PFR - Aerospace Engineering Sciences Senior Design Projects ...

University of Colorado Design/Build/Fly 

Buff-2 Bomber 

Aerospace Senior Design Report 

30 April 2009 

Project Final Report 

Jarryd Allison 

Shivali Bidaiah 

Daniel Colwell 

Ross DeFranco 

Mark Findley 

Eric Hall 

Ben Kemper 

Brett Miller 

Customer Dr. Brian Argrow 

Advisor Dr. Donna Gerren 

Advisor Dr. Kurt Maute

Project Final Report – CUDBF April 30 th , 2009 

ASEN 4028: Aerospace Senior Projects 

Table of Contents 

List of Figures ..................................................................................................................... vi 

List of Tables ........................................................................................................................x 

List of Acronyms................................................................................................................. xi 

List of Symbols ................................................................................................................. xiii 

1.0 Project Objectives and Requirements ............................................................................. 16 

1.1 Background ............................................................................................................... 16 

1.2 Project Goal ............................................................................................................... 16 

1.3 Project Objectives ...................................................................................................... 16 

2.0 System Architecture ...................................................................................................... 18 

2.1 Overview of Systems ................................................................................................. 18 

2.2 Competition Missions Concept of Operations ............................................................ 19 

2.2.1 Ground Mission: Assembly ................................................................................. 20 

2.2.2 Flight Mission 1: Ferry Flight ............................................................................. 20 

2.2.3 Flight Mission 2: Surveillance Flight................................................................... 21 

2.2.4 Flight Mission 3: Store Release/Asymmetric Loads ............................................ 21 

2.3 Mechanical and Electrical Design Requirements ........................................................ 22 

2.4 Overall System .......................................................................................................... 22 

2.4.1 Solid Model and Mass Breakdown ...................................................................... 22 

2.4.2 Electrical System Schematics .............................................................................. 24 

3.0 Development and Assessment of System Design Alternatives ....................................... 25 

3.1 Mission Sensitivity Analysis ...................................................................................... 25 

3.2 Aircraft Configuration ............................................................................................... 27 

3.2.1 (System Option #1) Flying Wing......................................................................... 27 

3.2.2 (System Option #2) Canard ................................................................................. 27 

3.2.3 (System Option #3) Conventional ....................................................................... 27 

3.3 Comparison of System Options .................................................................................. 27 

4.0 System Design-To Specifications .................................................................................. 30 

4.1 Aerodynamics Design-To Specifications ................................................................... 30 

4.2 Missions Design-To Specifications ............................................................................ 30 

5.0 Development and Assessment of Subsystem Design Alternatives .................................. 31 

5.1 Aerodynamics Subsystem Design Alternatives .......................................................... 31 

5.1.1 Aircraft Geometry ............................................................................................... 31 

5.1.2 Airfoil Selection .................................................................................................. 32 

5.2 Missions Subsystem Design Alternatives ................................................................... 34 

5.2.1 Wing Store Release Mechanism .......................................................................... 34 

5.2.2 Centerline Store .................................................................................................. 37 

5.2.3 Container ............................................................................................................ 39 

5.3 Propulsion Subsystem Design Alternatives ................................................................ 39 

5.4 Structures Subsystem Design Alternatives ................................................................. 42 

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5.4.1 Wing Construction .............................................................................................. 42 

5.4.2 Wing Span Reduction Method and Joint Location ............................................... 44 

5.4.3 Landing Gear Configuration ............................................................................... 45 

5.4.4 Main Landing Gear Material ............................................................................... 47 

6.0 Subsystem Design-To Specifications ............................................................................. 48 

6.1 Aerodynamics Design-To Specifications ................................................................... 48 

6.2 Missions Design-To Specifications ............................................................................ 48 

6.2.1 Wing Mounted Store ........................................................................................... 48 

6.2.2 Centerline Store .................................................................................................. 49 

6.2.3 Container ............................................................................................................ 49 

6.3 Propulsion Design-To Specifications ......................................................................... 49 

6.4 Structures Design-To Specifications .......................................................................... 50 

6.4.1 Aircraft Wing Requirements ............................................................................... 50 

6.4.2 Landing Gear Requirements................................................................................ 50 

6.5 Avionics Design-To Specifications ............................................................................ 51 

6.5.1 Transmitter ......................................................................................................... 51 

6.5.2 Telemetry System ............................................................................................... 51 

6.5.3 Microcontroller System ...................................................................................... 51 

7.0 Project Feasibility and Risk Assessment ........................................................................ 52 

7.1 Project Feasibility ...................................................................................................... 52 

7.1.1 Weight Budget and Feasibility ............................................................................ 52 

7.1.2 Cost Feasibility ................................................................................................... 52 

7.1.3 Aerodynamic Feasibility ..................................................................................... 52 

7.1.4 Propulsion Feasibility ......................................................................................... 54 

7.1.5 Payload Feasibility.............................................................................................. 56 

7.1.6 Assembly Feasibility........................................................................................... 56 

7.2 Risk Assessment ........................................................................................................ 57 

7.2.1 Aerodynamics ..................................................................................................... 57 

7.2.2 Avionics ............................................................................................................. 58 

7.2.3 Propulsion .......................................................................................................... 58 

7.2.4 Structures............................................................................................................ 58 

7.2.5 Missions ............................................................................................................. 58 

7.2.6 Microcontroller ................................................................................................... 58 

8.0 Mechanical Design Elements ......................................................................................... 60 

8.1 Aerodynamics Mechanical Design Elements ............................................................. 60 

8.1.1 Aircraft Geometry ............................................................................................... 60 

8.1.2 Airfoil Selection and Aerodynamic Twist ........................................................... 61 

8.1.3 Aircraft Incidence Angle ..................................................................................... 62 

8.1.4 Control Surface Sizing ........................................................................................ 62 

8.1.5 Stability Analysis ................................................................................................ 64 

8.1.6 Drag Analysis ..................................................................................................... 72 

8.2 Missions Mechanical Design Elements ...................................................................... 76 

8.2.1 Wing Store Design Element ................................................................................ 76 

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8.2.3 Box Design ......................................................................................................... 79 

8.3 Propulsion Mechanical Design Elements ................................................................... 80 

8.3.1 Motor Selection .................................................................................................. 80 

8.3.2 Propeller ............................................................................................................. 81 

8.4 Structures Mechanical Design Elements .................................................................... 82 

8.4.1 Wing Bending Model .......................................................................................... 82 

8.4.2 Wing Material Selection...................................................................................... 84 

8.4.3 Wing Stress Analysis .......................................................................................... 86 

8.4.4 Folding Wing System .......................................................................................... 88 

8.4.5 Landing Gear Positioning and Stability ............................................................... 89 

8.4.6 Longitudinal and Lateral Ground Stability: ......................................................... 90 

8.4.7 Main Gear Loading Analysis ............................................................................... 92 

8.4.8 Nose Gear Selection ............................................................................................ 95 

8.4.9 Motor Mount....................................................................................................... 96 

8.4.10 Motor Mount Loading Analysis ........................................................................ 97 

9.0 Electrical Design Elements .......................................................................................... 100 

9.1 Propulsion Electrical Design Elements ..................................................................... 100 

9.1.1 Propulsion Electrical Overview ......................................................................... 100 

9.1.2 Propulsion Batteries .......................................................................................... 100 

9.1.3 Electronic Speed Controller .............................................................................. 102 

9.1.4 Wire Gauge ....................................................................................................... 103 

9.1.5 Fuse .................................................................................................................. 104 

9.2 Avionics Electrical Design Elements ....................................................................... 104 

9.2.1 Avionics Electrical Overview ............................................................................ 104 

9.2.2 Payload Release Microcontroller ....................................................................... 105 

9.2.3 Transmitter/Receiver Selection ......................................................................... 108 

9.2.4 Servo Selection ................................................................................................. 108 

9.2.5 Eagle Tree Telemetry Capabilities .................................................................... 109 

10.0 Software Design Elements ......................................................................................... 112 

10.1 Aerodynamic Design Software............................................................................... 112 

10.2 Avionics Microcontroller Software ........................................................................ 113 

11.0 Integration Plan ......................................................................................................... 115 

11.1 Aircraft Overview .................................................................................................. 115 

11.2 Wing Sub-Assembly .............................................................................................. 116 

11.2.1 Wing Assembly............................................................................................... 116 

11.2.2 Vertical Tail Assembly.................................................................................... 116 

11.3 Structures Sub-Assembly ....................................................................................... 117 

11.3.1 Folding Wingtip Assembly.............................................................................. 118 

11.3.2 Landing Gear Assembly .................................................................................. 118 

11.4.1 Motor Mount Assembly .................................................................................. 120 

11.4.2 Battery Assembly ............................................................................................ 120 

11.5 Release Mechanism Sub- Assembly ....................................................................... 121 

11.6 Avionics Sub- Assembly ........................................................................................ 123 

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11.6.1 Receiver Assembly ......................................................................................... 123 

11.6.2 Servo Assembly .............................................................................................. 123 

11.7 Aircraft Assembly ................................................................................................. 123 

12.0 Fabrication and Integration ........................................................................................ 124 

12.1 Interior Sub-Assembly ........................................................................................... 124 

12.2 Exterior Sub-Assembly .......................................................................................... 125 

12.3 Wingtip Sub-Assembly .......................................................................................... 126 

12.4 Main Wing Sub-Assembly ..................................................................................... 127 

12.4 Full System Assembly ........................................................................................... 128 

13.0 Verification and Validation ....................................................................................... 130 

13.1 Subsystem Verification and Validation .................................................................. 130 

13.1.1 Missions Subsystem Verification and Validation ............................................ 130 

13.1.2 Propulsion Subsystem Verification and Validation .......................................... 131 

13.1.3 Structures Subsystem Verification and Validation ........................................... 133 

13.1.4 Avionics Subsystem Verification and Validation ............................................ 137 

13.2 System Verification and Validation ....................................................................... 138 

13.2.1Wingtip Lift Test ............................................................................................. 138 

13.2.2 System Flight Testing ..................................................................................... 138 

13.2.3 System Requirements Not Tested or Verified .................................................. 153 

14.0 Project Management Plan .......................................................................................... 154 

14.1 Organizational Responsibilities.............................................................................. 154 

14.2 Work Breakdown Structure ................................................................................... 155 

14.3 Construction and Testing Schedule Analysis.......................................................... 155 

14.3.1 Buff-2A Construction and Testing Schedule ................................................... 156 

14.3.2 Buff-2B Construction and Testing Schedule ................................................... 156 

14.3.3 Buff-2C Construction and Testing Schedule ................................................... 156 

14.4 Project Budget Analysis......................................................................................... 156 

13.4.1 Wing Budget ................................................................................................... 157 

13.4.2 Propulsion Budget .......................................................................................... 157 

13.4.3 Avionics Budget ............................................................................................. 157 

13.4.4 Missions Budget ............................................................................................. 158 

13.4.5 Travel Budget ................................................................................................. 158 

14.5 Specialized Facilities and Resources ...................................................................... 159 

14.5.1 RECUV Fabrication Lab ................................................................................. 159 

14.5.2 Boulder Aeromodeling Society Airfield .......................................................... 160 

14.5.3 AES Machine and Electronics Shop ................................................................ 160 

15.0 Lessons Learned ........................................................................................................ 161 

15.1 Manufacturing Lessons Learned ............................................................................ 161 

15.2 Testing Lessons Learned ....................................................................................... 161 

15.3 General Lessons Learned ....................................................................................... 161 

16.0 Acknowledgements ................................................................................................... 162 

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16.1 Professional Advisors ............................................................................................ 162 

16.2 Graduate Advisors ................................................................................................. 162 

16.3 Undergraduate Assistants ....................................................................................... 162 

16.4 Student Assistance ................................................................................................. 162 

16.5 Experienced RC Advisors ...................................................................................... 162 

16.6 Sponsors ................................................................................................................ 162 

17.0 References ............................................................................................................... 163 

18.0 Appendix: .................................................................................................................. 165 

Appendix A: Mission Sensitivity Code .......................................................................... 165 

Appendix B: Geometry for AVL Code ......................................................................... 165 

Appendix C: Wing Geometry Optimization Code .......................................................... 172 

Appendix D: Fit in the Box Code................................................................................... 174 

Appendix E: Weights of Competition Aircraft ............................................................... 175 

Appendix F: Performance Constraint Plot ...................................................................... 176 

Appendix G: Stability Analysis Code ........................................................................... 177 

Centerline Store ......................................................................................................... 177 

Four Stores ................................................................................................................ 179 

Two Stores on same wingtip ...................................................................................... 182 

Appendix H: Lift Distribution Code............................................................................... 185 

Appendix I: Landing Gear Analysis Code ..................................................................... 186 

Appendix J: Wing Loading Whiffle Tree Test .............................................................. 189 

Appendix K: PIC Code ................................................................................................. 190 

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List of Figures 

Figure 1: Centerline Store Payload (Bottle) ......................................................................... 18 

Figure 2: Wing Mounted Store (Rocket) .............................................................................. 19 

Figure 3: Flight Mission Lap Overview ............................................................................... 20 

Figure 4: Flight Mission 1 Profile ........................................................................................ 20 



Figure 7: Project Requirement Breakdown .......................................................................... 22 

Figure 8: Transparent Aircraft Overview ............................................................................. 23 

Figure 9: Aircraft Three-View ............................................................................................. 23 

Figure 10: Overall Aircraft Mass Budget ............................................................................. 24 

Figure 11: Aircraft Electrical Schematic .............................................................................. 24 

Figure 12: Determining Wing Geometry ............................................................................. 32 

Figure 13: Moment Coefficient and Lift Coefficient as a Function of Angle of Attack ........ 33 

Figure 14: Drag Polars for the Top Three Airfoils ............................................................... 33 

Figure 15: FBD and Summary of Equations Calculating Centripetal Force on Wing Stores . 34 

Figure 16: Preliminary Design of Magnetic Release Mechanism ......................................... 35 

Figure 17: Preliminary Drawing of Tab-Spring Payload System .......................................... 36 

Figure 18: Preliminary Drawing of Sliding Trigger Payload System (Left: Loaded; Right: 

Released) ............................................................................................................................ 37 

Figure 19: Metallic Wrap Centerline Store Release Mechanism .......................................... 38 

Figure 20: Isogrid Box ........................................................................................................ 39 

Figure 21: Single Motor Aircraft ......................................................................................... 40 

Figure 22: Dual In-Line Pusher Puller Motors ..................................................................... 40 

Figure 23: Dual Rear Mounted Motors ................................................................................ 41 

Figure 24: Dual Front Mounted Motors ............................................................................... 42 

Figure 25: Wing Construction Method Design Options ....................................................... 43 

Figure 26: Wing Span Reduction Design Options ................................................................ 44 

Figure 27: Wing Fold Joint Location Design Options .......................................................... 45 

Figure 28: Landing Gear Configuration Options .................................................................. 46 

Figure 29: Main Gear Material Comparisons ....................................................................... 47 

Figure 30: Performance Constraint Plot ............................................................................... 53 

Figure 31: Static Thrust Stand ............................................................................................. 55 

Figure 32: Assembly Feasibility .......................................................................................... 57 

Figure 33: Aircraft Geometry .............................................................................................. 60 

Figure 34: The Tip Airfoil (HS520) ..................................................................................... 62 

Figure 35: Location of the Fold on the Wing ....................................................................... 63 

Figure 36: Control Surfaces on the Aircraft ......................................................................... 63 

Figure 37: Longitudinal Stability Modes for the Bottle on the Airplane ............................... 65 

Figure 38: Longitudinal Stability Modes for Four Rockets on the Airplane ......................... 66 

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Figure 39: Longitudinal Stability Modes for Two Rockets on the Airplane .......................... 67 

Figure 40: Lateral Stability Modes for the Bottle on the Airplane ........................................ 69 

Figure 41: Lateral Stability Modes for Four Rockets on the Airplane ................................... 70 

Figure 42: Lateral Stability Modes for Two Rockets on the Airplane ................................... 71 

Figure 43: Velocity Magnitude around the Aircraft with the Bottle ...................................... 73 

Figure 44: Streamlines around the Aircraft with the Bottle .................................................. 74 

Figure 45: Velocity Magnitude around the Aircraft with the Rockets ................................... 74 

Figure 46: Streamlines around the Aircraft with the Rockets ............................................... 75 

Figure 47: Store Center of Gravity and Resulting Moment at Mechanism ............................ 76 

Figure 48: Wing Store Overview Detailing the Store Release Process ................................. 77 

Figure 49: Left: Store-Fixed Metal Tab; Right: Tru-Fire Trigger Assembly ......................... 78 

Figure 50: Isometric view of Centerline Store and Release Mechanism ............................... 79 

Figure 51: Box Isogrid Structure ......................................................................................... 79 

Figure 52: Box Drop Test Analysis Using COSMOSWorks ................................................ 80 

Figure 53: Neu Motor and Gearbox ..................................................................................... 81 

Figure 54: 14 x 7 APC-E Propeller ...................................................................................... 82 

Figure 55: Thin-Walled Ellipse............................................................................................ 84 

Figure 56: Balsa Ashby Chart .............................................................................................. 85 

Figure 57: Composites Ashby Chart .................................................................................... 85 

Figure 58: Wing Displacement Distribution ......................................................................... 87 

Figure 59: Von Mises Stress Distribution ............................................................................ 87 

Figure 60: Wingtip Hinge Design ........................................................................................ 88 

Figure 61: Integral Hinge .................................................................................................... 88 

Figure 62: Balsa Mounting Block ........................................................................................ 89 

Figure 63: Landing Gear Placement ..................................................................................... 90 

Figure 64: Longitudinal Stability ......................................................................................... 90 

Figure 65: Lateral Stability Angle Definition ....................................................................... 91 

Figure 66: Main Gear with Applied Loads .......................................................................... 92 

Figure 67: Beam Deflection Analysis .................................................................................. 93 

Figure 68: Beam Buckling Case .......................................................................................... 93 

Figure 69: Two View of the Landing Gear Structure ........................................................... 94 

Figure 70: COTS Nose Gear Assembly ............................................................................... 95 

Figure 71: Motor and Motor Pylon System .......................................................................... 96 

Figure 72: Motor Mount System Integrated into the Wing ................................................... 97 

Figure 73: Stress Analysis on Motor and Pylon Assembly ................................................... 98 

Figure 74: Maximum Stress on Motor and Pylon Assembly ................................................ 98 

Figure 75: Motor Pylon Strain and Deflection ..................................................................... 99 

Figure 76: Motor Mount Maximum Stresses ........................................................................ 99 

Figure 77: Propulsion Electrical Block Diagram ................................................................ 100 

Figure 78: Battery Pack Overview ..................................................................................... 102 

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Figure 79: Speed Controller .............................................................................................. 103 

Figure 80: 40 Amp Fuse .................................................................................................... 104 

Figure 81: Overall Avionics Diagram ................................................................................ 105 

Figure 82: USB Development Board ................................................................................. 105 

Figure 83: Wiring Diagram .............................................................................................. 106 

Figure 84: Completed Circuit Board .................................................................................. 107 

Figure 85: Circuit Board Schematic................................................................................... 107 

Figure 86: Transmitter and Receiver.................................................................................. 108 

Figure 87: Capabilities of Data Recorder (Seagull Pro Telemetry System) ........................ 110 

Figure 88: Airfoil Selection Flow Diagram........................................................................ 112 

Figure 89: Wing Geometry Determination Flow Diagram ................................................. 113 

Figure 90: Stability Determination Flow Diagram ............................................................. 113 

Figure 91: Drag Calculation Flow Diagram ....................................................................... 113 

Figure 92: Logic to arming microcontroller ...................................................................... 114 

Figure 93: Flowchart for releasing payloads ..................................................................... 114 

Figure 94: Assembly Flow Diagram .................................................................................. 115 

Figure 95: Main Wing Assembly ...................................................................................... 116 

Figure 96: Vertical Assembly ........................................................................................... 117 

Figure 97: Wing Sub-Assembly ....................................................................................... 117 

Figure 98: Folding Wingtip Assembly .............................................................................. 118 

Figure 99: Nose Gear Assembly ....................................................................................... 119 

Figure 100: Bottom View of Right Main Landing Gear .................................................... 119 

Figure 101: Motor Mount Assembly................................................................................. 120 

Figure 102: Battery Assembly .......................................................................................... 121 

Figure 103: Wing Store Release Mechanism .................................................................... 122 

Figure 104: Centerline Store Release Mechanism ............................................................. 122 

Figure 105: Aircraft Assembly .......................................................................................... 123 

Figure 106: Interior Landing Gear and Joiner Plate Assembly ........................................... 124 

Figure 107: Exterior Vertical and Main Gear Assembly .................................................... 125 

Figure 108: Wingtip Interior Sub-Assembly ...................................................................... 126 

Figure 109: Main Joined Wing Assembly .......................................................................... 127 

Figure 110: Full System Assembly .................................................................................... 128 

Figure 111: Inspect of the Isogrid Box Structural Corner after Drop Test .......................... 131 

Figure 112: Competition Battery Packs ............................................................................. 132 

Figure 113: Battery Voltage and Power Over Time ........................................................... 133 

Figure 114: Test Wing with Mounting Apparatus Top View ............................................. 134 

Figure 115: Test Wing with Mounting Apparatus Root View ............................................ 134 

Figure 116: Whiffle Tree Final Design .............................................................................. 135 

Figure 117: COSMOSWorks FEM Model of Tip Displacement ........................................ 135 

Figure 118: Whiffle Tree During Loading ......................................................................... 136 

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Figure 119: Wing Post-Failure ........................................................................................... 136 

Figure 120: Wing Tip and Hinge location Displacement vs. Loading Plot with FEM Model 

Predicted Displacement ..................................................................................................... 137 

Figure 121: PIC testing ..................................................................................................... 138 

Figure 122: Pictures of Buff-2A flight test #1 .................................................................... 139 

Figure 123: Nose gear failures from flight test #2 and #4 ................................................... 140 

Figure 124: Buff-2A motor failure during flight test #3 ..................................................... 140 

Figure 125: Elevator servo travel experienced on flight test #5 .......................................... 141 

Figure 126: Indicated airspeed versus time on flight test #6 ............................................... 142 

Figure 127: Actual versus predicted amp draw on flight test #6 ......................................... 143 

Figure 128: Competition lap flown on flight test #7 ........................................................... 144 

Figure 129: Flight pictures from flight test #8 .................................................................... 144 

Figure 130: Actual versus predicted amp draw on flight test #8 ......................................... 145 

Figure 131: Asymmetric loading for flight test #9.............................................................. 145 

Figure 132: Flight test checklist ......................................................................................... 146 

Figure 133: Frayed wire on NiMH battery pack ................................................................. 147 

Figure 134: Buff-2B airborne during flight test #12 ........................................................... 148 

Figure 135: Battery amp draw and voltage versus time on flight test #15 ........................... 149 

Figure 136: Battery power draw versus time on flight test #15 ........................................... 149 

Figure 137: Buff-2B carrying empty water bottle payload ................................................. 150 

Figure 138: Power draw for full water bottle payload flight ............................................... 151 

Figure 139: Buff-2C after takeoff on mission #1 at DBF competition ................................ 152 

Figure 140: Project Organizational Responsibilities ........................................................... 154 

Figure 141: Work Breakdown Structure ............................................................................ 155 

Figure 142: Predicted (Black) and Actual (Blue) Schedule ................................................ 156 

Figure 143: Project Budget Breakdown ............................................................................. 158 

Figure 144: Costs over Project Life Cycle ......................................................................... 159 

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List of Tables 

Table 1: Top Level Project Requirements ............................................................................ 17 

Table 2: Mission Score Sensitivity Results .......................................................................... 26 

Table 3: Aircraft Configuration Trade Comparison ............................................................. 29 

Table 4: Missions to be Completed by the Aircraft .............................................................. 30 

Table 5: Characteristics of the Top Three Airfoils Analyzed ............................................... 34 

Table 6: Characteristics of Aircraft Geometry ..................................................................... 61 

Table 7: Longitudinal Stability for the Bottle on the Airplane.............................................. 65 

Table 8: Longitudinal Stability for Four Rockets on the Airplane ........................................ 67 

Table 9: Longitudinal Stability for Two Rockets on the Airplane ........................................ 68 

Table 10: Lateral Stability for the Bottle on the Airplane ..................................................... 69 

Table 11: Lateral Stability for Four Rockets on the Airplane ............................................... 70 

Table 12: Lateral Stability for Two Rockets on the Airplane ............................................... 71 

Table 13: Drag Prediction on the Payload Calculated by Hand and in PowerFLOW ............ 72 

Table 14: Predicted System Drag for Flight Missions from PowerFLOW ............................ 75 

Table 15: Motor Selection ................................................................................................... 80 

Table 16: Propeller Options ................................................................................................. 82 

Table 17: Wing Skin Material Comparison.......................................................................... 86 

Table 18: Calculating the Load on Each Strut ...................................................................... 92 

Table 19: Battery Options ................................................................................................. 101 

Table 20: NiMH Battery Selection .................................................................................... 101 

Table 21: Speed Controllers Options ................................................................................. 103 

Table 22: Determination of output from multiplexer ......................................................... 106 

Table 23: Servo Selection for Control Surfaces ................................................................. 109 

Table 24: Servo Selection for External Stores and Nose Gear ............................................ 109 

Table 25: Static Thrust Test Results .................................................................................. 132 

Table 26: Project Budget Breakdown ................................................................................ 159 

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List of Acronyms 

AD: Analog Digital 

AES: Aerospace Engineering Sciences 

AGL: Above ground level 

AIAA: American Institute of Aeronautics and Astronautics 

AMA: Academy of Model Aeronautics 

AVL: Athena Vortex Lattice 

AWG: American Wire Gauge 

BAS: Boulder Aeromodeling Society 

BOTE: Back of the envelope 

CCP: Compare, Capture, Pulse width modulation 

CDR: Critical Design Review 

CFO: Chief Financial Officer 

CFRP: Carbon Fiber Reinforced Polymer 

CG: Center of Gravity 

COTS: Commercial off the shelf 

CUDBF: University of Colorado Design/Build/Fly 

DBF: Design/Build/Fly 

EEF: Engineering Excellence Fund 

EEF: Engineering Excellence Fund 

ENTJ: Extraversion-iNtuitive-Thinking-Judging 

EPS: Expanded polystyrene 

ESC: Electronic speed controller 

ESC: Electronic Speed Controller 

ESTJ: Extroversion-Sensing-Thinking-Judging 

FBD: Free Body Diagram 

FEM: Finite element method 

GFRP: Glass Fiber Reinforced Polymer 

GHz: Gigahertz 

GPS: Global Positioning System 

GPS: Global Positioning System 

INFJ: Introversion-iNtuitive-Feeling-Judging 

INFP: Introversion-iNtuitive-Feeling-Perceiving 

INTJ: Introversion-iNtuitive-Thinking-Judging 

ITLL: Integrated Teaching and Learning Laboratory 

ITLL: Integrated Teaching Learning Laboratory 

LBM: Lattice-Boltzmann Method 

LiPo: Lithium Polymer 

NACA: National Advisory Committee for Aeronautics 

NiCad: Nickel cadmium 

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NiMH: Nickel metal hydride 

PL: Payload 

PRJ: Project 

PWM: Pulse Width Modulation 

RAC: Rated Aircraft Cost 

RC: Remote Control 

RECUV: Research and Engineering Center for Unmanned Vehicles 

RFI: Radio Frequency Interference 

RPM: Revolutions per Minute 

SCF: System Complexity Factor 

SYS: System 

TO: Takeoff 

TOG: Takeoff Ground Distance 

UAS: Unmanned aerial system 

UIUC: University of Illinois Urbana-Champaign 

USB: Universal Serial Bus 

VI: Visual Interface 

WAAS: Wide Area Augmentation System 

xii



List of Symbols 

%W m : Weight Distribution Percent (Main Gear) 

%W n : Weight Distribution Percent (Nose Gear) 

A/C: Aircraft 

a: Semi-minor axis (Structures) 

Amps: Ampere 

AR: Aspect Ratio 

b: Semi-major axis (Structures) 

b: Span (Aerodynamics) 

C d0 : Coefficient of (Parasitic) Drag 

C L : Coefficient of Lift 

c root : Root Chord 

c tip : Tip Chord 

D: Drag Force 

E: Empty (Aerodynamics) 

e: Oswald Efficiency Factor 

E: Young’s Modulus 

f: Flight (Missions) 

ft: Feet 

g: Acceleration due to Gravity 

g: g-loading 

g: gram (Propulsions) 

I: Current (Propulsion) 

i: Imaginary Number 

in: Inch 

I zz : Area Moment of Inertia 

k: Spring Constant 

ksi: kilo-pounds per Square Inch 

L: Length of Strut 

L: Lift Force 

L: Liter 

L: Loading (Missions) 

lb: Pound Force 

l m : Distance between CG and Main Gear 

l n : Distance between CG and Nose Gear 

M: Moment 

mah: milli-Ampere-Hours 

mΩ: milli-Ohm 

n s : Number of Struts 

oz: Ounce 

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oz-in: ounce inch 

P(y): Lift Distribution 

P: Load Applied to Free tip of the Beam 

P: Power (Propulsion) 

P CR : Critical Applied Load 

P m : Main Gear Strut Loading 

P N : Impact Load Normal to the Runway 

P n : Nose Gear Strut Loading 

P s : Frictional Load applied due to Rolling Friction 

psi: Pounds per Square Inch 

q: Dynamic Pressure 

R: Radius 

rad: Radians 

Re: Reynolds Number 

s: Seconds 

S: Wing Area (Aerodynamics) 

sec: Seconds 

S TO : Takeoff Distance 

S TOG : Takeoff Distance 

S v : Winglet Area 

t: thickness (Structures) 

T: Thrust 

t: Time 

TX: Transmitter 

V: Shear Force (Structures) 

V: Velocity 

V: Voltage (Propulsions) 

V CR : Cruise Velocity 

W: Watts (Propulsions) 

W: Weight 

X: Span Location 

X cg : x CG location 

Y = distance along span 

Y cg : y CG location 

Z cg : z CG location 

α: Angle of Attack 

λ: Taper Ratio 

Λ c/4 : Quarter Chord Sweep 

Λ LE : Leading Edge Sweep 

Λ TE : Trailing Edge Sweep 

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λ v : Winglet Taper Ratio 

µ r : Coefficient of Friction 

ξ: Damping Ratio 

π: The Ratio of Circumference of a Circle to the Diameter 

ρ: Density 

ψ: Lateral Tip-over Angle 

ω n : Natural Frequency 

= Displacement 

’ = Displacement Slope 

xv



1.0 Project Objectives and Requirements 

Author: Daniel Colwell 

1.1 Background 

The annual AIAA Design/Build/Fly competition, sponsored by Cessna Aircraft Company and 

Raytheon Missile Systems, provides the opportunity to design an unmanned aerial system 

(UAS). The inter-university competition also provides teams an opportunity to represent their 

school on the international level. The 2008-2009 competition rules require each team to design 

an aircraft capable of completing a simulated surveillance and attack mission. The aircraft will 

carry multiple payloads including a 4 liter simulated fuel tank to provide an extended endurance 

required for surveillance as well as 4 wing stores to model attack capabilities. Teams must also 

store the UAS in a lightweight, low volume container and be able to quickly assemble the 

aircraft to simulate a situation where time and space are limited. 

