injector Design As stated in Section 2.6.5 studies were conducted by Mcalley in 2006 into the development of a liquid fuel system for pulsejet engines, with much work conducted on valveless engines. While his testing showed limited success, mostly due to poor injector placement, he developed numerous injector styles which showed produced useful results. The initial injector studies revolved around the design of multi-holed pin injectors, created from 1/8” stainless steel. Flow visualisation and engine tests were performed on varying size holes and hole positions. The initial tests produced 6 and 12 hole injectors with 1.7mm holes. Flow visulisation of the injectors can be seen in Figure 57 and Figure 58. It can be seen that the flow pattern from the 12 hole swirl injector aims to promote better mixing in the engine, however it was found that this injector design was unable to sustain thrust within the engine. The 6 hole opposed spray injector was successful in sustaining thrust, but only with forced air. It can be seen in the flow visualisation that flow is only exiting through the bottom injector. This 78

Section 3.12 Fuel Injector Design concluded that the injector hole size was too large, as the pressure drop was not great enough to force fuel out of the other injector holes.

Figure 57- 12 hole swirl injector

Figure 58- 6 hole opposed spray injector

Continued development with decreasing hole sizes progressively showed that a single or double hole injector, with hole size 0.51mm was the most effective injector design, as they provided high pressure injection, which promoted better air/fuel mixing.

After looking at the work undertaken by Mcalley, we undertook an investigation into the performance of a simple pin hole injector, to allow estimates to be made regarding required injector sizing and performance. Initial calculations based of work from Williams (1990), suggested that an orifice diameter of 0.1mm would be large enough to supply 300ml/min of petrol for a supply pressure of 3bar. We were unable to test these calculations as such small orifice sizes were not able to be manufactured. Also, 79

Chapter 3 Conceptual Design the calculations did not account for the dynamic behaviour of the engine, or the high operating temperatures. The high temperatures during operation would reduce the diameter of the orifice, causing a significant reduction in flow rate.

The operating characteristics of swirl injectors were also analysed, but in depth numerical calculations were not undertaken. Swirl injectors produce greater spray angles than pin hole injectors, due to the added tangential component of velocity. However, the performance of these injectors varies dramatically depending on the internal structure. Therefore, as we were not designing our own injectors, this exercise was done to obtain knowledge regarding general injector performance, to allow decisions to be made during engine testing, regarding the performance of the system.

Following this, a review of commercially available injectors was conducted. It was found that very few injectors met the specific requirements of both size and fuel flow rate, whilst being able to handle the required temperatures. Two main injectors were found, the first manufactured by BETE, who are a custom nozzle manufacture for a wide range of industries in the United States.

The BETE PJ series offered flow rates between 0.043 to 5.34 L/min from a supply pressure of between 3 and 10Bar. The injector offered was a cone spray type injector, as shown in Figure 59.

Figure 59-BETE PJ Cone Spray Injector

Cone spray injectors produce highly atomised fuel spray, in a 90 degree arc. This increases mixing compared with a single pinhole injector, producing droplet sizes of 80

Section 3.12 Fuel Injector Design under 50 microns (BETE 2008). From the data provided a maximum flow rate of 0.34l/min could be achieved at 5 Bar through a 0.51mm orifice. However despite this the size of the injector was still larger than desired, as the intake of the pulsejet only has a diameter of 29mm.

The other injector which was found was a similar design spray injector available through a local Australian dealer. The designs were constructed from 5mm stainless steel tube, with the injector nozzle fitting inside the tube diameter, as seen in Figure 60. The manufacturer also quoted droplet sizes of 25 microns.

Figure 60-5mm stainless steel injectors

From the data available it was seen that a seen that a single injector could only supply a maximum of 170m/min though a 1mm orifice. However due to the size of the injectors it was determined that 3 injectors could be placed down the length of the intake to achieve the required fuel flow rates. Due to the availability, small droplet size and significantly smaller physical size, these injectors were purchased.

Six injectors were purchased as the fuel flow rate was unknown. Two 1mm, two 0.8mm and two 0.6mm injectors as the flow rate was expected to be between 200 and 350ml/min. The different sizes would ensure that once the required fuel flow rate was 81

Chapter 3 Conceptual Design found, that the optimum injector configuration could be determined. A table of fuel flow rates for each size injector can be found in Table 11.

Table 11

Fuel Flow rate Orifice size

(5 bar)

0.6mm

0.101

0.8mm

0.138

1mm

0.178

82

Section 3.13 Conceptual Design Summary

3.13 Conceptual Design Summary From the previous analyses, a three viewed drawing of the aircraft can be developed. This is shown below in Figure 61.

Figure 61: Conceptual Design Three View

83

Chapter 0 .

84

4 Detailed Design 4.1 Fuselage Structure Design In this section, the detail structural layout and analysis of the fuselage is described. Under structural layout, the detailed components of the fuselage are described, while under structural analysis the fibreglass skin and longerons used are checked for strength.

4.1.1 Fuselage Structural Layout The detailed layout of the fuselage is illustrated above in Figure 62. It has an external fibreglass skin and internal support structure, comprising of bulkheads and longerons. A removable heat shield is located between the fuselage and the engine however this is not a structural component.

Removable

Fibreglass skin

heat

protection

section

Internal reinforcing

Landing gear notch

Figure 62: Overview of Fuselage Structural Layout

Fuselage Skin For the main surface of the fuselage, fibreglass material was selected. It provides a light weight yet stiff structure and is commonly used in similar UAVs.

The basis of the skin structure was 320 gsm, 0˚/90˚ aircraft grade glass cloth, with a 50:50 volume resin mix. Three layers of the cloth are used, as determined in section 85

Chapter 4 Detailed Design 4.1.2. The resin selected for use with the fibreglass was a high temperature version, in order to resist heat from the engine. The chosen resin is a vinyl ester type with a 177 ⁰C glass transition temperature.

Radio waves can propagate through fibreglass and thus the antenna for the radio controller is able to be placed internally inside the fuselage. This is not possible with a carbon fibre construction without a high risk of the signal being broken.

Removable aluminium section The removable aluminium section is of sufficient size that when removed, provides access to the majority of the fuselage. This is required to provide access to the fuel storage. The aluminium section also functions as a heat shield, reducing heat transfer to the fuselage.

Alternatively, smaller access holes could have been placed along the bottom of the fuselage. This would however make manufacture more difficult, provide inferior access to the area and not provide heat shielding.

Internal Reinforcing Structure As the design for an aluminium section provides a discontinuity in the fibreglass structure, reinforcement was designed to strengthen this area. Figure 63 shows how the bulkheads and longerons will be used to reinforced the fibreglass skin.

Bulkhead

Neutral

Longeron

Moment of Inertia Contribution of Longerons Figure 63: Fuselage Internal Reinforcing Structure

86

Section 4.1 Fuselage Structure Design The longerons have been added to improve the bending strength of the fuselage. Unidirectional carbon fibre tubing is used, as the fibers are orientated along the length of the tube, resulting in a very stiff longeron-bulkhead structure, and hence providing strong resistance to bending. Structural analysis of the longerons is conducted in the ‘Fuselage Structural Analysis’ section.

Three longerons are to be used, distributed at the top corners and lowest point of the rear bulkheads. The bulkheads perform the function of connecting the longerons to the fuselage skin. Bulkheads also providing stiffening to help the fuselage hold it’s cross sectional shape and prevent bulking. The specific structural contribution of the bulkheads has not been analysed, however it is expected they will significantly stiffen the fuselage.

Six bulkheads are to be used, including five main bulkheads and one nose bulkhead. The three rear bulkheads have a half-circular shape. The rearmost bulkhead assists with the tail fuselage join, while the remaining two provide strength around the wingfuselage interface. The bulkheads also provide segmentation between the fuel bags. Two full circle bulkheads sit in the back half the nose section. The main engine mount attaches to the bulkhead at the rear, while the forward bulkhead provides attachment for the front landing gear and electronics tray. The front most bulkhead in the aircraft sits toward the front of the nose and has a smaller radius, providing the second attachment point for the electronics tray.

Landing gear notch A landing gear notch has been incorporated into the fuselage shape at the required location of the main landing gear. This is required as the central part of the main landing gear is flat and a suitable contact surface is necessary. A landing gear tray was avoided, as the extra weight and complexity was considered unnecessary.

87

Chapter 4 Detailed Design 4.1.2 Fuselage Structure Selection To determine required specification of the fuselage skin and longerons, a stress analysis was performed. Component weights and estimated aerodynamic forces were used to estimate the force distribution on the fuselage and thus the stresses were determined. The details of the analysis can be found iAppendix I – Fuselage Stress Analysis. Table 12, shows the maximum stresses found at the most critical locations in the fuselage. Table 12: Fuselage Stress Analysis Results

Fibreglass

Longerons

Max Shear Stress

Max Shear Stress

Max Shear Stress

Max Bending Stress

3-ply (MPa)

2-ply (MPa)

1-ply (MPa)

(MPa)

13.88

20.83

41.65

70.91

For shear, the fuselage has been analysed for one, two or three layers of fibreglass on the basis of .25 mm per ply. While for bending, longerons of 7mm outer diameter and 5.5 mm inner diameter have been assumed

Fibreglass thickness selection While calculations suggest 1-ply of 80 gsm fibre-glass would be satisfactory, the calculations are based on idealised flight conditions, and do not consider bulking failure. In order to ensure that the plane is able to withstand a reasonable crash, 3 layers of 320gsm glass was selected. In addition to this, a layer of Kevlar was selected to be used in the nose area

Longeron selection Since the longerons do not solely carry the load of the aircraft, the fuselage carries some of the load, the size of longerons previously mentioned are sufficient for use in the aircraft. Due to availability, the final longerons selected had an outer diameter of 8 88

Section 4.2 Wing Design mm and an inner diameter of 6 mm, making them stronger due to the larger diameter and extra thickness.

4.2 Wing Design The aircraft wing was partially specified by the decisions made in conceptual design; however, more detailed specifications were required to enable construction of a final wing. The airfoil selected was a NACA 4412 to enable high lift and low drag at a varying range of angles of attack. The ailerons were sized to provide adequate performance in roll.

4.3 Wing Structural Design The wing structure was designed to withstand all operational loads anticipated within the aircrafts’ flight regime, with considerations made to ensure the structure is easy to assemble. All flight loads were determined from the United States Federal Aviation Regulations, section 23, (FAR-23) which governs the airworthiness standards for Normal, Utility, Acrobatic and Commuter aircraft. Material safety factors were obtained from the Australian Civil Aviation Safety Regulations (CASR), Section 101, which governs the use of Unmanned Aerial Vehicles and Rockets. The results are summarised below and have been taken into account in all calculations: •

Maximum wing load factor of 4.4 in accordance with FAR-23

•

A safety factor of 2.25 in accordance with CASR 101

•

A safety factor of 1.5 on composite materials in accordance with CASR 101

4.3.1 Lifting force profile The lifting force generated by the wing was initially determined so that the stresses in the wing could be designed for accurately. Two methods were initially utilised, to ensure that all conditions of flight could be analysed. In the analysis, the following assumptions were made. •

All lifting forces are due to the wing, the fuselage generates no lift.

89

Chapter 4 Detailed Design •

The aerodynamic lifting force is the most significant force on the wings, drag forces are not considered.

•

Both methods are approximations; they do not take into account wing tip vortices.

4.3.1.1 Shrenk’s Approximation The lifting profile was initially generated using Schrenk’s approximation. The method states that any untwisted planar wing’s spanwise load distribution shape can be approximated by the average of its actual planform shape and an elliptic wing shape of the same span and area (Schrenk, 1940), as illustrated in Figure 64.

Figure 64 - Schrenk's Approximation (ISOAR 2007)

The wing planform and elliptical planform are determined using Equation 13 and Equation 14 respectively. Equation 13

Equation 14

90

Section 4.3 Wing Structural Design The calculation of C(y) gives the overall area of the wing. In order to determine the lifting force, an approximation of the load to be supported by the wing is made. In this case each wing is assumed to support half the weight of the aircraft, multiplied by the load factor determined using FAR-23 calculations for V-n diagrams. Using the load estimation and the surface area a force factor Equation 15 can be calculated to determine the load distribution. Equation 15

The load created using this method was corrected for the weight of wing, giving the load profile as shown in Figure 65. It should be noted however that the wing is supported from the edge of the fuselage, such that all further calculations will only involve the area from 0.05m to 0.79m. Schrenk Approximated Lift Force 400 350

Load (N)

300 250 200 150 100 50 0 0

0.2

0.4

0.6

0.8

1

Spanwise Distance (m)

Figure 65 -Lifting Force Distribution Based on the load distribution calculated shear force diagram could be created from the integral of the load distribution between 0.05 and 0.79, including appropriate boundary conditions. Similarly the bending force diagram can be determined. The calculations were completed with the assistance of Matlab 7.4.

91

Chapter 4 Detailed Design Shear Forces 0 -20 -40

Force (N)

-60 -80 -100 -120 -140 -160 -180 -200 0.1

0.2

0.3 0.4 0.5 distance from root (m)

0.6

0.7

Figure 66- wing shear distribution Bending Force 60

50

Force (Nm)

40

30

20

10

0 0.1

0.2

0.3 0.4 0.5 distance from root (m)

0.6

0.7

Figure 67 - wing bending force distribution

4.3.1.2 Alternative Distribution Calculations Whist Schrek’s method provides a good estimate of the wing load distribution it has no theoretical reason for its accuracy. To better understand the aerodynamic forces applied to the wing, a more theoretical method was investigated. The method chosen is outlined in Theory of Wing Sections (Abbot 1959). The method considers the spanwise distribution to consist of two parts. The first part, known as the ‘basic distribution’ is based on the twist of the wing, which for in this case is zero. The second part is known as the additional distribution and is the lift due to the wing angle of attack. The method, uses tables to easily select parameters for the wing shape, and 92

Section 4.3 Wing Structural Design uses data from the aerofoil at different speeds to determine an accurate lift profile. The data required to find the spanwise lift distribution is given in table form, and can be interpolated based on the aspect ratio and taper ratio of the wing. From the data provided, and data gained from the Javafoil aerofoil simulator, and approximation fro the wing lifting for can be found using Equation 16. Equation 16

Where CL is the lift coefficient of the wing, as determined from Javafoil at a certain condition, and clal is determined from Equation 17. Equation 17

Based on these calculations the following lift distribution was determined for a cruise speed of 88m/s at zero angle of attack (Figure 68).

Corrected Cl Distribution (88m/s) 0.12 Corrected Cl

0.1 0.08 0.06 0.04 0.02 0 0

0.2

0.4

0.6

0.8

1

Spanwise Postion from Root (m)

Figure 68 - Corrected Cl Distribution

93

Chapter 4 Detailed Design By estimating the area of the wing area at span wise stations along the wing, the lift distribution was determined. The distribution is then corrected for weight, by estimating the shape of the wing as a trapezoid (Beer 2006).

Lift Distribution at 88m/s (Including Load Factor) 12

Weight Corrected Lift Distribution

Lift force (N/m)

10 8 6 4 2 0 0

0.2

0.4

0.6

0.8

1

Spanwise Distance from root (m)

Figure 69 - Lift Distribution at 88m/s This method offers the opportunity to analyse the loads on the wing at different points in the aircrafts flight profile. It also allows more design flexibility, as the effect of such factors as sweep and twist can be introduced to the analysis.

94

Section 4.3 Wing Structural Design 4.3.2 Spar Design Spars are utilised as the primary method for handling the bending and shear forces due to the aerodynamic lifting forces generated during flight. Spars were located at the 20% and 80% wing location (Avalakki et al, 2007). It was determined that two spars on the top and bottom surfaces of the wing would be the best option for transferring loads along the wing into the fuselage. Calculations for the bending stresses in the spars were undertaken by determining the moment of inertia for the four spar system. To determine the moment of inertia first the centroid of the spar system was calculated using Equation 18 and Equation 19. Equation 18

Equation 19

From the geometry of the aerofoil the moment of inertia along the wing (x direction) could be determined using Equation 20, assuming that only the distance between the spar and centroid caused inertia in the system. Equation 20

The analysis of the bending stresses in the wing at different positions was used to determine the cross sectional area required for each spar using Equation 21. Equation 21

95

Chapter 4 Detailed Design By analysing the bending stresses at the root of the wing, it was determined that an overall spar area of 1.25x10-5m was required. This was equivalent to four 1.25mm thick by 11mm wide spars manufactured from uni-directional carbon fibre with a tensile stress of 580MPa and a compounded safety factor of 3.75, based on the CASR regulations outlined in Section 4.3. The bending stresses were also analysed at spanwise sections along the wing, as shown in Figure 70

Optimum Spar Thickness 0.012

Spar Thickness (m)

0.01 0.008 0.006 0.004 0.002 0 0

0.1

0.2

0.3

0.4

0.5

0.6

0.7

Distance from root (m)

Figure 70 - Maximum spar thickness from root to tip of the wing This initial analysis assumed that the bending force was taken equally by each spar, however further analysis was undertaken due to the non-symmetric load conditions which exist on the wing. In general the centre of pressure for a wing is at approximately 25% of the chord. Therefore the percentage of the load carried by the front and rear spars was calculated using Equation 22 and E quation 23.