Dr. Brian Argrow, director of the Research and Engineering Center for Unmanned Vehicles 

(RECUV), has been the customer for CUDBF for the past 7 years. As a club, CUDBF has 

served as a means for aerospace engineering students to vertically integrate students from all 

levels of the curriculum to learn the design process. The CUDBF organization has improved 

each year and is now ready to compete for the first place position. Dr. Argrow’s main goal is for 

CUDBF to compete in the 2008-2009 Design/Build/Fly Competition, learn the practical concepts 

of aircraft design, and integrate underclassmen in order to ensure the future success and survival 

of the program. 

1.2 Project Goal 

The goal of CUDBF is to compete in the annual AIAA Design/Build/ Fly Competition and 

increase the potential for success for future teams. The team will achieve this goal by following 

all rules [1] assigned by the DBF director to pass technical inspection in competition. The team’s 

aircraft will be capable of completing all flight missions. Finally, the team will ensure future 

success by vertically integrating underclassmen into the design and fabrication process. 

1.3 Project Objectives 

The objective of this project is to design, build, test, and verify a remotely controlled aircraft 

capable of entering into the Design/Build/Fly competition. To accomplish this objective, the 

aircraft will be a fixed wing design with performance characteristics capable of completing all of 

the competition flight missions with a minimum range of 9,200 feet. The aircraft shall be able to 

accommodate all mission payloads, including the ability to detach each store independently and 

in any order. Deconstructed, the aircraft shall be able to be stored in at most two 4’x 2’x 2’ 

boxes and weigh no more than 55 pounds. The aircraft shall be capable of achieving a maximum 

take-off distance of 100 feet with an electric propulsion system powered by either NiMH or 

16



NiCad battery cells. Finally, a competition requirement of at least 4 underclassmen will be 

involved in the aircraft design process. Table 1 below shows a tabulated form of these top level 

project requirements. 

Table 1: Top Level Project Requirements 

Requirement Description Parent 

Requirement 

0.PRJ.1 The aircraft shall be designed to pass technical inspection by fulfilling 

Design/Build/Fly rules 

Customer 

0.PRJ.2 All payloads shall be integrated into the aircraft Customer 

0.PRJ.3 A mechanism shall be designed to release store payloads individually Customer 

0.PRJ.4 The aircraft shall fit disassembled into at most 2 containers no bigger Customer 

than 2’x2’x4’ 

0.PRJ.5 The aircraft shall weigh no more than 55 lb AMA 

0.PRJ.6 The aircraft shall have a minimum range of 9,200 feet Customer 

0.PRJ.7 The CUDBF team will incorporate at least 4 underclassmen in design 

activities 

Customer 

Objectives also include learning how to effectively interact and communicate with a team in 

order to achieve a common goal. Students will also learn about engineering design processes 

and manufacturing while also stressing systems integration between subsystems. These lessons 

learned during this project will provide valuable experience upon entering industry. 

17



2.0 System Architecture 

Author: Eric Hall 

Co-Author: Dan Colwell 

2.1 Overview of Systems 

The CUDBF aircraft will be a high performance aircraft capable of completing all competition 

missions in an optimized manner. In order to close the design envelope, the overall system and 

subsystems will approach each design with an iterative process attempting to design a system 

with higher efficiency. 

The competition requires two payloads to be flown in three separate flights. Each payload must 

be capable of being remotely released from the aircraft from the pilot’s transmitter. The first two 

mission flights require a 4 liter water bottle to be flown in configurations where the bottle is 

either flown empty or filled with water. The specific bottle is McMaster-Carr part 4322T6 and 

weighs 0.75 pounds empty. Filled with water, the total payload weight is 9.05 pounds. The 

bottle dimensions are 5.875 inches diameter and 11.25 inches length. The bottle can be observed 

below in Figure 1 [2] . 

Figure 1: Centerline Store Payload (Bottle) 

The payloads required for the third flight mission are four wing-mounted rockets. The specific 

rockets are Estes Patriot Rockets #2056. Each rocket is 21 inches in length and must be ballasted 

to 1.5 pounds. Two rockets must be on each wing half with the inboard rocket 24 inches from 

aircraft centerline and outboard rocket 30 inches from centerline. An Estes Patriot Rocket can be 

observed in Figure 2 [3] . 

18



Figure 2: Wing Mounted Store (Rocket) 

Before the aircraft can be flown at competition, officials from DBF will conduct a complete 

technical inspection of the aircraft to ensure the aircraft meets the design requirements and is 

safe to fly. The aircraft, transmitter, and all payloads will be loaded into the team’s boxes. The 

boxes will be measured to ensure that the maximum dimensions of the box are 2’x2’x4’ and then 

weighed. This weight becomes the rated aircraft cost (RAC) as seen in Equation 1. 

= + + + 

Equation 1: Rated Aircraft Cost 

The battery packs of the propulsion system will be weighed to ensure that no pack weighs more 

than 4 pounds and then will be visually inspected to verify that either NiMH or NiCad battery 

cells are being used. The propulsion electronic lines will be inspected to verify a 40-amp rated 

fuse is installed between the battery packs and the motors. The aircraft will then be loaded with 

the heaviest payload (full bottle) and will be lifted by the wingtips to ensure structural stability. 

The transmitter and receiver system will be activated on the aircraft to properly show adequate 

control of the aircraft’s control surfaces. Then the transmitter will be deactivated to show proper 

fail-safe procedure of the aircraft. The required fail-safe protocol is zero throttle, up elevator, 

right rudder, and right aileron. 

2.2 Competition Missions Concept of Operations 

The concept of operations consists of a ground assembly mission and three flight missions. The 

ground assembly mission must be completed before any flight missions can be attempted. 

During flight missions, the aircraft must adhere to the flight plan shown in Figure 3. The aircraft 

will take off from the starting line in under 100 feet, fly 500 feet downwind and make a 180 

degree turn. The upwind leg of the flight must be at least 1,000 feet and have a 360 degree loop 

before turning downwind back towards the starting line. 

19



Figure 3: Flight Mission Lap Overview 

2.2.1 Ground Mission: Assembly 

The ground mission begins with the aircraft, payloads, and transmitter fully restrained in the box. 

The box will be rotated on all sides to demonstrate restraint contents and will then be dropped 

from a height of 6 inches to show structural integrity. The team will then be timed on 

transitioning the stored aircraft to flight readiness with all payloads. The timed assembly factors 

into the Safety Complexity Factor (SCF) in Equation 2 below. 

= 

 

 

Equation 2: Ground Mission Score 

2.2.2 Flight Mission 1: Ferry Flight 

The ferry flight begins with the aircraft placed on the runway with the empty centerline store 

attached. The aircraft will take-off, fly two laps, and land. This mission profile can be observed 

in Figure 4 below. 

Figure 4: Flight Mission 1 Profile 

The total flight time of the aircraft will be recorded and factored into the mission score. The 

time begins when the aircraft throttles up for take-off and ends when the aircraft passes over the 

starting line in the air. The mission score can be observed in Equation 3. 

20



1 = 

 

 

Equation 3: Flight Mission 1 Score 

2.2.3 Flight Mission 2: Surveillance Flight 

The surveillance flight begins with the aircraft placed on the runway with the full centerline store 

attached. The aircraft will take-off, fly four laps, and land. The mission profile can be seen 

below in Figure 5. 


The mission score for this flight is equal only to the aircraft SCF. Equation 4 below shows the 

flight mission score. 

2 = 


2.2.4 Flight Mission 3: Store Release/Asymmetric Loads 

The store release flight begins with the aircraft on the runway with no payload attached and the 

rocket stores in the box. The time required to load the rocket stores to the aircraft will be used 

within the flight mission score. This relation can be seen in Equation 5 below. 

3 = 

 

 


The aircraft will take-off; fly one lap, then land. On the ground, the aircraft will taxi to a 

specified area and drop a store specified by the DBF officials. The aircraft will again take-off 

and repeat this process, finishing the mission by landing successfully after the fourth lap. This 

mission profile can be observed in Figure 6. 

21




2.3 Mechanical and Electrical Design Requirements 

The flow-down of requirements within the project have been delegated between the five primary 

sub-teams. Based on the needs of the project the primary sub-teams selected have been 

aerodynamics, propulsion, structures, missions, and avionics. Figure 7 below schematically 

illustrates the major requirements to be fulfilled by each sub-team. These requirements will be 

highlighted in detail in each subsystem’s design-to specifications. 

Figure 7: Project Requirement Breakdown 

2.4 Overall System 

2.4.1 Solid Model and Mass Breakdown 

The overall design of the CUDBF aircraft can be observed in Figure 8 and Figure 9 below. 

Figure 8 shows the aircraft transparent view in order to clearly show each subsystem and 

component integration and placement in the aircraft. Figure 9 below shows a classic three-view 

of the CUDBF aircraft with important dimensions highlighted. The integration and installation 

of each component will be highlighted later in this report. 

22



Figure 8: Transparent Aircraft Overview 

Figure 9: Aircraft Three-View 

The overall breakdown of the weights of the aircraft has been divided into individual 

subsystems. The total weight of the aircraft was 7.5 lbs. A visual weight breakdown can be 

observed in Figure 10. 

23



Propulsion, 

2.025 

Aircraft, 

4.225 

Missions, 

0.65 

Avionics, 0.6 

Figure 10: Overall Aircraft Mass Budget 

2.4.2 Electrical System Schematics 

For competition flights, the aircraft electrical system is comprised primarily of the transmitter, 

receiver, microcontroller, and propulsion system. During test flights, a telemetry system will 

also be integrated in order to gather data. The transmitter will be controlled by the pilot at all 

times and will communicate with the receiver onboard the aircraft. The receiver, powered by a 

devoted receiver battery, powers and controls all onboard servos. A microcontroller on the 

aircraft controls the payload release servos. A devoted propulsion battery will control the power 

to the propulsion motors. When integrated, the telemetry system will be powered by the receiver 

battery. 

Figure 11: Aircraft Electrical Schematic 

24



3.0 Development and Assessment of System Design Alternatives 

Author: Shivali Bidaiah 

Co-Author: Ben Kemper 

3.1 Mission Sensitivity Analysis 

In this design problem, an aircraft configuration had to be chosen to accomplish the mission 

goals and objectives. The missions and the scoring for each mission are described in the Concept 

of Operations, Section 2.2. A mission score sensitivity analysis was performed for each flight 

mission to determine the factors needed to weigh the system alternatives. This can be found in 

Table 2. The sensitivity parameters were the assembly time, load time, system weight and flight 

time. 

The assembly time is the time required to assemble the aircraft from the box to a flight ready 

state. This factor can vary based on the type of aircraft configuration chosen. The assembly time 

involves opening the storage container, removing the aircraft, stores, transmitter, and any 

required tools, assembling the aircraft, attaching the stores, returning any used tools to the 

container, and closing the container. 

The load time is the time required to load the payloads onto the aircraft i.e. time to load each of 

the four wing stores and centerline store. This time is independent of the aircraft configuration 

since the time required to load the payload onto the aircraft is dependent on the ground crew. As 

a result, this was not included as a factor in determining the aircraft configuration. 

The system weight is the combined weight of all stores, the aircraft, transmitter, containers, and 

assembly tools. System weight depends on the aircraft configuration, so it was considered to 

select the aircraft configuration. The aircraft flight time is the time for the aircraft to complete 

two laps. This is not included in the aircraft configuration choice. This is because the aircraft is 

required to meet a certain flight speed and therefore a flight time independent of configuration. 

The flight time is built into the performance sizing of the aircraft. 

From the four sensitivity parameters and the accompanying equations based on competition 

score, a set of four partial derivatives were created, one from each parameter. By combining the 

score weighting of each mission, an equation for the total flight score was created (Equation 6). 

A maximum mission score was assigned to each mission. The scores were 50, 75, and 100 for 

missions 1, 2, and 3 respectively. This is how performance at competition will be weighted. The 

total flight score will consist of the sum of the three individual flight scores. 

Equation 6: Total Flight Score 

25



From this equation, it was possible to determine a parameter’s influence on the total flight score 

by creating nominal values for each of the parameters and taking the partial derivative with 

respect to that particular design parameter. These nominal values are based on heuristics and are 

listed as follows: 

• 6 lb aircraft 

• 10 lb container 

• 2 lb transmitter 

• 14 lb payload 

• 10 sec load time 

• 30 sec assembly time 

• 120 sec flight time 

An example of this partial derivative of the flight score with respect to aircraft weight is shown 

in the following equation. 

Equation 7: Partial Derivative of the Flight Score with Respect to Aircraft Weight 

It was determined from the mission score sensitivity analysis that the assembly time of the 

aircraft was most sensitive to the overall score, followed by the load time, the aircraft weight and 

lastly the flight time. The results of the analysis can be seen in Table 2, and allowed the design 

to focus on maximizing the factors that most affect total overall score. 

Table 2: Mission Score Sensitivity Results 

Parameter Assembly Time Load Time Aircraft Weight Flight Time 

Percent Change -7.50 % -3.67 % -1.88 % -0.0488 % 

Order of Importance 1 st 2 nd 3 rd 4 th 

It is important to note that the drag of the aircraft was not counted for as a separate factor simply 

because the aircraft drag and weight are so closely related. More drag implies more thrust is 

needed for the aircraft to complete its mission. More thrust entails more batteries which in turn 

increase the overall system weight. 

26



3.2 Aircraft Configuration 

For this mission, the payload cannot be contained within the airplane. Because of this all 

configurations that have a fuselage were not considered as possible design solutions simply 

because a fuselage is unneeded and unused weight and space. Additionally, configurations that 

have multi-wing designs were eliminated since they add weight and are unnecessary for this 

mission. Possible design alternatives were thus narrowed down to three configurations: flying 

wing, conventional without a fuselage, and canard without a fuselage. The canard and 

conventional designs without a fuselage indicate that a boom replaces the fuselage which 

connect the nose and tail sections. 

3.2.1 (System Option #1) Flying Wing 

Pros: The absence of a tail provides a lighter airframe than other configurations. The lack of 

excess control surfaces creates less drag. Due to the absence of a tail, the aircraft is 

easier to compact and requires less pieces to construct (increases ground mission score). 

Cons: A large effort must be put into the aerodynamic design in order to make this aircraft stable. 

Longitudinally, the aircraft can easily become unstable. Although some sources exist, 

less literature is available on designing a flying wing aircraft. 

3.2.2 (System Option #2) Canard 

Pros: The canard surface at the front of the aircraft provides positive lift, decreasing the lift 

required by the wing. The canard increases lifting efficiency as opposed to decreasing. 

The presence of a canard decreases the time required to design the aircraft to be 

longitudinally stable. 

Cons: The canard construction increases the weight of the aircraft. This added weight makes the 

canard’s wing slightly bigger than the flying wing. The canard surface creates more 

pieces for the aircraft assembly time. 

3.2.3 (System Option #3) Conventional 

Pros: The conventional aircraft requires the least amount of design time and experience to 

design. The tail provides longitudinal stability. 

Cons: The tail introduces two negative factors: weight and negative lift. The excess weight from 

the tail requires a bigger wing to compensate; increasing the weight. The tail’s negative 

lift also increases the size of the wing. This leads to decreased efficiency compared to 

the other designs. The tail creates more pieces for the aircraft assembly time. 

3.3 Comparison of System Options 

To increase the project’s overall competition score, the aircraft must be lightweight, have low 

drag, and be assembled from the aircraft container as quickly as possible. The aircraft’s weight 

and wing area were estimated and aspect ratios were used to determine the drag effects on the 

27



aircraft. Finally the assembly time was estimated based on the numbers of steps required for 

assembly. 

The comparison between aircraft configurations began by defining the effects the additional 

surfaces have on the aircraft. The lift required by the wing and tail must equal the weight of the 

aircraft. The tail lift force from similar sized DBF aircraft is approximately 0.73 pounds 4 . For 

the canard, the surface provides positive lift while the conventional tail provides negative lift. 

The weight of the aircraft was estimated using 6 pounds for wing and avionics weight, 8 pounds 

for payload weight and 0.75 pounds for tail boom construction weight. The wing area was 

calculated by the lift required by the wing. Since the canard and conventional both required 

bigger wing areas than the flying wing and because the flying wing’s weight was only wing 

weight (no tail), it was decided that the canard and conventional aircraft weights needed to be 

increased. The weights of these configurations were increased and new wing areas were 

computed. This process was iterated until the weights and wing surface areas converged. 

Based on a minimum wing span of 5 feet, the aspect ratio was calculated. The aspect ratio is 

inversely proportional to the aircraft’s drag. Therefore, the flying wing and canard were both 

predicted to have similar drag characteristics while the conventional has the most drag. 

The assembly time of the aircraft was determined from the number of motions required to move 

the aircraft from the aircraft container to a flight ready condition. Since each aircraft must have a 

minimum wing span of 5 feet and the maximum box dimension is 4 feet, the wing cannot be a 

single piece. Three options are present: two separate wing halves joined by a spar, two wing 

halves folded by a hinge on the centerline, or folding wingtips. The folding wingtip design 

presents the best option due to the importance of the structural integrity of the aircraft’s center. 

Each configuration must have the folding wing tips, but the canard and conventional also have 

the large tail structure. This would also require some sort of attachment or folding mechanism. 

Additionally, servo connections to the control surfaces would introduce complexity which would 

increase the assembly time. The canard and conventional would therefore have similar assembly 

times, while the flying wing would have a lower assembly time due to its absence of a tail. 

The initial calculations performed for this design selection showed that the flying wing was the 

best suited configuration to optimize the overall competition score as seen in Table 3. 

. 

28



Table 3: Aircraft Configuration Trade Comparison 

Factor 

Factor 

Weight 

Flying Wing Canard Conventional 

Illustration 

Aircraft 

Weight 

Assembly 

Time 

20% 10 (15 lb) 9.49 (15.8 lb) 9.03 (16.6 lb) 

80% 10 (2 Joints) 6.67 (3 Joints) 6.67 (3 Joints) 

Final Score 100% 10 7.24 7.14 

Once the flying wing was selected, much concern was expressed in the inherent instability 

associated with the design of such an aircraft. Much analysis and design must be placed into this 

aircraft such that it is controllable in flight. However, this concern was mitigated with extensive 

analysis outlined in the following sections of this report. 

29



4.0 System Design-To Specifications 

Author: Jarryd Allison 

Co-Author: Ben Kemper 

4.1 Aerodynamics Design-To Specifications 

The aerodynamics subsystem was delegated the responsibility of the initial configuration of the 

aircraft. The Buff-2 Bomber was designed to be a fixed wing aircraft that could also be folded in 

order to fit within the 2ft x 2ft x 4ft storage containers. A flying wing was selected due to its 

small number of surface extremities (limited to 2 wings) that can be folded and deployed 

quickly. The aerodynamics team also selected the configuration that would meet the maximum 

100 ft takeoff requirement while still being able to carry all wing stores along with the external 

bottle/tank full of water. The subsystem also worked on reducing the overall weight of the plane, 

which favors the lightweight design of the flying wing. These requirements dictated the main 

tasks of the aerodynamics subsystem throughout the design process. 

4.2 Missions Design-To Specifications 

The Missions subsystem was then tasked with meeting the system requirements relative to 

holding the wing mounted and centerline stores along with storing the aircraft in the container. 

The aircraft must be able to fly three missions completely, seen below in Table 4. 

Table 4: Missions to be Completed by the Aircraft 

Mission Mission Specifications Scoring 

1 Empty 4L tank, 2 laps Equation 3 

2 Full 4L tank, 4 laps Equation 4 

3 Timed rocket loading, four laps, drop store 

at completion of each lap 

Equation 5 

The importance of the aircraft being quickly deployable with mounted stores that can easily 

attach to the aircraft must be stressed, as assembly time plays a role in each mission score 

calculation. In order to be successful, the aircraft must complete each mission with no stores 

being released while airborne. This presents quite a design problem for the Missions 

subsystems. Each store must release on command, yet must not fall off during flight. Free-play 

must also be minimized so as to not affect the aircraft performance. 

30



5.0 Development and Assessment of Subsystem Design Alternatives 

Author: Ross DeFranco 

Co-Author: Mark Findley 

5.1 Aerodynamics Subsystem Design Alternatives 

5.1.1 Aircraft Geometry 

The only known variable in terms of aircraft geometry was the wingspan, which was determined 

by payload placement requirements to be no less than 5ft. The aircraft geometry affects the 

stability of the aircraft, and therefore, the best sweep and taper were chosen. Additionally, the 

geometry of the aircraft was also driven by the requirement to fit in a 2ft x 2ft x 4ft box. 

Athena Vortex Lattice (AVL) [5] was used in conjunction with MATLAB [6] and the program 

AutoIT [7] to generate different geometry configurations, outlined further in Section 10.1. The 

static margin for every combination of leading edge sweep angle (between 0 and 25 degrees) and 

taper ratio (between 0 and 1.0) was analyzed in increments of 5 degrees and 0.1, respectively. 

MATLAB was used to generate geometry files compatible with AVL for every combination. 

This analysis helped narrow the selection of the optimal sweep angle and taper ratio. The code 

used to perform this analysis can be observed in Appendices B and C. 

This analysis was not performed separately for the quarter chord sweep angle and trailing edge 

sweep angle since they are dependent on the leading edge sweep angle. Figure 12 shows the 

variation in leading edge sweep angle and taper ratio as a function of static margin. 

Figure 12 indicates that sweep angles between 0 and 15 degrees produce a negative static margin 

for every possible taper ratio. A negative static margin indicates that the C.G. of the aircraft is aft 

of the aerodynamic center. As static margin increases, the responsiveness of the aircraft to pilot 

input decreases, but, longitudinal static stability increases. So, a negative static margin refers to 

poor longitudinal static stability, which is undesirable for a flying wing configuration since 

flying wings are inherently unstable. 

31



Static Margin 

0.2 

0.15 

0.1 

0.05 

0 

-0.05 

-0.1 

-0.15 

Static Margin as a function of Taper Ratio and Sweep Angle 

Sweep Angle of 25 o 





-0.2 Sweep Angle of 0 o 

-0.25 

-0.3 

1 2 3 4 5 6 7 8 9 10 

Taper Ratio * 10 -1 

Figure 12: Determining Wing Geometry 

The leading edge sweep angle was narrowed down to between 15 degrees and 25 degrees such 

that a positive static margin was assured. The choice of configuration is discussed in the 

Mechanical Design Elements section for aerodynamics. 

5.1.2 Airfoil Selection 

In order to choose the best suited airfoil for a flying wing configuration, several hundred airfoils 

on the UIUC database as well as the DBF Osborne databases were analyzed. All the airfoils 

were plotted in the airfoil analysis tool XFOIL [8] . The three best airfoils were chosen based on 

their drag polars and variance in coefficient of moment with angle of attack. The main driver for 

the airfoil selection was the coefficient of moment as a function of angle of attack. Since the 

configuration chosen for this aircraft is a flying wing design, the longitudinal static stability of 

the aircraft is very important. Flying wings have the tendency to be unstable in pitch; therefore, 

selecting an airfoil with optimal lift characteristics as well as a coefficient of moment very close 

to zero was extremely important. 

The variation in moment coefficient with angle of attack and the variation of the lift coefficient 

with angle of attack for the three best airfoils are shown in Figure 13. The top three airfoils 

shown are the HS520, Eppler216 and the HS602. 

32



Moment Coefficient and Lift Coefficient as a Function of Angle of Attack 

2 

Re: 844,493 

Mach: .0912 

Moment Coefficient and Lift Coefficient 

C m 

and C l 

, non-dimensional 

1.5 

1 

0.5 

0 

hs520 

e216 

hs602 

-0.5 

-5 0 5 10 15 20 

Angle of Attack, α, degrees 

Figure 13: Moment Coefficient and Lift Coefficient as a Function of Angle of Attack 

Figure 14 shows the drag polars of the top three airfoils. The HS520 and HS602 have very 

similar drag polars. The eppler216 airfoil has a greater increase in drag for the most increase in 

lift. Table 5 summarizes the airfoil characteristics of the top three airfoils. 

2 

Re: 844,493 

Mach: .0912 

Drag Polar 

Coefficient of Lift, C l 


1.5 

1 

0.5 

0 

hs520 

e216 

hs602 

-0.5 

0 0.01 0.02 0.03 0.04 0.05 0.06 0.07 0.08 0.09 0.1 

Coefficient of Drag, C d 


Figure 14: Drag Polars for the Top Three Airfoils 

33



Table 5: Characteristics of the Top Three Airfoils Analyzed 

Airfoil % t/c C l max C d (at Cl mx) C m (at cruise) Stall AOA (deg) 

Eppler216 10.40 1.78 0.0446 -0.1925 13.27 

HS520 8.84 1.42 0.0306 -0.0059 13.70 

HS602 10.21 1.38 0.0339 -0.101 12.60 

5.2 Missions Subsystem Design Alternatives 

The design alternatives for the Missions subsystem was broken into three separate elements: the 

wing mounted stores, the centerline store, and the aircraft container. Each of these were vital to 

the mission of the aircraft and its performance at competition. The design process included a 

complete understanding of the score sensitivities and the importance of store load times, an 

analysis of the loads acting on these stores, consideration of multiple design alternatives, testing 

of those alternatives, an educated selection based on the test results, and finally a successful 

implementation of the design choice. Based on established requirements, it was imperative the 

stores load quickly, release reliably, and do not release in flight. 

5.2.1 Wing Store Release Mechanism 

From the sensitivity analysis, the loading of these stores has a large effect on the aircraft’s score 

at competition. This loading time factors into both the System Complexity Factor score of the 

aircraft through its effect on assembly time, and wing store loading time is measured directly for 

the score in Mission 3 (the heaviest weighted mission). It was therefore important that the 

release mechanisms for these stores have the ability to load quickly, and this is a key design 

driver. It was also important the store release on the ground and not while in flight. To this end, 

the release mechanism must release 95% of the time while being able to constrain the store under 

flight conditions and loads. These flight conditions include a 3 g force in the aircraft’s Z-axis 

and a 1.73 g lateral force. These design-to loads were determined from Figure 15. 

Figure 15: FBD and Summary of Equations Calculating Centripetal Force on Wing Stores 

34



This was important information for the brainstorming phase when several conceivable ideas for 

releasing the stores were generated. These included a sliding pin design, a tab-spring design, a 

magnetic design, a rotating platform design, and a sliding trigger design. The sensitivity analysis 

was applied, and the pin and rotating platform designs were eliminated because of their projected 

high load times. The magnetic, spring-tab, and sliding trigger designs were found to be the two 

best designs because they could be loaded the fastest and both stood a high probability of success 

of being able to keep the store secure. 