96

Section 4.3 Wing Structural Design Equation 22

E quation 23

Analysis utilising this method concluded that four 1.25mm thick, and 12mm wide spars would be capable of resisting the bending stressed induced by the wing lifting forces.

The shear stress was also analysed at the root of the wing. Assuming that all the shear is taken by a single spar, a shear stress of 24Mpa was calculated using Equation 24. Equation 24

The shear strength of unidirectional carbon fibre is 70Mpa, therefore it was deemed unnecessary to analyse the shear at other positions along the wing.

4.3.3 Torsion Torsion in the wing structure was estimated by assuming that the lift force of the wing acts through the centre of pressure of the aerofoil, creating a moment around the static neutral point of the wing. The static neutral point is calculated by analysing the spars and the foam between the spars as a single structure. Using a method defined by Beer et al (2006), the structural neutral point for a 2-D body consisting of multiple materials can be approximated using a ratio of the materials Modulus of Elasticity, to create a scaled moment of inertia.

97

Chapter 4 Detailed Design The structural neutral point of the wing was determined at the root chord of the wing. The torsion was analysed in both takeoff and cruise scenarios between -5 and 15 degrees angle of attack. The centre of pressure from the aerofoil was determined using Designfoil software. This software re-creates the pressure profile over the aerofoil at various angles of attack. From this data the centre of pressure is determined using Equation 25. Equation 25

The centre of pressure is not affected by changes in Reynolds number. Figure 71 shows the change in position of centre of pressure with angle of attack.

Change in Centre of Pressure (Cp) of NACA 4412 with Angle of Attack (AOA)

Cp (x/c)

100 90 80 70 60 50 40 30 20 10 0 -5

0

5

10

15

20

AOA (degrees)

Figure 71 - Position of Centre of Pressure with AOA

Based on the structural neutral point and the centre of pressure, a level arm could be determined. The root chord was analysed at takeoff (70km/hr), climb (150km/hr) and cruise speed (300km/hr), at various angles of attack, as shown in the following figures. 98

Section 4.3 Wing Structural Design

Torsion vs Angle of Attack (Take-off) 4.5

Torsion Nm

4 3.5 3 2.5 2 1.5 1 0.5 0 0

5

10

15

20

AOA (degrees)

Figure 72 - Wing Torque at Takeoff (70km/hr)

Torsion vs AOA (Climb Speed) 25

Torsion (Nm)

20 15 10 5 0 0

5

10

15

20

AOA (degrees)

Figure 73 - Wing Torque at Climb Speed (150km/hr)

99

Chapter 4 Detailed Design Torsion vs AOA (Cruise Speed)

100 Torque N/m

80 60 40 20 0 0

5

10

15

20

AOA (degrees)

Figure 74 - Wing Torque at Cruise Speed (300km/hr)

The maximum torque of 77Nm occurs at 13 degrees angle of attack at 300km/hr. Although it is unlikely that these conditions will be experienced by the aircraft, it has been designed for to ensure the wing structure is capable of handling all conditions within the speed profile of the aircraft.

The wing torque was analysed using Equation 21, assuming that the fibreglass skin could take the torque in the wing.

The moment of inertia of the structure was calculated initially assuming a 0.5mm thickness. This equates to approximately 3 layers of 84gsm fibreglass. Stresses at the root of the wing were calculated resulting in a maximum stress of 23.2Mpa. This allows a 12.9 times safety factor, based on a tensile strength of 300MPa for fibreglass (FGI, 2007).

100

Section 4.4 Wing Connection

4.4 Wing Connection The wing connection was designed with three aims in mind. •

Ability of the structure to withstand normal flight loads

•

Be easy to assemble and disassemble

•

Easy to replace in the event of a crash

•

Protect the wing structure from damage during the event of a crash

For this reason it was determined that a twin beam system, as shown in was the most suitable method for this application. The system was designed to cope with ten times load factors, in accordance with a heavy landing where the wing comes into contact with the ground.

Figure 75- Wing Connection System

The wing system was designed so that it would fail in three stages during a crash. The aim of this was to progressively protect the more expensive and difficult to replace components. The first phase of failure was through a tensioning system. This protected the main load bearing beams from the initial Impact loads. Once failure of the tension system failed the loads were transferred to the wing connection beams, which were manufactured from uni-directional rolled carbon fibre. While these beams were relatively low in cost, it was found that obtaining the required diameter was difficult, 101

Chapter 4 Detailed Design with up to two months lead time quoted by some companies. Finally, in the event of a serious crash the beams would fail and the wings would detach from the main body.

The main beam section was designed to take the loads during standard flight operations. The problem can be analysed as a simple cantilever beam in bending. Force loads of 10x the normal flight loads were used to take into account landing, and to ensure deflection of the wing was kept to a minimum during normal 4G flight maneuvers. Two carbon fibre rods were used to support the wing. They were situated at 20% and 80% of the root chord, and extended to the start of the ailerons, where it was mounted into a 4mm ply rib. The ply rib was analyzed to take shear stresses during landing, based on a 10x load factor. Based on a shear stress of 50MPa for plywood, there was a 3.5x reserve factor, suggesting that plywood was the most suitable choice. From the stress analysis conducted it was determined from that two beams of 12mm and 10mm 1mm thick uni-directional carbon fibre were able to take the aircraft loadings, however further analysis showed that a deflection of 60mm in the wingtip would occur under maximum load conditions. This was deemed excessive and therefore the thickness of the main beam was increased to 2mm. this reduced the deflection by 2/3, which was seen as suitable for flight.

102

Section 4.5 Control Surface Sizing 4.5 Control Surface Sizing In this section the sizing of the control surfaces and servo motors to actuate them is presented. The aircraft had a rudder less design as sufficient control for utility flight can be achieved using the elevators and ailerons. Standard radio control servo motors were selected for use, with compact motors preferred.

4.5.1 Aileron Sizing The ailerons will provide roll control. Considerations relevant to their design are: •

It is preferable for ailerons to be located as far out-ward as possible - where they will provide maximum moment.

The same actuating effort will be

required regardless of position, as the effort depends only on the size of the surface being deflected. •

Guidelines state that suitable percent chord for the ailerons to occupy is between 35-20% (Raymer 2004).

•

The ailerons should occupy approximately 7% of the total wing area, according to Eger (1983).

Since the trailing edge of the wing is straight and the rear spar is straight, then for simplicity the aileron can be rectangular.

The required area for each aileron is: S = .38 m 2 2S aileron = .07 × S = .0266 m 2 S aileron = .0133 m 2

The required span for each aileron is: b aileron = .0133 m2 / .0694 m = 191.6 mm

103

Chapter 4 Detailed Design Thus the required dimensions of each aileron are approximately 70 mm by 200 mm, the parameters of the ailerons are tabulated in Table 13 and the ailerons are illustrated in Figure 76. Table 13 - Aileron Dimensions

S aileron (m2 – per aileron)

.0133

C aileron (mm)

70

b aileron (mm)

200

20% of root

20% of

chord (~69.4 mm)

root 40 mm 200 mm

Saileron

chord

Figure 76 - Aileron Dimensions

Functions in the radio controller allow for the ailerons to double as ‘flaperons’, where the motion of the surfaces is together instead of opposed, for additional lift at take-off. This will be integrated into the control system to provide added lift on takeoff and landing if it is considered necessary during testing.

4.5.2 Elevator Sizing The elevators were sized with a chord of approximately 20% of the horizontal tail. The elevators were designed to span as much of the horizontal tail as possible for maximum control authority. The elevators run from either side of the central section of the T-tail from the vertical tail outward.

The positioning of another elevator surface in the central part of the tail between the vertical tails was also considered. The extra complexity of manufacture and extra 104

Section 4.5 Control Surface Sizing servo required was thought to be too complex for the small increase in elevator ‘power’.

4.5.3 Servo Motor Sizing The servo motors to operate the control surfaces were sized based on the required torque to actuate the surfaces, which was found based on the dynamic pressure and the size of the control surfaces. The torques were calculated for the extended goal design speed of 250 km/hr, with greater velocity requiring more actuating torque and thus providing for additional reserve factor at lower flight speeds. The servo requirements can be seen in Table 14 Table 14- Servo Requirements

Surface

Deflection

Required

(degrees)

for

torque Selected

maximum servo

deflection (kg.m)

Torque selected

of Reserve Factor

servo (kg.m)

Aileron

20

2.0

Eagle E711

4.2

2.1

Elevator

15

1.2

JR NES 331

3.25

2.7

The servo motors supplied with the radio transmitter were sufficiently powerful to operate the control surfaces. However, their physical size was too large for the intended wing based. Therefore, smaller servos were sought.

The servo selected for the ailerons was the ‘Eagle E711’. It had the same width and depth as the servos supplied with the transmitter but was a ‘low profile’ servo with a much reduced height that enabled it to fit inside a wing mounting. They have dimensions measured to be 44 x 22 x 23 mm.

The servo selected for the elevators was the ‘JR NES 71’. It was a ‘micro’ servo, with measured dimensions of 30 x 13 x 28

105

Chapter 4 Detailed Design For the nose wheel steering, the servo supplied with the radio transmitter was deemed suitable as mounting space was not an issue and the servo torque was sufficient.

4.6 Pulsejet Engine Mount The engine mount system is required to take the loads of the engine in all operating conditions, as well as provide vibration and thermal isolation to the fuselage. Initial engine testing showed that the temperature of the engine could reach as an estimated 1000C if left for an extended period without any forced air cooling, requiring that thermal conditions be considered in the design of the mount.

Initial testing on the flight engine produced a maximum sustained thrust of 3.5kg on propane. As the engine performance at speed was not known, the peak thrust expected from the flight engine was safely defined at 4kg.

A safety factor of 2 was incorporated into the design to allow for any increased load due to the cyclic vibrations seen in the engine. It was also decided that the front engine mount would be designed to take the full thrust load of the engine, in the case that the rear mount was not aligned correctly.

4.6.1 Mounting Locations In order mount the engine securely, it was decided that the engine needed at least 2 rigid mounts. Initial concepts involved tabs welded to the engine; however it concern was raised in regards to the loading on the mounting locations. While the 2007 engines did not suffer any form of damage from the mounting tabs, the transition to 0.5mm sheet, from 1.2mm sheet significantly reduced the ability of the engine walls to take the bending and moment stresses. For this reason, it was decided that the spark plug nut was a better mounting location, as it is the strongest component of the engine, and hence least susceptible to warping or failure due to the engine mount loads.

106

Section 4.6 Pulsejet Engine Mount Following initial engine tests using the spark plug nut as the primary engine mount, a crack was discovered at the weld interface of the nut. It is believed that this occurred due to increased stresses as a result of the engine oscillating about the spark plug nut, and hence it was decided that another mounting point should be used.

It was decided that the addition of a mount forward of the combustion chamber would permit the mounting of the engine, while transferring the thrust and vibration loads to the combustion chamber. This extension is shown in Figure 77. Mounting the extension above the sparkplugs allows the engine to have a longer thermal path to the bulkhead, improving the thermal performance of the engine mount.

Figure 77 - The mounting extension on the front of the Chinese engine

4.6.2 Thermal Isolation The expected temperatures of the engine near the combustion chamber approach 1000C, which, if this temperature was not effectively isolated, would cause severe problems to the fibreglass and timber materials used in the aircraft. The maximum temperature for the plywood bulkhead is approximately 100C, after which the material begins to degrade. In order to ensure that the materials are not exposed to excessive temperatures, a thermal analysis was required. The model used considered both the effects of convection and radiation, in order to produce a realistic representation of the actual situation. 107

Chapter 4 Detailed Design

The emissivity of stainless steel varies dramatically with surface conditions, which, at the temperatures expected is not well known. Because of this, the value of emissivity was conservatively estimated at 0.5.

The convection coefficient also had to be estimated, due to the uncertainty of the conditions that will surround the mount. The value of 15 W/m2 K was estimated from common values given for forced and natural flows. It was assumed that the flow around the mount would be greater than natural convection, due to the movement of the engine and airflow, and as such, the high value in the common range of 3 to 20 W/m2 K was used (Mills, 1998).

4.6.3 Vibration Isolation Due to the operation of the pulse jet engine, a large amount of vibration is produced, which could lead to problems if not correctly isolated. No information was available in regards to the vibration tolerances of the electronic systems that will be present in the plane, so it was decided that the best vibration isolation that is reasonable obtainable should be used.

Engine tests on propane produced an operating frequency of around 200Hz, and hence this value was considered the dominant frequency to isolate.

As can be seen in Figure 78, the best isolation of a vibration force is achieved when the frequency ratio is as high as possible, and damping ratio is as low as possible. As such, it is desired to produce an engine mount and engine system that has the lowest natural frequency that is obtainable while maintaining enough structural strength to resist failure.

108

Section 4.6 Pulsejet Engine Mount

Figure 78 - Force transmissibility as a function of frequency ratio and damping ratio (Chan 2007)

The natural frequency of a spring-mass system is given by Equation 26 below: Equation 26

w=

k m

As such, to reduce the natural frequency, the spring stiffness must be as low as possible, as the mass of the engine is not changeable. Figure 78 shows that in order to achieve a reasonable level of vibration isolation of 50%, it is necessary to have a frequency ratio of around 2 for low damping systems, or around 3.5 for systems with high damping levels. To achieve a frequency ratio of 2, the engine system must have a natural frequency of around 100Hz, which, for a mass of 0.8kg, gives a spring stiffness of 315827N/m. For a spring stiffness of 315827N/m, the deflection that the spring would have under a thrust load of 40N is 0.13mm. Table 15 shows the expected spring stiffness and deflection over a range of frequency ratios.

109

Chapter 4 Detailed Design Table 15 - Spring stiffness and deflection under a 40N thrust load, for various frequecy ratios. The shaded reigion of the table will not provide any vibration isolation.

Frequency Ratio 0.5 1 1.5 2 2.5 3 4 5 6 8 10

Stiffness (N/M) 5053237 1263309 561471 315827 202129 140368 78957 50532 35092 19739 12633

Deflection (mm) 0.01 0.03 0.07 0.13 0.20 0.28 0.51 0.79 1.14 2.03 3.17

The actual vibration levels produced by the engine were not available, as during testing the vibration intensity is affected by the significant mass (around 14kg) of the mounting trolley. Recoded data suggests that the engine and trolley has a variation in thrust of around ±0.3kg about the mean thrust. In order to obtain more accurate vibration quantification, the test stand was approximated to be a spring-mass system with an input force equal to the thrust variation, approximated as a sinusoidal force. Data obtained from the manufacturer of the load cell and an ANSYS simulation of the trolley was used to estimate a spring coefficient for the system. From this an estimate of the natural frequency of the trolley-engine system was found to be around 940Hz. This gives a frequency ratio of around 8. Using Figure 78, and the assumption that the trolley system was almost critically damped due to friction, it was determined that the vibrations transmission is in the order of 15%. Thus, the actual force variation generated by the engine is estimated to be in the order of +-2kg. The calculations used to determine this can be found in Appendix E – Engine Mounting Calculations

4.6.4 Vibration Isolation Method: In order to develop the optimal design, it was necessary to consider various methods which would allow for vibration isolation. 110

Section 4.6 Pulsejet Engine Mount Buckling beam The use of a ‘buckling beam’ style vibration isolation system was considered to provide the spring support. This approach exploits the non-linear spring co-efficient of a buckling beam, and would hence be able to provide a high level of vibration isolation, without excessive deflection. The problem that limits the use of this design is that under part thrust conditions, where the average loading is less than that at peak thrust, the spring stiffness will be much higher than is required to isolate vibrations. As a result, this approach would only provide vibration isolation under full thrust conditions, and is hence not suitable for the application.

Vibration damping A soft foam or rubber vibration bed was considered in order to provide or supplement the isolation. Initial research produced no materials that could provide stiffness as low as is required, but still operate in the thermal conditions near the engine. The use of the rubber as a supplement to the mount deflection is however reasonable. In order to investigate this approach, a sample of the most suitable rubber obtainable from a retail outlet was obtained. As the outlet could not provide material data for the product, the young’s modulus of the material was estimated from a basic deflection test to be in the order of 5.6 MPa. A problem associated with the use of a rubber is the increased damping of the system will increase force transmissibility, and hence its use is not ideal if a spring system is possible. The material was also flame tested to ensure that it did not pose a significant fire hazard, which it did not.

Spring and Linkage System The use of a linkage system and a conventional coil spring was also considered for this task. This approach could easily allow the required deflection, and would not have any structural issues due to cantilever bending. The major concern that was associated with this design is the complexity of the system, which requires 4 separate parts and 2 pivot axis. The primary concern was related to the thermal conditions at the top pivot

111

Chapter 4 Detailed Design point. It was thought that a pivot point that is tight enough to stop vibration within the joint when cold may bind when the joint gets hot due to thermal expansion.

Cantilever beam The use of a cantilever beam spring system is the simplest and most conventional method considered. This approach would provide the required stiffness for vibration isolation, while also providing thermal resistance. The potential downside to this approach is that the ideal vibration isolation may not be possible whilst retaining thermal isolation and a low mass. This approach was the most promising of the options considered, and hence was investigated further.

4.6.5 Engine Mount Materials The material used for the engine mount is required to be able to support the loads of the engine at high temperatures for a reasonable period of time. Initial investigations showed that few materials were available that were able to provide the required structural strength at the expected operating temperatures, as well as provide good thermal isolation. The best materials for thermal isolation, predominantly ceramics, would be very difficult to manufacture, and would be prone to brittle failure, which is undesirable, as it would most likely result in a crash of the model, and a significant fire hazard. Aluminium alloys were not considered due to melting points commonly in the order of 600C. The materials reviewed for selection were restricted to mild and stainless steels on the grounds of manufacturing and durability requirements. Table 16: Material Selection for Engine Mount Matweb 2008

Material

Thermal

Yield strength

Max Service

Density

conductivity

(MPa)

Temperature

1018 Mild Steel

~50

275

(not given)

7.87

310 Stainless Steel

~20

310

1150

8.00

304 Stainless Steel

~20

215

925

8.00

302 Stainless Steel

~20

275

925

7.86

112

Section 4.6 Pulsejet Engine Mount

The material properties of stainless steel at elevated temperatures can be estimated from Figure 79.

Figure 79 - Yield stress relative to room temperature as a function of temperature for 301,302,304,321,347 annealed stainless steels (USADOD, 1998)

The increased thermal conductivity of the mild steel will result in significantly reduced thermal isolation, and as such is not suitable for the application. Of the stainless steels considered, 310 stainless steel provides the best performance characteristics. 310 stainless steel is also readily available, and hence is the material chosen for the manufacture of the engine mount.

4.6.6 Final Design This design incorporates the engine extension to reduce stresses on the spark plug nut. As a result of the extension, it also produced much better thermal analysis results than the initial design. The final design is shown in Figure 80.

113

Chapter 4 Detailed Design

Figure 80 - The final engine mount design. The modification made to the front of the engine is shown in green.

Thermal Analysis Analysis of the final engine mount design showed that the temperatures at the mounting points were at quite safe levels. The addition of the extension to the combustion chamber provides the majority of the thermal isolation, reducing the temperature at the interface with the engine mount to around 175C. If further thermal isolation were required, a layer of ceramic material could be placed between the two parts at the interface, and would significantly reduce the temperature at the bulkhead.

Figure 81 - Thermal analysis results of the engine mount.

114

Section 4.6 Pulsejet Engine Mount The results of the thermal analysis suggest that there will not be thermal damage the bulkhead, and hence the design is considered to be satisfactory.

Vibration Analysis Using a preliminary ANSYS structural model, the spring co-efficient of the mount was derived from the deflection due to the loading. This value was then used to calculate the natural frequency of the engine system. The natural frequency was found to be 31 Hz, resulting in a frequency ratio of 6.3. This value will reduce vibration to around 10% of the initial value.

This result was considered to be acceptable for the application, and hence no further design changes were required on the basis of vibration isolation.

Structural Analysis Structural analysis of the engine mount showed that stresses in the critical locations are below the yield stress of the material under the 80N design load. The peak reported stresses are singularities at the constraint points, and not of concern.

Figure 82 - Stress distribution within the initial design under an 80N load

115

Chapter 4 Detailed Design

In order to determine if the stress behaviour of the engine changes under dynamic conditions, a flexural dynamic analysis was performed. In this analysis, a load of 40N was applied, with a variation of 20N at 200 Hz about this mean value. The stress distribution is shown in Figure 83, showing that all critical locations are significantly below the material capabilities. As with the static analysis, the peak reported stresses are singularities at the constraint points, and not of concern.

Figure 83 - Stress distribution within the design under dynamic loading of 40N +- 20N

Verification In order to verify the thermal and vibration analysis results, the engine mount was used throughout the testing stage of the Chinese engine. During this testing, it was found that the temperatures during operation were never high enough to cause damage to the plywood mounting plate, even on extended runs. Due to technical issues with the load cell, vibration data was not obtained, and hence the vibration analysis results could not be verified.

116

Section 4.6 Pulsejet Engine Mount 4.6.7 Engine Modal Analysis A modal analysis of the engine was performed using ANSYS software to determine if specific mounting systems would cause excessive vibration transmission to the fuselage of the plane. It was determined that mounting the engine at the spark plug nut and the end of the tail resulted in oscillation modes at 197 and 208 Hz, with a shape similar to the first mode of a simply supported beam (Figure 84). This was reason for concern, as the operation of the engine may excite this mode, resulting in excessive vibration of the engine.

Figure 84 - 208Hz vibration mode of the engine, mounted at ends

Due to the presence of this vibration mode, the rear engine mount was moved to inwards from the tip of the tail, to the joint between the exhaust cone and the final pipe. An analysis of this mounting method did not produce any modal frequencies in the operation range, and as such it was considered the best solution.

117

Chapter 4 Detailed Design 4.7 Pulsejet Launch System In order to fly the aircraft powered by a valveless pulsejet engine, it is necessary to develop a method of safely starting the engine while mounted to the fuselage. This requires that the aircraft is supplied with propane, compressed air, and spark during the start up process. Due to safety concerns, the operators are required to be positioned several metres away from the aircraft, to ensure safety even in the occurrence of a fire. As such, the aircraft will be started and released from a distance of at least 5 metres.

As propane, air and spark are only required for the initial start up, these systems are integrated into the launch system, and not attached directly to the plane. After the engine is started and the plane is released, the propane and compressed air injectors will be left behind.

4.7.1 Launch process The following is the launch process required to start the pulsejet engine and release the plane on the runway.

Start the engine on propane: Propane and compressed air are supplied via the supply lines. Spark is supplied from the generator circuit, located near to, but not in the way of the plane. Transition to onboard liquid fuel: Throttle on the RC unit is increased as the propane flow is decreased. Completely stop Propane and air: turn off both propane and air. This must be done before release to prevent flames coming in contact with the tail of the plane. Remove spark: The spark plug is hooked up to the spark generation box with a quick release coupling. Once the engine has settled, the spark circuit is turned off. The wires can then be pulled off from a remote location, releasing the coupling from the aircraft.

118

Section 4.7 Pulsejet Launch System Release: Once the engine is running on liquid fuel, the plane is released from the stand immediately. This is done by releasing the pin that joins the plane to the launch stand. This leaves the air and propane injectors behind. If complete automation was required, this could be done with an electric solenoid.

4.7.2 Launch Stand Components The launch stand consists of four main components which are required to allow each step of the launch process. These components are discussed in the following section.

Release mechanism The launch stand is required to restrain the plane until the engine is started and settled. The supply lines will provide the restraint force required to resist engine thrust during start up. A tab welded on the supply lines and tabs welded on the engine intake are connected with a 4mm rod. The rod is pulled releasing the coupling and allowing the plane to move under thrust.

Figure 85 - Release tab attached to the intake of the engine, and release tab on the launch stand

Spark Plug wires The ignition circuit requires both an active wire, and a ground wire. The active wire is attached to the head of the spark plug using an alligator clip with external insulation. This is required to prevent sparking against the engine mount. The ground wire is attached to the engine mount with another alligator clip. The use of alligator clips allows the wires to be disconnected remotely, by simply pulling on the wires. Care must be taken to ensure that the spark generator has been 119

Chapter 4 Detailed Design turned off before pulling the wires, to ensure that the connectors do not touch and short the discharge coil.

Supply lines The supply lines are required to provide the engine with propane fuel and compressed air. Additionally, they act as injectors and the restraint system. These lines are 5mm ID steel tubing, with fittings on the inlet end to allow connection to the compressed air and propane gas supplies.

Stand The stand must be able to support the thrust force of the plane during the start up process. Due to the likelihood of starting the plane on a sealed runway, it is necessary to use weights to hold the stand down, and prevent it from moving. The stand structure is shown in Figure 86 below. Bricks or other non-flammable heavy objects are to be placed on the three legs of the structure, preventing it from moving.

Figure 86 - Launch stand for pulsejet flight

120

Section 4.8 Electrical and Electronic Components

4.8 Electrical and Electronic Components The electronic and electrical components installed inside the fuselage include the pump and its related components, and radio equipment.

4.8.1 Pump and related components A Flightworks 200 C fuel pump was used to supply the engine with high pressure liquid fuel. The pump is compact, lightweight and designed for use on an aircraft. The dimensions of the pump are shown in Figure 87.

Figure 87: Flight Works Fuel Pump (FlightWorks Inc. 2008)

Pump battery A separate battery will be used for the pump to ensure that the pump operates reliably without interference during the flightA ‘FlightPower’ 7.4 V (3 cell lithium polymer) battery was selected with a capacity of 3750 mAh according to the packaging. This allowed for multiple flights on a single charge, and ample reserve capacity to avoid excessive voltage drop during discharge of the cells.

Throttle To regulate the voltage of the pump and thereby throttle the fuel supply, an ‘Electronic Speed Controller’ (ESC) was required. The ESC provides the pump with a Pulse Width 121

Chapter 4 Detailed Design Modulated (PWM) signal, allowing variation in pump power. ESCs are lightweight and will form the only link from the radio control circuit to the engine control circuit.

As the pump is DC and requires only a small amount of power thus the smallest available brushed ESC was selected for use

4.8.2 Radio Controller The radio controller selected was based on the number of channels required and ease of programming.

A minimum of five channels were required for the following

functions: •

Throttle

•

Aileron (Left)

•

Aileron (Right)

•

Elevator

•

Nose wheel steering

Two ailerons channels were required as the ailerons were to have separate servos and be programmable for use in unison as ‘flaperons’ to assist take-off. The controller selected was the Spektrum DX7. Although it’s 7 channels exceeded the requirements, it was selected for its more powerful programming capabilities than most 4-5 channel radios, and its 2.4 GHz radio frequency which did not require the use of frequency pins. The radio system was supplied with a receiver and additional antenna that formed part of a loom including a small 4 x AA style NiCad battery package to power the servo motors. This loom was to be installed inside the nose of the aircraft, with extension cables used to connect the servo motors. (Model Flight 2007)

122

Section 4.9 Ducted Fan 4.9 Ducted Fan 4.9.1 Purpose of fan For the first flight tests of the aircraft it was decided that an alternate propulsion system was to be used in order to ensure the reliability of the propulsion. A ducted fan was chosen for this first flight as it had similar propulsive capabilities and it was estimated that the system would have the same weight as the pulsejet engine and fuel system.

4.9.2

Selection of fan system

The fan system that was to be selected needed to have the capacity to produce the same thrust as the pulsejet engine, it also was required to be able to have enough power for 10 minutes of flight whilst having a system weight similar to that of the pulsejet engine system. The fan also was required to be capable of being mounted to the aircraft.

Ducted Fan The fan selected was the compact Schubeler DS-51-DIA HDT electric ducted fan, which has a 90mm inner diameter and 51 cm2 swept area. The fan is manufactured of carbon fibre wrapped around aluminium, and the duct is carbon-fibre sandwiching a honeycomb core. The outer diameter of the fan is sufficiently small that it would fit inside the rounded cut away inside the bulkheads for the pulsejet engine. (SchuebelerJets 2008) The fan can be seen below in Figure 88

Figure 88 - Schubeler Ducted Fan

(Schuebeler-Jets 2008)

123

Chapter 4 Detailed Design Other reasons as to the selection of this ducted fan are that it offers the most static thrust output for the fan size compared to others and that the fan is supplied balanced, which was thought to reduce the set-up time required.

The thrust range of the fan is 10-33+ N, closely approximating the design thrust range expected from the valveless pulsejet engine. The fan can produce up to 38 N static thrust with a suitably low-revving, high torque motor (Schuebeler-Jets 2008), approximating the extended thrust goals for the valveless pulsejet engine.

As the fan is supplied balanced there shouldn’t have be any balancing issues associated with the high angular velocity, however the fan will be tested statically before attempted use in flight to ensure the fan is balanced and to verify the thrust output. The parameters of the ducted fan can be seen below in Table 17 Table 17 - Ducted Fan Parameters

Inner Diameter (mm)

90

Outer Diameter (mm)

94.5

Thrust Range

10-33+ N

Speed Range (rev/min)

35,000 – 45,000+

Maximum motor diameter (mm)

38

Fits motor shaft size (mm)

5

Mass of Fan (kg)

.058

(Schuebeler-Jets 2008)

Electric Motor The motor selected for use was the Lehner 1950/11 D, a brushless electric motor manufactured by Lehner Motoren Technik. Similar motors are shown in Figure 89.

124

Section 4.9 Ducted Fan

Figure 89 - Lehner electric motors

(Ductedfans.com 2008)

This motor was selected on the basis that it would fit inside the motor cavity provided in the fan, it would produce the revolutions per minute required by the fan in order to produce the required thrust and the motor was available.

The parameters of the motor can be seen in Table 18 Table 18 - Parameters of Lehner 1950 Electric Motor

(Fine Design RC 2008)

Mass (kg)

0.355

Maximum Angular Velocity

50000

(rpm) Maximum power input (kW)

3.0

Maximum current input (A)

100

Batteries The batteries selected for use were 2 sets of 5-cell Lithium-Polymer batteries in series for a total of 4 batteries. The battery packs that were selected can be seen below in Figure 90

125

Chapter 4 Detailed Design

Figure 90 - ZIPPY-R battery pack

(Hobbycity.com 2008)

These battery packs in such a configuration were selected based on; the required voltage to operate the motor, the required capacity to fly for 10 minutes, cost & availability and capability for the required current draw.

Voltage equivalent to 10 cells (or 37 volts) were required to operate the fan. However 10-cell battery packs were found to be extremely bulky and difficult to package – they consisted of 2 5-cell packs joined together lengthways and therefore would be difficult to package between the bulkheads in the airframe, subsequently it was thought that two 5-cell battery packs connected in series would be easier to package inside the airframe.

The amount of capacity required for 10 minutes of flight was conservatively estimated based on full throttle settings which indicated a requirement for around 10 Ah. Thus each battery packs in parallel needed to have a capacity of around 5 Ah.

Based on the first three considerations, ‘ZIPPY-R 4800 mAh 25C 5S1P’ battery packs were selected. The mass and dimensions of these packs are 636 grams and 170 x 43x 45 mm respectively. Each battery pack was capable of a claimed maximum current draw of 125 A, so these battery packs are capable of the required current draw. (Hobbycity 2008)

126

Section 4.9 Ducted Fan 4.9.3

Modifications to the airframe for Ducted Fan Testing

A ducted fan requires relatively consistent airflow into the duct. This required the addition of cover to reduce the turbulence induced by the sudden drop from the nose of the aircraft to the pulsejet compartment. A mounting for the ducted fan past this cover was also designed to provide a suitable mounting point without excessive additional drag. The ducted fan mounted in the airframe is shown in Figure 91.

Figure 91 - Ducted fan mounted in the airframe

Ducted Fan Mounting A fan mount was developed based on three considerations: •

Maximum length for a smoothly ramped intake,

•

Minimise the pitching moment contribution of the fan, and

•

Minimise the amount of weight added towards the rear of the airframe.

To ensure the maximum length for a smoothly ramped intake, the fan was mounted towards close to the rear of the airframe. It was positioned just forward, and central of, the two vertical tails. This allowed for an intake with a length of 280 mm, described in the next section.