5.2.1.1 Magnetic Design 

The concept of using magnets to attach the store had several drivers. The first and largest is that 

this design had the potential for putting a large amount of the release mechanism weight into the 

store itself. The store must weigh 1.5 lbs. Therefore, much of the release mechanism could be 

directly integrated into the store itself (rather than into the aircraft). With the magnetic system, 

this works well as the magnets are the heaviest portion of this release mechanism and will be 

mounted to the store. Another advantage of this system is that competition rules stipulate that 

the release mechanism have no moving parts within the store. Magnets are an effective solution 

for this because they are solid state. A concept of this can be observed in Figure 16. 

Figure 16: Preliminary Design of Magnetic Release Mechanism 

Another driver was the amount of free play present in the release mechanism. Free-play would 

be disastrous to the stability of the aircraft as additional forces and vibrations in flight would 

affect all other systems. It was predicted that the magnetic design had the potential of keeping 

the store rigidly attached in all directions. The magnetic attraction to the underside of the wing 

keeps the store restrained in the aircraft’s Z direction. Because magnetic attraction is 

significantly weaker in the shear direction, restraining tabs will be placed on the underside of the 

wing that keeps the store from sliding off by restraining it in the X and Y directions. 

5.2.1.2 Tab-Spring Payload Design 

The tab-spring payload release mechanism was designed in order to deploy two rockets using 

one servo while still allowing fast loading. In order to decrease load time, it was decided that a 

spring system would need to restore a beveled aluminum strip. Therefore, the pin could retract 

35



from the insertion of the rocket and restrain itself automatically. In order to release two stores 

from one servo, it was determined that one servo would control two pins by string. This would 

allow the servo to retract one pin which is experiencing the tension but the other pin would 

remain stationary since no load is applied through compression. The servo actuator would need 

to produce a torque equal to the spring constant times the gap width multiplied by the length of 

the servo arm. A concept of this can be observed in Figure 17. 

Figure 17: Preliminary Drawing of Tab-Spring Payload System 

5.2.1.3 Sliding Trigger Design 

The sliding trigger design is based off of the coordinated movement of two pins to load and 

release the store. A tab with a hole is rigidly fixed to the store. To secure the store to the 

aircraft, this store-fixed tab is inserted into the bottom of the aircraft. As it slides in, it pushes 

against the rotating pin arm (seen in the figure below) which in turn rotates the release pin. 

Further insertion causes the rotating pin to rotate until it is perpendicular to the vertical store tab. 

At this moment, a spring will force the release pin back into its original position. This will 

prevent the rotating pin from coming back to its original position and the store will remain 

secured to the underside of the aircraft. To release the store, a servo actuator pulls the release pin 

out of its locked position and the weight of the store causes the rotating pin to rotate to its 

original position and release the store. A concept of this option can be seen in Figure 18. 

36



Figure 18: Preliminary Drawing of Sliding Trigger Payload System (Left: Loaded; Right: Released) 

This design has the potential to hold a great deal of weight due to the manner in which these pins 

are aligned. It is therefore considered to be very reliable in both securing the store to the aircraft 

and releasing it on cue. Without fixed guides to easily maneuver the store into proper alignment 

into the system, load time is projected to be slightly higher. This mechanism will also require 

tight manufacturing tolerances. However because of both its reliability and projected quick load 

times, this design was selected for the final design. 

5.2.2 Centerline Store 

The expectations for the wing mounted store also apply to the centerline store. The design for 

this store must be able to support the larger load of the filled wattle bottle at 9 lbs. Load time 

remains a major design driver. Multiple designs were considered and three alternatives made the 

final analysis. These designs involve support on both sides of the water bottle for security and 

reduction of free-play. The first two designs involve the tab-spring and sliding trigger concepts 

applied to both ends of the centerline store because it is the weight of approximately 6 wing 

stores. The third design alternative involves that concept of a metallic sheet that is wrapped 

around the bottle and released by pressing the top ends of the sheet inward. Magnets for the 

centerline store were quickly ruled out because it was determined early on that they would not be 

able to support the store weight and load during maneuvering flight. 

5.2.2.1 Forward and Rear Mounting System 

The centerline store may also be secured using either the sliding trigger or tab-spring methods 

discussed for releasing the wing mounted stores. The only difference would be that the 

centerline store would require two of each particular mechanism, one located towards the front 

and one towards the aft of the centerline store because of both its heavy weight and it not having 

a useful contributing moment. The advantage of this is construction reproducibility and 

therefore simplicity (even though it requires an additional release mechanism). Releasing both 

mechanisms simultaneously so that the loaded centerline store does not jam in the wing joiner 

37



will be a design concern but is manageable. A servo placed on the joiner between the two 

release mechanisms could be integrated with tight tolerances so as to release both mechanisms at 

the same instant. This design alternative was selected for the centerline store because of its 

reproducibility with the other wing release mechanisms and its ease to manufacture. 

5.2.2.2 Metallic Wrap 

In this design, the centerline store is supported by a metallic spring tab designed to support the 

load of the full 4L water bottle in all directions. The tab is counter sunk into the store and 

manufactured tight enough to ensure that the bottle will not move under any flight condition. A 

servo is attached to two rakes by a lever arm and a pin. The rakes rest over the tips of the spring 

tabs. Two U-bolts help to keep the rakes in proper postion and alignment over the holes in wing 

joiner. 

Figure 19: Metallic Wrap Centerline Store Release Mechanism 

To load the centerline store, the spring tabs are inserted through the bottom of the aircraft. They 

must be aligned with the cutout holes in the wing root. As the tabs are inserted, they will deflect 

inward due to their tip shape. Once inserted far enough, they will spring back nearly to their 

original outboard locations. They do not spring back all the way out to apply pre-stress to the 

tabs and root, helping to minimize store free-play. To release the centerline store, servo rotates 

and pulls the two rakes overlaying tabs inward. These rakes will then engage the outward edges 

of the spring tabs and pull them inward. The U-bolts over the rakes push and keep the rakes 

from rotating upward as the servo pull them in. This configuration allows the removal of the 

arms on the servo to prevent accidental deployment during flight. 

The sliding trigger release mechanism was chosen for both the wing and centerline stores. After 

being prototyped, it was shown that this mechanism did in fact have the greatest reliability when 

it came to securing the store on the aircraft. It was determined that this function was the most 

38



mission critical design parameter because if the stores detach during flight, that mission would 

be considered a failure at competition. This mechanism also has the ability to load quickly if 

loading guides are placed on the store and aircraft. The sliding trigger design is the most reliable 

mechanism from preliminary design while maintaining an ability to load quickly. 

5.2.3 Container 

It is important that the container be light, be able to be unloaded quickly, and that its contents be 

securely stored and not shift or be damaged after a 6” drop. It was initially determined that the 

best approach would be to use two containers. Due to the large weight of the centerline store 

payload, a small solid box was used to contain the bottle. To prevent large dynamic loads 

experienced during the drop, the bottle was solidly supported on all sides. The second box, 

being the maximum size, contained the aircraft, rocket payloads and transmitter. This container 

was constructed using an isogrid skeleton in order to survive the drop at a light weight. 

Figure 20: Isogrid Box 

5.3 Propulsion Subsystem Design Alternatives 

The primary design choice for the propulsion system was whether to utilize a single motor or 

dual motors. The advantage of using a single motor is the reduced weight of a second motor, 

gearbox, speed controller, and the additional wiring necessary for a second motor. The weight of 

a second motor, gearbox, speed controller, and wiring is approximately 0.5 pounds. A conceptual 

design of the aircraft with a single motor is shown below in Figure 21. 

39



Figure 21: Single Motor Aircraft 

However, in order to get the thrust required to meet the 100ft takeoff distance requirement with a 

single motor, a propeller size of 20 in was necessary. In order to accommodate a 20 in propeller 

with adequate ground clearance, the landing gear would need to be at least 11 in long. This 

would allow for 10 in or half the propeller blade to spin under the wing with 1 in of ground 

clearance. This added landing gear size would require more structure and thus added weight and 

drag from the long landing gear. A system level requirement states that the aircraft must fit 

within a 2 ft x 2 ft x 4 ft box. With 11 in landing gear, an aircraft thickness, and 13 in winglets 

that fold 15 in in from the tip of the wing, it would be impossible to fit the aircraft vertically in a 

box under 2 ft tall. This drove the design of the propulsion system to utilize dual motors. 

Other disadvantages of a single motor include less directional stability due to the large amount of 

p-factor coming off one motor. In order to use a single motor rated for the required power 

necessary, the motor itself would weigh 0.25lbs more. That means the overall weight savings 

from using a single motor in this case would only be 0.25lbs. Because of these characteristics, a 

dual motor configuration was chosen. 

Once dual motors were selected, the exact configuration of the motors was analyzed. Three 

configurations were analyzed. These included dual inline pusher puller motors, dual rear 

mounted motors, and dual front mounted motors. 

The advantage of using dual inline motors is that if an engine were to fail, asymmetric thrust 

would not be an issue. A conceptual design of dual inline motors is shown in Figure 22. 

Figure 22: Dual In-Line Pusher Puller Motors 

The major disadvantage of using dual inline motors is that it would add length to the aircraft. The 

front motor would need to be placed in front of the nose of the aircraft, adding an additional 3 in 

40



to the aircraft length. The aircraft is already 22.5 in from the nose to the most rearward location 

of the wing after the wingtip is folded. If any additional length were added in front on the nose, 

the aircraft would not be less than 2 ft long, and therefore would not fit into the box. 

Another big disadvantage would be that the propellers would interfere with the water bottle 

placement on the CG. In order to keep the aircraft static margin the same during all flight 

missions, the payload CG must be placed on the aircraft CG. The wing chord at the root is 20 in, 

and the length of the water bottle with aerodynamic fairings is 15 in long. Depending upon 

where the aircraft CG ends up, placing the water bottle on the CG may be impossible without the 

propeller striking the bottle. 

As for dual rear mounted motors, the primary advantage is that the wake from the propellers 

does not interfere with the lift and efficiency of the wing. A conceptual design of dual rear 

mounted motors is shown below in Figure 23. 

Figure 23: Dual Rear Mounted Motors 

The major disadvantage of having dual rear mounted motors is that there is additional ground 

clearance required for rotation on takeoff. As the aircraft nose pitches up for takeoff, the rear of 

the aircraft moves down about the CG. Estimating that the CG was approximately ½ way down 

the chord, or 10 in, and assuming a 10° angle of attack on rotation, the rear propellers would 

move down 1.5 in. This means that the landing gear would need to be 1.5 in taller when 

compared to front mounted motors. This added landing gear length and structure would add 

unnecessary weight and drag. 

Two other disadvantages to having dual rear mounted motors are weaker structure and 

interference with other aircraft systems. The airfoil chosen for the wing becomes very thin 

towards the rear. If the motors needed to be mounted on the rear, the structure would need to be 

strengthened significantly to handle the weight of the motors, as well as the stress caused by the 

force from the thrust, and any torque from the motors. As for interference, there are many 

components in the rear of the aircraft. All of the control surfaces will be at the rear of the aircraft. 

In addition, all of the servos and wiring required to move them will need to be integrated in the 

rear. The two main landing gear struts will also need to be placed at the rear of the aircraft. The 

41



attachment and integration of these parts will be difficult if they are all placed in the rear of the 

aircraft. 

The propulsion configuration chosen incorporates dual front mounted motors. This configuration 

requires the least additional ground clearance, reducing landing gear size, weight, and drag. 

Another major advantage to mounting the motors in the front of the aircraft is it helps move the 

CG forward, increasing stability. Each motor assembly is approximately 0.5lbs, a significant 

portion of the aircraft total weight. Since the aircraft center of gravity needs to be in front of the 

center of pressure, and the larger the static margin the more stable the aircraft is, having the 

weight of the motors in the front of the aircraft is the best for stability. 

For the same reason, the battery pack will be placed at the front of the aircraft. The weight of the 

battery pack was estimated at 1.0lbs, so placing the batteries towards the front of the aircraft was 

deemed essential for stability. By utilizing dual front mounted motors, the amount of wiring 

required to run from the batteries to the motors will be decreased, eliminating the extra wire 

weight. Finally, these motors fit without interfering with other aircraft systems. Unlike the rear 

of the aircraft, there are almost no systems in the front of the aircraft aside from the main landing 

gear in the center. For these reasons, the propulsion configuration chosen incorporated dual front 

mounted motors. 

A conceptual design of the aircraft using dual front mounted motors is shown in Figure 24. 

Figure 24: Dual Front Mounted Motors 

5.4 Structures Subsystem Design Alternatives 

5.4.1 Wing Construction 

Competition scoring is directly affected by aircraft weight and minimizing the overall system 

weight is a top priority in order to maximize the score. The two driving parameters for the 

selection of the wing construction method were wing weight and overall design complexity. 

Wing weight is defined as the overall weight derived from the material, adhesives, and fasteners 

required for each construction method. A minimal wing weight was desired to minimize the 

total aircraft weight. 

Design complexity is composed of two construction method factors: the time of manufacturing 

and the ease of reparability. Both of these factors are primarily driven by the wing construction 

42



complexity such as number of components, joints, and manufacturing steps. It was deemed 

important to reduce the time of manufacturing to increase the number of wings produced for 

testing, flights, and back-up replacements. The ease of repair was deemed important to increase 

each wing’s lifetime and allow for quick repairs if needed on-site or at competition. Higher 

preference was awarded to wing weight over design complexity due to its direct affect on the 

competition scoring. While design complexity is undesirable during the project construction and 

testing phases, it was identified as a secondary parameter because it does not directly affect 

competition scoring. 

The two main wing construction methods considered for the final design were balsa/foam-core 

composite wings and a traditional balsa rib and spar construction. A foam-core with balsa-skin 

sheeting relies on the balsa skin to carry the majority of the wing bending loads while the foam 

provides general wing structure and rigidity. The traditional balsa rib and spar construction 

technique applies the wing loads on the spar while the Monokote skin only provides an 

aerodynamic surface. A third available construction option was a composite shell technique 

made from fiberglass or carbon fiber. However this construction method was not considered due 

to the excessively large weights associated with composites and the higher complexity in 

manufacturing. The two wing construction techniques considered for final selection are 

displayed in Figure 25 9 . 

Figure 25: Wing Construction Method Design Options 

After examining the characteristics of both wing construction methods, it was determined that 

the weight would be similar between both methods. While the rib and spar method required 

fewer wing-critical materials, the addition of strengthened mounting locations for the wing stores 

would increase the overall wing construction weight to approximately the same as the balsa and 

foam-core composite. Thus the overall design complexity was used as a tie-breaker between the 

two construction techniques. It was determined that both the time of manufacturing and wing 

repair ease were vastly superior on the balsa-foam composite due to the fewer number of 

components and manufacturing steps. The rib and spar construction method inherently requires 

many components including the wing spar, joint locations, and various ribs sizes across the 

length of the wing span. The complexity and number of wing rib components required for a 

tapered and swept wing alone would dramatically increase each wing’s manufacturing time. The 

rib and spar system, in a light weight design, will also be vulnerable in crashes and require 

43



extensive repair time to account for varying sized parts throughout the wing. Due to the time 

savings during the construction and repair of a balsa-foam composite wing, this construction 

method was selected as the final design selection 

5.4.2 Wing Span Reduction Method and Joint Location 

The existence of both the span-wise minimum wing span and the maximum aircraft container 

dimensions made it impossible to employ a traditional solid single-piece wing. Some form of 

wing span reduction is necessary in the design to meet both span and container requirements. 

The primary decision parameters for wing span reduction included minimizing assembly time 

(the most significant mission score component), and design/manufacturing complexity. The 

most sensitive score component at competition is derived from assembly time, and any reduction 

in total assembly time is given the highest priority and preference. The design and 

manufacturing complexity is important for reducing failure opportunities and improving 

performance and construction consistency. 

The two primary methods considered as design options for total wing span reduction were wing 

folding and wing disassembly. Wing folding involves the use of hinges at a fold joint to allow a 

wing section to either fold upwards or downwards to reduce the span-wise length. Wing 

disassembly splits the wing into multiple sections that would come apart for container storage 

and reassemble into a complete wing. Both the wing folding and wing disassembly design 

concepts are provided within Figure 26. 

Figure 26: Wing Span Reduction Design Options 

The wing folding solution presents many advantages over wing disassembly, particularly with 

respect to the important aircraft assembly time derived from the mission sensitivity analysis. 

With a hinged folding wing, the handler only needs to repeat one simple swinging motion to 

reposition the wings into flight ready position. Another advantage to a folding wing is that any 

wiring across the fold location does not need to separate and reconnect repeatedly. A 

disassembled wing not only requires time for reassembly of various individual components, but 

has increased assembly time and complexity to realign components and reconnect any 

44



disconnected wiring. The simplicity and single motion input from the aircraft assembler was 

expected to significantly decrease the assembly time of a folded wing design over a disassembled 

wing design. The significance of assembly time on the mission score in conjunction with the 

simplicity and quick assembly of the wing folding system resulted in it being chosen as the final 

wing span reduction method. 

Various locations along the wing for the fold joint location were considered, including near the 

wing tips, near the wing root, and a single side wing fold. Design concepts for each of the wing 

fold joint locations are provided in Figure 27. The wing tip or mid-span wing folds place the fold 

joint in the outer ends of the wing span. Folding at or near the wing root would place the fold 

joints on or near the aircraft centerline. A single side folding placed the fold joint in one side of 

the wing, sparing the other side from any fold joints. A fold joint at the wing root will encounter 

the largest bending moment, and the significant risk to the overall aircraft structural integrity 

disqualified this design option. A fold joint on one side was disqualified due to the asymmetric 

design issues and the significant strain placed on that wing half. Wing fold joints located in the 

outer sections of the wing was chosen primarily because it encounters the smallest wing loads 

and moments. Another reason for choosing wing-tip folding is that it creates the smallest folded 

region out of the three options. 

Figure 27: Wing Fold Joint Location Design Options 

5.4.3 Landing Gear Configuration 

Determination of the aircraft landing gear configuration was primarily driven by four landing 

gear parameters: ground stability, ground authority, take-off rotation, and drag. Factors such as 

landing gear weight and integration were not included in the determination process due to their 

dependence on the overall aircraft design and configuration. A preliminary landing gear height 

of 7 inches was estimated by accounting for adequate centerline store and propeller ground 

clearance below the wing. The four main decision drivers do not directly meet mission 

requirements, but all play significant roles in the completion of the various missions. Ground 

stability is crucial during the ground components of Mission 3 where an asymmetric wing store 

configuration is possible and any aircraft that tips over will fail that mission attempt. Ground 

authority is important once again in Mission 3’s ground component when the pilot must taxi the 

aircraft to a specified location for rocket drops. Take-off rotation is crucial in every mission as 

the aircraft must take off within 100 feet every time. Lastly, drag is an important factor in 

45



Mission 1, where two laps will be timed for the score and in Mission 3 where flight endurance is 

crucial. 

The three landing gear configurations considered for the design included tricycle, bicycle, and 

tail-dragger style landing gear and are shown in Figure 28. A tricycle configuration has one nose 

gear and two wheels on a main gear just aft of the aircraft center of gravity. A bicycle 

configuration has one nose gear and one nose wheel inline on the centerline of aircraft, and 

sometimes employs wing-tip outriggers for lateral balance. The tail-dragger configuration is 

similar to an inverted tricycle gear, where a tail gear is used and the two wheels on the main gear 

are just forward of the aircraft center of gravity. 

Figure 28: Landing Gear Configuration Options 

After examining the three landing gear configuration’s characteristics in each of the four main 

parameters, the tricycle configuration was determined to be the best all around performer. The 

tricycle configuration is widely regarded as the most ground stable due to the aircraft center of 

gravity’s location just forward of the main gear. The tricycle configuration is not as vulnerable 

to tip over from take-off side gusts. Tail-dragger and bicycle configurations are vulnerable since 

they must pivot about one wheel. Ground authority on a tricycle landing gear is simpler than the 

tail-dragger which has nose over and ground loop issues. The tricycle gear’s ground stability 

also outperforms the bicycle gear which must balance on the center-line mounted landing gear. 

The bicycle gear is also particularly vulnerable to tip over with the large expected lateral CG 

shifts. The tricycle gear is the only one of the three that has the capability to rotate during the 

46



take-off ground roll and generate additional positive angle of attack for take-off. The drag 

penalty on the three large surface tricycle configuration is higher than the two surface bicycle 

gears and slightly higher than the more aerodynamically efficient tail-dragger configuration. 

However, due to the drag penalties only affecting one timed mission and general aircraft 

endurance, it was decided that the superior performance in the other three parameters far 

outweighed the tricycle gear’s slight drag penalty. Due to the tricycle gear’s superior 

performance in ground stability and authority, a unique ability to rotate on the ground among the 

three, and only slight underperformance in drag, it was chosen as the final selection for landing 

gear. 

5.4.4 Main Landing Gear Material 

The main landing gear material was chosen by examining the weight and radius required in order 

to satisfy the landing gear design-to-specifications. These properties were compared for various 

potential gear construction materials including aluminum, steel, fiber glass, and carbon fiber. 

The results are shown in Figure 29. The material that required the smallest radius, thus 

generating the least drag, and smallest weight was carbon fiber. However after examining 

potential landing gear construction and mounting methods, it was determined that carbon fiber 

was not an ideal landing gear solution for this aircraft. The next best solution was aluminum, 

which still had an ideal low weight but with a slightly larger radius. Along with its low weight, 

aluminum’s malleability made it an attractive material for landing gear shaping and mounting. 

For these reasons aluminum was chosen as the main landing gear strut material. 

0.25 

Landing Gear Strut Radius Sensitivity to Sweep Angle and Youngs Modulus 

2500 ksi (Fiberglass Composite) 

10,000 ksi (Aluminum) 

20,000 ksi (Carbon Fiber w/ Epoxy) 

30,000 ksi (Medium Alloy Steel) 

0.2 

0.15 

Radius (in) 

0.1 

0.05 

0 

0 10 20 30 40 50 60 

Angle Sweep of Strut (Degree) 

Figure 29: Main Gear Material Comparisons 

47



6.0 Subsystem Design-To Specifications 


Co-Author: Brett Miller 

6.1 Aerodynamics Design-To Specifications 

Requirements for the aerodynamics subsystem were derived from customer requirements as well 

as from the mission goals. The customer required that the aircraft take off in no more than 100 ft 

and fly at conditions in a 5000 ft density altitude. Additionally, it was required that the minimum 

wing span shall be no less than 5 ft and the airplane is required to be stable in both the 

asymmetric and symmetric load cases. 

Since the overall goal of the team is to win the competition, performance requirements were 

derived from past competition winners. The aircraft was required to cruise at 100 ft/sec, stall at 

40 ft/sec, and perform a 2 g maneuver at 80 ft/sec. Heuristic data is valid because performance 

specifications such as cruise speed, and stall speed have not varied over the past years’ 

competition winners regardless of mission or overall aircraft configuration. 

6.2 Missions Design-To Specifications 

The design-to specifications for the Missions subsystem can be broken into three separate 

elements: the wing mounted stores, the centerline store, and the aircraft container. Each of these 

is vital to the mission of the aircraft and its performance at competition. 

The design process includes an analysis of the loads acting on the stores, consideration of 

multiple design alternatives, testing of those alternatives, an educated selection based on the test 

results, and finally a successful implementation of the design choice. 

6.2.1 Wing Mounted Store 

Based on the sensitivity analysis, the loading of these stores will have a large effect on the 

aircraft’s score at competition. This loading time factors into both the System Complexity 

Factor score of the aircraft through its effect on assembly time, and wing store loading time 

directly affects the score in Mission 3 (the heaviest weighted mission). It is therefore important 

that the release mechanisms for these stores have the ability to load quickly. Although a specific 

time requirement has not been established, multiple iterations should be utilized to achieve the 

fastest loading design. 

It is also important the store release when desired and not while in flight. To this end, the release 

mechanism must release 95% of the time while being able to constrain the store under flight 

conditions and loads. These flight conditions include a 3g force in the aircraft’s Z-axis and a 

1.73 g lateral force. These design-to loads were determined from the following equations (which 

in turn were derived using Error! Reference source not found.). 

48



Equation 8: Summary of Equations Calculating Centripetal Force on Wing Stores 


The centerline store experienced a change in requirements between PDR and CDR in the fact 

that it must be releasable. Therefore, the expectations for the wing mounted store also apply to 

the centerline store. The design for this store must be able to support the larger load of the filled 

wattle bottle at 9 lbs. 

6.2.3 Container 

The containers have many objectives to achieve. The first of which is that it must accommodate 

the payloads, the aircraft, and the transmitter. These items must be secured in the containers. 

The containers themselves must also be able to be secured so that it may be handled without 

opening. The containers must be able to protect themselves and their contents from a 6” drop 

onto pavement. If any contents are shifted or damaged this test results in failure. It is also 

important that the design facilitate a fast loading assembly of the aircraft after being secured in 

the containers. The time it takes to do this is factored directly into the assembly time of the 

aircraft at competition, which drives the SCF. 

6.3 Propulsion Design-To Specifications 

The design of the propulsion system was restricted by several competition requirements. The 

propulsion system must be electrically powered and all parts must be COTS. For safety, the 

battery chemistry must be either NiCad or NiMH. The maximum battery pack weight is limited 

to 4lbs. The amount of current flowing from the battery is limited by a 40 amp fuse. The 40 amp 

fuse must be implemented in such a way so that no single part of the propulsion system sees 

more than 40 amps. 

The biggest requirement for the overall propulsion system is that the aircraft must be able to take 

off in under 100 ft with a 5,000 ft density altitude and no wind. Once airborne, the aircraft must 

have a range long enough to complete all three flight missions. The two most battery power 

challenging missions are mission 2 and mission 3. Mission 2 requires four laps flown with a fully 

loaded water bottle (9.0 lbs). Mission 3 requires four takeoffs, four landings, and four laps, 

49



dropping one rocket in between each lap. The first lap will be flown at 12 lbs, and the last lap 

will be flown at 7.5 lbs, with a 1.5 lb rocket being dropped in between each landing and takeoff. 

The batteries must be designed for a four lap endurance plus a 30 second power reserve. 

6.4 Structures Design-To Specifications 

6.4.1 Aircraft Wing Requirements 

The aircraft wing is required to sustain a +3.5g load while carrying the centerline store and a -2g 

load while carrying the tip stores. The required g loads were determined by including an 

additional 1 g of load to the worst case expected scenario for each payload. The worst case wing 

loading scenarios were derived from the expected in-flight bank turns the aircraft must make 

during each of the missions. The addition of the g loading margin creates a 1.4 safety factor for 

the centerline store flight scenario and a safety factor of 2 for the tip store flight scenario. The 

safety factor added to the worst expected loading is an added margin for unexpected factors 

including in-flight wind gusts and ideal assumptions made for the wing material and 

construction. 

The aircraft wing is required to fit within the aircraft container while still meeting the wingspan 

and payload integration requirements. This requires the wing be foldable or collapsible for incontainer 

storage due to the 48” x 24” x 24” dimensions specified by the competition. To 

guarantee the aircraft wing does not exceed the maximum container dimensions, the aircraft’s 

maximum dimensions on each side shall be 1” shorter than the container’s dimensions. With the 

maximum wing dimensions bound to 47” x 23” x 23”, a margin of safety was assured for this 

dimension requirement. This margin accounts for the thickness of the walls as well as some free 

space on any side of the aircraft in the container. 

The wing structure is also required to integrate the wingtip and centerline stores such that they 

are securely restrained in the worst case flight loading with some safety factor. The worst case g 

loading is expected to be a 2.5g bank turn and 1g of safety factor was added to that loading. A 

3.5g loading generates 28 lbs at the centerline store mount and 10.5 lbs at each wingtip store 

mount. The additional 1g included in the design wing loading creates a safety factor of 1.4 for 

each payload’s mounting point. 

6.4.2 Landing Gear Requirements 

The aircraft landing gear and supporting aircraft structure is required to survive the impact loads 

involved during landing. A successful mission requires that the aircraft survive take-off, flight, 

and landings. The landing gear requirement dictates that each gear strut survive a dynamic 

loading assumed to be approximately 3 times the heaviest aircraft weight configuration. Landing 

the plane on one gear in the heaviest configuration was deemed the worst case landing scenario. 

The 3g single gear strut loading includes a safety factor of 1.2 during the heaviest aircraft 

configuration and a safety factor of 2.25 in a no payload configuration. An additional 25% 

50



safety factor applied to the final gear design was required to account for additional loads in 

landing and for the ideal assumptions made in the gear materials and construction. 

6.5 Avionics Design-To Specifications 

6.5.1 Transmitter 

The transmitter is required to control all control surfaces, the propulsion system, and the 

releasable external stores. The releasable stores are required to be dropped one at a time. The 

transmitter is also required to have a fail-safe mode that is automatically selected during loss of 

transmit signal. During fail-safe the aircraft receiver must select throttle closed, full up elevator, 

full right or left aileron, and full right rudder. 