The pitching moment contribution of the fan was calculated, and it was found that if the fan was positioned above the fuselage then the nose down contribution of the fan due to the high thrust line was such that the elevators did not have enough power to pitch the aircraft up during take-off. Therefore the fan has been positioned as low in the airframe as possible to minimise the pitching the moment contribution. 127

Chapter 4 Detailed Design

Several designs for the actual mount were prepared and considered from plywood. All the designs took into consideration the provided fan mounts, which were supplied with the fan itself. To minimise the amount of weight added at the rear of the airframe, the final mount consisted of two wood ‘tabs’ attached with epoxy to sides of the fuselage with the fan was attached using 4 screws to the suppled mountings. The fan mount can be seen below in Figure 92.

Figure 92 - Ducted Fan Mounting Tabs

Ducted Fan Cover In order to achieve optimal performance from the ducted fan, it is required that the incident airflow is as undisturbed as possible. Ideally, this would be achieved by placing the fan in free stream air. However, due to constraints on the height of the thrust line, this is not a possible option. As such, it was necessary to create a duct that would provide the fan with the best possible quality of incident air. To direct flow into the fan, it was necessary to have a downward sloping cutout, due to the fan being mounted much lower than the top of the fuselage. The lowest angle that could be achieved, utilizing the entire length of the cover, was approximately 2 degrees. This angle was chosen as it was thought that the smallest angle possible would help prevent separation at non-smooth points on the surface, produced in the manufacturing process. The final design is shown in Figure 93.

128

Section 4.10 Final Stability Analysis

Figure 93 - The cover design for the ducted fan

The cover is made from 0.5mm aluminum sheet in order to obtain the lowest weight possible. Initial plans were to cut the cover from foam. However the complexity of the cutting due to the presence of the bulkheads and longerons made the use of metal sheet more suitable. The cover is secured with 6 screws along the lower side edges, as well as sliding in under the fan duct. The cover installed on the airframe is shown in Figure 94.

Figure 94 - The cover installed on the plane

4.10

Final Stability Analysis

A final analysis of the aircraft stability was conducted to ensure the flight quality of the aircraft on the ducted fan.

129

Chapter 4 Detailed Design 4.10.1 Longitudinal Moment Analysis To ensure the aircraft can be trimmed and to take into account the effects of engine power on aircraft stability, the longitudinal moment derivatives of the aircraft were determined using approximations developed by Roskam (1985f). The wing component of this was determined by applying corrections for the aspect ratio, sweep and twist to the aerofoil zero lift pitching moment coefficient. The fuselage component was found by applying Hoak (1978)’s empirical equation and approximating the fuselage as a constant area (this gives an overestimate of the magnitude of this moment which is conservative in determining whether the aircraft is able to be trimmed). The contribution of the tail to this value was neglected as the tail produced no lift at zero aircraft angle of attack (without any control surface deflection). The pulsejet and ducted fan configurations provide a small difference in moment in the power-on configuration due to differing thrust lines and maximum thrust levels. The base moment coefficient at zero angle of attack was hence determined as between -0.109 (pulsejet power on) and -0.107 (power off).

The effective overall aerodynamic centre of the aircraft was then determined for both stick free and stick fixed conditions. The stick fixed, power off value was determined earlier in Section 3.5.2. To determine the stick free position, the effectiveness of the elevator in changing the lift of the tail was first determined. This entailed determination of the effects of changing the aircraft angle of attack on the hinge moment of the elevator, changing the deflection angle on the hinge moment, and changing the elevator deflection ion the angle of attack. These were determined through empirical and statistical methods developed by Roskam (1985f). The effects of changing the angle of attack of the aircraft on the thrust produced by the ducted fan were then approximated by using Roskam (1985f)’s empirical approach, with a ducted fan approximating a propeller in this respect. The effects of the normal force produced by the ducted fan on the aerodynamic centre was also calculated but was so small that it could have been neglected. The thrust produced by the pulsejet is relatively independent of angle of attack and so it has a negligible effect on the Cmα term being

130

Section 4.10 Final Stability Analysis considered. The aerodynamic centres and centres of gravity under different flight conditions can be seen below in Figure 95. Centre of Gravity and Aerodynamic Centre Excursion Diagram 8 7 6 Stick Fixed,No Power

W(kg)

5

Stick Free, No Power Stick Fixed, Ducted Fan Power

4

Stick Free,Ducted Fan Power Centre of Gravity (Ducted Fan)

3

Centre of Gravity (Pulsejet) 2 1 0 0.00%

20.00%

40.00%

60.00%

80.00%

100.00%

120.00%

%MAC

Figure 95: Centre of Gravity and Aerodynamic Centre Excursion Diagram

With the data on the effectiveness of the elevators calculated for the stick-free aerodynamic centre and the moment at zero angle of attack calculated earlier, trim ability diagrams for the aircraft were drawn for three key configurations. The trim ability diagram for the ducted fan configuration with power off can be seen below in Figure 96.

131

Chapter 4 Detailed Design Cm-Cl graph (Power Off) 2

1.5

Cl

1 Engine Off Engine Off, Elevator -10 degs Engine Off, Elevator +10 degs Alpha=0 Alpha=12 (Stall)

0.5

0 -1.2

-1

-0.8

-0.6

-0.4

-0.2

0

0.2

0.4

0.6

0.8

-0.5

-1 Cm(ducted fan CG)

Figure 96:Cm-Cl Graph (Power On)

This shows that with power off, elevator deflection downwards of 7 degrees is required to trim the aircraft at maximum lift coefficient. In cruise flight, a smaller elevator deflection is required, in the vicinity of 4 degrees. With ducted fan power on, the trim ability diagram can be seen below in Figure 97. Cm-Cl graph (Power On) 2

1.5

1

Cl

Engine On Engine On, Elevator -10 degs 0.5

Engine On, Elevator +10 degs Alpha=0 Alpha=12 (Stall)

0 -2

-1.5

-1

-0.5

0

0.5

1

-0.5

-1 Cm(ducted fan CG)

Figure 97: Cm-Cl Graph (Power On)

132

Section 4.10 Final Stability Analysis With power on, the ducted fan based aircraft requires around 14 degrees of elevator deflection downwards to trim at maximum lift coefficient. At cruise, the required deflection is 7 degrees with power on. Below in Figure 98 the trim ability diagram for pulsejet flight with the centre of gravity at its most forward point can be seen. This is sufficient as all other centre of gravity positions will produce a trim ability diagram between this and the diagram produced for power off, ducted fan flight (with a small offset to the left in the case of pulsejet power on). Cm-Cl graph (Pulsejet) 2

1.5

Cl

1 Pulsejet Pulsejet, Elevator -10 degs Pulsejet, Elevator +10 degs Alpha=0 Alpha=12 (Stall)

0.5

0 -2

-1.5

-1

-0.5

0

0.5

1

-0.5

-1 Cm(ducted fan CG)

Figure 98: Cm-Cl Graph (Pulsejet)

This shows similar results to under ducted fan, power on conditions as the distance between the centre of gravity and the aerodynamic centre is within several percent of the mean aerodynamic chord. The elevator deflection to trim the aircraft for maximum lift is 15 degrees while the trim deflection at approximately level flight is 7 degrees.

The overall result of this trim ability analysis shows that if manufacturing allows an elevator deflection of at least 15 degrees downwards and 4 degrees upwards (seen from the lower end of the diagrams) then the aircraft will be trimmable and capable of flying at any constant angle of attack the aerofoil is capable of. It also shows that a more efficient aircraft could be obtained by inclining the horizontal tail at an incidence angle equal to around 5 degrees elevator deflection downwards. This has been noted

133

Chapter 4 Detailed Design for future designs, however as manufacturing of the tail had commenced by this point, this alteration could not be performed.

4.10.2 Roll Stability Analysis Before final manufacturing, the aircraft’s roll stability derivative was determined. Although roll stability is not a necessity for any aircraft, if the aircraft is too unstable in roll it will lead to a continually tightening turn until the aircraft is flying at 90 degrees to level. If this occurs too quickly the aircraft will be significantly harder for the pilot to control. The roll stability contributions from the wing-fuselage, horizontal tail and vertical tail were calculated using empirical methods outlined by Roskam (1985f).

The wing and fuselage contribution took into account the wing geometry and the location and size of the wing relative to the fuselage. The sweep and aspect ratio effects produced a stable contribution which slightly outweighed the unstable contribution from the low wing configuration.

The effect of the horizontal tail of the aircraft produced a significant stabilising effect on the roll stability of the aircraft as a whole. This combined with a location above the fuselage with significant sweep and low aspect ratio to produce a stabilising effect around four times larger than the wing. The vertical tails also produced a small stabilising effect due to the coupling of roll with yaw. The contributions of each component can be seen below in Figure 99.

134

Section 4.10 Final Stability Analysis Roll Stability Contributions 0

-0.02

-0.04

ClB

-0.06 Vertical Tails Horizontal Tail

-0.08

Wing and Fuselage -0.1

-0.12

-0.14

-0.16

Figure 99: Roll Stability Contributions

As the overall roll stability derivative of the aircraft is negative, the aircraft is stable in roll, providing for a good test platform. For a future design to better fill a target drone role, an aircraft that is slightly unstable in roll would provide better turning performance but as a proof of concept design, stability in roll is suitable.

4.10.3 Ground Performance The first consideration for suitable ground performance is that the front landing gear must take a reasonable fraction of the aircraft’s weight when on the ground so that steering is effective. The fraction of weight applied to the front and rear landing gears was calculated through a simple force balance between the weight of the aircraft and two reactions at the landing gear locations. For the locations of front and rear landing gears outlined earlier and the centre of gravity point for the ducted fan, the front landing gear takes 22% of the weight of the aircraft. For the pulsejet and full fuel configuration (as it will be on the runway), the front landing gear takes 28% of the aircraft weight. Both these values were within a reasonable range for the landing gear configuration and would allow suitable steering effectiveness.

For the aircraft to be able to rotate for takeoff climb, a suitable pitch angular acceleration must be provided by the elevators at the point where lift equals weight of the aircraft. To calculate this angular acceleration, Equation 27 seen below outlined by Roskam (1985g) was used. 135

Chapter 4 Detailed Design

Equation 27:Pitch angular acceleration

For this analysis the angle of attack of the aircraft was measured as 3 degrees (from the CAD model) and the elevator deflection allowed was only up to 10 degrees (providing a margin of safety). The moment of inertia was estimated from the CAD model available. This enabled the angular acceleration of the aircraft in ducted fan and pulsejet configurations to be determined at the instant of takeoff. In ducted fan configuration, the angular acceleration was 4.3 deg/s2 while pulsejet configuration only provided 1.5 deg/s2.This was a worrying result as these values are significantly lower than those recommended by Roskam (1985g). However, as these are still positive (allowing a rotation upwards) the aircraft will still takeoff, just at a slower climb than ideal. For the pulsejet configuration, the engine may require throttling at the point of takeoff to improve rotational performance (if the thrust is dropped to 1kg then the rotational acceleration is increased to 2.2 deg/s2). Ideally, the location of the rear landing gear would be brought further forward to improve rotation while still allowing suitable front landing gear loads.

The position of ideal balance between requirements would been around 900mm (50mm further forward of the position previously considered) from the nose, providing 17% weight on the front landing gear and 6.3 deg/s2 of rotation in the ducted fan configuration. This would also allow 23% weight on the front landing gear and 3.5 deg/s2 of angular acceleration in the pulsejet configuration, allowing a better takeoff and margin for calculation error. However, as at the stage these calculations were performed the fuselage plans were already finalised and sent away for manufacture with the landing gear location set these modifications could not be made.

136

5 Airframe Manufacture The aircraft design weight was required to be kept to a minimum to allow greater performance from the aircraft. A conscious effort was made throughout the entire design and build process to minimise weight, while still ensuring the structural integrity of the final product. The use of composite materials was chosen, to produce a high strength airframe with minimal weight. All composite work was completed by the group, with the guidance, supervision and assistance mechanical engineering workshop staff as well as aircraft modelling enthusiasts. This working environment was critical to the success of the project on the whole, as all members of the group had little to no manufacturing experience. The airframe was manufactured in a number of sections including the wings, empennage, fuselage and internal structure. On completion of all sections, the airframe was completed through assembly.

5.1 Available Manufacturing Methods The available manufacturing methods were a key driver throughout the entire design process. It was important for the group to know, and be familiar with, the manufacturing techniques that were available to them. It was also important that all designs were able to be manufactured correctly and under budget.

The primary tool available for precisely manufacturing a wide range of parts is a CNC machine. The machine allows CAD images to be imported and cut to shape. A CNC machine was available through the mechanical engineering workshop, and others were sourced external to the university.

The fuselage can be manufactured in two main ways, using either male or female moulds. In either case, the manufacture of a plug is required, to form the desired fuselage shape. The plug will commonly be made from either wood or foam depending on the application. A male mould system allows the final product to be formed straight from the plug. This is a relatively quick process, but results in a poor external surface finish. A female mould system uses the plug to produce a female mould, which is then used to create the final product. This is a longer process, but can produce a final 137

Chapter 5 Airframe Manufacture product with a high quality surface finish. Female moulds also allow parts to be remade quickly.

A number of different techniques were considered for the manufacture the body of the wings. Two techniques commonly used are CNC and hot wire cutting. A primary advantage of hot wire cutting over CNC cutting is the time flexibility available. This would increase the flexibility of the manufacturing process (no machine needed) and would reduce production costs. The disadvantage of wire cutting over CNC cutting is a reduction in cutting accuracy.

The most common technique used for the production of composite structures similar to those being considered in this project, is hand lay-up production. This technique was suitable for applying a fibreglass skin to the surface of the wing core, as well as for constructing the outer structure of the fuselage. Hand lay up is relatively simple, requiring manual addition of a reinforcing fibres and resin. The main reasons for using this technique are that it can be completed in a home workshop, without requiring a high skills base.

A range of other manufacturing approaches and techniques were suggested to the group from a range of sources including modelling enthusiasts and business experts. The principal approaches suggested did not differ greatly from those previously considered. However, numerous ‘tricks of the trade’ designs and techniques were suggested for individual components of the design. These suggestions were considered by the group.

5.2 Wing Construction The wing structure comprised of a foam core, fibreglass skin with carbon fibre and plywood reinforcement. The wings are connected to the fuselage using two unidirectional carbon fibre rods as can be seen in Figure 101.

The core of the wings, made from 24kg/m3 density polystyrene foam, was shaped using a hot wire cutting method. Foam blocks were cut to the appropriate size, with 138

Section 5.2 Wing Construction wing and tip profiles used to achieve the NACA 4412 aerofoil shape. The end profiles were cut from a kitchen laminate, using a CNC machine. The profiles were used to guide the hot wire to achieve the desired wing dimensions, including the taper and sweep.

Due to the nature of hot wire cutting, imperfections were present in the final cut product. The addition of taper made cutting more difficult as it required different cutting speeds along the length of the wing. Two sets of wings were cut, with the best set used for the aircraft and the second set used for fibreglassing practice.

A

lightweight, mouldable plaster paste was used to achieve the final desired aerofoil shape and surface finish.

Three vertical plywood ribs were installed into each wing. One at the root, one at the tip and one just inboard of the aileron. The ribs were made from 4mm plywood and were cut using a CNC machine. The root and centre rib had holes pre drilled, to allow the ribs to hold and support the wing to fuselage connection carbon rods. The foam structure of the wing was cut into two sections, to allow the centre rib to be installed. Two holes were drilled down the length of the wing between the root and centre rib, to accommodate the wing-fuselage carbon rods. A third hole was also drilled to allow servo wiring to pass through the centre of the wing. Epoxy resin was used to re-join the wing structure into a single piece. The sectioned wing is shown in Figure 100.

Figure 100 - Rib Installation in Wings

Reinforcing spars were added to support the basic structure of the wing at 20% and 80% of the root chord. Slots were cut out of the foam core to accommodate the spars

139

Chapter 5 Airframe Manufacture and ensure they were assembled flush to the surface. Epoxy resin was used to bond the spars to the foam core. The full structure of the wings is shown in Figure 101.

Figure 101 - Wing structure schematic

The prepared wings were covered with two layers of 0/90° 85gsm fibreglass cloth. The middle layer of glass was added at 45° to increase the torsional strength of the wings. A 50% volume of resin to fibreglass was maintained to minimise the total volume used. The wings were vacuum bagged during drying to ensure a well bonded and smooth finish, as well as to remove any excess resin. Finally, the wings were heated at 40 degrees for two hours to allow the resin to cure and reach maximum strength.