6.5.2 Telemetry System 

The telemetry system must record several flight characteristics including: Motor voltage, amps, 

motor RPM, angle of attack, G forces, servo positions, altitude, and airspeed. 

6.5.3 Microcontroller System 

The aircraft’s onboard microcontroller must be able to take signals from the receiver, process 

which order is being commanded by the pilot, and determine when to drop the payload stores. 

51



7.0 Project Feasibility and Risk Assessment 

Author: Eric Hall 

Co-Author: Dan Colwell 

7.1 Project Feasibility 

7.1.1 Weight Budget and Feasibility 

The estimated weight of the aircraft was determined from the sum of the weights of all the 

systems components. From the combination of the system weights stated earlier, the estimated 

maximum takeoff weight of the aircraft is approximately 15 pounds. This weight reasonably 

satisfies the aircraft maximum weight requirement (0.PRJ.5). 

This weight estimation can be further validated by comparison to similar aircraft. A database of 

aircraft from the 2007-2008 DBF competition shows an average payload weight to empty aircraft 

weight ratio (W PL /W E ) of 1.27. The 2008-2009 competition payload weight of 8.33 pounds 

determines the aircraft should have an empty weight of approximately 6.54 pounds. This 

estimate also satisfies 0.PRJ.5. 

7.1.2 Cost Feasibility 

In order to guarantee the project’s material costs are less than the project’s budget, a base 

estimate of the prices of each component was collected. The estimated budget for construction 

and travel can be observed in Section 14.4. The estimated budget for the project is $12,130. 

Based on the funding provided by the Aerospace Engineering Sciences Department ($4,000), the 

Engineering Excellence Fund ($2,000), and Lockheed Martin ($10,000), the budget for 

constructing the aircraft is feasible even with a 25% margin. 

7.1.3 Aerodynamic Feasibility 

Based on the design to specifications, the aircraft was sized to a takeoff distance of 90 feet with a 

10 ft margin, 40 ft/sec stall speed, 100 ft/sec cruise speed and 2g maneuver at 80 ft/sec. The 

performance sizing plot is shown in Figure 30. Performance sizing was checked with Dr. 

Gerren, a team advisor [10] . 

52



Figure 30: Performance Constraint Plot 

From the performance sizing, the design area that meets all requirements was determined. In the 

plot shown, the shaded area indicates the optimal design area. The equations used to determine 

the weight to power ratio as a function of stall speed, takeoff distance, cruise speed and 

maneuver are shown in Equation 9, Equation 10, and Equation 11 respectively: 

= 1 

2 

 

Equation 9: Weight-to-Power Ratio for Stall Speed 

In Equation 9, is the ambient density, is the maximum lift coefficient and V stall is the stall 

speed. 

= ∙ ∙ ∙ 

∙ 

1.44 ∙ 

Equation 10: Weight-to-Power ratio for takeoff. 

In the performance sizing equation for takeoff distance, is the takeoff distance, is the 

density, is the maximum lift coefficient, is the landing speed, is the acceleration 

due to gravity and is the wing loading. 

 

53



 

= 

 

 

∙ ∙ 

 

 

 

 

1 

 

 

 

+ 

 

 

∙ ∙ ∙ ∙ 

∙ 

 

∙ 

∙ 

 

 

 

 

Equation 11: Weight-to-Power ratio for cruise speed. 

In Equation 11, is the dynamic pressure based on the ambient density, and cruise velocity, 

is the parasitic drag, 

 

is the takeoff thrust to cruise thrust ratio, is the Oswald efficiency 

factor of the wing, is the aspect ratio of the wing, 

 

is the ratio of the cruise weight to the 

takeoff weight, is the wing loading and 

is the cruise speed. A detailed list of the 

assumptions and values used to size this aircraft can be seen in Appendix F. The equations used 

to do the performance sizing were obtained from Roskam [11] . 

The design point chosen from this design area is also marked in the plot. This design point gave 

a wing loading, W/S of 2.1 lb/ft 2 and a weight to power ratio, W/P of 0.047 lb/lb ft/s. Using an 

estimated gross takeoff weight of 15lb, the wing area necessary from the wing loading was 

determined to be 7.14 ft 2 . The power required from the weight to power ratio was determined to 

be 530 W. The performance sizing shown demonstrates that a number of designs that meet all 

the performance requirements for stall, takeoff and cruise are feasible, and an optimized design 

can be created that would maximize the cruise speed while keeping the weight to power ratio 

small. 

7.1.4 Propulsion Feasibility 

To ensure that the propulsion system can meet all the design to specifications, a feasibility study 

was conducted. The two major requirements to verify are the 100 ft takeoff distance and the 

maximum 4 lb battery weight. The rest of the requirements can be verified by inspection. In 

order to make the 100 ft takeoff, the thrust required from the motors was calculated. The 

equation for takeoff distance is in Equation 12. 

S 

LO 

= 

g * ρ * S * C 

L 

2 

1.44 * W 

*{ T − [ D + µ * ( W 

r 

− L)} 

Equation 12: Takeoff Distance Calculation 

To simplify the equation, it is assumed that drag, D is very small compared to thrust and that the 

coefficient of friction, µ of the landing gear wheels is negligible. This yields the simplified 

version solved for thrust in Equation 13. 

r 

54



T 

= 

2 

1.44* W 

g * ρ * S * C L 

* S LO 

Equation 13: Simplified Takeoff Distance Calculation 

Where W is the fully loaded weight of the aircraft (15 lbs), ρ is the density for a 5,000 ft density 

altitude, S is wing area, g is gravity, C L is takeoff lift coefficient, and S LO is takeoff distance. The 

aircraft stalls at a C L of 1.38 and an angle of attack of 12.6°. Therefore, assuming a takeoff 

rotation of 10°, takeoff C L was assumed to be 1.1. To account for the small drag and rolling 

friction assumption, a 10% margin was added to the takeoff distance. Therefore, S LO was chosen 

to be 90ft. With these variables, the thrust required is 6.0lb, or 3.0lb per motor. To ensure that 

this thrust could be met, a static thrust stand was created shown in Figure 31. 

Figure 31: Static Thrust Stand 

The thrust stand was designed so that as the motor generated thrust, it would pull against a load 

cell to record the static thrust. The load cell was connected to a LabView VI to record the thrust 

data [12] . During static thrust tests, an Eagle Tree telemetry system was used to record voltage, 

current, power, and motor RPM [13] . The motor set up utilized to test the thrust available was the 

Neu [14] 1107 2Y motor with 3300 rpm/V at 19.0 V, 40 amps, and an 8,000 ft density altitude. To 

make up for the extra rpm/V of the Neu 1107 motor compared to the Neu 1110 2Y motor, the 

voltage was lowered below that of the estimated battery pack voltage to compensate. Several 

different propeller sizes were tested. The amount of thrust produced by this single motor varied 

between 4.30 lb and 5.87 lb of thrust depending upon the propeller size and pitch. This 

confirmed that the amount of thrust required to make the 100ft takeoff requirement is feasible. 

The other major design to specification of the propulsion system is the maximum battery weight 

of 4 lbs. To check the feasibility of this requirement, the amount of power required to fly four 

laps at maximum weight was estimated. The average current draw was assumed to be 20 amps at 

55



20V. At this power draw, an 18 cell battery pack would last for 4.5 minutes. This would be 

enough time to complete four laps with a minimum 30 second reserve. After glancing at several 

NiCad and NiMH batteries, a conservative weight of 30g per battery was chosen. The battery 

pack weight was calculated to be 1.2lbs. This is significantly less than the 4lb maximum battery 

pack weight, making it feasible to meet all requirements. 

7.1.5 Payload Feasibility 

Payload deployment capabilities are important in order to successfully complete the competition 

missions. Currently two parallel concepts are being designed; a mechanical pin concept and a 

magnetic plate concept. In order to restrain the payload in flight, both systems need to withstand 

a 4.5 lb vertical load while restraining a 1.5 lb horizontal load. This is not a concern for the 

mechanical system as the payload system would need to fail structurally for the payload to 

prematurely detach. For the magnetic system, the payload would need to overcome the magnetic 

force between the payload and the deployment system. In flight, a maximum force of 4.74 lb 

could be encountered, requiring a magnetic force equal to or greater than this force. Two D8X8 

neodymium magnets installed in the payload provides an attractive force of 12 lb, more than 

enough to restrain the payload in worst case flight conditions. Deploying the mechanisms is 

solely dependent on the force applied by an actuator. Standard digital servos are capable of 

providing enough torque to deploy both mechanisms. 

Landing gear ground stability for the worst asymmetric payload configuration was analyzed 

through lateral center of gravity (C.G.) shifting. An aircraft assumed to weigh 6 pounds under a 

symmetric configuration will have an aircraft C.G. at the lateral aircraft center. The worst case 

payload configuration was determined to be two wing stores on one wing (27 inches from 

aircraft center), while the other wing carries no wing stores. The combined two wing stores will 

generate a moment of 81 lb-in about the aircraft center, forcing the total system C.G. to shift 

towards the wing tip by 9 inches from its initial location. With a design margin of 1.5 the rear 

landing gears will be 13.5 inches left and right of the center line with 27 inches between them. A 

distance of 27 inches will assure adequate ground stability for the aircraft even under the worst 

asymmetric loading conditions while providing more than enough space for the 6 inch wide 

centerline fuel tank. 

7.1.6 Assembly Feasibility 

The box will be aligned such that the 4’ direction is along the aircraft wingspan, while the height 

and length of the aircraft will be aligned in the 2’ directions. To get the aircraft to fit into a single 

box, the wings will be designed to fold 15” in from the tip, reducing the folded wingspan to 38”. 

As for height, the aircraft will be designed with landing gear of 7”. This is driven by the need to 

store the 6” diameter water bottle under the wing and by the need for propeller ground clearance. 

Above that, the wing will only be a few inches thick to house the motors and batteries. The 

propeller could be positioned horizontally parallel to the wing while packed in the box. The 

winglets are not expected to exceed a height of 15”. As seen in Figure 32, the required vertical 

56



clearance is available with the wing-tips folded: This was determined using the code provided in 

Appendix D 

Figure 32: Assembly Feasibility 

The wing area was determined to be 7.14 ft 2 from the performance sizing. Using an upper bound 

of 25° leading edge wing sweep, and a nominal taper ratio of 0.5, the leading edge of the wing at 

the fold is 19 in behind the nose. This means the aircraft will be 23 in long when folded in the 

box. This was the upper bound because ½ in thickness for the walls of the box was assumed. It is 

also important to note that even with a 14 in propeller (the largest under consideration for the 

project), the propellers can still be mounted on the front of the aircraft with adequate clearance 

from the wing, and still remain behind the nose of the aircraft. Based upon this information, it is 

entirely feasible to fit the flying wing inside a 2’ x 2’ x 4’ box by only folding the wingtips. 

7.2 Risk Assessment 

The largest risks involved with designing the Buff-2 Bomber are enough to warrant concern and 

additional analysis in order to alleviate as much danger of total aircraft failure. The major risks 

from each subsystem were further analyzed to determine the overall importance, probability, and 

mitigation. 

7.2.1 Aerodynamics 

Insufficient stability of the aircraft would result in a loss of pilot control and erratic flight. 

Included in this risk is the possibility of aircraft damage during takeoff and landing if the 

unstable nature of the aircraft prompts the plane to pitch or roll when operating at low velocities. 

Should the aircraft be unstable, the control surfaces would be unable to compensate and a crash 

is imminent. This was mitigated through extensive analysis and research into airfoils and design 

of flying wings. Even with the amount of time devoted to aerodynamic design, the probability of 

instabilities is still at a medium level. However, because the consequences are so great, that this 

aerodynamic risk possibilities are still undergoing continuous analysis. 

57



7.2.2 Avionics 

Should the pilot, while operating an in-flight aircraft, lose radio communication to the vehicle 

due to Radio Frequency Interference (RFI) the aircraft will crash. However, this probability is 

low because of the transmitter selection. Because the operating frequency is 2.4 GHz, no other 

radios can operate on the same channel and interfere with the communication with the Buff-2 

Bomber. Also the selected transmitter has a programmable failsafe. If for some reason the 

communication with the vehicle fails, the receiver will automatically cut throttle and deflect the 

control surfaces to bring the aircraft down. 

7.2.3 Propulsion 

The propulsion sub-team must be most worried about whether or not their selected motor and 

battery configuration can propel the aircraft far enough to complete each mission. If the total 

range is insufficient to complete four laps at maximum weight, the consequence is no score and 

ultimately no chance of winning the competition. Fortunately this is a lower level probability as 

much analysis and testing has been done to accurately calculate the actual thrust provided from 

the motors. Further mitigation will involve dynamic testing of the entire propulsion system. 

7.2.4 Structures 

The wingtip hinge design is necessary for the aircraft to fit into the storage container. However, 

“breaking” the wing along the span increases the possibility of a catastrophic failure during flight 

that results in competition failure and possible endangerment of bystanders. Because of the 

mechanical complexity, the wingtip hinge presents the most probable and dangerous risk the 

aircraft faces. However, much of the risks associated with this design aspect can and will be 

mitigated with extensive ground testing. The overall strength of the hinge system and its ability 

to restrict free-play will be optimized before the system is tested in flight. 

7.2.5 Missions 

Tasked with designing the release mechanism of the wing mounted stores, the missions sub-team 

was able to determine their most significant risk to mission success is the possibility of the 

release mechanism jamming or failing during the mission. This possibility is mid-to low range 

simply because extensive testing will be done on the prototype and final design to ensure that the 

mechanism can release the stores with a 95% success rate. Another primary risk is if the rules, 

which undergo constant upgrades and explanations, may soon not allow magnetic mechanisms to 

hold the payloads to the wing. Although a low risk, this will be mitigated by not abandoning the 

mechanical design system and holding it as a backup to the magnetic system. 

7.2.6 Microcontroller 

Implementing the microcontroller creates additional risks to the overall avionics subsystem. The 

microcontroller is only enabled when the pilot gives the appropriate input on the transmitter; 

however, it is possible that the microcontroller receives a false reading and activates while the 

plane is in flight. The probability of this situation occurring will be fairly low as safety features 

58



will be coded into software and extensive testing will be performed to determine the likelihood 

of the microcontroller receiving a false signal. 

59



8.0 Mechanical Design Elements 


Co-Author: Ben Kemper, Mark Findley 

8.1 Aerodynamics Mechanical Design Elements 

8.1.1 Aircraft Geometry 

The analysis that was done on the effect of static margin with varying sweep angles and varying 

taper was used to determine the optimal geometry given the requirement for the aircraft to fit in 

the box. In order to have the most room along the wing for control surfaces, the span was chosen 

to be 68 inches. This satisfies the minimum wing span requirement of 60 inches. Figure 33 

shows the aircraft geometry: 

Figure 33: Aircraft Geometry 

The analysis done showed that the static margin increased as the sweep angle increased between 

20 and 25 degrees. The best possible configuration that meets the requirement to fit in a 4’x2’x2’ 

box is a leading edge sweep angle of 23 degrees and a taper ratio of 0.5. 

Because this project involves flight missions with asymmetric loading in the lateral direction, the 

addition of a vertical is necessary. The vertical is the equivalent of a vertical stabilizer and 

provides yaw control. Winglets were added to each wing tip to serve the purpose of a vertical 

stabilizer. The winglets were sized using the tail volume coefficient method. 

The tail volume coefficient method is based off of the sizing of past aircraft. This sizing is based 

off of the area of the wing, S w , the wing span, b, the area of the vertical tail, A vt , and the distance 

between the c. g. of the aircraft and the aerodynamic center of the vertical tail. The tail volume 

coefficient for most aircraft falls in the range of 0.03 to 0.06. A larger tail volume coefficient 

60



means that the aircraft will be more directionally stable. With the main geometry of the wing 

selected, the only parameter to determine is the sizing of the vertical tail. A tail volume 

coefficient of 0.04 was used giving the two vertical tails an area of 0.69 ft 2 . The vertical tails are 

shaped with a taper ratio of 0.5 and a straight trailing edge. This drove the vertical tails to be 

13.25 inches tall. Table 6 summarizes the geometry of the aircraft. 

Table 6: Characteristics of Aircraft Geometry 

Parameter 

Value 

Leading Edge Sweep, Λ LE 23 degrees 

Quarter Chord Sweep, Λ c/4 19 degrees 

Trailing Edge Sweep, Λ TE 7.3 degrees 

Span, b 68” 

Wing Area, S 7.14 ft 2 

Aspect Ratio, AR 4.5 

Root Chord, c root 20.16” 

Tip Chord, c tip 10.08” 

Taper Ratio, λ 0.5 

Winglet Height 13.25” 

Winglet Taper Ratio, λ v 0.5 

Wing Area, S v 0.69 ft 2 

8.1.2 Airfoil Selection and Aerodynamic Twist 

The airfoil chosen for the root of the aircraft was the HS602. This airfoil was chosen for the root 

primarily because it has a thickness to chord ratio of 10.21%. Because this aircraft does not have 

a fuselage, the wings will house the batteries, motor mounts, release mechanism mounts, the 

servos and the wiring. In order to ensure that these components will be housed in the wing, the 

root must have a reasonably thick airfoil. 

From a stability standpoint, it is ideal to have an airfoil with a moment coefficient of 0. Of all the 

airfoils analyzed, the airfoil that had the smallest moment coefficient was the HS520 airfoil 

(Figure 34). Since the HS520 has a thickness to chord percentage of only 8%, it is not the 

optimal choice for the root. In order to make a fair trade between the most desirable thickness to 

chord percentage that would house most of the components and the airfoil with the best moment 

coefficient, an aerodynamic twist was implemented into the design. The root section has the 

HS602 airfoil and the tip has the HS520 airfoil with a linear twist between the root and the tip of 

the wing. The winglets have the NACA 0010 airfoil; a symmetric airfoil with a 10% thickness to 

chord ratio. 

61



Figure 34: The Tip Airfoil (HS520) 

8.1.3 Aircraft Incidence Angle 

By definition, the incidence angle on a fixed-wing aircraft is the angle between the chord line of 

the wing root where the wing is mounted to the fuselage and the longitudinal axis of the fuselage. 

However, on this aircraft, due to the absence of a fuselage, the incidence angle is the angle of 

attack. A slight incidence angle is necessary to ensure the necessary takeoff rotation needed 

during takeoff. To determine the angle of incidence necessary, the takeoff distance was estimated 

using Equation 14 

S 

LO 

2 

1.44W 

= 

ρgC 

T 

L max 

Equation 14: Takeoff Distance Equation 

The maximum lift coefficient in the takeoff distance equation corresponds to the angle of attack 

or in this case, the incidence angle necessary to meet the takeoff distance requirement. From the 

coefficient of lift vs. angle of attack curve for the airfoil selected, at an angle of attack of 0, the 

corresponding Cl is ~0.1, and at the stall angle of attack of 13º, the Cl is ~1.4. The lift coefficient 

needed to meet the takeoff distance requirement of 100 ft is ~0.7, which corresponds to an 

incidence angle of 5 degrees. As a result, a 5 degree incidence angle was implemented into the 

design to achieve the rotation and ground roll necessary on takeoff. 

8.1.4 Control Surface Sizing 

For this airplane, conventional control surfaces (ailerons, elevators and rudders) were chosen. 

The aileron sizing was driven by the requirement to fit the airplane in the box. The ailerons were 

sized based on the location of the fold on the wings. It is ideal to use the least number of servos 

possible in order to reduce weight. In order to eliminate the weight of an additional servo, it is 

desirable to limit the length of the aileron by the location of the fold. Figure 35 shows the 

location of the fold and the aileron size based on the fold: 

62



Figure 35: Location of the Fold on the Wing 

The percentage of the chord for the control surfaces is important to ensure that the control 

surfaces produce the necessary control authority. An upper bound on the chord percentage is 

about 30% because flow separation occurs on the wing for larger percentages. The ailerons were 

sized to 15” from the wingtip and 30% along the chord. 

The rudders were sized so that maximum possible control could be achieved. The rudder was 

sized to be the entire length of the vertical with a 0.25” clearance so it would not interfere with 

the ailerons. The rudder is 13” long along the winglets and 30% along the chord. 

Since pitch control is a concern for flying wing configurations, it is important to ensure that 

maximum pitch control is available. The elevator was sized by the ailerons. The elevator is 

located between the ailerons, inside of the fold along the wing. The elevator has a constant chord 

percentage of 15%. Figure 36 shows the three control surfaces employed by this design: 

Figure 36: Control Surfaces on the Aircraft 

63



In order to guarantee that the designed control surfaces could produce the deflections necessary 

to trim the aircraft, analysis was done in AVL. Three cases were setup to determine if the 

necessary deflections were feasible. Under the worst case scenario (2 rockets on one wing during 

takeoff), the maximum deflection necessary from the rudder was 10 degrees, 15 degrees from the 

aileron and 15 degrees from the elevator. The maximum possible deflection is approximately 25 

degrees. Flow separation occurs over the wing around 25 degrees of deflection. It is clear that 

there is enough margin for deflection for all the control surfaces before flow separation occurs. 

8.1.5 Stability Analysis 

Stability analysis was done in AVL and in MATLAB, the code can be found in Appendix G. 

Non-dimensional stability derivatives were obtained from AVL for three different scenarios that 

mirror flight missions. The first scenario simulated a case where the centerline fuel tank is flown 

on the airplane for cruise conditions, the second scenario was four rockets on the airplane in 

cruise, and the third was the worst case scenario of two rockets on one wing tip. To get a clear 

understanding of the aircraft’s stability it is important to analyze stability in the longitudinal and 

lateral domains. The longitudinal domain refers to the stability of the aircraft about the axis of 

pitch. Lateral stability refers to the stability of the aircraft about the yaw and roll axis. 

Stability modes in the longitudinal domain are the phugoid and short period and stability modes 

in the lateral domain are roll subsidence, dutch roll and spiral divergence. Stability derivatives 

represent an incremental change in a force or moment acting on the aircraft to a corresponding 

change in any of the following variables: velocity, angle of attack, side slip angle, bank angle, 

pitch angle, pitch rate, roll rate and yaw rate. Because aircraft dynamics is fairly complex, it is 

necessary that the equations of motions are linearized to assess stability. 

The non dimensional stability derivatives outputted from AVL were dimensionalized in 

MATLAB to analyze dimensional stability for different modes in the longitudinal and lateral 

directions. Matrix methods were used to solve the set of differential equations to obtain the 

stability of the aircraft. The aircraft is modeled as a dynamic system where the system receives 

input in the form of control surface deflections implemented by the pilot. Equation 15 shows the 

longitudinal equations of motion. The plant that describes this system is represented by Equation 

16. 

0 − cos 

 

= + − sin 

 

0 

0 0 1 0 

0 0 0 

0 − 

= 

0 0 

 

0 − 1 0 

0 0 0 1 

Equation 15: Equations of Motion in Matrix Form for Longitudinal Stability 

 

= 

 

Equation 16: Representation of the Dynamic System for Longitudinal Stability 

64



The eigenvalues of the A matrix are the poles to the dynamic system. 

This analysis was performed for the three cases described previously in the longitudinal and 

lateral directions. The longitudinal case will be presented first. The poles plotted for the airplane 

with the bottle is shown in Figure 37: 

0.1 

0.08 

Longitudinal Stability Modes for the Bottle on the Airplane 

Real Roots 

Short Period 

0.06 

0.04 

Imaginary 

0.02 

0 

-0.02 

-0.04 

-0.06 

-0.08 

-0.1 

-7 -6 -5 -4 -3 -2 -1 0 1 

Real 

Figure 37: Longitudinal Stability Modes for the Bottle on the Airplane 

The natural frequency and the period of each of the roots for the longitudinal case for the bottle 

are summarized in Table 7: 

Table 7: Longitudinal Stability for the Bottle on the Airplane 

Bottle on the Airplane-Longitudinal 

Mode Roots Time Constant (Real Roots) 

Natural Frequency (Complex Conjugates) 

Period (sec) 

ζ 

Real Root -6.04 Time constant 6.04 (1/sec) 0.17 -- 


Short Period 0.03±0.10i ω n 0.20 (rad/sec) 31.60 0.005 

65



The s-plane is a visual representation of the roots of the system to characterize the behavior of 

the system. Poles in the right half of the plane are unstable where as poles in the left half of the s- 

plane indicates stability. 

As seen in Figure 37 and Table 7, the real roots are located on the negative real axis. These roots 

are stable and have time constants of 6.04 s -1 and 0.17 s -1 . The short period is slightly unstable. 

The period of the short period roots were analyzed to determine if this is a concern. The period 

for both short period roots is 31.60 seconds. Since the period for natural instability in the short 

period mode is as slow as 31.60 seconds, this indicates that the pilot has 31 seconds to input 

control surface response to the system. As a result, this is not an issue. 

The analysis in the s-plane for the airplane with four rockets is shown in Figure 38: 

0.15 

0.1 

Longitudinal Stability Modes for Four Rockets on the Airplane 

Real Roots 

Short Period 

0.05 

Imaginary 

0 

-0.05 

-0.1 

-0.15 

-0.2 

-8 -7 -6 -5 -4 -3 -2 -1 0 1 

Real 

Figure 38: Longitudinal Stability Modes for Four Rockets on the Airplane 

The natural frequency and period of each of the roots in the longitudinal case for the scenario 

with four rockets on the airplane are summarized in Table 8: 

66



Table 8: Longitudinal Stability for Four Rockets on the Airplane 

Four Rockets on the Airplane-Longitudinal 



Period (sec) 

ζ 



Short Period 0.03±0.11i ω n 0.24(rad/sec) 26.23 0.005 

For this flight case, there were two real, stable roots which had time constants of 7.72 s -1 and 

0.19 s -1 . The short period roots were slightly unstable in this case as well. The period for the 

short period roots was calculated to be 26.23 seconds. Again, this is not a concern since the pilot 

has enough time to compensate. 

The poles plotted in the s-plane for the worst case scenario, i.e. the airplane with two rockets on 

one wing is shown in Figure 39: 

0.15 

0.1 

Longitudinal Stability Modes for Two Rockets on the Airplane 

Real Roots 

Short Period 

0.05 

Imaginary 

0 

-0.05 

-0.1 

-0.15 

-0.2 

-12 -10 -8 -6 -4 -2 0 2 

Real 

Figure 39: Longitudinal Stability Modes for Two Rockets on the Airplane 

67



The natural frequency and period of each of the roots for the longitudinal case for two rockets on 

one wing are summarized in Table 9: 

Table 9: Longitudinal Stability for Two Rockets on the Airplane 

Two Rockets on the Airplane-Longitudinal 



Period 

(sec) 

ζ 



Short Period 0.04±0.13i ω n 0.24 (rad/sec) 26.23 0.005 

As seen in the preceding figure, there are two real, stable roots. The time constants for these 

roots are 10.50 s -1 and 0.19 s -1 . The short period is unstable in this scenario as well, but, it has a 

period of 26.23 seconds which is enough time for the pilot to input control. 

For lateral stability, the modes concerned are spiral divergence, dutch roll and roll subsidence. In 

aircraft design, there exists a tradeoff between spiral divergence and dutch roll stability. The 

equations of motion for lateral stability analysis are shown in Equation 17 and the plant that 

models the dynamic system is represented in Equation 18: 

cos ( − ) 

 

= 

0 

 

0 1 0 0 

0 

0 0 0 

0 1 0 − 

= 

⁄ 

 

0 0 1 0 

0 − ⁄ 0 1 

Equation 17: Equations of Motion in Matrix Form for Lateral Stability 

= 

Equation 18: Representation of the Dynamic System for Lateral Stability 

The visual representation of the poles in the s plane for the three cases described previously are 

shown in Figure 40, Figure 41, and Figure 42. 

68



Lateral Stability Modes for the Bottle on the Airplane 

0.1 

Roll Subsidence 

Spiral Divergence 

Dutch Roll 

0.05 

Imaginary 

0 

-0.05 

-0.1 

-0.15 

-0.2 

-0.8 -0.7 -0.6 -0.5 -0.4 -0.3 -0.2 -0.1 0 0.1 

Real 

Figure 40: Lateral Stability Modes for the Bottle on the Airplane 

The natural frequency and period of each of the roots for the lateral case for the bottle on the 

airplane are summarized in Table 10: 

Table 10: Lateral Stability for the Bottle on the Airplane 

Bottle on the Airplane-Lateral 



Period 

(sec) 

ζ 

Roll Subsidence -0.71 Time constant 0.71 (1/sec) 1.41 -- 

Spiral Divergence 0.0001 Time constant -0.0001 (1/sec) 10,000 -- 

Dutch Roll 0.00±0.11i ω n 0.11 (rad/sec) 58.68 -- 

For the case with the airplane carrying the bottle, the airplane was stable in roll subsidence, 

neutrally stable in spiral divergence and barely unstable in dutch roll. The time constant for the 

roll subsidence mode was determined to be 0.71 s -1 , and the time constant for spiral divergence 

was -0.0001 s -1 and the period for dutch roll was determined to be 58.68 sec and 10,000 seconds 

for the spiral divergence. This is clearly not a concern for stability. 