Figure 102 - Servo Installation

Prior to glassing, balsawood reinforcement was added either side of the aileron hinge line. The ailerons themselves were cut from the main structure of the wing after fibreglassing. Hinges were installed inside the balsa reinforcement, and the required servo, pushrod and servo horn were installed on the bottom side of the wing as illustrated in Figure 102.

5.3 Empennage Construction The entire empennage section, comprising two vertical and one swept horizontal surface, was produced in four separate sections. The cores for each section were hot 140

Section 5.4 Fuselage Construction wire cut from polystyrene blocks, using the same technique as for the wings. However blue foam, of higher density was used, as it was found to be more stable under hotwire cutting. The two horizontal surfaces were attached using epoxy resin to create the horizontal tail section (Figure 103a). The addition of control surfaces and glassing of the remaining three sections was completed in the same way as the wings. The glassed horizontal tail is shown in Figure 103b.

A half aerofoil shape was cut into the ends of the vertical tail sections to allow the horizontal tail to allow a smooth transition from vertical to horizontal sections. The two verticals surfaces were first attached directly to the fuselage (Figure 103c) using an epoxy/micro-balloon mix and the horizontal tail was then attached to the verticals. As the empennage section was permanently fixed to the fuselage, much care was taken to ensure all surfaces were aligned correctly before attachment.

a

b

c

Figure 103 - a) Horizontal tail joined as a single piece, b) horizontatal tail after glassing, c) installation of vertical tail onto fuselage

5.4 Fuselage Construction The final shape of the fuselage was based around the design of the pulsejet engine. It allowed for streamlining of the installed engine, sufficient fuel storage for 10 minutes of flight, as well as room for all other auxiliary systems. The final design was carefully planned to ensure that manufacturing could be completed with minimal complications. Wing flanges were important for both aerodynamic performance and for ease of manufacturing. The smooth transition from fuselage to wings was made to ensure that mould release was able to be achieved without damaging the final 141

Chapter 5 Airframe Manufacture product. A recess was also designed in the under-belly of the fuselage, which allowed the rear landing gear to be mounted flush with the body.

The information in the CAD package was imported into the CNC milling machine, allowing the aircrafts plugs to be shaped accurately. Two plugs were made, one for the port and starboard side of the aircraft fuselage. The plugs were made from Jelutong, which is a timber commonly used for modelling due to its consistent texture, allowing a smooth surface finish to be achieved in minimal time. After machining, each plug was painted and sanded multiple times, until the surface was determined to be sufficiently smooth for producing the moulds. Figure 104 shows the port plug straight after machining.

Figure 104 - Fuselage plug

The first step in creating the moulds from the plugs is to ensure that the moulds will be able to be removed from the plugs after the glass and resin has been applied. This was achieved by applying multiple coats of wax to the plug surface, with a final coat of PVA release agent. Following this, a layer of gel coat (Figure 105) was added over the surface of the plugs, allowing the same smooth surface finish of the plugs to be achieved in the moulds. This coat formed the inside surface of the moulds, which inturn forms the outer surface of the fuselage. Once the gel coat had set, 6 layers of 300gsm chop strand fibreglass mat were added, thereby forming the structure of the mould. The combination of the chop strand mat and vinyl ester resin softened the fibreglass, making it significantly easier to mould to the desired shape. The vinyl ester resin was also used to reduce setting times, compared to those required for most epoxy resins. Once set, the fibreglass moulds were released from the plugs and heated 142

Section 5.4 Fuselage Construction to allow the resin to cure, thereby significantly increasing its strength. Wet and dry sandpaper, as well as surface polish were again used on the mould to obtain the smoothest surface finish possible. Further layers of wax and PVA release agent were then added to the mould surface in preparation for laying up the final fuselage.

Figure 105 - Gel coat being applied to plugs in preparation for creating the moulds

A thin coat of an epoxy resin/micro-balloon mix was chosen for the outer surface of the fuselage for a number of reasons. It provided the smoothest surface finish, helped to fill tight corners to allow the correct shape to be produced and also made the sanding and cleaning of the final surface easier. Onto this coat, one layer of 0/90° 85gsm and then three layers of 0/90° 320gsm aircraft grade fibreglass cloth were added. The middle layer of 320gsm cloth was added at 45° to increase the structures torsional stiffness. A high temperature vinyl ester resin, with glass transition temperature of 177 ⁰C was used in the fuselage to help protect it from the heat produced by the pulsejet engine. Fibreglass rovings were added to the corners of the flanges in the engine cut-away section for increased stiffness.

143

Chapter 5 Airframe Manufacture

Figure 106 - Fuselage

After being released from the moulds, the two halves of the fuselage were joined using a layer of 0/90° 85gsm fibreglass on the outside, and a layer of 0/90° 320gsm Kevlar cloth on the inside. A layer of Kevlar was also applied to the nose of the fuselage for added impact resistance. The resultant fuselage is shown in Figure 106. Although it was expected that significant weight saving could have been achieved through the use of carbon fibre in the fuselage structure, this was avoided to help prevent radio interference between the R/C controller and receiver. Depending on the degree of interference, the radio signal can be lost mid flight. This was an unacceptable risk and hence was avoided.

5.5 Internal Fuselage Construction The primary aim of the fuselage internal structure was to stiffen the structure, while also allowing for a range of component mountings. The three aft most bulkheads were carefully shaped carefully to allow for the installation of the pulsejet and corresponding heat shielding. A total of six bulkheads were added and are shown in Figure 107. Each bulkhead was shaped using a CNC machine from 4mm plywood, covered in two layers of 0/90° 85gsm fibreglass.

The bulkheads were joined to the fuselage structure using an epoxy resin/microballoon mix. The shape and layout of the fuselage was designed to allow fuel bags and 144

Section 5.6 Internal Access batteries to be attached directly to the fuselage itself, removing the need to have additional mounting platforms. The details of each bulkhead are outlined below: -

Bulkheads One and Two were installed to allow for the addition of the electronics tray and to attach the front landing gear. The electronics tray was used as the mounting point for the receiver and power supply for the remote control equipment. The use of the tray allows for vibration isolation to be added later, as is required for pulsejet flight.

-

Bulkhead Three was used as the mounting point for the pulsejet vibration isolation mount, as well as to provide significant structural rigidity to the fuselage, which was weakened due to the engine cut away section.

-

Bulkheads Four and Five were mounted inline with the leading and trailing edge of the wings respectively, adding structural support, while still allowing space for the installation of the pulsejet fuel system components, namely the fuel bags.

-

Bulkhead Six, positioned at the aft end of the aircraft, was a primary point of reinforcement for the two vertical tails.

Holes were drilled in each bulkhead and unidirectional carbon fibre longerons were inserted the length of the fuselage as shown in Figure 107. The longerons provided additional bending strength to the fuse and were secured using epoxy resin.

Figure 107 - Location of bulkheads (blue) and longerons (red)

5.6 Internal Access Due to the design and layout of the pulsejet fuel system, internal access was required for a large portion of the aircraft fuselage. This was achieved with two access areas. 145

Chapter 5 Airframe Manufacture The largest of these two access areas ran over half the length of the fuselage and was required for installation and maintenance of the fuel bags. The fuselage was designed with a cut-away to allow access, requiring a separate cover to make the design as streamlined as possible. The cover was made from 0.5mm aluminium sheeting and was moulded by the mechanical engineering workshop staff. The cover was fixed to the fuselage using a total of 6 screws, three on the port and starboard sides respectively.

The nose hatch was required for access to the electronics tray, GPS logger and front landing gear. A removable panel was therefore cut from the nose of the aircraft to achieve this access. The front two bulkheads were then used as securing points, to allow the hatch to be easily added and removed, with only two screws. Figure 108 shows the aircraft with and without the two access panels installed.

a)

b)

Figure 108 - The aircraft showing the both access panels a) removed and b) attached

5.7 Propulsion System Two propulsion systems were installed at different times on the aircraft. The ducted fan was installed for use during initial flight test, to ensure the flight readiness of the airframe, while the pulsejet system was installed in preparation for final pulsejet flights.

5.7.1 Ducted fan The assembly of the ducted fan motor was critical in achieving successful operation. The rotor-blade-housing system, shown in Figure 109, is manufactured with extremely fine tolerances, with only a fraction of a millimetre separating the blade, spinning at 45,000rpm, from its carbon-fibre housing. The external mounting bracket also required 146

Section 5.7 Propulsion System careful assembly, as even the smallest deviations caused enough deformation of the housing to produce interference with the blades.

The ducted fan was mounted to the rear of the fuselage, as close as possible to the centreline of the aircraft to minimise the pitching moment produced. This position was also chosen in order to ensure the aircraft behaviour is similar to when the pulsejet is used, as the thrust lines are at similar heights. The fan itself was fixed to two mounting blocks, which were inturn fixed to the fuselage. The mounting blocks were held in place with epoxy resin, and the fan was secured to the mounting with four screws, as was allowed for in the design.

Finalising the mounting angle and position of the fan was important to ensure the thrust line of the fan was along the centre line of the aircraft. Therefore, a number of minor adjustments were made to the surface of the mounting blocks during installation so as to achieve the optimal mounting conditions.

Figure 109 – Schubeler ducted fan (Schubeler Jets, 2008)

5.7.2 Pulsejet Installation of a pulsejet engine is achieved through two mounting points, located at the front and rear of the engines. The front mount (Figure 110), located on the end cap of the engine, next to the spark plug, attaches to the third fuselage bulkhead using three screws. This mount provides vibration and heat isolation to the fuselage which is critical for pulsejet operation. The rear engine mount is a tab welded to the exhaust of 147

Chapter 5 Airframe Manufacture the engine, in line with the rear bulkhead. The tab extends to the bottom of the bulkhead, where it is fixed in place with a rubber mount, to damp vibration.

Figure 110 - Front pulsejet engine mount

Mounting of the pulsejet engine also requires the installation of the engines fuel system. This comprises two fuel bags, a pump, ESC, battery and fuel injectors. The fuel bags are soft mounted to the bottom of the fuselage in between bulkheads 3, 4 and 5, 6. The pump, ESC and battery were all installed in between bulkheads 4 and 5, and were again soft mounted to protect them from the engine vibration. Finally, the injectors were able to be secured to the sheet metal which covers the majority of the fuselage.

5.8 Landing Gears and Wheels The rear landing gear was a carbon-fibre unit, Model ‘F3A large’, sourced from Bolly Propellers. The landing gear was attached to the fuselage using five nylon bolts, which were designed to shear on heavy impact, to protect the rest of the aircraft, in particular the tail. The design of the fuselage allowed the landing gear to be mounted flush with the surface of the under-belly.

The basic structure of the front landing gear was a pre-manufactured item, made from spring steel. A hole was drilled into the fuselage to allow the landing gear to be mounted inside the fuselage to the second fuselage bulkhead. A servo, also mounted to the bulkhead, was connected to the landing gear. Once the servo was programmed 148

Section 5.9 Control System Installation into the radio control system, it allowed the aircraft to be safely manoeuvred on the runway. The steering system is shown in Figure 111.

Figure 111 - Front landing gear steering system

5.9 Control System Installation As previously mentioned, allowances were made during manufacture of both the wings and empennage for wiring to actuate the control surfaces. A total of five servos were used, using a total of four channels on the controller. The two wing servos used two channels, allowing them to act both dependently or independently as either ailerons or flaps. The two tail servos used a single channel, as did the front landing gear. All wiring for the servos was connected directly to the radio control receiver, which was mounted to the electronics tray.

The ESC selected was mounted in the centre of the fuselage, between the wings, and was connected directly to the ducted fan, allowing the speed of the fan to be controlled. The ESC and two lithium polymer batteries were all soft mounted to the fuselage, to protect them during high impacts.

149

Chapter 5 Airframe Manufacture

150

6

Testing

Significant testing on both the engine and aircraft systems were conducted. The results can be found in the following sections.

6.1 Engine Testing Three successful testing stages have been completed over the duration of this project. All testing was conducted at ‘Bunker 5’ leased from the Federal Government, near the RAAF Edinburgh. Over the testing period, 100 successful tests on three different pulsejet engines were conducted. Measurements of thrust, and fuel consumption were made at various stages during testing. Footage was recorded via a digital camera, and all data was recorded on a laptop computer.

The testing setup used was developed by previous students for their studies into valveless pulsejets. A schematic of the system layout can be seen Figure 112.

Figure 112 - Test System Layout (Coombes et al 2007)

151

Chapter 6 Testing 6.1.1 Phase One Testing Phase one of testing intended to test the engine control systems and establish baseline performance of the two FWE engines. As part of this, the effect of injector position on engine performance is investigated for both engines. The fixed length FWE, which was previously unable to sustain operation, was tested with the pipe flares developed. All tests were conducted using propane gas for fuel.

Auxiliary system implementation and engine familiarisation Initial testing quickly showed that all auxiliary systems such as the fuel and ignition had been implemented correctly. During the two days of testing, the safety systems and procedures planned and implemented performed as intended, with no safety concerns encountered. The time spent conducting live tests provided the operators with valuable experience about the operational characteristics of valveless pulsejet engines, which could only be learnt during live testing. Initial problems were encountered when attempting to obtain self-sustaining internal combustion with the engines. A rethink of the starting procedure with relation to the physics behind the system allowed a repeatable and reliable starting procedure to be found.

In all tests where self-sustaining operation was achieved, the data logger showed engine thrust decreasing with time, as shown in Figure 113. This was believed to be as a result of the gas bottle cooling, and therefore not being capable of providing the same fuel flow to the engine. This was undesirable and ultimately affected the results. All tests over approximately 30 seconds in length were dramatically affected by this phenomenon. This made it almost impossible to accurately test the throttle range of the engines.

152

Section 6.1 Engine Testing

Figure 113 – Reducing thrust during extended operation

Test the importance of injector position The tests undertaken suggested a relationship between intake position and engine performance. Figure 114 shows a reduction in thrust as the fuel injector was placed further along the length of the intake pipe, closer to the combustion chamber.

Thrust (kg)

Effect of Injector Postion 1.2 1 0.8 0.6 0.4 0.2 0 0

20

40

60

80

100

120

140

Injector Postion (mm)

Figure 114 – Effect of fuel injection position on engine performance

This result is thought to be caused by differences in air-fuel mixing in the engine. With the injector placed closer to the opening of the intake, the fuel and air had a larger time to mix than an injector placed closer to the combustion chamber. More testing was required to confirm this result in later tests. 153

Chapter 6 Testing Effect of flared pipes on valveless pulsejet engine performance The impact the addition of flares had on the engine was clear. Without any flares, the engine was incapable of producing self-sustaining thrust. However, with the addition of the flares, the engine immediately obtained continual operation and thrust. As the design of the flares allowed the lengths of the intake and exhaust pipe to be varied, a range of configurations were able to be tested. With this set-up, the engine obtained self-sustaining thrust in almost every test undertaken, proving the theory behind the flare design, and their importance in pulsejet design. The results from the different configurations however showed very little correlation between the engine performance and the lengths of the exhaust and intake. The erratic behaviour of the engine can be seen in Figure 115. Effect of intake and exhaust length on thrust 1.8 1.6

Thrust (kg)

1.4 1.2 66 series

1

68 series

0.8

78 series

0.6 0.4 0.2 0 0

20

40

60

80

100

120

Exhaust extension (mm)

Figure 115 – Effect of exhaust and intake length on engine performance

The range of thrusts produced by engine 2 varied from as low as 0.7kg up to 1.7kg, which was less than the 1.8kg obtained from engine 1. This was unexpected due to the larger physical size of engine 2. It is believed that the low thrust is a result of poor tuning of the exhaust and intake pipes to the combustion chamber. Comparisons to more successful designs suggests that the large diameter exhaust allows combustion gasses to escape the combustion chamber too easily, and therefore reduces the 154

Section 6.1 Engine Testing intensity of the pressure waves produced. The results from these tests clearly suggest that there is more to engine performance than physical size.

Conclusion The systems implemented for the tests all performed as desired. Cooling of the propane gas bottle was thought to impede the engine performance as the bottle cooled and lost pressure, and no solution to this problem was able to be found in the time available. Some trends were found with regards to injector placement within the intake pipe; however more testing was required to confirm the results. The flares added to engine 2 allowed self-sustaining thrust to be obtained. The flares were thought to enhance the engines ability to produce thrust, even though the engine was not tuned correctly. The poor tuning of engine 2 meant the engine produced less thrust than engine 1, which was smaller in size. The thrust produced by both engines tested was too low to be capable of powering a UAV in its present state. As a result, further engine development was required. A full testing document is supplied in Appendix J.