69



0.06 

0.04 

Lateral Stability Modes for Four Rockets on the Airplane 



Dutch Roll 

0.02 

Imaginary 

0 

-0.02 

-0.04 

-0.06 

-1 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 

Real 

Figure 41: Lateral Stability Modes for Four Rockets on the Airplane 

The natural frequency and period of each of the roots for the lateral case for the four rockets on 

the airplane are summarized in Table 11: 

Table 11: Lateral Stability for Four Rockets on the Airplane 

Four Rockets on the Airplane-Lateral 



Period 

(sec) 

ζ 


Spiral Divergence 0.00 -- -- -- 

Dutch Roll 0.00±0.05i ω n 0.05(rad/sec) 125.7 -- 

The lateral stability for the case with four rockets on the airplane was not much of a concern 

either. This is because the airplane was stable in roll subsidence, neutrally stable in spiral 

divergence and very slightly unstable in dutch roll. The time constant for the roll subsidence 

mode was 0.91 s -1 , and the period for dutch roll was determined to be 125.7 seconds. 

70



0.08 

0.06 

Lateral Stability Modes for Two Rockets on the Airplane 



Dutch Roll 

0.04 

0.02 

Imaginary 

0 

-0.02 

-0.04 

-0.06 

-0.08 

-1.2 -1 -0.8 -0.6 -0.4 -0.2 0 0.2 

Real 

Figure 42: Lateral Stability Modes for Two Rockets on the Airplane 

The natural frequency and period of each of the roots for lateral stability for two rockets on one 

wing are summarized in Table 12: 

Table 12: Lateral Stability for Two Rockets on the Airplane 

Two Rockets on the Airplane-Lateral 



Period (sec) 

ζ 


Spiral Divergence 0.001 -0.001 10,000 -- 

Dutch Roll 0.00±0.065i ω n 0.05 (rad/sec) 96.7 -- 

The case with two rockets on one wing was also analyzed for lateral stability. This was the worst 

case scenario, as mentioned before. Stability for this case was also not an issue because the 

airplane was stable in roll subsidence, neutrally stable in spiral divergence and very slightly 

unstable in dutch roll. The time constant for the roll subsidence mode was 1.21 s -1 , and the period 

for dutch roll was determined to be 96.7 seconds 

71



As shown in the preceding paragraphs, stability for the designed aircraft configuration will not 

be an issue and therefore will not be an obstacle in accomplishing the mission goals. 

8.1.6 Drag Analysis 

With payloads that don’t seem very aerodynamic such as the water bottle, it is necessary to 

ensure that there is enough available thrust to overcome the drag predicted. The drag analysis 

was performed in a program called PowerFLOW [15] . PowerFLOW is a computational fluid 

dynamics tool used to aid in the analysis and visualization of internal and external flows using 

the Lattice-Boltzmann Method (LBM). External 3-dimensional, incompressible flow was used 

for the simulations for drag estimates obtained from this tool. PowerFLOW consists of two 

modules used to set up a simulation and one used to analyze flow results. The module used to 

setup the simulation case is called PowerCASE [16] and the module used to visualize the flow 

results is called PowerVIZ [17] . 

The drag on the airplane was predicted for the case with four rockets on the airplane, the bottle 

on the airplane, the bottle, the rocket, and the airplane without payloads. The drag results from 

the bottle and the rocket in a simulation volume were compared to hand calculations to 

determine if the results produced by PowerFLOW were reasonable. The results from the drag on 

the payload alone as predicted by the software and as calculated by hand are shown in Table 13: 

Table 13: Drag Prediction on the Payload Calculated by Hand and in PowerFLOW 

Payload Drag (By Hand) Re C D Drag (PowerFLOW) 

Rocket 0.082 lb 527473 0.75 0.081 lb 

Bottle 1.64 lb 998681 0.85 0.531 lb 

As seen in the preceding table, it is clear that the drag estimate was very close for the rocket but 

for the bottle, the hand calculation was predicted to be much larger than the result from 

PowerFLOW. The purpose of calculating drag by hand was to verify the results from 

PowerFLOW and evaluate the discrepancy, if any. The reason that the hand calculation for the 

drag on the rocket was so similar to the result obtained from PowerFLOW is because the drag 

coefficient for the exact rocket being used was found to be 0.75 [18] . 

On the other hand, for the bottle, the drag coefficient was estimated as a cylinder from Horener 

[19] . The bottle being used on the airplane is not a perfect cylinder. This is responsible for the 

difference in drag estimates for the bottle. Because the drag prediction for the rocket done by 

hand was accurate (since the correct drag coefficient was used) and this value was different from 

PowerFLOW by 0.001 lb, it is valid to assume that the predictions from PowerFLOW are 

accurate with the caveat that the cases were set up correctly. 

72



The following figure, Figure 43, shows the velocity magnitude around the aircraft with the bottle 

during steady level cruise flight. The blue region seen around the bottle indicates that the flow is 

significantly slower. Also, from the picture it is clear that the non-streamlined shape of the bottle 

creates a significant amount of turbulent flow around the body. The green region around the 

entire system shows that the velocity around the entire configuration is reasonably high for this 

case (red refers to the highest velocity). It is also interesting to note that the disrupted flow 

caused by the bottle does not interfere with the wing. This emphasizes that the selection of the 

flying wing configuration was also optimal for aerodynamics. Had there been a tail present, it is 

possible that the disrupted flow from the bottle might have affected the flow seen by the tail. 

Figure 43: Velocity Magnitude around the Aircraft with the Bottle 

Figure 44 shows the streamlines around the aircraft with the bottle. As expected, the streamlines 

around the trailing edge of the bottle are not extremely smooth due to the blunt nature of the 

bottle. 

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Figure 44: Streamlines around the Aircraft with the Bottle 

The velocity magnitude around the airplane with four rockets on it is shown in Figure 45. Unlike 

the previous case described, the rockets are streamlined and therefore the disrupted airflow 

around them does not appear to be as severe. The flow around the trailing edge of the rockets is 

the slowest and the flow over the entire system is reasonably high (yellow/orange refers to higher 

speeds). 

Figure 45: Velocity Magnitude around the Aircraft with the Rockets 

The streamlines for the airplane carrying four rockets show that the flow around the entire 

system is smooth. This is expected since the airplane and the rockets are streamlined. This is 

seen in Figure 46: 

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Figure 46: Streamlines around the Aircraft with the Rockets 

The drag estimates for the cases with the payloads on the airplane and the airplane without 

payload is summarized in Table 14: 

Table 14: Predicted System Drag for Flight Missions from PowerFLOW 

Flight Mission Reynolds Number Drag 

No Payload on Airplane 708923 1.03 lb 

4 Rockets on Airplane 708923 1.43 lb 

Bottle on Airplane 708923 1.57 lb 

2 Rockets on Airplane 708923 1.30 lb 

The analysis was set up to simulate cruise at expected flight conditions in the competition site, 

Tucson, Arizona. The maximum drag expected is 1.57 lb which is for the case with the bottle on 

the airplane. The results make intuitive sense, since the bottle is not streamlined and therefore 

disrupts the flow around it, causing more drag than the rockets. 

It is expected that the thrust available is more than required to overcome the predicted drag. This 

will be described in the propulsion section. Additionally, this drag prediction will be tested in 

flight for comparison purposes. This is detailed in the test and verification section. 

75



8.2 Missions Mechanical Design Elements 

8.2.1 Wing Store Design Element 

Although the basic operation of this mechanism was discussed in preliminary design, several 

other design considerations had to be made to incorporate the system into a working release 

mechanism. The most important of these considerations was the moment that the store CG 

would generate on the release mechanism (shown in Figure 1). The CG was placed here to keep 

the longitudinal aircraft CG unchanged while the stores are loaded and unloaded. The design 

area was limited by physical attachment points located on the underside of the wing. 

Figure 47: Store Center of Gravity and Resulting Moment at Mechanism 

To allow the wing release mechanism to operate under these conditions, a sliding trigger system 

was placed at the forward portion of the allowable design area. At this point, the mechanism 

produced a reaction force in the positive Z direction. Towards the rear of the store, the store 

mount will actually press up against the underside of the aircraft. Here, a small neodymium 

magnet was placed in a pylon that was attached to the store. This magnet wa attracted to an iron 

plate in the wing. This will help to decrease load time, keep the store directional aligned with the 

aircraft and keep the store secure in negative g-loading flight conditions. To further ensure the 

store did not release prematurely, a 1/8” aluminum tab was added to the back of the wooden 

store pylon. This tab rests on the inside of the reinforced aircraft skin during negative g-loads. 

A diagram of the release system can be seen in Figure 48. 

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Figure 48: Wing Store Overview Detailing the Store Release Process 

To release, the trigger connecting to the forward connection point releases the store-fixed tab 

with a hole for the trigger to slide through. The store then rotates nose down on the rear pylon 

due to its CG location. The forward portion of the pylon will increase in distance from the 

aircraft as the store pivots. This increased distance decreases the magnetic attraction to the iron 

plate embedded in the wing. The store then falls forward as the 1/8” tab slides out of the rear of 

the wing. The store clears the aircraft before the nose of the store travels 6” and hits the ground. 

The detailed design of the release mechanism consisted of the following: A Tru-Fire bow 

release trigger was cut and sized to fit into the aircraft at ½ in tall. Each trigger was rated to 

support 100 lbs, far greater than any flight loads. Two 1½ in aluminum L-wedges were placed 

on either side of the Tru-Fire trigger and bolted in place. Two of these were located in each wing 

with a HS-125MG servo between the pairs. When an inboard or out board store needed to be 

released, the servo arm pulled on a Kevlar string attached to each Tru-Fire trigger, causing that 

trigger to release the store. Due to the use of a string that only allowed the transfer of tension, an 

inboard deployment did not affect an outboard store. The rear store pylons were made of birch 

and contained a magnet and an aluminum tab secured to the top of the pylons with epoxy. The 

aluminum tab protruded 1/8 in over the rear of the pylon. Each rocket was ballasted to 1.5 lbs 

77



using 1/32 in diameter copper BBs and the longitudinal CG of the rocket was aligned with that of 

the aircraft. Competition rockets employed 1/8 in aluminum fins to ensure that they did not 

break on store drops. 

The two connection points on each store were located as far apart from one another as possible 

on the underside of the wing which effectively reduced the moment holding the store in place. 

This meant that the inboard stores experienced a smaller load on the front metal tab and rear 

pylon than the outboard stores due to the greater chord length of the wing. This also meant that 

each store was unique due to inboard/outboard and left wing/right wing combination 

considerations. Each store pylon location was located ½ in from the trailing edge of the aircraft 

and remained clear of aileron movement. 


The centerline store utilized two Tru-Fire triggers located on the front and aft sides of the store. 

An HS-77BB servo actuator was placed between the two mechanisms and connected to the 

release mechanisms with Kevlar string. The servo, string length, and mechanisms were 

calibrated to ensure that they release at the same instant, reducing the chance of a jamming 

scenario. 

. 

Figure 49: Left: Store-Fixed Metal Tab; Right: Tru-Fire Trigger Assembly 

Each Tru-Fire trigger was mounted to the centerline joiner plate using ½ in aluminum L-wedges 

and bolts. 1¼” metal tabs with a hole for the trigger were secured to the store with aluminum 

hose clamps wrapped and tightened about the circumference of the bottle. To release, the 

centerline servo will rotated counter-clockwise until both triggers released their tabs 

simultaneously. Figure 50 shows the centerline release design. 

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Figure 50: Isometric view of Centerline Store and Release Mechanism 

8.2.3 Box Design 

Due to the fact that the box weights were counted in the total system weight, they must be made 

as light as possible while at the same time strong enough to withstand a 6 inch drop during the 

competition technical inspection. Balsa was again chosen as the primary construction material 

for the aircraft box, due to its high strength to weight ratio. In order to increase the box strength 

even further, the isogrid structure shown in Figure 51 lent its inherent strength to the box and 

was able to distribute loads throughout the entire container. For the bottle box, thick hollowed 

foam was used for its ability to restrain the payload on all sides and withstand the shock during 

the drop test. 

Figure 51: Box Isogrid Structure 

79



Foam inserts mounted around the spaces reserved for the aircraft components further ensure no 

damage came to the aircraft or its elements during the drop test. Together, the entire balsa and 

foam structure was projected to weigh 3.1lb. Once the structure and materials were finalized, 

COSMOSWorks [20] was used to determine whether or not the structure could feasibly pass the 

6in drop test without significant damage to the box, or damage to any aircraft components stored 

inside. Figure 52 shows the results of the computer design which showed that the box would 

indeed survive the drop. In case, during testing, the box did fail due to imperfect manufacturing 

(especially on the corners), wedges made from a composite material added to the box sides 

would further increase the strength without significantly adding to overall container weight. 

Figure 52: Box Drop Test Analysis Using COSMOSWorks 

8.3 Propulsion Mechanical Design Elements 

8.3.1 Motor Selection 

The Neu 1100 series motors are custom built electric brushless motors capable of handling 

between 50 and 1500W of power. They are designed to spin as fast as 60,000 rpm. The motor 

needed to be able to handle 400W of continuous power, with 800W short bursts for takeoff. The 

motor also needed to have a low enough rpm/V such that the motor will not exceed 60,000rpm 

with 21.6 V. Table 15 shows the motor model options considered. 

Table 15: Motor Selection 

Neu Motor RPM / V Continuous 

Watts 

Maximum 

Watts 

Weight (g) 

1107 2.5Y 2750 300 600 90 

1110 1.5Y 3350 400 800 114 

1110 2Y 2500 500 1000 114 

1112 1.5Y 2900 600 1200 138 

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The Neu 1110 2Y motor was chosen because it meets all the requirements at the lightest weight. 

In order to increase the torque provided by the motors and increase the efficiency at cruise, the 

motor is geared using a gearbox. The two options available for the Neu 1110 motor are a 4.4:1 

and 6.7:1 gearbox. According to Electricalc [21] ,amp draw by the Neu 1100 motor using a 4.4:1 

gearbox is 60 amps compared to a 24 amp draw with a 6.7:1 gearbox. Since the propulsion 

system was designed to utilize up to 40 amps per motor, the 6.7:1 gearbox was eliminated. 

Therefore a 4.4:1 gearbox was chosen to provide more torque and increase the efficiency of the 

motor at cruise. The motor and gearbox can be seen in Figure 53. 

Figure 53: Neu Motor and Gearbox 

8.3.2 Propeller 

The propeller choice was driven by the requirement to produce enough thrust in order to make 

the 100ft takeoff while providing the most efficiency at cruise. The limiting factor in the overall 

propeller size was ground clearance. The main landing gear were designed to be 7” long, 

allowing the 5.75” water bottle to be stored under the wing while maintaining approximately 1” 

of clearance from the ground. Due to the 5° incidence angle from the nose gear, the ground 

clearance available for the propeller is 8”. In order to reduce the overall landing gear size and 

weight, the size of the propeller was limited to 14”. This allows the propeller to spin freely while 

still maintaining 1” of clearance from the ground. The 1” of ground clearance from the propellers 

are sufficient even during a crosswind landing since the propeller is inboard of the landing gear. 

Thin electric propellers with diameters from 12-14” and of all different pitch were considered. 

The lower the propeller pitch, the more static thrust is generated on takeoff. However, the higher 

the propeller pitch, the more efficient the propeller is at cruise. The propeller sizes shown in 

Table 16 were tested at the maximum theoretical power of 800W (20V and 40 amps per motor 

(80 amps total)) which can be obtained from the battery pack. It is important to note that while 

the battery pack can deliver up to a total of 80 amps (40 amps per motor), the propulsion system 

was sized based upon the assumption that a maximum of 60 amps can be obtained. The battery 

pack can only deliver 80 amps immediately after charging (while the battery pack is still hot). 

However, due to the unknown waiting time at the competition flight line, the battery pack may 

not be used for up to 30 minutes of it being charged, consequently reducing the maximum amp 

draw to 60 amps. In order to generate the required thrust of 8.0lbs total at only 60 amps, each 

motor must be able to produce 5.33 lbs at 40 amps. 

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Table 16: Propeller Options 

Propeller 

Diameter (in) 

Propeller 

Pitch 

Maximum 

Thrust (lb) 

12 6 5.37 

12 12 4.30 

13 6.5 5.06 

13 10 4.78 

14 7 5.87 

14 12 4.65 

The only propellers that met the minimum thrust requirement were the 12in x 6 and the 14in x 7 

propellers. The 14in x 7 propellers, shown in Figure 54, was chosen because it provides the 

maximum amount of static thrust. With a designed maximum power draw of 1200W, dual 

motors utilizing 14in x 7 propellers can produce up to 8.81 lbs of static thrust, enough to make 

the 100ft takeoff distance. 

Figure 54: 14 x 7 APC-E Propeller 

8.4 Structures Mechanical Design Elements 

8.4.1 Wing Bending Model 

Analysis for the deflections and stresses acting on the wing due to lifting loads was performed 

using hand calculations and confirmed by COSMOSWorks simulations. The distributed force 

acting on the beam was estimated using the Treffitz plot and strip forces representing a 3.5g load 

determined from AVL. Many of the structures analysis was checked with Dr. Maute, a team 

advisor [22] . The near parabolic shape was inserted into MATLAB and a 2 nd order polynomial 

function was best fit to the lift distribution. This function can be seen in Equation 19. 

() = −0.001 + 0.0215 + 0.807 

Equation 19: Lift Distribution Estimation 

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Using the 4 th order integration method on Equation 19 produces the following functions for the 

total displacement, displacement slope, moment, and shear force. These can be seen in Equation 

20, Equation 21, Equation 22, and Equation 23, respectively. The code for this can be found in 

Appendix H. 

() = − 0.001 

3 + 0.0215 + 0.807 + 

2 

Equation 20: Shear Force Distribution 

() = − 0.001 

12 + 0.0215 + 0.807 + + 

6 

2 

Equation 21: Bending Moment Distribution 

() = − 0.001 

60 + 0.0215 

24 + 0.807 

6 + 2 + + 

Equation 22: Displacement Slope Distribution 

() = − 0.001 

360 + 0.0215 

120 + 0.807 

24 + 6 + 2 + + 

Equation 23: Displacement Distribution 

Applying boundary conditions of a fixed restraint at the wing root and a free end at the wingtip, 

the following boundary conditions are produced. 

2 = 0; 2 = 0 

(0) = 0; (0) = 0 

Equation 24: Boundary Conditions for Wing Bending 

Using these boundary conditions, the integration constants from the previous equations are 

computed in Equation 25. 

= 0.001 

3 2 

= 0.001 

12 2 

− 0.0215 

2 2 − .807 2 

− 0.0215 

2 2 − .807 

2 − 2 

= 0; = 0 

Equation 25: Constants of Integration 

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To determine the moment of inertia for the airfoil cross-section, a few assumptions were 

made. To simplify the cross-sectional shape, the foam was neglected and a thin-walled ellipse of 

balsa was selected as seen in Figure 55. 

Figure 55: Thin-Walled Ellipse 

Since the taper and sweep are not considered, the mean aerodynamic chord of 16.6 inches was 

selected as the semi-major axis. The semi-minor axis was defined as the airfoil thickness at the 

mean aerodynamic chord. The average thickness-to-chord ratio of 0.9 was used to give the semiminor 

axis a value of 0.375 inches (0.75 for the entire thickness). Using the balsa thickness of 

1/32 inch, Equation 26 produces the moment of inertia [23] of approximately 0.7 in 4 . 

= 4 1 + 3 

 

Equation 26: Moment of Inertial for Thin Walled Ellipse 

8.4.2 Wing Material Selection 

The wing design materials were selected to minimize overall weight while still meeting the 

minimum wing design-to-specifications. A foam-core made of EPS foam (Expanded 

Polystyrene Foam) was selected after the foam-core and skin composite construction technique 

was selected over the traditional rib and spar construction. Additional analysis was done in order 

to select an optimum skin material from balsa, fiber-glass, or carbon fiber. The most significant 

resource employed in the skin material selection process was a Young’s modulus vs. Density 

Ashby plot [24] .The two significant Ashby charts are displayed within Figure 56 and Figure 57, 

the balsa and composites Ashby charts respectively. 

84



Figure 56: Balsa Ashby Chart 

Figure 57: Composites Ashby Chart 

The wing root stress was computed using the Treffitz plot and Equation 19 once more. After 

determining the minimum required material strength from the wing bending analysis, a minimum 

85



skin thickness was calculated for each of the three materials. Using the minimum thickness, 

wing area, and the material density, a wing skin weight was estimated. The estimated wing skin 

weights for each material are provided within Table 17. 

Table 17: Wing Skin Material Comparison 

Property EPS Foam Balsa Fiberglass Carbon Fiber 

(w/ the grain) (E-Fiber) (with epoxy) 

Tensile Strength 16 psi 566 psi 250,200 psi 602,000 psi 

Young’s Modulus 435 psi 261 ksi 1,595 ksi 14,500 ksi 

Density 0.0006 lb/in³ 0.0032 lb/in³ 0.0939 lb/in³ 0.065 lb/in³ 

Weight per wing 0.280 lb 0.0834 lb 0.7737 lb 0.5356 lb 

It should be pointed out that the Young’s modulus for both fiber-glass and carbon fiber are 

orders of magnitude higher than balsa, resulting in a minimum required thickness far thinner than 

what could actually be applied with epoxy. Both fiber-glass and carbon fiber thickness were 

then estimated as the thinnest possible for the team to apply, approximately 0.06 inches. After 

computing the skin weights from these thicknesses, it was determined that the much higher 

density of both composite materials far outweighed the balsa. These thicknesses were the 

minimum required to meet the strength requirements, and due to physical practicalities the 

composite materials are over designed for this aircraft. Since wing weight is a significant design 

driver, balsa sheeting was selected because it satisfies the strength requirements and weighs the 

least of the three materials. 

8.4.3 Wing Stress Analysis 

Using the Young’s Modulus for the balsa skin, the maximum deflection of the wing is 0.62 

inches. This value was then compared later to a COSMOSWorks FEM model in order to 

validate a proper COSMOSWorks study. 

The COSMOSWorks simulations were simulated using the half wing span from the Solidworks 

models [25] . This model varies from the hand calculations due to the balsa sheeting and foam 

core analysis (rather than just the balsa). Other changes include the removal of control surfaces 

and the addition of the propulsion motor reinforcement due to its large size. The restraint for the 

study was defined as fixed at the root and the load was defined by the parabolic 3.5 g lift 

distribution shown in Equation 19. The COSMOSWorks results for URES displacement is 

0.4957 inches. The discrepancy between the hand calculation and COSMOSWorks result is 

25%. This error can be attributed to extra stiffness in the wing due to the addition of the foam 

core. Equation 20 below shows the displacement distribution results for the entire wing half. 

86



Figure 58: Wing Displacement Distribution 

Upon acceptance that the run case in COSMOSWorks produced accurate results, stress 

distributions can be used to determine the maximum stress acting on the wing and if unexpected 

stress concentration are developed. Figure 59 below shows the Von Mises stress distribution for 

the wing half. 

Figure 59: Von Mises Stress Distribution 

87



The maximum stress from COSMOSWorks produces a stress of 951.1 psi at a stress 

concentration close to the wing root. This is expected due to the maximum bending moment at 

the root. Using the balsa tensile yield strength of 2,900psi, the minimum safety factor of the 

wing is 3.05. The wing also shows a stress drop across the motor mount. 

8.4.4 Folding Wing System 

The ability of the aircraft to fit into the 2ft x 2ft x 4ft box is feasible only with a well designed 

folding wingtip that can withstand the forces along the wing, minimize vibrations, and fold 

enough to allow the aircraft ample storage room within the required space. The final wing 

folding design can be seen in Figure 60. 

Figure 60: Wingtip Hinge Design 

First, the hinge selected. A traditional cabinet hinge, a concealed Half-Mortise/Integral hinge, 

was used for its ability to provide a degree of pre-stress to the wingtip to minimize excessive 

movement during flight. Figure 61 shows the integral hinge [26] . 

Figure 61: Integral Hinge 

When the hinge is folded, the spring forces the moving parts into the recessed cavity in the door 

leaf. This allows the hinge to pull the wingtip into the main section of the wing. Both the tip 

and main section are lined along the edge by rubber sheeting. When pressed together, free-play 

in the hinge system is mitigated by the pre-stress and frictional force enacted by the rubber 

bushing. 

88



The hinge is connected to a long balsa spar that runs through most of the length of the wingtip. 

This spar rigidly connects the wingtip to the main body, and distributes the load along the wing. 

It is screwed into the integral hinge and glued into the underside of the wingtip foam. 

Since the door leaf has a relatively small surface area and is metal, it is not feasible to glue this 

into the foam itself. Rather, the hinge is screwed into the balsa mount, which is then glued into 

the foam. The balsa mount increases the surface area where the forces on the hinge are 

distributed, increasing the overall load the entire wingtip hinge system can withstand. The balsa 

mount itself is shown in Figure 62, and the actual placement in the wing can be seen in Figure 

60. 

Figure 62: Balsa Mounting Block 

Finally, in order to increase the strength of the hinge in the direction that it opens as well as 

provide more pre-stress to the rubber sheeting, a small clasp was fastened to the trailing edge 

side of the hinge, seen in Figure 60. This limits movement even further and strengthens the 

entire system to withstand the loads expected. 

8.4.5 Landing Gear Positioning and Stability 

The Buff-2 Bomber must be stable and controllable while on the ground because of the 

possibility of extended ground taxi during the missions. To that end, the main landing gear was 

placed far apart on the aircraft to provide longitudinal and lateral stability as seen in Figure 63. 

The nose wheel is 1.5 inches behind the nose and the main gear wheels were 13 inches aft of the 

nose and 17 inches outboard of the aircraft centerline. The landing gear placement is show in 

Figure 63. 

89



Figure 63: Landing Gear Placement 

8.4.6 Longitudinal and Lateral Ground Stability: 

Longitudinal stability is determined by the angle from the main gear to the aircraft center of 

gravity as seen in Figure 64. 

Figure 64: Longitudinal Stability 

In order for stability to be assured, this angle must not be less than 15°, or else the plane will 

tend to nose into the ground, prompting a propeller strike on landing or even during taxi. The 

aircraft may also tip backwards under the same conditions, rendering ground control impossible. 

With the current landing gear placement, the angle for the Buff-2 Bomber is 20°, giving the 

vehicle ample stability in the longitudinal direction. 

Lateral stability depends on the angle ψ which is seen in Figure 65. In order to be statically 

stable (i.e. when not in motion) the angle ψ must be under 90°. Because the Buff-2 Bomber has 

a lateral stability angle of 85° under the worst loading case (two rockets on one wing), the 

aircraft achieves lateral stability in all circumstances (lateral stability was calculated using AAA 

[27] ). 

90



Figure 65: Lateral Stability Angle Definition 

Once the ground stability was guaranteed, the total load applied on each gear strut was 

determined using the following equations. The determined gear loading values are provided in 

Table 18, including the final weight percentage that each strut carries during the worst loading 

scenario. 

= 

+ 

Equation 27 - Nose Gear Load 

= 

 

( + ) 

Equation 28 - Main Gear Load 

% = 

 

100 

Equation 29 - Weight Distribution Percent on the Nose Gear 

% = 

 

100 

Equation 30 - Weight Distribution Percent on the Main Gear 

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Table 18: Calculating the Load on Each Strut 

Parameter Variable Value Units 

Length from main gear to CG l m 4.00 inches 

Length from nose gear to CG l n 7.50 inches 

Takeoff Weight W to 14.00 lbs 

Number of main gear struts n s 2 none 

Load on the nose gear strut P n 4.87 lbs 

Load on each main gear strut P m 4.57 lbs 

Percent of takeoff weight carried by the nose gear %W n 34.78 % 

Percent of takeoff weight carried by the main gear %W m 65.22 % 

8.4.7 Main Gear Loading Analysis 

The strength of the main gear for the Buff-2 Bomber was determined by analyzing both a beam 

deflection and beam buckling case for the horizontal and vertical forces seen during landing. 

These forces were due to the friction between the wheel and the landing strip (calculated with a 

coefficient of friction 0.2 of to be 9lbs) and the impact force seen when the aircraft first touches 

down (found to be 54lbs), as Figure 66 clarifies. Note that the main gear struts are angled at 80°, 

increasing the overall length to 7.1 inches to preserve the 7 inches clearance of the aircraft. P N is 

the impact load normal to the runway, and P S is the frictional load applied due to rolling friction. 

This angle was done to extend the landing gear span while at the same time anchoring the gear to 

a thick portion of the wing to prevent the structure from tearing out during landing. 

Figure 66: Main Gear with Applied Loads 

Figure 67 shows the beam deflection case Free Body Diagram (FBD). 

92



Figure 67: Beam Deflection Analysis 

The maximum allowable deflection of the beam, set to 0.5in, allows the beam to be analyzed to 

determine the minimum radius of a number of materials, including the final selected Al 2024. 

Knowing the deflection as well as the supposed 3g maximum load on the landing gear (assuming 

heaviest aircraft weight of 14lbs), Equation 31 was twice integrated to solve for the required 

radius of the aluminum strut. The final solution is seen in Equation 32. These equations were 

obtained from Vable [28] . 

Equation 31: Second Order Moment Differential Equation 

Figure 68 shows the beam buckling case FBD. 

Equation 32: Minimum Radius for the Deflection Case 

Figure 68: Beam Buckling Case 

93



A simply supported beam was analyzed, fixed at one end with a load P applied at the free tip of 

the beam. From Vable, the critical load experienced by the beam (i.e. the maximum amount of 

force the column can take before it buckles) is known and seen in Equation 33. Simply 

rearranging the critical load equation, knowing the moment of inertia for a cylinder, the radius 

can be solved for as in Equation 34. The code is provided in Appendix I. 

Equation 33: Beam Buckling Critical Load 

Equation 34: Minimum Radius for the Buckling Case 

The analysis of these different beam cases under the same applied forces proved the deflection of 

the landing gear to be of the higher concern. Thus Equation 32 showed the minimum diameter 

for the Al 2024 material properties was 0.2 inches. In order to both add a safety factor while at 

the same time easing the manufacturing of the landing gear, the final selected landing gear 

diameter is the readily available 0.25 inch diameter Al 2024. 

Once the landing gear was sized, it needed to be integrated into the aircraft structure in such a 

way that the risk of tearing out of the wing was mitigated. The strut is to be attached to a 

plywood plate using two small brackets that strap to the strut structure. The two brackets, seen 

in Figure 69, will then be screwed into the plywood plate. 

Figure 69: Two View of the Landing Gear Structure 

94



The plywood plate distributes the force received on the landing gear into the foam body. This 

prevents puncture through the body and gives the strut a solid anchor into the aircraft. In order 

to minimize the potential of the screws shearing out of the plywood mount, a T-nut will be 

employed on top of the screws. This forces the screws into a metal-on-metal connection, 

strengthening the bond, and also distributing the load acting on each screw over a larger area of 

the plywood. 

8.4.8 Nose Gear Selection 

The nose gear was required to attach to a servo to control steering. Due to this required 

connection, a common commercial off the shelf nose gear strut was selected. The spring steel 

nose gear strut, along with servo control, is mounted to an aluminum T-bar wing joiner as shown 

in Figure 70. This nose gear strut is angled forward to move the mounting T-bar further back in 

the wing. The strut is 7 inches in height and mounts to the same aircraft wheels as the main gear. 

The steel material makes this nose gear strut stronger than the main gear struts with some degree 

of flexibility and allows for a thinner diameter. Due to the superior strength properties of steel 

and the higher quality associated with COTS gear struts, it was determined that this gear strut 

was adequate for the landing design-to-specifications. 

Figure 70: COTS Nose Gear Assembly 

95



8.4.9 Motor Mount 

The motor and propulsion systems were designed to attach to the foam wing. A mounting 

system that integrates into the wing was designed in order to accomplish this. This mounting 

system consists of four major components. The motor is the first component and attaches 

directly into the motor pylon, the second component, as shown in Figure 71. Simple M2.6 

machine screws and aluminum strap are used to secure the main motor body to the pylon. A 

motor pylon was required in order to maintain necessary propeller clearance on a swept wing. 

The motor pylon is to be made from 0.02 inch 2024 aluminum sheet formed such that it was 6.5 

inches long and 1.2 inches in width. 

Figure 71: Motor and Motor Pylon System 

The motor and pylon assembly then integrates into the third component which is a balsa hard 

mount that is installed into the wing structure. The wing itself makes the fourth and final 

component of the motor mounting system. In addition to accommodating the motor, the 

aluminum pylon in the airflow also acts as an effective heat sink for both the attached motor and 

speed controller. The integration of the motor pylon into the balsa hard mount and into the wing 

itself is shown in Figure 72. This balsa insert was designed to distribute the propulsion system’s 

load into a large surface area of the wing foam-core and balsa skin. The motor pylon requires a 

hard point to bolt onto because the foam does not provide adequate structural support. The pylon 

and balsa insert are bolted together using four light ¼ inch nylon bolts through the holes seen in 

Figure 72. Lastly, the balsa insert is glued into a slot cut from the wing before the wing is 

sheeted with the balsa skin. Wire access to the motor is provided by a tunnel bored through the 

foam and balsa to the pylon. 

96



Figure 72: Motor Mount System Integrated into the Wing 

8.4.10 Motor Mount Loading Analysis 

To ensure that the motor mount is capable of supporting the stresses encountered in flight, a 

loading analysis was done using COSMOSWorks. The wing and balsa attachment was 

determined to be strong enough due to the use of both glue and the wing skin to hold the balsa 

insert in place. A stress analysis of the motor on the pylon was done to assure the design of the 

pylon was structurally sufficient. A static load of 6 pounds was applied from the motor to the 

pylon, with the nylon bolts serving as the only restraint. The 6 pound load is the highest 

expected thrust from any propeller configuration with the motor. The results of the 

COSMOSWorks analysis can be seen in Figure 73. A maximum stress of 18.7 ksi was located at 

the nylon bolt holes shown in close-up within Figure 74. However the yield stress of aluminum 

2024 is 58 ksi, which creates a minimum 3.1 safety factor in the pylon design. A strain analysis 

was also done on the entire motor pylon assembly, shown in Figure 75, and it was confirmed that 

all strain deflections were negligible in magnitude. 

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Figure 73: Stress Analysis on Motor and Pylon Assembly 

Figure 74: Maximum Stress on Motor and Pylon Assembly 

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Figure 75: Motor Pylon Strain and Deflection 

Next, another COSMOSWorks analysis was done on the stresses applied to the motor, pylon, 

balsa insert, and nylon screw assembly to factor in the balsa mount’s material strength. The 

results from the entire motor mount assembly analysis are provided in Figure 76. It was 

determined that the maximum stress encountered on the balsa insert was only 210 psi. Balsa 

wood has a minimum compressive strength of 377 psi, so there is a minimum 1.8 safety factor in 

this component. The maximum stress encountered among the four nylon screws was 4.5 ksi. 

The yield stress for the nylon bolts is 20.1 ksi, creating a minimum 4.5 safety factor in the bolt 

components. Each component was determined to have adequate safety factor built into the 

component design, and the motor mount assembly was determined to be structurally sound. 

Figure 76: Motor Mount Maximum Stresses 

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9.0 Electrical Design Elements 

Author: Brett Miller 

Co-Author: Jarryd Allison, Ross DeFranco 

9.1 Propulsion Electrical Design Elements 

9.1.1 Propulsion Electrical Overview 

The diagram shown in Figure 77 shows an overview of the propulsion electrical system required 

to produce thrust to power the aircraft. 

Figure 77: Propulsion Electrical Block Diagram 

For the pilot to control the amount of thrust, a signal is sent from the transmitter to the receiver. 

The signal flows from the receiver, where it is split in two different directions, and flows to each 

speed controller. The speed controllers then draw power from the battery pack and through the 

40 amp fuse. The power then splits and flows to each motor, which spins the propeller, and that 

produces thrust. 

9.1.2 Propulsion Batteries 

The battery selection was one of the most important processes for the aircraft. The battery pack 

itself will end up being the heaviest single components in the aircraft, with an estimated weight 

of 1.2lbs. The battery chemistry was limited to NiCad or NiMH chemistry by competition rules. 

A trade study was conducted comparing top of the line 1400mah NiCad and NiMH batteries to 

determine the best choice for the aircraft as shown in Table 19. Lowest weight, lowest internal 

resistance, and the highest recommended maximum continuous discharge were considered for 

each cell. 

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Table 19: Battery Options 

NiCad Sanyo 

KR 1400AE 

NiMH Intellect 

IB 1400 

Weight (g) 31 27 

Internal Resistance (mΩ) 10 4 

Maximum Continuous 

Discharge (amps) 

10 20 

In all categories, the NiMH battery chemistry outperformed the NiCad battery for this 

application. Therefore, NiMH chemistry battery cells were chosen for the aircraft. 

With the battery chemistry chosen, several specific batteries were analyzed in a trade study to 

choose the exact battery cell. The battery needed to be able to discharge up to 40 amps of current 

continuously, with 80 amps in short bursts. The most important factor in choosing the battery 

was energy density. Four top of the line 1.2V NiMH batteries were analyzed in the trade study 

provided in Table 20. The Elite 1500 battery was chosen because it had offered the most 

capacity at the least weight. 

Table 20: NiMH Battery Selection 

Image 

IB 1400 Elite 1500 Elite 1700 GB 2000 

Capacity (mAh) 1400 1500 1700 2000 

Weight (grams) 27.5 25 28.5 36.5 

Density (mAh /g) 50.9 60 59.6 54.8 

The battery pack was sized in order to meet two requirements; 100ft takeoff and four lap range 

with full water bottle payload weight. In order to generate the required thrust in order to make 

the 100ft takeoff, the battery pack must be able to generate 1200W of power. The battery pack is 

capable of discharging at up to 80 amps if it is taken directly off the charger and used within 10 

minutes of being charged while it is still hot. However, due to the unknown timing of the 

competition flight line, there may be up to a 30 minute wait from when the battery is taken off 

the charger until takeoff. That would limit the maximum amp draw to approximately 60 amps. 

Using the equation P=I*V, the battery pack voltage for takeoff must be at least 20V. Assuming 

€ 

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a battery voltage of 0.9V/cell on takeoff, there must be at least 22 batteries in order for the 

aircraft to make the 100ft takeoff distance. 

Next, the amount of battery cells required to make the four lap range with the full water bottle 

payload was calculated. The cruise power draw required for the full water bottle payload was 

assumed to be 350W. With an average lap time of 1:00 per lap, the battery must be able to run at 

cruise power for 4:00. Although the Elite 1500’s have 1500mah of capacity, the batteries cannot 

discharge this amount when discharged under high amp draw. Therefore, it was assumed that 

only 1000mah of actual usable power could be drawn from the battery pack before the power 

would drop off below a usable level. The rest would be dissipated as heat from the internal 

resistance of the battery cells. 1000mah of usable power means the batteries can be discharged at 

1 amp for an hour, or 15 amps for the required 4:00. While the Elite 1500’s are 1.2V batteries, 

the battery voltage drops while under such high amp draw. It was assumed that each battery 

would have an average voltage of 1.0V. In order to get 350W of power at a maximum average 

amp draw of 15 amps, the equation P=I*Vwas used to calculate the required battery pack 

voltage under load. The battery pack voltage must be at least 23.33V. Therefore, there must be at 

least 24 batteries in order to make the four lap range at full weight. From this analysis, the 

optimal battery pack size was determined to be 24 cells. With 24 cells, the battery pack weighs 

1.3lbs. A picture of the battery pack is shown in Figure 78. 

Figure 78: Battery Pack Overview 

The battery pack was designed like this in order to fit in the very nose of the aircraft. This allows 

the center of gravity to be moved as far forward as possible, increasing the aircraft static margin 

and overall aircraft stability. 

9.1.3 Electronic Speed Controller 

Electronic speed controllers are required for each motor in order to control the amount of power 

flowing from the battery pack to the motor. The speed controllers must be able to handle 33.6V 

and up to 40 amps each. The best speed controllers suggested for Neu Motors are Castle 

Creations Phoenix series. Within the Phoenix speed controller line up, three different speed 

controllers were considered. 

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Table 21: Speed Controllers Options 

Speed Controller Maximum Maximum Weight (g) 

Voltage Amps 

Phoenix 45 19.2 45 30 

Phoenix HV-45 50 45 60 

Phoenix 60 35.0 60 58 

Because of the high voltage requirement, the Phoenix 60 speed controller was chosen. The 

Phoenix 60 was chosen over the Phoenix HV-45 speed controller because it was 2g lighter and is 

rated for higher current. The higher current rating means the speed controller will run cooler and 

will run less risk of overheating. An image of the speed controller is shown in Figure 79. 

Figure 79: Speed Controller 

9.1.4 Wire Gauge 

In order to connect the battery pack at the center of the aircraft to each motor, approximately 1ft 

of wire must be run in each direction. This wire must be able to sustain the amount of current 

flowing without overheating or causing a significant voltage drop, while keeping weight down. 

The wire gauge sizes considered were 16AWG, 14AWG, and 12AWG. The higher the wire 

gauge, the lighter the wire. However, the lower the wire gauge, the more amps the wire is rated 

to handle. This means the wire will not get as hot, and there will be less voltage drop across the 

wire at a given voltage. 

The wire gauge was chosen based upon current rating, voltage drop, and weight. 16AWG wire is 

rated for 22 amps, while 14AWG wire is rated for 32 amps, and 12AWG wire is rated for 41 

amps. These ratings are for continuous current, and can be exceeded for short power surges. 

Each motor may draw up to 20 amps of continuous current and up to 40 amps in short bursts for 

takeoff. 

Next, the weight of one foot of wire and the voltage drop over that span were calculated for each 

wire. The voltage drop was multiplied by the average current in order to get the average power 

lost in the wire. This value was compared to the equivalent weight of a battery required to 

provide that amount of power. The extra weight of the wire was compared to the equivalent 

weight of the extra batteries required to provide that power. 

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With this analysis complete, 14AWG wire was chosen because it proved to be the lowest weight 

option. The 32 amp continuous rating will ensure that the wire will not overheat, and short 40 

amp bursts will be handled by the wire. 

9.1.5 Fuse 

Competition rules state that the propulsion system must contain a 40 amp fuse placed such that 

no portion of the propulsion system sees more than 40 amps of continuous current. A Cooper- 

Bussmann [29] ATC style 40 amp blade fuse was chosen to meet this requirement. The fuse is 

rated for 40 amps of continuous current, with up to 80 amps in short bursts less than 10 seconds. 

The fuse will be placed just behind the battery pack right before the power splits to both motors. 

This will allow the fuse to be easily accessible from the top of the wing for easy arming and 

disarming. An image of the fuse is shown in Figure 80. 

Figure 80: 40 Amp Fuse 

9.2 Avionics Electrical Design Elements 

9.2.1 Avionics Electrical Overview 

The overall responsibility of the avionics subsystem is to implement the communication system 

between the pilot and the plane and to use a telemetry system to record the necessary flight 

characteristics. 

The communication system will consist of a transmitter and receiver that will be able to control 

the servos that are connected to the control surfaces, electronic speed controllers, microcontroller 

and the releasable payloads. Figure 81 illustrates that six channels will be required for the 

transmitter and receiver. Four channels will be utilized to control the aircraft, one will be needed 

for the payload release system, and the last channel will be used to control the propulsion system. 

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Figure 81: Overall Avionics Diagram 

In order to verify that the plane meets certain project requirements, it is necessary to have an 

onboard telemetry system capable of measuring several flight characteristics. The team will be 

using a commercially manufactured onboard telemetry system produced by Eagle Tree Systems 

that was purchased by the previous CUDBF team. The Eagle Tree telemetry system includes a 

small, less than 1.5 oz, flight data recorder that will be placed on the plane. Post flight, the data 

stored on the data recorder will be able to download to a computer via the USB port. The Eagle 

Tree telemetry system is also capable of producing real-time data; however, at this time the 

CUDBF team feels it is unnecessary to utilize this feature. 

9.2.2 Payload Release Microcontroller 

With limited experience designing a developmental board that would allow a microcontroller to 

be programmed, a commercial off the shelf USB development board was purchased, as seen in 

Figure 82. 

Figure 82: USB Development Board 

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The board features a PIC18F2553 microcontroller, which has one compare, capture, and PWM 

(CCP) pin. The CCP pin is used to read in the PWM signal from the receiver. Due to the fact 

that the board only had one CCP pin, only one signal could be read from the transmitter, thus the 

board was used for proof-of concept. In order to read all of the required incoming signals, a 

multiplexer chip was implemented. On the board there are digital input/output pins that can be 

set high or low in software and the state of these pins can be read into the multiplexer chip with 

the corresponding signal from the transmitter. In other words, each signal (rudder, elevator, 

aileron, and switch) will be one of the inputs into the multiplexer and the digital I/O pins will be 

the other input. The output of the multiplexing chip is then connected to the CCP pin and 

depending what the PWM signal is the desired servo will be controlled. Table 22 illustrates how 

the output of the multiplexer is determined and Figure 83 illustrates how everything is wired. 

Table 22: Determination of output from multiplexer 

A (I/O Pin) B (I/O Pin) Switch Aileron Elevator Rudder Output 

L L L X X X L 

L L H X X X H 

L H X L X X L 

L H X H X X H 

H L X X L X L 

H L X X H X H 

H H X X X L L 

H H X X X H H 

Figure 83: Wiring Diagram 

The power distribution board on the diagram illustrates that the PWM signal outputted by the 

microcontroller cannot power the servos; a 5 V battery must be used to provide power to the 

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servos. Also it is important that the electrical devices are properly grounded, to prevent the 

devices from blowing up. 

With the assistance from one of our graduate advisors, Josh Fromm, the team was able to design 

a fully functioning circuit board. The completed circuit board is shown in Figure 84. The circuit 

board layout can be observed in Figure 85. 

Figure 84: Completed Circuit Board 

Figure 85: Circuit Board Schematic 

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9.2.3 Transmitter/Receiver Selection 

As mentioned in the previous section, the transmitter and receiver must be capable of supporting 

six channels; four control channels and two switches. In addition, the transmitter needs to be 

lightweight as it is included in the total weight score at competition. Also to meet requirement 

0.PRJ.7 the transmitter must have a fail-safe mode that is automatically selected during loss of 

transmit signal. During fail-safe the aircraft receiver must select throttle closed, full up elevator, 

and full right or left aileron. 

After in-depth research the team has selected a transmitter and receiver that will fulfill all the 

necessary requirements. The transmitter selected is the Spectrum DX 6i with the BR6000 

receiver [30] , both shown in Figure 86. Together these components have an overall weight of 2.0 

pounds and operate on a frequency of 2.4GHz. 

Figure 86: Transmitter and Receiver 

With the appropriate transmitter and receiver selected, it is important to understand how all of 

the components will be connected to ensure functionally of the communication system. The 

rudder, throttle, aileron, and elevator commands are controlled by the main four channels. For 

extra control over the elevators, the second elevator servo is linked through the fifth channel, the 

flaps channel. The microcontroller arming switch is controlled through the final channel, the 

gear channel. The overview of the wiring can be seen in Figure 85. To save on complexity of 

the wiring inside the aircraft, all Y-harnessing is done through the circuit board. 

9.2.4 Servo Selection 

The servos for the control surfaces and releasable payloads must provide adequate torque and be 

lightweight as the overall weight of the aircraft is an important design factor. The servo 

selection was divided into two groups: aircraft control servos and external store servos. All 

servos were selected based off the torque they can produce using 6 volts, which is the voltage 

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provided by the receiver. The selected servos are from Hitec ® RCD [31] ; Hitec was selected over 

other manufacturers due to superior servo selection and performance. 

The required torque for the controls surfaces were found for the worst case scenario, cruise speed 

and maximum deflection of 25 degrees. The servos selected for each control surface is displayed 

in Table 23. For the nose gear, a servo was selected that would provide adequate torque to allow 

the nose gear to steer the aircraft while on the ground. 

Table 23: Servo Selection for Control Surfaces 

Control Surface Required Torque (oz-in) Selected Servo Servo Torque (oz-in) 

Aileron (2) 38.4 HS-125MG 48.6 

Elevator (2) 47.8 HS-225MG 66.65 

Rudder (2) 22.2 HS-82MG 47.22 

Nose Gear 31 HS-82MG 47.22 

The required torques for the external stores were found for the minimum needed to release the 

stores. 

Table 24: Servo Selection for External Stores and Nose Gear 

Required Torque (oz-in) Selected Servo Servo Torque (oz-in) 

Wing Stores 30 HS-125MG 48.6 

Centerline Store 60 HS-77MG 76.3 

9.2.5 Eagle Tree Telemetry Capabilities 

An overview of what the data recorder is capable of measuring during flight is shown in Figure 

87. This data recorder will be vital to the testing and verification plan for next semester. 

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Figure 87: Capabilities of Data Recorder (Seagull Pro Telemetry System) 

The specific components that will be used for these measurements are described in the following: 

Dual Channel A/D Input Board (DUAL – AD) 

• Allows for team to add analog sensors, such as angle of attack sensor. The analog to 

digital converter has a 15 bit resolution and has two channels. 

Electric Expander -100 Amp (ELEC-EXP-100) 

• Measures motor battery-pack current and voltage. 

G-Force Expander (GFORCE-38) 

• Capable of measuring G-force up to +/- 38 G’s. Also measures dual axis acceleration. 

GPS Expander Module (GPS-STD) 

• Key measurements include: Latitude and longitude, GPS altitude, ground speed, and 

distance to pilot (WAAS compatible). 

Optical RPM Sensor (OPT-RPM) 

• Measures the revolutions per minute of the motor. 

Pitot Static System 

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• Measures airspeed and altitude and verifies measurements from GPS. 

Servo Current Logger (Servo-CURR-LOG) 

• Measures the current draw of each individual servo (0-5 Amps continuously with 

-0.01 Amp resolution). It also measures the servo position and whether or not an error 

occurs in the signal to the servos. 

Thermocouple Expander with CHT Probe Kit (THERM-EXP-CHT) 

• Supports two Type K thermocouple probes and measures the temperature of battery pack 

and motors to be measured. 

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10.0 Software Design Elements 

Author: Brett Miller 

Co-Author: Mark Findley, Jarryd Allison 

The Buff-2 Bomber had two main software components. The first allowed the team to iterate 

their wing design in order to optimize the aircraft’s wing geometry. The second controlled the 

aircraft’s payload through the microcontroller. These two components detailed in the following 

sections below. 

10.1 Aerodynamic Design Software 

The aerodynamics sub-team focused on minimizing weight, configuring the aircraft to be able to 

fit within the dimensions of the storage container, and most importantly being able to fly all 

missions. It was important to choose a geometry without negatively affecting the stability of the 

aircraft. Therefore, analysis was performed to select the optimal geometry configuration, while 

still adhering to project requirements. 

First the airfoil was selected, and the program XFOIL was used to determine certain airfoil 

properties such as lift characteristics, drag, and the moment coefficients. Flying wing airfoils 

documented in the University of Illinois, Urbana-Champaign airfoil database were assessed 

along with the Osborne Design/Build/Fly airfoil collection. From this collection, the drag polars 

were used to select the airfoils to be used. 

Figure 88: Airfoil Selection Flow Diagram 

To analyze the effect on static margin for a given center of gravity location with varying leading 

edge sweep angle and taper ratio, MATLAB code was then created to iterate between many 

different geometries at once, and output files that were compatible with AVL. The output 

geometries from MATLAB were run through the AutoIT program which entered the geometries 

along with commands into AVL. This process yielded the stability derivatives for every 

combination produced in the MATLAB iteration, and these stability derivatives were then reentered 

into a different MATLAB code to determine the static margin of the selected geometry. 

This MATLAB code stored the different sweep, taper, and static margin values, and the results 

plotted in order to observe how sweep and taper affect the static margin. In this way, the ideal 

static margin and final geometry of the Buff-2 Bomber was selected. Figure 89 shows this flow 

diagram. 

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Figure 89: Wing Geometry Determination Flow Diagram 

The stability derivatives calculated from AVL were then used to determine the dimensional 

stability of the aircraft under different loading conditions for the chosen geometry. MATLAB 

code was created to convert the non-dimensional derivatives provided by AVL into dimensional 

derivatives. These were then entered into Equation 15 and Equation 16 along with the inertias 

from SolidWorks for each loading case and mission. The eigenvalues of the system were then 

plotted to determine the stability of the spiral divergence, dutch-roll, and roll subsidence in the 

lateral direction, along with short and long period in the longitudinal direction. The flow 

diagram is seen in Figure 90. 

Figure 90: Stability Determination Flow Diagram 

Finally, drag analysis was needed to establish that the propulsion subsystem provides enough 

thrust for the aircraft to perform as expected. The geometry of the aircraft developed in AVL 

was created as a three-dimensional model in SolidWorks. The model was then subjected to 

Powerflow, and the drag on the body determined. 

Figure 91: Drag Calculation Flow Diagram 

10.2 Avionics Microcontroller Software 

The following explains the software that was programmed to the microcontroller. All coding 

was written in C and MPLAB was used to compile the code and generate a hex file, so it could 

be programmed to the circuit board via a USB cable. 

The microcontroller was programmed such that the pilot had to arm the microcontroller when he 

was ready to release the payload. To activate the microcontroller, the pilot simply flipped a 

switch on the transmitter. The switch on the transmitter produced two different PWM signals 

each with a period of 20.0 ms. One position of the switch produced a signal with a 10% duty 

cycle and the other position produced a signal with a 5% duty cycle. The signal was read into the 

microcontroller and based on the duty cycle, the microcontroller either did nothing (no 

deployment required) or allowed signals to be sent to the payload servos (when deployment was 

required). The logic for arming the PIC can be seen in Figure 92. The microcontroller was able 

to differentiate between the duty cycles by using the capture module associated with the CCP 

pin. In software, it was written such that a flag was set when the CCP pin read a rising edge of 

the signal. After setting a flag, a timer was initiated. Once the microcontroller found the falling 

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edge of the signal, it stopped the timer. The timer value reflected how many instruction cycles it 

took the microcontroller to read the high time of the incoming signal. Knowing that the 

microcontroller operated at a frequency of 48 MHz and executed 4 instructions per clock cycle, 

the timer value was converted to a time. 

Figure 92: Logic to arming microcontroller 

Once the microcontroller was armed it then reads the incoming signals, using the CCP pin as 

before, from the aileron, elevator, and rudder. To avoid accidental store release, there is a time 

delay of three seconds between arming the microcontroller and the microcontroller reading the 

incoming signals. The inputs for these control surfaces are similar to how the switch works. On 

the transmitter there are two sticks that produce various PWM signals depending on their 

position. For instance, moving the left stick to the far left causes the rudder signal to have a high 

time of 1.0 ms or duty cycle of 5% and moving it to the far right gives the rudder a signal with a 

high time of 2.0 ms or duty cycle of 10%. Based on the signals, the appropriate payload is 

released. The flowchart of this process is shown in Figure 93. Every time the microcontroller is 

armed it goes through this sequence once and then returns to the arming routine. Code for the 

Microcontroller can be observed in Appendix K. 

Figure 93: Flowchart for releasing payloads 

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11.0 Integration Plan 

Author: Mark Findley 

Co-Author: Eric Hall 

11.1 Aircraft Overview 

The Buff-2 Bomber is composed of four main mechanical sub-assemblies: wing, structure, 

propulsion, and avionics. The following chapter details the integration plan of the major 

components into the four sub-assemblies and then the integration of the sub-assemblies into the 

complete aircraft. Figure 94 below shows the assembly flow diagram for the construction of the 

CUDBF aircraft. 

Figure 94: Assembly Flow Diagram 

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11.2 Wing Sub-Assembly 

The wing sub-assembly consists of the main wing assembly and the vertical assembly. 

11.2.1 Wing Assembly 

The main wing assembly is composed of a core of 1pcf EPS foam. The foam cores will be 

outsourced due to the geometric twist of the wing. This will ensure that the foam cores are the 

correct dimensions. The balsa motor mount blocks are then glued into the foam using epoxy. 

Also, the main gear mount plywood mount will be glued into the core before sheeting. The foam 

core will then skinned with 1/32” light weight balsa. The control surfaces will also be sheeted 

with 1/32” light weight balsa after a balsa insert is glued into the foam for the control horn. The 

ailerons and elevators will then be attached to the wing halves. Both wing halves will be 

constructed in parallel and can be seen below in Figure 95. 

Figure 95: Main Wing Assembly 

11.2.2 Vertical Tail Assembly 

The foam cores for the verticals will also be outsourced in order to achieve the desired 

tolerances. The verticals and rudders will be sheeted the same way as previously mentioned. 

After they are sheeted, the carbon support tubes will be inserted and glued in place. These will 

then be glued to the balsa vertical spar mount. The rudders will then be attached to each winglet. 

Both vertical tails will be constructed in parallel. A vertical tail can be seen in Figure 96. 

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Figure 96: Vertical Assembly 

The wing assemblies and the winglet assemblies will then be joined. The two wing halves with 

winglets will then be joined with the root wing joiner to form the wing sub-assembly. The wing 

sub-assembly is shown in Figure 97. 

11.3 Structures Sub-Assembly 

Figure 97: Wing Sub-Assembly 

The joint required to fold the wing must be structurally sound since a considerable amount of the 

wing folds. This is also a very critical element of the aircraft. The wing tips need to fold and 

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unfold properly for this aircraft to be successful. High tolerances need to be implemented when 

joining the wing tip to the main wing. 

11.3.1 Folding Wingtip Assembly 

The wing tip mounts are glued into the foam and balsa skinned main wings. The hinge is then 

screwed into the wing tip spar and glued into the wing tip. The hinge is then screwed into the 

wing tip mount. The rubber bushings are also glued into the adjoining surfaces. This assembly 

is shown in Figure 98. 