6.1.2 Phase Two Testing Phase two of testing intended to investigate the performance of the engine modifications, as well as test the performance of the Chinese style engine. The primary modification tested was the expanding tail pipe for the adjustable FWE engine, which was tested at various engine lengths to determine best engine performance. The ceramic coating on the fixed length FWE engine was also tested to determine any changes in performance.

Effectiveness of an expanding tail pipe on an FWE engine After the results of the first round of testing, the adjustable FWE engine configuration failed to produce 2kg of thrust. In order to improve the thrust of this engine an investigation into the effect of an expanding exhaust section was conducted. Tests were conducted to compare the performance of the test engine from 2007, to the 155

Chapter 6 Testing equivalent engine, with an expanding exhaust section. Short lengths of 50mm diameter pipe were cut to allow for fine tuning of the exhaust section.

Figure 116 - The adjustable FWE engine with expanding tail section and 100mm extension.

Initially the expansion section was added to the original FWE engine, with an additional 100mm extension section added to the exhaust, as shown in Figure 116. The engine length was then increased in 50mm and 100mm intervals, to determine the most suitable operating conditions. Based on the results, the most effective length was then selected, and 25mm sections were added and removed in order to find the optimal length of the engine. The results are summarised in Table 19. Table 19 - Expanding Exhaust Test Results Section

Length

(mm) Max Thrust (kg)

100

150

300

325

350

375

400

450

650

1

3

3.25

3.3

3.55

3

3

3.2

2

From the tests it was concluded that a total exhaust length of 1080mm was optimum for this configuration, producing 3.55kg of thrust, achieving both the project goal of producing an engine with 3kg of thrust and extension goal of achieving 3.5kg of thrust. The results suggest that the engine is also less sensitive to changes in exhaust length, which is beneficial from an engine design point of view.

Testing of the expansion section also revealed that the engines were extremely throttle able, with numerous engine configurations still maintaining self sustained combustion to as low as 0.5kg of thrust. This is a significant achievement, as a throttle able engine will give an aircraft more control during flight.

156

Section 6.1 Engine Testing The effect of the intake length was planned as an additional investigation, however during the final testings, a failure of the main fuel cut-off valve occurred, which jeopardised the safety of the remaining tests, hence the tests were aborted.

Test Chinese Engine for thrust measurement The Chinese engine, which was manufactured after the mixed success of the first round of testing, was tested to determine the engines maximum thrust, most effective injector position, and throttle range.

Ten tests were performed on the Chinese engine, with the main aim of determining a suitable injector position for producing maximum thrust. The fuel consumption and throttle range of the engine were also of interest.

Seven different injector positions were tested, with the results shown in Figure 117. It was noticed during testing that as the injector position was moved into the engine, the engine became more throttle able, with the engine throttling as low as 0.8kg on several occasions. The engine was also easier to ‘flame out’, which suggested the fuel consumption of the engine was less, however data could not be collected regarding this due to issues encountered with the load cell.

Thrust vs Injector Position 4

Max Thrust (kg)

3.5 3 2.5 2 1.5 1 0.5 0 0

25

50

75

100

Injector Position from Intake Mouth (mm)

Figure 117 - Affect of injector position on engine thrust

157

Chapter 6 Testing Maximum thrust was achieved with the injector positioned at 32mm inside the engine intake, with a maximum thrust of 3.62kg achieved. The results from this test can be seen in Figure 118.

Figure 118 - Thrust Results

A fuel consumption test was performed at this intake position, yielding a specific fuel consumption of 5kg/kg/hr. It is expected that better results can be achieved by moving the injector position further into the intake.

Effect of Ceramic coating on Engine Performance In an attempt to reduce the temperature of the engine, the interior of the larger FWE engine was ceramic coated by Ceramic Coats Australia. The ceramic coating was intended to reduce heat transfer to the metal, in an attempt to reduce the need for heat shielding on the aircraft.

Four tests were performed on the ceramic coated engine, with thrust results of 1.3kg achieved, which was similar to the results from the first stage of testing. From video footage it could be determined that the engine ran visibly cooler, however it was noticed that damage occurred to the ceramic coating after several runs (Figure 119). It is anticipated that the ceramic coating applied was not suitable for the temperatures 158

Section 6.1 Engine Testing achieved in the pulsejet. Further investigation is required to determine if ceramic coating is a feasible option for reducing the temperatures experienced by the engine structure.

Figure 119 - Visible damage to ceramic coating

Conclusion The second phase of testing was successful, with two engines achieving the projects extended goal of 3.5kg of thrust. Specific fuel consumption figures of 5kg/kg/hr were also promising, as they showed improvement on both results from the studies of Coombs et al, who recorded 7.1kg/kg/hr at 1.6kg of thrust, and Enics pulsejets, which advertise figures of 6.6kg/kg/hr for their engines.

The testing program proved that an expanding exhaust is capable of improving the performance of an engine; however significant work must still be conducted to determine a method for optimizing the length and expansion of these sections.

Finally the tests showed that ceramic coating is an option for reducing the heat transfer out of the engine, with similar thrust results achieved for the engine with and without ceramic coating. However the damage to the coating suggests that more research into the most suitable coating for the engine must also be conducted before this is added to the flight weight engine. The added weight of the engine must also be determined.

159

Chapter 6 Testing A full report of all tests can be found in Appendix H – Test Log Books

6.1.3 Phase Three Testing Phase three of engine testing was aimed at testing the liquid fuelling system developed and readying the engine for use on the UAV platform. Various configurations of injectors, fuels and engines were trialled throughout the testing phase.

Test the performance of the Chinese engine on liquid fuel Initial testing was performed with the injectors placed down the intake in a similar layout as with propane testing. While this setup is not ideal due to both excessive restriction of the intake flow, and poor spray pattern distribution, the layout allowed the modification of the position without altering the engine. This was done to establish a reasonable idea of required injector position. This testing indicated that transition from gas to liquid was not feasible, as the engine would not maintain operation as gas flow was reduced. It was however, possible to start the engine on liquid fuel only, providing that the engine had been brought up to operating temperature on propane.

It was later determined that transitioning could occur with the injectors placed through the side of the intake pipe. It is thought that the inability to do so in initial testing was due to the overheating of the injectors while the engine warms on propane. This would dramatically reduce the flow rate of the injectors to a level that is unable to sustain the engine operation. It was noted from later testing that the fuel flow rates during operation were significantly less than those achieved during flow bench testing, which is most likely also due to injector heating.

The most effective method of startup determined from testing is as follows: •

Turn on ignition circuit and air supply

•

Turn on gas supply system and adjust to start engine

•

Turn off air supply and ignition circuit 160

Section 6.1 Engine Testing •

Allow the engine to warm until the combustion chamber begins to glow (~5 seconds)

•

Set gas flow to medium throttle, turn on liquid at medium throttle

•

Simultaneously reduce gas flow and increase liquid flow, ensuring not to over fuel the engine.

Test the effect of injector position on performance Testing was performed on the Chinese engine to determine the optimal injector position on liquid fuels. Flow requirements developed from equating energy quantities between fuels suggested that 3 injectors were required for operation. The injectors obtained were of various sizes, denoted ‘6’, ‘8’ and ‘10’, with the number relating to the size of the orifice in tenths of a millimeter.

Initial tests established that having the larger injectors closest to the combustion chamber resulted in the best performance, both in thrust and fuel consumption. This is thought to be due to less wastage of the fuel, as the larger injector is further from the mouth of the intake. It is thought that having the larger injector furthermost from the combustion chamber may allow for easier starting, however transitioning to liquid fuel after operation is established on propane makes this irrelevant. The injector layout is shown in Figure 120

Figure 120 - liquid fuel injectors placed mid way along the intake tube

161

Chapter 6 Testing The effect of injector position was investigated by drilling holes along the intake tube at regular intervals, allowing the injectors to be moved as required. The remaining holes were sealed by fireproof fabric with a metal backing held in place with hose clamps.

Best performance (both thrust and fuel consumption) was achieved with the injectors placed midway along the intake tube. This is unlike the propane tests, where maximum thrust was obtained with the injector at the mouth of the intake. It is possible that excessive losses in fuel are the reason for reduced thrust at the outermost position, as the pump and injector set was unable to over-fuel the engine during testing.

A larger pump was purchased in order to provide a larger fuel flow rate. However, this did not provide significant improvements in engine performance. It is believed that both pumps are able to supply the engine with all the fuel it required, but not enough to make the engine flame out. To determine if fuel supply limitations were affecting thrust output, a fourth injector was added to increase fuel flow. This did not produce any more thrust, but did increase fuel usage, and hence it was determined that extra fuel was not required.

In the interests of improving fuel atomization, the injectors were positioned in an opposing configuration, such that the sprays would interfere with each other. This is setup is shown in Figure 121. This also did not produce positive results, and hence it was determined that the best performance was with the three injectors placed midway along the intake tube, on the same side. Comparison of performance with the various injector setups is shown in Figure 122.

162

Section 6.1 Engine Testing

Figure 121 - Opposed injector configuration

Figure 122 - Performance of the Chinese engine with different injector placements

Test the fuel mixtures the Chinese engine In order to achieve the best engine performance, testing was done to determine the best fuel mixture for the engine. Standard unleaded petrol was used as the bas fuel for most testing, with additions of other fuels tested. A single test on kerosene was performed, however, it was noted that the engine produced a lot of fuel vapor during the test. Due to this, and a lack of any performance improvements, kerosene was considered to be unsuitable for the application.

Methanol was trialed as an additive, as on vaporization, a relatively large cooling effect is present. It was thought that this may improve engine performance by increasing the 163

Chapter 6 Testing density and hence mass of air drawn into the engine each cycle. Conversely, the addition of Shellite should increase the heat produced during combustion, and hence potentially improve performance by creating larger combustion pressures.

The best performance was obtained from the addition of 10% methanol to standard unleaded petrol. This addition produced similar thrust, but reduced the fuel consumption of the engine significantly. The relatively consistent thrust levels suggest that the engine is not very sensitive to the degree of variation between the fuels trialed. Engine performance on the trialed fuels is shown in Figure 123.

Figure 123 - engine performance on various fuels

Test the performance of the FWE engine with expanding tail section on liquid fuel Testing was performed to determine the ability of the FWE engine to operate on liquid fuel. The injector setup used was that which provided the best performance with the Chinese engine, and the engine configuration was that which provided the best performance on propane fuel (350mm extension on the tail section).

During the first and only test performed, sustained operation on straight unleaded petrol was achieved, producing a maximum thrust of 1.5kg. During the test a significant amount of un-burnt fuel vapor was released into the air and was therefore

164

Section 6.1 Engine Testing considered a serious safety risk. Hence, no further liquid fuel testing was performed with the FWE engine. The excessive vapor release from the FWE engine on liquid fuels was a factor that had been expected from the background research stage. It is thought that the FWE style combustion chamber does provide the correct environment to support, which was a primary reason for investigation into the Chinese engine style.

Engine Geometry changes Due to the relatively poor performance achieved on liquid fuel, the engine was modified with the aim of obtaining more thrust. To do this, the engine was cut at both the intake and exhaust. The pipe was then either extended by wrapping with metal sheet, or shortened by replacement of the cut section with a smaller piece.

It was found that modifications to the engine generally resulted in a reduction in thrust from the engine. It is believed that this results from the interference on the flow due to the irregularities in the engine walls after modification. This is especially evident when the original configuration was retried and produced less thrust than was obtainable previously. Based on this testing, it is thought that the intake length should be increased slightly over the original length, as the extra 5mm produced approximately 0.5kg of extra thrust compared to the original sized engine after cutting. This is shown in Figure 124, which compares the original performance of the engine with performance after it was cut, as well as the thrust variations with various exhaust lengths. It was also noted that longer exhaust lengths were unable to achieve sustained operation on propane fuel, but would sustain on liquid. Further testing is required in order to determine if a greater extension of the intake will produce better performance, and if the performance degradation post modification can be rectified by re-welding the joins to the original.

165

Chapter 6 Testing

Figure 124 - Performance of the Chinese engine for various lengths

The following theories have been developed to explain the reduced performance of the engine on liquid fuels: •

The slower burn rate of the propane fuel means that combustion is still occurring in the expanding section of the exhaust. This combustion creates a positive pressure gradient, which helps to prevent separation. The expansion angle in the engine is relatively large, so on liquid fuel, where the combustion occurs more so in the combustion chamber, the pressure gradient is not present, and separation occurs, causing large losses in the system. Faster combustion may also increase the speed of the gasses at the expansion point, further increasing the likelihood of separation.

•

Changes in burn rate have affected the mean temperatures in the exhaust, and hence changed the acoustic length. This therefore means the intake is not tuned correctly to the exhaust, reducing engine performance. This theory is derived from the results of the tests after modification of the engine.

Conclusion Phase three of engine testing produced positive results, with operation achieved on both the FWE and Chinese engines. Thrust levels obtained while operating on liquid 166

Section 6.1 Engine Testing fuel were significantly down on those achieved on propane fuel. The Chinese engine was found to run well on a variety of liquid fuel mixtures, with the best performance achieved with the addition of 10% methanol to unleaded petrol. Liquid fuel operation of the FWE engine was obtained, however excessive fuel misting was produced, posing a safety risk.

The optimal injector setup was found to be with three injectors placed midway along the intake, with the larger injectors closest to the combustion chamber. The maximum thrust achieved was 2.25kg, with a thrust specific fuel consumption of 4.8kg/kg/hr.

Modifications to the Chinese engine suggest that improved performance would be achieved by increasing the length of the intake. Further testing is required to confirm this, as a thrust reduction occurred as a result of cutting the engine. Airframe testing has however, shown that 2.2kg of thrust is enough to successfully power the UAV, and as such greater engine performance is not specifically required in order to be implemented for flight. In current form, the engine is capable of fulfilling the goal of flight, however, it is expected that more thrust would be required in order to achieve the speed goals.

A full report of all tests can be found in Appendix H – Test Log Books.

167

Chapter 6 Testing

6.2

Aircraft Testing

The aircraft testing consisted of three phases: ground tests to test components, ducted fan tests and pulsejet tests. The phases of testing are summarised in Figure 125.

Figure 125 - Aircraft testing flow chart

6.2.1

Wing Structural Testing

The wings were tested for structural strength to confirm they were capable of carrying the aircraft’s weight up to the maximum load factor, the deflections of the wings under loading was measured and compared to the calculated values, this test also determine if there were design problems of the wing fuselage interface.

Figure 126 – Load zones for wing structural testing

To ensure the strength of the wings, they were tested with applied loads approximating those expected on the aircraft. This was completed using sandbags applied to the wings with a distribution attempting to match the lift distribution that would be encounter during flight. Loads started at 9kg per wing and increased to 18kg 168

Section 6.2 Aircraft Testing which correspond to a 4.4g manoeuvre. Results of the wing structural test can be seen below in Figure 127.

Figure 127 - Experimental Wing Deflection

The results suggested that all bending could be contributed to the wing beams and the testing rig. Based on calculations including the aircraft stiffness, it was determined that the mounting mechanism was suitable for flight.

6.2.2

Electrical Component Testing

The servo motors that were installed on the aircraft were tested for smooth and consistent operation and deflection using a ‘servo tester’. The servo tester is a connection between a battery source and the servo, with a dial that can be used to turn the position of the servo motor. The servos that were installed worked smoothly and with consistent deflection..

The radio transmitter and receiver were testing twice before the range test at the flight location. Firstly the servos were connected to the receiver circuit outside the airframe, where they operated effectively. The servos and receiver circuit was then installed inside the airframe, where longer extension and junction cables were during the tests the transmitter, receiver and servos again operated effectively. 169

Chapter 6 Testing

The servo motor attached to the nose steering wheel was tested. The motor’s torque was found to be sufficient to turn the front wheel stationary or at speed. The maximum amount of deflection was set, based on estimates of the amount of turn that would provide reasonable but not excessive steering sensitivity. Limits were set for maximum servo deflection on the controller to match the maximum deflection possible on the control surfaces without ‘grinding’. The maximum deflections possible on the control surfaces are shown below in Table 20 Table 20 - Maximum control surface/servo motor deflection

6.3

Surface

Maximum deflection (degrees)

Ailerons

15

Elevators

20

Aircraft Pre-flight Tests 6.3.1

C.G. Test

To ensure that the aircraft was stable in the longitudinal direction, the centre of gravity was experimentally determined, and compared to the expected values from the ‘C.G. Excursion diagram’ shown previously in Section 3.5.2. This required the aircraft to be loaded with the systems required in the airframe for flight, before the airframe was mounted on a ‘pin joint’ that allowed the centre of gravity to be determined when the airframe became statically stable, shown in Figure 128.