Figure 98: Folding Wingtip Assembly 

11.3.2 Landing Gear Assembly 

The landing gear assembly consists of the nose gear assembly and the left and right main gear 

assemblies. The main landing gear assembly consists of an quarter inch aluminum 2024 strut 

screwed to the main gear mount with steel straps. The screws screw into tee nuts in the main 

gear mount. The left and right main gear can be constructed in parallel. The strut is screwed the 

main gear mount after the wing is skinned. The main wheels are secured on the aluminum strut 

with ¼” collets. The nose gear assembly is made up of several parts to provide for the steering 

of the aircraft on the ground. It is mounted inside of the wing just behind the battery box. The 

nose gear assembly is shown in Figure 99 and the main landing gear assembly is shown in 

Figure 100. 

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Figure 99: Nose Gear Assembly 

Figure 100: Bottom View of Right Main Landing Gear 

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11.4 Propulsion Sub-Assembly 

All electronic components for the propulsion system are COTS. 

11.4.1 Motor Mount Assembly 

The motor and gear box assembly are secured to the motor pylon. The motor pylon is then 

attached to the main wing by means of the motor block mounts already in the wing. It is attached 

with four ¼ in nylon screws. The top piece of the motor block mount is then replaced back into 

the wing. The speed controller is wired to the motor using bullet connectors and is connected to 

the battery extensions and the receiver extension wire as the whole motor assembly is being 

attached to the wing. Figure 101 shows the installed motor mount assembly. 

Figure 101: Motor Mount Assembly 

11.4.2 Battery Assembly 

The battery pack consists of up to 24 Elite 1500 cells. These cells are soldered together in a “V” 

configuration in order to fit into the very nose of the wing. The battery pack is housed in a box 

made of 3/16” balsa. The battery pack is connected to the battery extension wire by a Deans 

connector. A hatch in the top of the wing allows the battery pack to be removed to be recharged. 

The battery assembly can be seen in Figure 102. 

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11.5 Release Mechanism Sub- Assembly 

Figure 102: Battery Assembly 

11.5.1 Wing Store Release Mechanism 

The main components of this system are the release servo, the Tru-Fire triggers, the metal plates, 

the pull pins, and the rails. This assembly is shown below in Figure 103. The wing store release 

mechanism is installed into the wing tip after it is sheeted but before the wing tip spar is glued 

into place. This allows for proper alignment of the two components into the wing tip. The 

mechanism locations are described in Section 8 with the key points being that their placement 

was measured to keep the centerline of the inboard wing store 24 1/4” from the centerline of the 

aircraft and each store 6 1/8 “ apart to meet competition based design requirements. The rear 

pylon location is reinforced and measured to be ½” from the rear of the wing to avoid the rocket 

fins from interfering with the ailerons. The mechanisms are surrounded by tape during 

construction to prevent glue from the balsa skin from with and seizing the mechanisms and 

servo. 

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Figure 103: Wing Store Release Mechanism 

11.5.2 Centerline Store Release Mechanism 

The centerline release mechanism consists of one servo and two arm hooks. The servo is 

mounted in the center of the root wing joiner and is shown below in Figure 104. It is important 

to note that the plate supporting the centerline store also joins the wings to reduce weight. Like 

the wing release mechanisms, the centerline store is surrounded by a balsa box during 

construction to prevent glue from interacting with and seizing the mechanisms and servo. 

Figure 104: Centerline Store Release Mechanism 

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11.6 Avionics Sub- Assembly 

11.6.1 Receiver Assembly 

The receiver and receiver battery are mounted just behind the nose gear mount and just forward 

of the root wing joiner. The receiver is wired to each servo, including the release mechanism 

servos. 

11.6.2 Servo Assembly 

The servos for the rudders, ailerons, and the elevator are installed after the wing assembly is 

completed. The servos are installed and connected to the control horns with their respective 

linkages. 

11.7 Aircraft Assembly 

The aircraft assembly is shown in Figure 105. 

Figure 105: Aircraft Assembly 

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12.0 Fabrication and Integration 


Co-Authors: Mark Findley, Eric Hall 

12.1 Interior Sub-Assembly 

Figure 106: Interior Landing Gear and Joiner Plate Assembly 

To begin building the aircraft multiple interior subsystems needed to be built in tandem in order 

to integrate these systems into the wing. The nose gear was started by first cutting a 10 inch 

section of aluminum T-bar. Once cut, all sharp edges were filed down and the end corners 

rounded to ease assimilation into the wing halves. A hole sized to fit the steering servo was cut 

into the aluminum bar, and to reduce weight, holes were drilled one inch apart long the shorter 

perpendicular section of the aluminum bar. The servo was set into the T-bar and secured by 

drilling holes into the aluminum bar at the appropriate locations. The COTS landing gear system 

(to include servo connection arm, hard plastic holders, and spar) was then installed into the 

aluminum bar by drilling holes into the aluminum which holds the spar holders via four small 

screws. The spar was then sent through the holders and held in place with the servo connection 

arm. The COTS landing gear was then attached to the servo and the system tested using the 

transmitter. 

The PIC microcontroller was built to attach to the receiver and control all avionics systems of the 

aircraft as well as the release mechanisms. Once the PIC was selected, code was written for the 

microcontroller and tested with the PIC on a breadboard. Once the breadboard prototype was 

completed, a circuit board was designed to house all of the necessary electronic components in a 

compact fashion, so it could easily be integrated into the plane. The components were soldered 

onto the circuit board and the microcontroller tested to ensure that all connections were secure. 

Each release mechanism consisted of a Trufire bow release trigger held in place using small 

sections of aluminum L-bar. To alter the triggers, the hand grip was removed on the band saw 

and the screws that hold the two halves of the system together also removed. The screw holes 

were then extended using a drill and the release trigger bolted to the aluminum holders. In the 

actual releasing part of the system, the spring was manipulated by installing a small screw into 

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the release trigger directly behind the spring. This screw was then tied to a servo, and when 

pulled, activated the mechanism and released any stores that were attached to it. 

The centerline mount was built from a large sheet of plywood. Once cut, the plywood sheet was 

sanded smooth on all sides. Holes were cut into the mount to hold two release mechanisms and 

the servo in between which controlled both mechanisms. Great care was taken to ensure that the 

servo arm lay in the same plane as the release mechanism’s tops, so small balsa blocks were 

sized to raise the servo up to the necessary height. Making sure that the wiring remains long 

enough to thread through the wings, the servo was screwed into the balsa blocks and the release 

mechanisms into the plywood mount. The servo arm was then tied to each release system and 

the centerline mount tested with the centerline store. 

12.2 Exterior Sub-Assembly 

Figure 107: Exterior Vertical and Main Gear Assembly 

The aircraft verticals were built by first removing the rudders using a bandsaw and carefully 

measuring the exact locations where the rudder is located. Once removed, the vertical and 

rudder were sheeted with balsa wood attached with Gorilla Glue © . When the glue set, the pieces 

were removed from their molds and any excess balsa removed. The verticals and rudders were 

then sanded to remove excess glue. At this point, the verticals were shaped to be easily fitted to 

the wingtip of the aircraft. Thus an airfoil hollow which stops halfway through the vertical was 

cut into the bottom of the vertical. This allowed the vertical to sit atop the wingtip snugly. 

Small balsa blocks were attached to the tops of the two pieces and sanded to a curved shape to 

improve aircraft aerodynamics. A cavity was created in the verticals to hold an interior-mounted 

servo which controls the rudder, and at the same time the rudders were monokoted. Holes were 

drilled into the rudder and small hinges glued into the holes, with like holes being drilled into the 

vertical. The vertical was then itself monokoted and the exposed rudder hinges glued into the 

vertical taking care not to cover the hinge with excess glue, which impinged on rudder 

movement. The servo was then attached to the rudder and the system tested. 

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The metal motor mounts were fabricated by printing out a template from SolidWorks © and then 

gluing the paper to thin aluminum sheeting. The shape was cut out using a bandsaw and the 

rough edges filed to prevent injury. Holes were then drilled into the metal in the appropriate 

places marked on the template. Once this was completed, the mounts were bent on a sheet metal 

bender. 

Given their unique shape, the main landing gear could not be purchased and thus had to be built 

from scratch. To begin construction, plywood mounts for the main gear were cut, the corners 

rounded, and blind nuts secured into the mounts to attach the main gear spars to the mounts. 

Al2024 spars were store-bought and the necessary bend locations marked along the spars. Each 

gear spar was bent by hand using a strong table mounted clamp and a hammer, making sure to 

bend the sections attached to the mount at a 90 degree angle, the longest sections at a 10 degree 

angle, and the section which holds the wheel at a 90 degree angle. Once the spar was bent, 1/8 

inch aluminum sheeting was cut into 1.5 inch long strips ¼ inches wide in order to secure the 

main gear strut to the plywood mount. The strips were then hammered over a piece of the 

aluminum spar to bend them into a horseshoe-like shape. Holes were then drilled into the strips, 

which were bolted into the blind nuts on the plywood mounts, securing the spars to the mounts. 

The main gear was then completed by attaching wheels and collars at the base of the spar. 

12.3 Wingtip Sub-Assembly 

Figure 108: Wingtip Interior Sub-Assembly 

While the interior and exterior sub-assemblies were being completed, other team members were 

beginning work on the aircraft wing. To begin fabricating the verticals, the main wing was first 

separated from the wingtip at the fold location using a hot wire cutter. After the cut was made, 

the aileron was removed. The aileron was then sheeted, sanded, and monokoted. Small holes 

were then drilled along the aileron leading edge and small hinges similar to those used in the 

rudder where glued into the aileron. The wingtip was then subjected to a series of hollows using 

a routing bit and a drill press. Holes were made for the wiring, balsa spar, and release 

mechanisms (to include the release mechanism mentioned above as well as the small metal plates 

that attract the magnets located in the wingtip stores and the servo). Once the holes were 

created, the release mechanisms were glued into place and balsa covers glued on top of the 

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apparatus to both protect the device as well as maintain the integrity and shape of the wing. A 

small balsa plate was glued into the servo cavity which allows the servo to be screwed into place. 

Wiring for the rudder, aileron, and release mechanisms was placed in the hole created for it, and 

enough run through the wingtip to attach the vertical, aileron, and wingtip to the main wing. The 

balsa spar and hinge system was then glued into place on the wingtip. Great care was taken to 

ensure that the balsa spar and hinge were properly placed and aligned within the wingtip such 

that no twist or different angles of attack are introduced to the final aircraft. 

Once the spar, wiring, and release mechanisms were installed, the release mechanisms were tied 

to the servo with Kevlar string and paper was taped over all hollows to prevent glue from 

spreading into the hollows. The wingtips were then sheeted with balsa in the same manner as the 

verticals. The wiring and release mechanisms were tested after sheeting to ensure no glue spread 

to the system. Rubber sections were then cut to fit the airfoil shape at the wingtip folding edge 

(this provided some degree of pre-stress for the hinge system). Balsa sheets were first fitted to 

the wingtip edge, and the thin rubber airfoil shapes glues to the wingtip. The location where the 

wingtip stores connect to the aircraft via magnets was then hollowed out to allow the rocket 

magnet to directly contact the installed metal. Next, a small hollow was created for the aileron 

control servo and the servo installed. Like holes were then drilled into the trailing edge which 

matches those created for the aileron. The aileron hinges were glued into place, servo linkages 

were attached to the aileron, and the control surfaces tested. 

12.4 Main Wing Sub-Assembly 

Figure 109: Main Joined Wing Assembly 

Once the cut along the aircraft fold location was made, the main wing could be built in tandem 

with the wingtips. The elevators were removed first using a bandsaw, and the elevators sheeted 

and sanded. These control surfaces were then monokoted and holes drilled along the leading 

edge. Small hinges similar to the ones used in the rudder and aileron were then installed, and the 

elevator was then set aside until later. In the wing, sections where the motors are placed were 

removed from the wing and balsa mounts integrated. The mounts were made by shaping blocks 

carefully to maintain wing shape and tightly hold the metal motor mounts which directly house 

the motors. Once these were glued into place, the wing halves were sheeted. While the wings 

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were drying, balsa mounts, which tightly hold the wingtip hinge systems to the main wing, were 

shaped to fit into the tips of the main wings. After the wings were dry, the excess balsa was 

removed (to include the balsa sheeting covering the motor mount holes) and the wings sanded 

while cavities to house the balsa hinge mounts as well as the wiring, nose gear, and centerline 

mount were created. While removing excess glue, the glue that inevitably formed around the 

nose gear was removed such that the nose gear can rotate unimpeded. The hinge mounts were 

glued into place, the wiring fed throughout the wing halves, and the halves joined with the 

centerline mounts and nose gear spar securing the two halves together. After this glue dried, the 

excess was removed by sanding. Hollows in the wing were then created to house the batteries 

and the receiver as well as the circuit board, which together control the aircraft avionics systems. 

After the hollows were created, the wiring was fed to the motor locations and into the main 

hatch, and battery cables linked to the battery cavity. Holes were cut into the bottom of the wing 

to house the elevator servos, which were mounted onto plywood mounts to secure the servos to 

the aircraft. Additionally, cavities were cut to house the main landing gear to the bottom of the 

airplane. Holes that match the hinge locations on the elevators were drilled into the trailing edge 

of the main wing and the elevators attached. The servos were linked to the elevators and the 

system tested to ensure functionality. 

12.4 Full System Assembly 

Figure 110: Full System Assembly 

To complete the aircraft assembly, the wingtip wiring was soldered to the wiring running 

through the main wing assembly and the hinges screwed into the balsa hinge mounts. The circuit 

board which houses the microcontroller and receiver was connected to the wiring within the 

aircraft main hatch. The verticals could then be attached to the wingtips. Small cavities were 

shaped in both the wingtip and vertical to house two small L shaped aluminum pieces per 

vertical. Glue was then applied to the verticals and wingtips and the verticals attached to the 

aircraft. Although the vertical sits nicely atop the wingtip airfoil shape, there remained a section 

on the bottom of the aircraft where the vertical sits which required a fairing. A small balsa piece 

was shaped and attached to maintain an aerodynamic profile at the bottom of the vertical/wingtip 

connection. The main landing gear was then glued to the bottom of the aircraft. Once full 

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systems functionality was tested, the entire top and bottom of the aircraft was monokoted the 

appropriate colors. Motors were attached to the metal motor mounts and secured to the aircraft 

using nylon bolts, and a perpendicular spar fitted to the nose gear and a wheel attached. The 

aircraft CG was then tested to ensure it is within limits. Finally, the aircraft was ground tested 

before flight. 

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13.0 Verification and Validation 

Author: Ben Kemper 

Co-Author: Ross Defranco 

13.1 Subsystem Verification and Validation 

13.1.1 Missions Subsystem Verification and Validation 

13.1.1.1 Release Mechanism Testing 

Both the wing stores and the centerline store needed to be tested for 95% deployment reliability. 

Payload release testing consisted of loading the rockets and releasing the rockets remotely 

repeatedly. Each individual release mechanism was fully tested as a subsystem before it was 

integrated into the Buff-2B and Buff-2C models. Further testing was done to ensure release 

reliability for both store types by loading them rapidly and then releasing them with only the 

powered servos, PIC, and transmitter. Both types released reliably on 25 out of 25 attempts after 

construction, validating the 95% release rate. After further wear and testing, the Kevlar string 

connecting the servo on the left outboard wing release mechanism broke. This required cutting 

into the aircraft for repair. After further testing and analysis, it was determined the CA used to 

secure the string to the metal pin made the single strand of Kevlar brittle and more likely to snap. 

This was redesigned on the Buff-2C model by using a screw so that CA was not required and 

braiding 3 strands of Kevlar together. Strings were seen to stretch initially after use on the Buff- 

2C model (one requiring tightening), but broken strings never again occurred. 

It was also important to ensure that the stores would remain fixed to the aircraft during flight. 

This was first tested on the ground by performing shake tests on both the aircraft and stores and 

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pulling/pushing on the store in each of the 6 coordinate directions. When properly loaded, the 

final payloads remained fixed to the aircraft throughout all of these conditions. The centerline 

store offered slightly more free-play than the wing mounted stores, but was determined to be safe 

and acceptable for flight on the C model. The centerline store on the B model had added foam 

inserts to prevent possible free-play during flight. In all, the final integrated release system 

proved reliable for release and attachment. 

13.1.1.2 Aircraft Container Drop Testing 

Both the foam box and balsa isogrid box were dropped multiple times in order to verify the 6 

inch drop requirement. Both boxes were loaded with all contents, lifted such that no part of the 

box was below 6 inches, and released. Following each test, both boxes were inspected for 

structural damage and the contents were checked to make sure nothing shifted during the drop. 

Both boxes were dropped 7 times over the course of the project and at least once on each side. 

For the foam box, the only visible damage to the box was foam compression on the surface. This 

damage was superficial and not structural. Similarly, the balsa only experienced surface 

scratching due to the concrete surface the box was being dropped on. Again, the box did not 

sustain damage and its contents did not shift. An inspection of the container’s corner can be seen 

in Figure 111. 

Figure 111: Inspect of the Isogrid Box Structural Corner after Drop Test 

13.1.2 Propulsion Subsystem Verification and Validation 

13.1.2.1 Static Thrust Testing 

Static thrust testing was conducted in order to ensure the aircraft had sufficient power to make 

the 100ft takeoff with the heaviest payload. No values were predicted before this test because of 

the known inaccuracy in the program Electricalc (the program used to determine the static 

thrust). Instead, the purpose of this test was to experimentally measure static thrust as a function 

of power draw for different diameter and pitch propeller sizes. No propellers less than 12” in 

diameter were tested because the ground clearance provided for the water bottle allowed for a 

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12” propeller without any additional landing gear size and weight. Propeller sizes up to 14” in 

diameter were tested incase additional thrust was needed, since the takeoff incidence angle 

allows for propellers of up to 14”. A motor was run at 800W, the maximum theoretical power 

that could be delivered by the battery pack to each motor (20V and 40 amps per motor, or 80 

amps total). 

Table 25: Static Thrust Test Results 

Propeller Diameter (in) Propeller Pitch Maximum Thrust (lb) 

12 6 5.37 

12 12 4.30 

13 6.5 5.06 

13 10 4.78 

14 7 5.87 

14 12 4.65 

This test confirms the expected result that higher diameter propellers as well as lower pitch 

propellers produce more static thrust than lower diameter and higher pitch propellers. In order to 

get the required thrust of 8.0lbs of static thrust at 1200W, each motor must produce 5.33 lbs of 

thrust at 800W. Therefore, this test confirms that the only usable propellers are 12in x 6 and 14in 

x 7 propellers. 

13.1.2.2 Battery Endurance Testing 

In order to determine aircraft endurance, the battery pack was discharged on the ground in a 

mission profile similar to the aircraft when in flight. It was determined that mission two, 

described in 2.2, would require the most battery endurance since the aircraft will be carrying the 

weight of the full water bottle load for four laps. In order to simulate this mission, the battery 

was first discharged at full power for ten seconds to simulate takeoff. The power was then cut in 

half for an additional ten seconds to simulate climb, and then reduced to cruise power until the 

battery pack voltage dropped below 0.9V per cell. The battery pack used in this test can be seen 


Figure 112: Competition Battery Packs 

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The cruise power was estimated to be 350W. Assuming an average battery pack voltage of 1.0V 

per cell (24.0V) during the discharge, an approximate amp draw of 15 amps was required to 

generate 350W of power. While the batteries are designed to hold 1500mah of power, the 

batteries cannot discharge this amount when discharged under such high amp draw (10C). 

Therefore, it was assumed that only 1000mah of actual usable power could be drawn from the 

battery pack before the power would drop off below a usable level. The rest would be dissipated 

as heat from the internal resistance of the battery cells. With an average amp draw of 15 amps 

and 1000mah of usable power, the calculated endurance of the battery pack at cruise speed 

would be 4:00. 

The plots in Figure 113 show the battery pack voltage and power consumption as a function of 

time. 

34 

Battery Voltage 

1200 

Power Usage 

32 

30 

1000 

Battery Voltage (V) 

28 

26 

24 

22 

Power Usage (W) 

800 

600 

400 

20 

18 

200 

16 

0 50 100 150 200 250 300 350 400 

Time (seconds) 

0 

0 50 100 150 200 250 300 350 400 


Figure 113: Battery Voltage and Power Over Time 

The results of the test indicate that the battery pack voltage dropped below 0.9 V/cell at 4 

minutes 10 seconds into the test, and then the power dropped below the 350 W required for 

cruise flight. This was slightly higher than the 4 minutes predicted because the battery pack 

provided about 1050 mah of usable power, above the 1000 mah used in the initial assumptions. 

Based upon the assumption of one minute per lap, this test verifies that the Buff-2 Bomber can in 

fact fly four laps in cruise with the heaviest payload configuration. The battery pack continues to 

provide sufficient power to sustain a slow descent until 4 minutes 30 seconds into the test, when 

the power drops below 200W. This extra power margin was designed to be used as reserve 

power during flight. 

13.1.3 Structures Subsystem Verification and Validation 

A whiffle tree test was performed to determine whether or not the Buff-2 Bomber wing can 

sustain the expected loads seen in flight. The wing was loaded using a whiffle tree system and 

mounted upside down to simulate lift in flight. A right wing half was built according to normal 

construction spec in order to correctly determine the actual strength of the wing. The wing half 

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was mounted via a plywood wing joiner and aluminum spar (as in the actual plane) to a 

2”x4”x30” piece of hardwood and glued into place. The hardwood mount was then bolted to a 

large sheet of plywood and weights were applied to the large plywood sheet such that the wing 

hangs off of an overhang. A top view of the constructed test wing with the mounting apparatus 

is shown in Figure 114 and a root view is provided within Figure 115. 

Figure 114: Test Wing with Mounting Apparatus Top View 

Figure 115: Test Wing with Mounting Apparatus Root View 

The whiffle tree was designed using the lift function derived from the AVL lift distribution plot. 

The lift distribution function is seen in Equation 35. Using the ratio of lift loads at specified span 

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locations, the required whiffle tree configuration was designed as seen in Figure 116, where ∆ 

corresponds to a distance of 5 inches. 

() = −0.001 + 0.0215 + 0.807 

Equation 35: Estimated Wing Loading Function 

Figure 116: Whiffle Tree Final Design 

The results from the test were then compared to a predicted max tip displacement of 0.62 inches 

at 3 g loading, and with a COSMOSWorks FEM model prediction of 0.69 inch max tip 

displacement. The COSMOSWorks FEM model and loading analysis are shown in Figure 117. 

Figure 117: COSMOSWorks FEM Model of Tip Displacement 

The whiffle tree then distributed a single load applied on the bottom rung to model the actual lift 

curve along the wing. The whiffle tree system was composed of multiple straps and metal beams 

placed onto the wing and taped into place to avoid slippage during testing. Weight was applied 

in 5 lb increments to the bottom rung of the whiffle tree to a loading of 3 g’s (22.5 lbs), which 

the wing was designed to withstand. The deflection at this point was then measured by using a 

ruler to measure the difference between the wing tip and an above the wing beam reference. 

This beam reference was a spar mounted above the wing onto the 2”x4” board where 

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measurements can be taken from a fixed location. As the wing deflected, the distance was 

viewed on the measuring stick and recorded for data collection. After applying the 3 g load to 

the wing, additional weight was added in increments of 5 lb, until the wing ultimately broke at 

the wing root. The deflection of the wing at each weight increment beyond the 3 g load was also 

recorded to determine the ultimate failure load. Photographs of the whiffle tree test are shown in 

Figure 118 and Figure 119. 

Figure 118: Whiffle Tree During Loading 

Figure 119: Wing Post-Failure 

The measured wing tip displacement as a function of loaded weight along with the estimated 

accuracy is provided in a plot within Figure 120. The displacement at the hinge location was 

also recorded in order to determine if the majority of the deflection was occurring within the 

wing-tip section itself. As seen within the plot, both the wing tip and the hinge locations 

deflected in similar fashion, implying that the wing itself maintained its shape and deflected as a 

whole. This is in agreement with the visual inspection of the wing during and after the wing 

loading, which observed the wing deflecting as one piece. The wing displacement data is 

provided within the Appendix J. 

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The wing tip displacement was also compared with the COSMOSWorks FEM prediction 

together in Figure 120. The measured values are in red asterisks and the FEM predicted 

displacement is shown in black line. The model agrees with the measured deflection closely, 

especially below a lift load of 30 lbs. The 3g, or 23.5 lbs, predicted deflection of 0.69 inches 

agrees exceptionally well with the measured value of 0.71 inches, thus validating the FEM model 

of the wing used within COSMOSWorks. The maximum breaking load of the wing was 

recorded as 56 lbs, approximately 7.5 g’s; far exceeding the design required 3 g minimum wing 

loading. The recorded maximum displacement before failure was 2.25 inches at the wing tip, 

and 1.03 inches at the hinge location. 

Total Deflection at Tip and Hinge Locations, (in) 

2.5 

2 

1.5 

1 

0.5 

0 

Wing Loading Recorded Tip and Hinge Displacements with Error Bars 

Wing Tip Deflection (Max = 2.25") 

FEM Modeled Wing Tip Deflection 

Hinge Point Deflection (Max = 1.03") 

-0.5 

0 10 20 30 40 50 60 

Loaded Weight = Total Lift of Wing Half, (lbs) 

Figure 120: Wing Tip and Hinge location Displacement vs. Loading Plot with FEM Model Predicted 

Displacement 

13.1.4 Avionics Subsystem Verification and Validation 

In order to verify that the microcontroller was ready to be used in competition, two tests were 

conducted. The first test was to determine if the battery could power the circuit board for the 

duration of the missions. The battery was composed of three 1.5V watch batteries soldered 

together. A code was programmed to the PIC (Programmable Integrated Circuit) to continuously 

run. It was determined that the battery could supply sufficient power for 90 minutes and since 

the missions were only ten minutes in duration, the battery was deemed reliable. The second test 

was to determine if the microcontroller could receive false signals and release the payloads if the 

plane were in flight. The microcontroller was programmed with the complete flight code and the 

137



transmitter was set so the microcontroller was disarmed. A power source was used for the 

transmitter, so the transmitter would not lose power during testing and provide false results. The 

setup of this experiment is shown in Figure 121. 

Figure 121: PIC testing 

The PIC was programmed such that if it got past the arming routine and starts to read the control 

surface signals, an LED would turn on. After 90 minutes of testing, no false signals were read 

by the microcontroller. To ensure communication between the transmitter and the PIC was still 

working, the arming/disarming switch was flipped at the end of the test. When the switch was 

moved to the “arm” position at the end of the test, the LED turned on, confirming no 

communication loss, indicating that the PIC was programmed correctly for the test. With the 

results of these two tests, it was determined that the microcontroller was indeed safe to use for 

competition. 

13.2 System Verification and Validation 

13.2.1Wingtip Lift Test 

Before any flight testing can be completed, had to be lifted off the ground by the wingtips. This 

test simulates a 2.5g load at the root of the wing, and is demonstrated to the judges upon 

technical inspection at competition. This test demonstrates that the aircraft’s structure is flight 

worthy. Each model of the Buff-2 successfully passed the wingtip lift test. 

13.2.2 System Flight Testing 

13.2.2.1 Flight Test #1 

The purpose of flight test #1 was to assess the general aerodynamic flying qualities and test the 

proof-of-concept of the flying wing design with the aerodynamic prototype, Buff-2A. As a safety 

precaution for the first flight, the aircraft was initially ballasted with weight in order to increase 

the static margin to 10% to ensure the aircraft was stable. The goal was to have the pilot takeoff, 

fly for no more than two minutes, and land the aircraft. The purpose of this flight test was purely 

to get qualitative data on aircraft handling characteristics from the pilot. 

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The first flight of the aerodynamic prototype was a success. The pilot reported that the aircraft 

was indeed stable, and the decision was made to remove the ballast to return the aircraft to its 

normal static margin of 5%. A long ground roll was noticed on takeoff. This was attributed to an 

analysis error made early on in the year. The relationship between the Cl in the takeoff distance 

equation and aircraft incidence angle was not recognized, which resulted in a lower aircraft Cl 

than predicted during initial analysis. To account for this lack of lift, an incidence angle of 5° 

was added in order to increase the takeoff C L to 0.6 in order to make the 100 ft takeoff 

requirement. More about this change can be found in section 8.1.3. Pictures from the first flight 

are shown in Figure 122. 

Figure 122: Pictures of Buff-2A flight test #1 

13.2.2.2 Flight Test #2 and #4 

The purpose of these flight tests were to gather more qualitative data on aircraft handling and to 

accustom the pilot to the aircraft before riskier tests were undertaken. Both of these flight tests 

had successful takeoffs and airborne maneuvering portions. However, both of these flight tests 

ended with the nose gear breaking on landing. The main reason for the nose gear failure was that 

the aircraft was designed to sustain normal main-gear-first landings of up to 4.0 g’s. However, 

the nose gear was not designed to sustain the loads encountered by fast nose-gear-first landings. 