Figure 128 - C.G. Test Setup

170

Section 6.3 Aircraft Pre-flight Tests The mounting point could be moved in the longitudinal direction to determine the point at which no pitching moment would result from the distance between the weight force and the reaction at the pinning point.

Figure 129 - C.G. Test Photo

The centre of gravity was determined through this method shown in Figure 129, and it was located at 840 mm from the nose of the aircraft (ducted fan configuration), which equates to 97% of MAC, giving the aircraft a static margin of 10% MAC. This is slightly rearward of the expected position of the C.G. in ducted fan configuration by 41.6 mm.

This provides sufficient stability for an average human pilot. Ballast will be added at the front of the aircraft to provide a margin of error if the aerodynamic centre of the aircraft was away from the expected value in first flights.

However for the flight testing the batteries that are used to power the ducted fan were separated and were moved forward, while a GPS system was added in the nose of the aircraft these changes to the aircraft moved the C.G. forward by approximately 40mm

6.3.2 Other pre-flight checks

Radio Control Range Test At the flight location, to check that radio control was both free from interference at the aircraft and free from interference due to other radio signals, range tests were performed on each flight day.

171

Chapter 6 Testing The transmitter was positioned at up to 100 m from the aircraft, and the throttle control was positioned to 100 % for 2 seconds with the aircraft orientated straight on and at 45 degrees to the left and to the right.

For each orientation the ducted fan ran smoothly, indicating good radio contact between the transmitter and receiver showing a successful range test.

Weight distribution on the landing gear The weight distribution of the aircraft was measured, and it was found that 20 % of the weight of the aircraft was being supported by the nose wheel while 80% was being supported by the main landing gear. This agreed with expected and acceptable values.

Fasteners and Pre-Flight Assembly Before each flight all fasteners were checked, and all other connections were also checked. The servo motors and control surfaces were again checked for operation at the flight location before each flight.

6.3.3

Location for flying

Ducted Fan Tests Ducted fan flight tests were arranged to be conducted at the Gawler airfield where the Adelaide Soaring Club is located. The Soaring Club has provided their airfield for use for several University of Adelaide projects in the past and this aided in the choice of this airfield.

The airfield has runways that exceeded requirements with two 1 kilometre long runways, consisting of a 50 m wide central gravel runway with several long tarmac runways of approximately 5 m width at the sides within the gravel runways. The airfield was available for early morning tests before others needed to use the runway.

Pulsejet Tests Pulsejet flights test were arranged to be conducted at the Adelaide Model Aerosport club airfield at Monarto. Jet models are regularly flown at this airfield.

172

Section 6.3 Aircraft Pre-flight Tests The airfield has a 400 m tarmac runway which exceeds the design take-off. There is a large clear area around the runway providing for excellent visibility.

The loud operation of pulsejet engine restricts use in urban built up areas, however at the Monarto airfield there is no housing nearby, and the adjoining properties include motorcycle and rally racing facilities where loud noise production is not an issue. Contact with the club suggested that Sunday would be the best flying day, once the aircraft was checked for its heavy model certification, which the pilot was able to do.

6.3.4

Pilot

Murray Scott who has over 20 years experience with model aircraft including jet models kindly offered to fly the aircraft for the tests.

The expertise of a

knowledgeable and experienced model aircraft pilot was extremely helpful in completing flight tests.

6.3.5 Engine and Flight Tests

Static Ducted Fan Test The ducted fan was statically tested to check effective operation and to determine the maximum thrust possible. The fan was controlled using the full set-up as intended for the airframe. The test configuration involved the fan mounted on the airframe, the aircraft was attached to a digital strain gauge.

The first objective of the static test of the ducted fan was to check the basic functionality of ducted fan and if the operation was smooth and free from inference. On the first running, the fan ran, and it ran smoothly and appeared to be balanced, suggesting correct assembly and no balancing issues.

Further objectives were to determine the maximum thrust possible to check these against manufacturer’s data, and if possible account for some installation effects. With the ducted fan cover in place, the fan produced 2.2 kg of static thrust on maximum throttle, less than the 3 kg expected. Without the ducted fan cover in place the fan produced 2.4 kg of thrust, which was thought to be due to the fan being able to 173

Chapter 6 Testing access greater air in static running, despite this the ducted fan cover would still be used in flight tests as it would greatly reduce turbulence and interference drag around the fuselage and into the ducted fan in flight.

To determine the frequency reached by the fan, a spectrum analysis was performed which found the fan to be running at a maximum of 37,500 rev/min even on full throttle settings, below the maximum angular velocity of 45,000 rev/min expected. The amount of thrust the fan produced was consistent with the manufacturer’s data for this frequency.

It was determined that the electronic speed controller (ESC) was restricting the maximum frequency of the ducted fan. The ESC was found to limit the with 11 pole motor to 37,500 rpm, confirming the results from the spectral analysis.

Calculations suggested that the aircraft would be able to take-off on 2.2 kg of thrust, and an ESC capable of allowing the motor to reach higher frequency could not be sourced from Australia.

Ground Roll Test The ground roll test was the first test performed at the flight test location. The objective of the ground roll test was to see if the aircraft travelled straight and was controllable on the runway. The pilot was to accelerate gradually and the behaviour of the aircraft was to be observed. The pilot would stop accelerating if the aircraft veered off course. Figure 130 shows this test in progress.

Figure 130 - Ground Roll Test at Gawler Airfield

174

Section 6.3 Aircraft Pre-flight Tests The results of this test suggested that the nose steering was too sensitive, this was therefore reduced and the test was completed successfully.

Touch and Go Test The purpose of a touch and go test is to practise a large amount of landings in a short amount of time. Landings are generally considered the most dangerous part of flying, and are quite risky for a new aircraft. This test, involving repeated landings and takeoff, allows a large amount of information about an aircraft’s dynamics and behaviour to be determined quickly. The test was to be performed at a very low altitude to allow the stability and controllability of the aircraft to be tested in a situation where minimal damage could result in the case of a failure. The touch and go test was attempted on the same day as the ground roll test.

For this first flight test, in order to ensure extra longitudinal stability 0.5 kg of ballast was added to the nose of the aircraft.

The results of the touch and go were that the airplane made one successful take off in the required 70m distance. The plane drifted to the right so power was removed. Subsequently the plane had a heavy landing in grass to the right of the runway. Thus multiple landings were not able to be completed.

Inherent longitudinal stability appeared to be satisfactory – the aircraft did not pitch up as soon as it took off, so this indicates there are no major longitudinal stability issues. The duration of the flight was 7 seconds. It was thought that the plane could have continued climbing and would not have had the heavy landing, if power had not removed. The power was only removed as a touch and go test was being attempted – it was decided the next test would involve an immediate flight quality test attempt.

There was a small amount of structural damage from the heavy landing. Nylon bolts used to attach the main landing gear sheared off during landing. This was a design feature of the aircraft to ensure the aircraft performed a belly landing, in order to protect the tail from structural damage. 175

Chapter 6 Testing Flight Quality Test The objective of the flight quality test was to verify the stability, controllability and manoeuvrability of the airframe. The main result of the test was that the airframe flew successfully,

and

the

airframe

was

stable,

controllable

and

satisfactorily

manoeuvrable. The duration of the flight was 4 minutes.

The method of the test was to perform a take-off, check manoeuvrability in roll by performing figure eights, determine whether the inherent power off stick-fixed pitch response was nose up or nose down, perform some speed runs and then attempt a powered landing on the first flight, and a power off (or gliding) landing on the second flight this was to be done to test whether the aircraft could land with no throttle to determine whether for the pulsejet flight the engine would be able to be switched off during flight to reduce the heat.

A GPS data logger was purchased for use for this flight and it was positioned in the bottom of the nose below the electronics tray. A plot of the flight path taken from the GPS data logger is shown in Figure 131

Figure 131 - Plotted flight path from GPS logger

The data from the flight tests can be seen below in Table 21. Table 21 - Flight data from GPS logger

Duration

4 minutes 176

Section 6.3 Aircraft Pre-flight Tests Maximum altitude

Total distance covered Top Speed

190 m (240 m absolute)

5.61 km 147 km/hr

The results of specific tests and other comments are detailed below:

Take-off The aircraft took off successfully with a take-off distance of 70 m, which was within the expected range. On review of video from the flight, it was found that the amount of rotation of airframe required to lift-off was quite substantial and this was thought to be a function of the rear landing gear. As the position of the landing gear notch was fixed, in order to reduce the likelihood of the airframe tipping over on the ground, the distance from the CG to the main landing gear wheels had been increased by reversing the direction of orientation this gear; this was thought to be the cause of the issue.

The aircraft also veered off to the left from the tarmac part of the runway to the gravel part, this was thought to be due to over sensitive nose gear steering which was a common problem on jet powered model aircraft flying off tarmac runways after being used to grass airfields.

Figure eights & roll The figure eights were performed successfully, first at half aileron deflection and the second at maximum aileron deflection. The lack of a rudder did not hinder the yaw control of the aircraft. The aircraft was at an altitude sufficiently high that it was difficult to observe the aircraft clearly enough to time the roll motion and record the roll-rate.

Pitching moment on power-off

177

Chapter 6 Testing Untrimmed, the aircraft displayed a desirable nose up moment when power was removed. Once trimmed, the aircraft was neutral and didn’t pitch when power was removed.

Top Speed A top speed run was not attempted, however from the GPS the recorded top speed was 150 km/hr.

Landing The landing was quite hard, and the right wingtip hit the ground which broke the central beam between the wings. There was no other structural damage. The heavy landing was not unexpected, as it was only the second landing for the aircraft.

The landing distance was approximately 100 metres and the flaps were not used. It was thought that flaps could be considered for use in the up position to provide washout in order to have less tip stall to assist in a softer landing.

Other comments The glide angle of the aircraft was shallow (which is desirable), and the pilot stated that for much of the flight, between tasks, the aircraft was gliding on low power.

For level flight some roll trim was required (possibly to counter-act the torque exerted by the ducted fan) and the aircraft displayed a self pitch up quality where the pilot ran out of pitch trim on the controller, this may have been wind related.

It was noted that visibility of the airframe was low due to the white and blue (unpainted) colours that blended into the sky. Red, orange or yellow was suggested as preferable colours for visibility with different colours on different surfaces to assist visualisation of the orientation of the aircraft.

178

Section 6.4 Pulsejet Flight Test As noted for take-off, the nose gear steering may have been too sensitive which is a common problem on asphalt runways.

6.4

Pulsejet Flight Test

In order to fly the airframe with the valveless pulsejet engine on liquid fuel the following had been prepared: •

Flight engine demonstrated to produce sustained, continuous, operation with excellent ability to be throttled.

•

Flight engine capable of 3.5 kg of thrust (extended goal) on propane, and 2.25 kg of thrust on liquid fuels.

•

Development of an optimal liquid fuel blend and injectors to best atomise and supply this fuel.

•

A liquid fuel system with fuel bags, connectors, and pump which has been demonstrated to produce continual, uninterrupted fuel flow.

•

A valveless pulsejet and airframe interfacing consisting of an engine mount to resist heat and vibration being transferred, rear engine mount and installation for the fuel system and injectors into the airframe.

•

Airframe demonstrated to be capable of take-off, stable, controllable flight and landing when propelled by a system that produces 2.2 kg of static thrust.

•

Larger version of the flight engine had been designed to produce 3.5 kg of thrust, for extended aircraft goals.

Due to restrictions related to occupational health and safety in being able to start the engine and the impending fire ban season the expected outcomes of a pulsejet powered flight test will be described.

Based on the similar thrust of the alternative propulsion system and the valveless pulsejet, and the continuous operation and ability to be throttled available from the valveless pulsejet, it is expected that the aircraft would be able to fly successfully and based on the fuel consumption and the amount of fuel carried onboard the airframe, it would be capable of sustaining flight for 10 minutes with thrust supplied by this valveless pulsejet engine. The larger engine capable of producing 3.5 kg of thrust 179

Chapter 6 Testing would propel the aircraft to over the goal of 200 km/hr, possibly reaching the extended goal of 250 km/hr. In this way, the pulsejet testing would be split into separate sprint and endurance tests.

6.5

Discussion of experimental results

In this section the testing program with the airframe has been presented. Preliminary tests focused on checking components and determining parameters of the airframe like the centre of gravity and weight distribution across the nose and main landing gear.

Flight tests of the airframe demonstrated that it was capable of smooth, controllable and manoeuvrable flight. The developments prepared to fly the airframe with the valveless pulsejet engine on liquid fuel have been summarised, along with the expected results.

180

7 Management A project management role was initiated during the preliminary stages of the project. The aim of this position was to ensure the project remained on time, on budget and to ensure that all risks were managed appropriately. The project was managed using several tools, including Gantt charts, Microsoft Excel for maintaining control of the project budget, and a risk management strategy, which was utilised to identify, control and monitor all risks associated with the project.

During the initial stages of the project, a basic outline of deliverables and target dates were developed to help create a suitable project timeline. These dates were set as milestones using Microsoft Office Project. As specific goals and tasks were issued, the project plan was developed. The timeline was developed utilising ‘start no later than’ and ‘finish no later than’ constraints. This ensured that most tasks relied on the completion of another, to ensure the project continued to progress successfully.

The group was given an allowance of $2,000 per student from the School of Mechanical Engineering, totalling in $12,000. This was used to cover the manufacturing costs of the project. To determine an approximate cost of the project, a projected budget was prepared. Risk factors were included into the budget to include factors such as rebuilding of the aircraft, express shipping of key components and high manufacturing costs. Based on the results of this study it was determined that more sponsorship was required to ensure the project would be completed successfully. An initial attempt to achieve sponsorship from a number of defence related companies showed reasonable success, with presentations prepared for BAE Systems, and successful recruitment of ASC and Australian Aerospace, who offered the group $2,000 and $1,000 respectively. A grant was also prepared for the Sir Ross and Keith Smith Fund, who generously offered to sponsor the group to the sum of $12,000, this ensured the projects financial stability.

181

Chapter 7 Management The development of a prototype pulsejet powered UAV has numerous risks associated with it. These risks include developmental and manufacturing risks, which affect the projects critical path line, as well as occupational health and safety risks, which are associated with the testing of the engines and handling of fuel. As mentioned earlier risks were also included into the financial plan, in order to determine a suitable project budget.

7.1 Time Management The timeline for the project was managed using Microsoft Project to maintain control of both internal and School set deliverables. Gantt charts were used to determine the allowable length of tasks, as well as the personnel who were responsible for that task.

Time management became exceptionally crucial during the manufacturing stage of the project. Significant delays were experienced during the manufacture of the fuselage, due to the use of an external contractor. This delayed the project by over one month, before it was taken and manufactured at university by the project team with the assistance of Bill Finch.

The Gantt charts used throughout the project can be found in Appendix J- Gantt Charts. An example of the overall project timeline can be seen in Figure 132.

182

Section 7.1 Time Management

ID

Tas k Name

Dur ation

Start

Finish

1

Res earch and Benchm ark ing

2

Pulsejet Backg roun d

31 days? Mo n 3/12/07 Mo n 14/01/08

7

Air craft Research

16 days? Mo n 3/12/07 Mo n 24/12/07

11

Air craft Prelim Design

43 days? Tue 25/12/07 Thu 21/02/08

18

Pulsejet Desig n

82 days? Mo n 21/01/08 Tue 13/05/08

33

Air craft Detaile d De sign

67 days?

43

Pulsejet Testi ng

69 days? Mo n 21/04/08 Thu 24/07/08

48

Manufacturing

135 days ?

Tue 1/04/08 Mo n 6/10/08

49

Win gs

109 days ?

Tue 1/04/08

56

Fus elage

67

Tail

26. 5 days ? Thu 14/08/08

73

Air craft

11. 5 days ?

Fri 19/09/08 Mo n 6/10/08

80

Air craft T estin g

8 d ays?

Tue 7/10/08 Thu 16/10/08

Dec '07 Jan '08 Feb '08 Mar '08 Apr '08 May '08 Jun '08 Jul '08 Aug '08 Sep '08 Oct '08 Nov '08 3 10 17 24 31 7 14 21 28 4 11 18 25 3 10 17 24 31 7 14 21 28 5 12 19 26 2 9 16 23 30 7 14 21 28 4 11 18 25 1 8 15 22 29 6 13 20 27 3 31 days? Mo n 3/12/07 Mo n 14/01/08

Fri 22/02/08 Mo n 26/05/08

Fri 29/08/08

81 days? Tue 27/05/08 Tue 16/09/08 Fri 19/09/08

Figure 132-Compact Gantt Chart

183

Chapter 7 Management 7.2 Financial Management The project required a level of finance in excess of a standard final year project and so additional sources of income were required. As such, additional sponsors were sought to fill the budgetary shortfall. Thanks to the generous support of private companies Australian Submarine Corporation and Australian Aerospace the project received additional inputs of $2000 and $1000 respectively. Despite this capital the project still required a major sponsor to fill requirements and The Sir Ross and Sir Keith Smith Fund’s generous support provided an additional input of $12000 to the project. This enabled a cash budget of $16000 for the project with $12000 in kind support from the University of Adelaide.

As all manufacturing was completed in house much of the cash expense of the project was reduced, however this increased the hours worked by the group members (Table 22), so this was taken into account when producing the final cost analysis of the project. Table 22- Hours Worked By Group Members

Karn Michael Mitchell Nick Ryan Terry Total

Hours 446.5 484 565 799 631 613 3538.5

Total Cost $29,767 $32,267 $37,667 $53,267 $41,646 $40,867 $235,479

The cost breakdown of the aircraft can be seen in Figure 133, This represents the cash spent by the group on components and external manufacturing. In total XXX was spent on workshop manufacture, which can be contributed to time spent manufacturing the fuselage in house.

Including approximate times spent by the students on design manufacture and testing brings the total expenditure of the project cost to $245 000. 184

Section 7.3 Risk Management

$580.30

$1,083.57 Engine Manuf acture and Testing Costs Fuselage Manufacture

$1,920.68

Wing and Tail Manufacturing Ducted Fan $2,353.85

Electronics + Control Systems

$543.66

Figure 133-Cost Breakdown

7.3 Risk Management A risk assessment was made to determine key areas of the project which could have an effect on the final outcome. Based on the analysis a risk management plan was developed based on Australian Standard AS/NZS 4360:1999. The risk register developed can be seen in Appendix K- Risk Register

It can be seen from the risk register that the project has a high level of risk associated with its development. Throughout the project numerous other risks were identified which effected the timeline of the project. The risk register was invaluable in identifying future risks and ensuring they were avoided or mitigated appropriately.

185

Chapter 7 Management

186

8 Conclusion and Future Work This project has aimed to design and build a pulsejet UAV. The goals set for the project were ambitious, and the project has been considered a success. This section contains an analysis of project goals, concerns encountered during the project as well as outlining potential future work and recommendations.

8.1 Review of project goals The various primary and extended goals of the project were the basis of all the work performed. The goals were completed to varying extents, with most uncompleted goals within the capability of the UAV, but not obtained due to various reasons. 1. To modify and manufacture a valveless pulsejet, with the aim of producing 3kg of thrust, with an engine weight of 1.5kg or less. This goal will be quantified by the output received from the thrust measurement stand constructed during the 2007 Project. The development of the Chinese style pulsejet engine occurred over the duration of the project. The resulting engine was capable of producing well over 3kg of thrust, as tested on the thrust measurement stand, The engine weight was also well below the requirement, at just 760gm. 2. Develop a liquid fuelled system for a pulsejet engine and integrate a flight weight version into the UAV design. A liquid fuelling system for the Chinese pulsejet engine was developed over the engine testing period. The fuel system was designed to be integrated into the fuselage, and is capable of providing fuel to pulsejet during all flight stages. 3. Based on the desired pulsejet specifications; design, develop and build a lightweight UAV capable of sustaining flight for 10 minutes with thrust supplied by a valveless pulsejet engine. 187

Chapter 8 Conclusion and Future Work The airframe was designed, developed and built based on the anticipated performance of the pulsejet engine. The airframe was flight tested with a ducted fan in place of the pulsejet, which provided similar thrust and weight characteristics to the pulsejet engine. From this test it was concluded that the airframe is capable of sustaining 10 minutes of flight on the developed pulsejet engine. 4. Achieve a cruise speed of over 200km/h. As measured by onboard GPS or a similar system. Due to limitations in thrust achievable on the ducted fan, this goal was not obtainable. A maximum speed of 147km/h was achieved with only 2.2kg of static thrust, significantly less than the 3.5kg used to produce the performance goals. It is expected that the non-static thrust performance of the pulsejet will allow for a cruise speed of significantly higher than was obtained with the ducted fan. 5. Achieve a flight time of 10 minutes Due to time constraints, an endurance flight was not conducted. The flight test conducted had a duration of over 4 minutes, and was completed with less than 30% of the battery capacity. This goal was not obtained, but is well within the capability of the airframe in its current state. 6. Gain a better engineering perspective on the workings of pulsejets, with the aim of developing different engine design alternatives. The dynamic behaviour of the pulsejet engine was the focus of extensive investigation for the duration of the project. At the completion of the project, a solid understanding of the behavioural aspects of the engine has been developed, which was used to develop the performance of the Chinese engine.

188

Section 8.1 Review of project goals 8.1.1 Extension Goals Completion of the ambitious extension goals was also completed to varying degrees over the project. 1. Achieve 3.5kg of thrust from a valveless pulsejet engine. Peak thrust of 3.6kg was achieved from the Chinese engine on propane fuel during testing. 3.5kg of thrust was also achieved from the adjustable FWE engine after the development of the expanding tail section. The performance of both the developed engines is sufficient for the completion of this extension goal. 2. Achieve a cruise speed of 250 km/h or above. Due to limitations in thrust achievable on the ducted fan, this goal was not obtainable. If a correctly designed duct was implemented on the airframe, and an upgraded speed controller was purchased, it is likely that the maximum speed of the UAV would be significantly improved. 3. Increase flight time of the proposed liquid fuelled pulsejet UAV to over 15 minutes. The thrust specific fuel consumption of the pulsejet engine was in the order of 60% of the value used for performance calculations. Considering this, in cruise conditions, it is quite likely that 15 minutes of flight time would be obtainable for the Chinese engine. 4. Manufacture an alternative engine design for future development. The investigations into the Chinese style engine for use on the UAV lead to the manufacture of the alternatively designed engine. In order for improved performance in a UAV application, the engine requires further development, particularly in the area of performance on liquid fuels.

189

Chapter 8 Conclusion and Future Work 8.2 Project Concerns Manufacturing Experience The manufacture of the PENGUIN was completed primarily with composite materials. Due to a lack of previous manufacturing experience, the group were required to either outsource the manufacturing, or obtain guidance from industry professionals. Initially, outsourcing was attempted, however significant delays were incurred, which threatened to prevent the overall success of the project. It was therefore decided by the authors to complete all other manufacturing ”in-house”.

Airframe Weight When working with composite materials, significant experience and care is required to ensure all manufacture is completed within the designed weight allowances. The airframe was designed with a maximum takeoff weight of 8kg. The importance placed on manufacturing the airframe as light as possible, without affecting the structural integrity, resulted in a final takeoff weight of 6.5kg; almost 20% below maximum. This gave the authors flexibility during final installation of componentry, to ensure characteristics of the aircraft such as the centre of gravity were optimised for flight.

Engine Performance The design thrust for the aircraft was set at 3kg for an aircraft takeoff weight of 8kg. The maximum thrust obtained from the flight engine, running on liquid fuel, was 2.25kg; 25% below design. Concerns were initially raised regarding the ability of the aircraft to take off with such power. However, the lower final weight of the airframe inturn reduced the thrust required for takeoff. The ducted fan, installed to test the performance of the airframe, produced a maximum thrust of 2.2kg during static testing. The successful flight achieved on the ducted fan strongly suggested that pulsejet flight would also be successful.

Pulsejet Flight

190

Section 8.3 Future Developments and Recommendations The successful flight of the aircraft using pulsejet propulsion was not achieved during the duration of this project primarily due of a range of occupational health and safety reasons. Significant restrictions were placed on the authors, relating to both the high temperature and noise level produced by the engines, which delayed the final installation and flight using the pulsejet engine. Further restrictions were also encountered due to fire concerns at the airfield, which would require the presence of the Country Fire Service. However, all flight systems were designed in preparation for flight, to minimise delays once all other safety concerns had been dealt with.

8.3 Future Developments and Recommendations This project has shown that a pulsejet powered UAV is a feasible option for target drone applications. There are nonetheless several areas of development for any future designs following on from the work undertaken in this project. These have been split into the areas of airframe design, manufacturing and engine development.

Airframe Design Design of the airframe was significantly limited by the imposition of stall speeds and takeoff distances suitable for wheeled launch. Although this was required to reduce the complexity of what was a proof of concept aircraft, significant improvements in cruise performance could be achieved through a launching system and alternate landing method such as a parachute. These would allow a lower maximum lift aerofoil to be chosen and for maximum cruise speed to occur closer to the maximum L/D for the aircraft and hence significantly reduce drag. The aircraft had a large longitudinal centre of gravity range, depending on fuel loads. This could be reduced by pressurising the fuel or allocating more vertical space for fuel bags. This would enable a smaller tail and hence more aerodynamically efficient aircraft. The effect of downwash on the tail design was significant, requiring a modified T-tail to move the horizontal tail away from the downwash of the wing. The relatively aft location of the wing due to the rearward engine weight also contributed to this. Considering the wing location, a canard design should be considered for any future 191

Chapter 8 Conclusion and Future Work designs with a similar engine and fuel location, improving aerodynamic efficiency and reducing downwash effects. The possibility of including a deployable payload such as flares or chaff should be considered for target drone applications. This would allow a more flexible and effective target drone and could probably be achieved with only a small reduction in cruise speed. An autopilot system could be integrated within the design to improve the aircraft’s usefulness as a target drone. This would allow the PENGUIN to be used without direct control input, improving the safety of use. Any future pulsejet engines composite UAVs should be designed with heat shielding of the airframe a prime consideration in pulsejet mounting. The compromises inherent in this design lead to the tail mounting close to the pulsejet and suitable heat shielding being improvised for the situation. A canard or H-tailed design would have reduced this danger and allowed the aluminium heat shielding to protect the fuselage only. Stress analysis of composite components was learned by the project group during this project. It is recommended that more subjects incorporating this matter are included in the School of Mechanical Engineering curriculum to better prepare students for any projects including composites and their future employment.

Manufacture Fuselage manufacturing was outsourced to a subcontractor and was subsequently met with significant delays. After more than two months, the mould produced by the contractor were taken back in-house and project members completed manufacturing of the fuselage. It is recommended that future projects use in-house manufacturing techniques for any major components to avoid the delays inherent in subcontractor usage.

Engine Development Studies undertaken during this project have shown that augmenters can improve thrust of pulsejet engines by up to 100% without increasing fuel flow required. However, detailed design was not considered applicable to the airframe designed.. 192

Section 8.3 Future Developments and Recommendations Future projects using pulsejet engines should consider the possibility of using augmenters during airframe design due to the significant thrust specific fuel consumption improvements achievable. Further development of liquid fuelling systems to improve the tuning of the engine used and hence improve thrust and efficiency should be considered by future projects. The engines initially tuned for propane used in this project showed a thrust reduction of X% when run on liquid fuels due to differing burn rates. To reduce the impact of engine performance during liquid fuelled operation, an investigation into the operation and installation of a liquid draw propane system should be undertaken. By using a lightweight pressure cylinder, drawing fuel in its liquid state, and using the available engine heat the fuel to its gaseous state, could eliminate the need for a conventional liquid fuel system and the associated reduction in thrust. Ceramic coating of the flight engine should be further investigated to reduce the heat transfer from engine to airframe. This was briefly considered by this project but neglected due to time constraints. Further development into a better theoretical understanding of pulsed combustion, building on the achievements of this project is another area that requires additional development. This would enable better tuning of liquid engines and probable further performance increases. It would be advantageous to develop a system for starting the engine effectively without direct input. This would improve the safety of the aircraft by reducing the risk to operators in the case of a pulsejet failure.

193

Chapter 8

194

References

References Abbott, I, von Doenhoff, A and Stivers, L 1945, Summary of Airfoil Data, National Advisory Committee for Aeronautics (NACA), Washington Abbott et al, 1959, Theory of Wing Sections, Dover Publications Inc. Agency for Toxic Substances and Disease Registry (ATSDR). 1995. Toxicological profile for fuel oils. Atlanta, GA: U.S. Department of Health and Human Services, Public Health Service AMT 1998, Technical info of the KB-70 and the used pulse jet engine, Advanced Micro Turbines, , last viewed 8/5/08 AMT

Netherlands

B.V,

2008,

AMT

Netherlands

retail

pricelist,

Avalakki, N.U, Bannister J, Chartier B.J.J, Downie T.M, Gibson B.A.A, Gottwald C.A, Moncrieff P.I, Williams M.S, 2008 ,ISOAR Search and Resuce UAV, University of Adelaide Beck,

E.,

2005,

Dyna-Jet

Pictures

and

Video,

Beer, F., Johnston, R., & DeWolf, J, 2006. Mechanics of Materials: 4th Edition. US: McGraw-Hill. Blair, G and Cahoon, M 2006, Best bell, Race Engine Technology magazine, September 2006 edition, High Power Media, Somerset Blast Wave Jet Corporation 2004, Plans For Pulsejet Engine, Blast Wave Jet Corporation, last viewed 8/5/08, Caltex Australia Petroleum Pty Ltd, Product Glossary, viewed May 7, 2008. Civil Aviation Saftey Regulations, 1998 Civil Aviation Safety Regulations Section 101, Unmanned Aerial Vehicles and Rocket Operations, Civil Aviation Safety Authority. ConocoPhillips Company, 2004, Glossary of Terms, viewed May 8, 2008.

195

References Coombes, J., Hollands, M., Jones, J., Matthewson, T., and Smith, R., 2007, Level IV Design Project 2007: Design and Build of a Pulsejet Engine and Thrust Measurement Stand Defence Update, 2007, Advanced PulseJet Vertical Lifters, < http://www.defenseupdate.com/events/2007/summary/mdm07_vertical.htm, last viewed 5/5/08 at 12:34pm> Ducted Fans.com 2008, ‘Lehner motors’, United States, viewed 28/10/08, . Eger S 1983, Proektirovanie Samoletov, Moscow. Enics, 2006a, ЭНИКС Продукция, (Translation: ‘ENICS PRODUCTS’), 2006, Enics, 2006b, Воздушная мишень E95M, (Translation: ‘Ariel Target E95M’), 2006, , last viewed on 5/5/08 at 2:39pm Euromodels,

2005,

Jeti

6

Amp

Microprocessor

Motor

Controller,

Experience Festival 2008, Pulse jet engine – History, Experience Festival, last viewed 8/5/08,

Falconet,

2001,

Federal Aviation Authority, 2008, Federal Aviation Regualtions, Section 23 (FAR-23), Fine Design RC 2008, ‘Lehner motors’, United States, viewed 28/10/08, . Goebel,

G

2003,

The

V-1

Flying

Bomb,

Axis

History

Factbook,

, last viewed 8/5/08 Hoak, D et al, USAF Stability and Control Datcom, Flight Control Division, Air Force Flight Dynamics Laboratory, 1978 Hobbycity.com 2008, ‘ZIPPY-R 4800 mAh 25C 5S1P’, United States, viewed 28/10/08, .

196

References Irvine, J, Irvine Aeropulse Labs 2007, Pressurised Liquid Fuel Injection in Valveless Pulsejets Jennings, G 1987, Two-stroke Tuner’s Handbook, HP Books, Tucson Kailasanathan, K, ‘Experimental Investigation of pulsejet engines’ North Carolina State University Mcalley, CT, 2006, ‘Liquid Fuel Development of a Pulsejet Engine’, North Carolina State University Mills, A 1998, Heat Transfer, Prentice Hall, New Jersey Model Flight, 2007, ‘Spektrum DX7’ last viewed 29/10/08 Munson, H. & Young, D. & Okiishi, H. 2006, Fundamentals of Fluid Mechanics, 5th Edition Munson, K, 2001, Jane’s All The Worlds UAV’s, Janes Information Group Naughton, R 2001, The ‘Aerial Target’ and ‘Aerial Torpedo’ in Germany, Monash University,

last

viewed

8/5/08,

Parsch, A., 2003, Directory of U.S. Military Rockets and Missiles, Appendix 1: Early Missiles

and

Drones,

KDH,

Press

Digital,

2007,

eBonTek

DL-3200BT

Bluetooth

GPS

Datalogger,