The reason for the nose-gear-first landings was due to the pilot still getting accustomed to 

landing the aircraft. The Buff-2 Bomber was designed with a relatively small wing and high stall 

speed in order to reduce aircraft size, fit in the box, and reduces overall weight. However, these 

design decisions indicate that the aircraft must land at a relatively high speed (25 mph when 

empty) and the pilot needed some practice flying the aircraft in order to make purely main-gearfirst 

landings. The nose gear failures that occurred during these two flight tests led to a redesign 

of the nose gear shown in section 8.48. Pictures of the nose gear failure upon landing are shown 


139



Figure 123: Nose gear failures from flight test #2 and #4 


The purpose of this flight was to test the takeoff distance using an added 5° incidence angle. The 

initial flights of the Buff-2A aircraft utilized a LiPo battery in order to test aerodynamic 

characteristics of the aircraft without the limitations inherit to NiMH batteries. Therefore, the 

takeoff distance will not exactly match the performance of the aircraft with a 24-cell competition 

battery pack. The aircraft was able to takeoff in just 25ft, proving that the new incidence angle 

did significantly reduce takeoff ground roll. 

During the flight, the aircraft experienced a left motor failure in flight, causing the aircraft to 

enter a spin and consequently crash. This caused the aircraft to enter a spin and crash. 

Fortunately, the resulting damage was minor: broken propellers, damaged motor mounts and 

damaged nose gear. 

Figure 124: Buff-2A motor failure during flight test #3 

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In order to isolate the cause of the motor failure, the propulsion system was bench tested on the 

ground after the flight. It was determined that there were no issues with the motors, speed 

controllers, or wiring. The problem was determined to be a receiver issue. The receiver used on 

this flight was a 72 MHz receiver instead of a 2.4 GHz receiver, which will be used at 

competition. During bench testing, there was significant interference noticed with the 72 MHz 

receiver. The possible cause of the crash was suspected to be from the faulty receiver which 

caused the motor to cut out in flight. The 72 MHz receiver was replaced with the 2.4 GHz 

receiver. There were no receiver issues after replacing the receiver. 


Flight test #5 was another flight dedicated to getting the pilot used to flying the aircraft. During 

this flight, the aircraft experienced servo travel on the right elevator. The right elevator was 

deflected up about 20° while the pilot commanded neutral elevator. The red arrow shown in 

Figure 125 shows the deflection experienced in flight. The pilot aborted landing twice before 

being able to force the aircraft on the ground during the third attempt. After the flight, it was 

determined that the servo travel occurred because the servo was over-torqued. 

Figure 125: Elevator servo travel experienced on flight test #5 

During analysis, the worst case scenario torque calculated for the elevator servo was 92 oz/in. 

This would only occur at the aircraft top speed of 70 mph, and a full control surface deflection of 

25°. However, it was determined that the pilot would never use full control deflection at top 

speed. Therefore, a lower torque servo was used in order to save weight. The servo torque of 

48.6 oz/in was used since it still allows for 20° of deflection at 60 mph. However, this torque 

proved to be insufficient through flight testing. Therefore, the next size up in servos was chosen, 

providing 66.7oz/in of torque. This allows for full 25° elevator deflection at up to 60 mph, and 

up to 20° of elevator deflection at the top speed of 70 mph. 

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After the pilot had five flights to practice flying the Buff-2 Bomber, and several initial design 

problems had been fixed, the team felt comfortable installing the expensive Eagle Tree telemetry 

system on the aircraft. The purpose of flight test #6 was to collect numeric flight data for the first 

time on the empty aircraft. The goal was to gather data on takeoff speed, landing speed, top 

speed, and average power draw and compare these to the predicted values. 

Takeoff and landing speed were calculated using the standard value of 1.3*V stall . Without any 

payload, the takeoff and landing speed was predicted to be 25 mph. The actual takeoff and 

landing speeds were 27 mph and 23 mph respectively. The slight discrepancy in actual versus 

predicted speed is due to the fact that RC aircraft takeoff and landing speeds are based upon pilot 

judgment, and nearly impossible to fly at the exact chosen speeds. 

After climbing to a safe altitude, the pilot attempted a brief (less than 10 second) full power 

acceleration to see if the aircraft could reach the predicted top speed of 100 ft/s (68 mph). The 

aircraft reached a top speed of 70 mph, exceeding the 100 ft/s top speed requirement set in the 

PDD. The aircraft wasn’t flown at the top speed for an extended period of time because with RC 

aircraft, there is a possible risk that the aircraft becomes unstable at such high speeds. A plot of 

indicated airspeed versus time is shown in Figure 126. The airspeed is accurate to within 5 mph. 

This discrepancy is due to imperfect mounting of the pitot tubes on the aircraft. The top speed 

run occurred 130 seconds into the flight and lasted for only 10 seconds, allowing the aircraft to 

accelerate from 35 mph to 70 mph (plus or minus 3mph). 

70 

Indicated Airspeed vs. Time 

60 

Indicated Airspeed (mph) 

50 

40 

30 

20 

10 

0 

0 20 40 60 80 100 120 140 160 180 200 


Figure 126: Indicated airspeed versus time on flight test #6 

Finally, average current draw was predicted to be 20 amps. The average current draw throughout 

the flight was 15.5 amps. A plot of the predicted versus actual current draw is shown in Figure 

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127. The blue line represents the predicted value, and the red line represents the actual value. 

These values are accurate to within +/- 1 amp as specified by Eagle Tree. 

60 

50 

Power Draw 

Actual 

Predicted 

Amp Draw (amps) 

40 

30 

20 

10 

0 

0 20 40 60 80 100 120 140 160 180 200 


Figure 127: Actual versus predicted amp draw on flight test #6 

The propulsion team assumed that the pilot would fly at the same cruise power setting regardless 

of aircraft weight, and would end up flying faster with the empty aircraft. As it turned out, the 

pilot decided to fly at the same cruise speed, and therefore flew at a lower power setting during 

empty flights. Overall, flight test #6 was a very successful flight in which takeoff speed, landing 

speed, top speed, and average current draw were all tested and shown to be very close to the 

predicted values. 

13.2.2.6 Flight Test #7 and 8 

The goal of flight test #7 and #8 was to fly the aircraft with a full rocket payload. In flight test 

#7, the aircraft was flown with a payload of two symmetric rockets (3.0lb of total payload) in 

order to assess the aerodynamic flying qualities and aircraft performance at a higher weight. 

Another goal of flight test #7 was to fly a competition style lap and verify it using the GPS 

receiver. The GPS coordinates were recorded and plotted on Google Earth. An image of the 

competition lap is shown in Figure 128. 

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Figure 128: Competition lap flown on flight test #7 

After the successful completion of flight test #7, without any pilot complaints, flight test #8 was 

conducted using the full four rocket payload (6.0 lbs of total payload). The goal of this flight was 

to collect data on takeoff speed, landing speed, and average current draw. Takeoff and landing 

speed were predicted using 1.3*Vstall. With the four rocket payload, takeoff and landing speed 

were predicted to be 35 mph. Takeoff speed was 35 mph, and landing speed occurred at 33 mph. 

Figure 129: Flight pictures from flight test #8 

The average current draw was 21.9 amps, slightly above the 20 amps predicted. A plot of the 

predicted versus actual current draw is shown in Figure 130. The blue line represents the 

predicted value, and the red line represents the actual value. The small discrepancy can be 

attributed to variations in pilot flying style. The average speed during the flight was 44.8 mph, 

very close to the 43 mph average speed during the empty flight. This data confirms that the pilot 

chose to fly at a constant airspeed, rather than a constant power setting. Ultimately, all these 

values proved to be close to the predicted values. Flight test #8 ended with a broken nose gear 

upon landing. 

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60 

50 

Power Profile 

Actual 

Predicted 


40 

30 

20 

10 

0 

0 50 100 150 200 250 300 350 


Figure 130: Actual versus predicted amp draw on flight test #8 


The purpose of flight test #9 was to fly the aircraft under asymmetric loads. This flight would 

test the asymmetric load for the smallest lateral CG shift. The configuration was two rockets: one 

inboard on one wing, and the other rocket outboard on the other wing. This resulted in a lateral 

CG shift of 0.9”. A diagram of the loading is shown in Figure 131. 

Figure 131: Asymmetric loading for flight test #9 

During the taxi test before the flight, the pilot noticed a significant lack of ground control. The 

nose gear twisted in one direction, and the servo was unable to turn the wheel in the other 

direction. To compensate for the lack of steering, it was decided to use elevator on takeoff to 

reduce the load on the nose gear and steer using the rudders. This method is commonly used by 

small aircraft for soft field takeoffs. However, this caused the aircraft to lift off at 25 mph instead 

of the predicted 30 mph takeoff speed. In addition to being slow, the aircraft took off at a very 

steep angle. This caused the aircraft to stall shortly after liftoff. The aircraft then rolled the 

opposite direction of the pull of the asymmetric load, and then crashed. This resulted in the loss 

of the Buff-2A, the aerodynamic prototype. 

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The cause of the crash was due to a weakened nose gear servo from previous flight test crashes. 

This resulted in the aircraft unable to steer on takeoff, resulting in a slow takeoff and subsequent 

stall. 

Two major lessons were learned from this flight test. The first is that all aircraft components 

should be carefully inspected before every flight. A flight test checklist was created in order to 

verify that critical systems are checked before flight. The flight test checklist created is shown in 

Figure 132. Another major lesson learned was to “Fix the problem, not compensate for it.” This 

lesson was applied throughout the rest of the project. 

Figure 132: Flight test checklist 


The purpose of flight test #10 was to assess the general flying qualities of the new aircraft, the 

Buff-2B. In addition, this flight test would be the first flight test utilizing the NiMH battery pack. 

A series of events contributed to a crash on landing. First, the transmitter used for previous flight 

tests was under recall. The transmitter that was substituted for this flight did not have the dual 

rate control surface deflections programmed into it. Dual rate control surface deflections allow 

the pilot to have full control surface deflection for takeoff and landing, and then limit deflection 

once airborne in order to avoid over-controlling. Since the aircraft was not carrying any payload, 

the pilot decided it would be better to select the limited control surface deflections, assuming he 

would still have sufficient control for takeoff and landing. Another issue was that the ailerons on 

the first aircraft had been trimmed slightly up for some additional nose up control. The ailerons 

could not be digitally trimmed up on the transmitter being used. Finally, a wire on the battery 

146



pack was frayed during loading. This meant that the battery pack could not deliver full power 

during flight. A picture of the damaged battery pack is show in Figure 133. The green box shows 

an intact wire, while the red box shows the frayed wire. 

Figure 133: Frayed wire on NiMH battery pack 

The aircraft approached nose low on landing, the pilot lacked full control surface authority, the 

usual nose up trim, and was unable to provide full power for a go around. This resulted in the 

aircraft striking the ground extremely hard and nose gear first. Fortunately, only minor damage 

was suffered. The propellers and motor mounts required replacement, and additional screws 

needed to be replaced on a wingtip hinge. Despite hitting the ground nose gear first, there was 

absolutely no damage to the nose gear. This hard impact proved that the nose gear design is 

effective at withstanding hard nose gear first landings and proved to be a vast improvement from 

the original design. 

13.2.2.9 Flight Test #11 and 12 

The purpose of flight test 11 and 12 was to get the Buff-2B airborne following the crash. Flight 

test 11 utilized a LiPo battery. This allowed the pilot to assess the aircraft handling qualities 

without introducing any propulsion battery issues into the mix. After a successful landing, the 

LiPo battery was replaced with the repaired NiMH battery pack. The goal of flight test 12 was to 

test the aircraft handling characteristics of the Buff-2B with the NiMH battery pack. The pilot 

reported he had sufficient power throughout the flight. This confirmed that the frayed battery 

wire caused the lack of power on flight test 10. After these successful flight tests, the aircraft was 

cleared to start flying actual competition missions. A picture of the Buff-2B airborne during 

flight test #12 is shown in Figure 134. 

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Figure 134: Buff-2B airborne during flight test #12 

13.2.2.10 Flight Test #13, 14, 15, and 16 

The purpose of flight test #13 was to assess the aerodynamic handling qualities of the Buff-2B 

with the empty water bottle payload (1.0 lbs of payload). Another goal was to trim the aircraft 

out for future flights with the water bottle payload. 

After the successful flight test #13, the competition course was marked out in order to begin 

simulating competition mission #1. Competition mission #1 consisted of two laps flown as fast 

as possible with the empty water bottle payload. The goal of flight test #14, 15, and 16 was to fly 

the competition course as fast as possible, compare lap time to the predicted value of 1:00 per 

lap, and verify the predicted power draw with the empty water bottle payload. The following 

times represent the time it took to fly two laps from takeoff until crossing the finish line of the 

second lap while airborne. 

Flight test #14 = 2:02 



These flight times closely matched the predicted value of 1:00 per lap, with the pilot improving 

after each flight attempt. The average power draw was recorded on flight test #15 and was 198 

W for the entire flight. This was very close to the predicted value of 200 W. Figures of current 

draw and battery voltage over time are shown in Figure 135. Battery power usage throughout the 

flight is shown in Figure 136. 

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35 

Amp Draw 

25 

Battery Pack Voltage 

30 

24 


25 

20 

15 

10 

5 

0 

0 20 40 60 80 100 120 140 160 180 200 


Voltage (V) 

23 

22 

21 

20 

19 

18 

17 

16 

15 

0 20 40 60 80 100 120 140 160 180 200 


Figure 135: Battery amp draw and voltage versus time on flight test #15 

500 

Power Usage 

450 

400 


350 

300 

250 

200 

150 

100 

50 

0 

0 20 40 60 80 100 120 140 160 180 200 


Figure 136: Battery power draw versus time on flight test #15 

The largest observed discrepancy is the takeoff current draw was much lower than predicted. The 

reason for this was the pilot decided it was unnecessary to use full power on takeoff with a very 

light aircraft, since takeoff distance is a function of weight squared. This is better for the battery 

pack life as it reduces the heat generated on takeoff. The maximum current draw on takeoff was 

only 30 amps, compared to 60 amps that can be drawn when needed. Aside from this 

discrepency in flying style, all other predictions were very close to the predicted value. Some 

photographs from flight test #13-16 are shown in Figure 137. 

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Figure 137: Buff-2B carrying empty water bottle payload 

13.2.2.11 Flight Test #17 

The purpose of flight test #17 was to load the aircraft with ½ the weight of the full bottle, in 

order to get the pilot comfortable flying with a heavier load. The payload consisted of 4.0 lbs of 

ballast placed in the external battery box, a payload designed for Buff-2B flights with the Lipo 

battery. The pilot felt this was necessary due to the extreme change in payload weight from the 

empty bottle (1.0 lbs) to full bottle (9.0 lbs) flight. Since the full bottle weighs more than the 

entire aircraft, it was important to test the CG shift and flight characteristics at a lower weight. 

No quantitative data was measured on this flight since it does not simulate any competition 

missions. The flight test was successful, and the pilot felt comfortable moving on to the full 

bottle payload. Unfortunately, some ground damage occurred after the flight. The full bottle 

flight was pushed back until the next flight test. 

13.2.2.12 Flight Test #18 and #19 

The purpose of flight test #18 was to fly the aircraft with the empty water bottle payload as a 

pilot warm up for the full bottle flight. After a successful landing on flight #18, the pilot was 

ready to attempt the full bottle flight. 

The goal of flight test #19 was to fly the aircraft with the full water bottle payload (9.0 lb 

payload weight). The two requirements that needed to be tested on this flight were the 100 ft 

takeoff requirement under full weight, and the average power draw at full weight. 

The pilot noticed steering issues during taxi tests before the flight. The weight of the full water 

bottle compressed the nose gear significantly, making it difficult to steer the aircraft. The pilot 

decided it would be safer to slowly accelerate rather than push to make the 100 ft takeoff 

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distance and risk going off the runway and destroying the aircraft. The flight was not scrubbed 

because we could still gather the average battery power draw once airborne, and the pilot could 

still get experience flying at the maximum payload weight. 

The pilot was able to get airborne in 150 ft by slowly accelerating up to takeoff speed. Once 

airborne, the pilot did some maneuvering, flew one competition lap, and landed. The average 

power draw was 332 W, less than the 350 W of power draw. This was very significant because 

this proved that the aircraft could indeed make the four lap range with the full water bottle 

payload. A plot of power usage versus time is shown in Figure 138. 

800 

Power Usage 

700 

600 


500 

400 

300 

200 

100 

0 

0 20 40 60 80 100 120 140 160 


Figure 138: Power draw for full water bottle payload flight 

13.2.2.13 Flight Test #20 

The goal of flight test #20 was to fly the aircraft with what would be the best case single rocket 

asymmetric load; one rocket inboard on one wing, and no rockets on the other wing. This led to a 

lateral CG shift of 4.0”. 

On takeoff roll, there was a strong pull to the left (the side on which the asymmetric rocket was 

placed). The aircraft almost went off the runway, so the pilot applied full up elevator in order to 

lift off before going off the runway. The aircraft rolled more than 90°, stalled, and impacted the 

ground. The left wingtip was broken in half, along with damaged motor mounts and propellers. 

The cause of the accident was traced to the fact that the asymmetric load on one wingtip caused 

one of the main gear to compress significantly, causing the aircraft to pull to the left and almost 

go off the runway. The full up elevator applied in order to make the aircraft takeoff only 

exacerbated the turn and led to the stall. The pilot also regretted not using opposite aileron during 

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the takeoff roll and rotation. For any future asymmetric loads, the pilot decided to use some 

opposite aileron input to counteract the asymmetric load. 

13.2.2.14 Flight Test #21 

Flight test #21 was the first flight of the Buff-2C aircraft. The purpose of this flight test was to 

get the aircraft airborne and find out if there were any major issues before competition. The pilot 

reported the plane was trimmed, and that the aircraft had plenty of thrust at the lower density 

altitude. 

13.2.2.15 Flight Test #22 (Competition flight #1) 

The goal of this flight test was to successfully complete competition mission #1 at the DBF 

competition. The requirement tested was to complete two laps under 2:00. On this flight, the 

pilot pushed the aircraft harder than he ever had, and the time to complete two laps from takeoff 

to crossing the finish line was 1:43, 15 seconds faster than before. The flight ended in a good 

landing. A picture of the Buff-2C shortly after takeoff is shown in Figure 139. 

Figure 139: Buff-2C after takeoff on mission #1 at DBF competition 

13.2.2.16 Flight Test #23 (Competition flight #2) 

The goal of this flight test was to successfully complete competition mission #2 at the DBF 

competition. The two requirements to be tested were the 100ft takeoff with full payload weight, 

and the four lap range with full payload weight. The aircraft was able to lift off at approximately 

90ft, verifying the 100ft takeoff requirement at a 5,000ft density altitude. This was the distance 

that the aircraft was designed to takeoff in. After takeoff, the pilot noticed he was holding a lot of 

up elevator, and decided to reduce up elevator in order to avoid a stall. This caused the nose of 

the aircraft to drop significantly resulting in a crash. 

The primary reason for this crash was lack of pilot experience flying with the full water bottle. It 

was deemed a very risky flight, and there was very little time available to test fly with this 

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payload. The team learned the value of flight testing and gaining flight experience is very 

important to project success. 

13.2.3 System Requirements Not Tested or Verified 

Two requirements were not tested. The first was the four lap aircraft range. This was supposed to 

be tested during competition flight #2, however the aircraft crashed shortly after takeoff. While 

this requirement was not directly tested, the four lap range was verified using a variety of 

subsystem tests. 

The average lap time was verified to be 1:00, and the average power draw during the full payload 

flight was tested to be 332W. The battery pack was tested on the ground during endurance 

testing to provide 350W of power for 4:10, before using the reserve power. This was greater than 

the power required, and for a longer period of time than it takes to fly four laps. Therefore, the 

four lap range was able to be verified using a variety of ground tests. 

The major requirement that could not be tested or verified was the ability of the aircraft to 

successfully fly under asymmetric loads. This test was attempted twice. Flight test #9 resulted in 

the loss of the Buff-2A, while flight test #20 resulted in major damage to the Buff-2B. Both of 

these failures were a result of lack of steering on take-off roll, and not necessarily due to the 

asymmetric load. The asymmetric load flight was going to be attempted during competition on 

the Buff-2C (which utilized wheels that did not compress as much, and a stronger nose gear 

servo for steering). However, the team ran out of time and was unable to attempt this flight test. 

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14.0 Project Management Plan 

Author: Daniel Colwell 

Co-Author: Shivali Bidaiah 

14.1 Organizational Responsibilities 

This project’s organization begins at the top with the customer, Dr. Brian Argrow. The 

customer’s requirements were provided to the project manager. The specialty engineers (safety, 

systems and software) take the requirements from the project manager and relay them to the 

subsystem technical leads. The safety engineer also oversees both the fabrication and testing 

engineers to ensure no test or manufacturing process presents a safety issue to a team member’s 

well being. Each technical subsystem consists of a lead engineer and multiple engineers, 

including underclassmen. The project’s webmaster and CFO operate in conjunction with the 

project manager. Figure 140 illustrates these concepts. 

Figure 140: Project Organizational Responsibilities 

The team member’s positions were determined with the help of the Myers-Briggs personality 

tests. The test resulted in 2 ENTJ (Field Marshal), 1 INFP (Healer), 2 INTJ (Mastermind), 2 

ESTJ (Supervisors), and 1 INFJ (Counselor). It was determined that the 2 masterminds (who 

also had the most DBF experience) would become the team’s project manager and systems 

engineer. A supervisor was chosen as the safety engineer to ensure maximum safety during the 

154



project’s manufacturing and testing phases. Subsystem technical leads were chosen based on 

experience and interest in the given field. 

14.2 Work Breakdown Structure 

The breakdown of each subsystem’s work structure was determined based on technical 

responsibility. The program manager and system engineer focused on team organization and 

subsystem integration, respectively. Each subsystem has specific technical objectives which 

must be researched and designed to. The technical subsystems communicated with each other 

with the assistance of the project manager and systems engineer. Figure 141 is a diagram of this 

project’s work breakdown structure. 

Figure 141: Work Breakdown Structure 

14.3 Construction and Testing Schedule Analysis 

At the end of the fall semester, a detailed construction and testing schedule was drafted in order 

to organize the team towards completing the major goals of this project. In hindsight, the 

predicted schedule and actual schedule differd greatly. Both the actual and predicted timelines 

can be seen in Figure 142. 

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Figure 142: Predicted (Black) and Actual (Blue) Schedule 

14.3.1 Buff-2A Construction and Testing Schedule 

The aerodynamic model, the Buff-2A, was constructed and test flown on time. Due to the time 

required to make numerous nose gear repairs, the duration of the Buff-2A’s flight testing was 

carried out longer than originally anticipated. The tests that the team wished to complete with 

the Buff-2A therefore extended past its predicted time and the work spent on repairs caused 

delays for other aspects of the project. 

14.3.2 Buff-2B Construction and Testing Schedule 

The initial construction of the Buff-2B was delayed for two reasons. First, the numerous repairs 

to the Buff-2A drew attention away from early construction on the Buff-2B. Second, the foam 

cores, which initially were to be ordered before the Winter Break, were instead ordered after. 

Waiting for these materials to arrive delayed the start of construction. The construction duration 

also took longer than anticipated. It was originally assumed that lessons learned from the 

aerodynamic model would allow the construction of the Buff-2B to be faster. Instead, 

construction techniques needed to be learned in order to construct the elements not included in 

the Buff-2A. This led to a much longer construction time than originally expected and thus 

delayed the flight schedule. 

14.3.3 Buff-2C Construction and Testing Schedule 

The construction of the Buff-2C was delayed due to the delay of the Buff-2B. The construction 

duration of the Buff-2C was similar to the predicted duration. This unfortunately left little test 

flight time for the Buff-2C before competition. 

14.4 Project Budget Analysis 

The overall project budget was estimated from past projects as well as quotes provided by 

suppliers. The budget was divided into two equally important components of aircraft 

construction and travel as both are needed for the project. 

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The funding available for the project was provided by the Aerospace Engineering Sciences 

Department at the University of Colorado, the Engineering Excellence Fund, and Lockheed 

Martin Corporation. Each internally funded project receives $4,000 funding from AES. Since 

this funding was inadequate for the entire project, additional funding was sought from EEF and 

Lockheed Martin. The team applied for and received funding from EEF on the order of $2,000. 

Lockheed Martin, a longtime supporter of CUDBF and RECUV, also supported the project by 

providing a generous donation of $10,000. Funding under these sponsors brought the CUDBF 

budget to a maximum of $16,000. 

Aircraft construction was composed of four major categories: wing, propulsion, avionics, and 

missions. A breakdown of the entire project budget can be observed in Figure 143 on page 158 

and Table 26 on page 159. The total estimated cost of the project was $15,162 including a 25% 

margin added to all spending. This placed the project under the $16,000 limit by $838. The 

actual cost of the budget was $13,955. 

13.4.1 Wing Budget 

The major components for the wing construction were the foam wing cores and the balsa 

sheeting. Based on previous contact with FlyingFoam.com, the estimated cost for custom cut 

foam cores is approximately $1,500 for 12 sets of wings. Balsa sheeting will be purchased from 

Specialized Balsa and is quoted to cost $957. The reinforcements and other structural 

components will cost approximately $450 while adhesives cost an estimated $300. The total cost 

to build all the wings is approximately $3,291. Due to an overestimation of the foam core costs 

from FlyingFoam.com, the total cost of the wing construction was cheaper than expected. The 

actual cost of the wing construction was $3,107. 

13.4.2 Propulsion Budget 

Propulsion components were comprised of the battery cells, motors, gearboxes, and ESCs. 

Competition battery cells were been ordered from CheapBatteryPacks.com with a discount value 

of 25% for a total of $180. An additional LiPo battery was ordered to allow for longer test 

flights for $200. Neu Motors agreed to a 50% discount on all parts, bringing the motor and 

gearbox costs to $375 and $300, respectively. A 30% discount was received from Castle 

Creations, allowing the purchase of ESCs to cost $686. With the addition of propellers and 

wiring, the total propulsion team cost was estimated to be $1,841. The actual costs of propulsion 

were much higher than anticipated. Backup motors, batteries, propellers, and speed controllers 

were ordered to replace broken items during testing. Additional shipping costs also inflated the 

expenditures. The total budget for propulsion was $3,128. 

13.4.3 Avionics Budget 

The avionics system was comprised of the transmitter, receiver and servos. The team needed to 

purchase a DX6i transmitter and BR7000 receiver from Spektrum for $320. Servos and 

connectors were purchased from ServoCity.com for about $680. The total budget for avionics 

was estimated to be $1,000. The actual costs of avionics was $1,551. This was due to ordering a 

157



backup BR6000 receiver, and multiple new servos to replace malfunctioning servos during the 

testing phase. 

13.4.4 Missions Budget 

The team needed to purchase payloads as well as construct the payload release system. Estes 

rockets were ordered from ACSupplyco.com for a discounted price of $10 per rocket. Eight 

rockets were ordered in order to have a full flying set as well as replacements. Two water bottles 

were ordered from McMaster-Carr for a total price of $50. The components for the release 

system were ordered from McMaster-Carr for approximately $80. Finally, balsa material was 

ordered from Specialized Balsa for an estimated price of $200. The total cost of the missions 

and payload is estimated to be $410. The actual missions costs was $681. Similar to avionics, 

the missions team needed to order new servos in order to replace malfunctioning servos. 

13.4.5 Travel Budget 

The total cost of travel consists of food, hotel, and transportation costs. The team will depart 

Boulder on Thursday, April 16 th and leave Tucson on Sunday, April 19 th . A $40 food budget is 

required to be supplied to each person per day by law. Each student traveling will therefore 

require $160 for the entire trip. A hotel reservation for a two bedroom at the Days Inn in Tucson 

is $72 per night, totaling $108 per person for the entire stay. Therefore taking 16 students to 

competition will cost $4,288. 

Renting a large van for the duration of the trip will cost approximately $300. The RECUV trailer 

will be towed to the competition site by Eric Hall’s GMC Sierra 150. The drive from Tucson to 

Boulder is 921 miles. After adding travel within Tucson of 180 miles, the total round trip will be 

approximately 2,022 miles. Assuming gasoline prices of $2.50 per gallon and gas mileage of 

12.5 miles per gallon, the total gas cost is estimated to be $808. The total cost of the trip 

combining student and transportation costs is currently estimated to be $5,388. 

Cost Breakdown 

Propulsion 

22% 

Travel 

36% 

Avionics 

11% 

Missions 

5% 

Wing 

Construction 

22% 

Administrative 

4% 

Figure 143: Project Budget Breakdown 

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Table 26: Project Budget Breakdown 

Item Predicted Cost ($) Actual Cost ($) 

Propulsion 1,841.00 3,128.71 

Wing Construction 3,291.00 3,107.62 

Administrative 200.00 486.35 

Missions 410.00 681.40 

Avionics 1,000.00 1,551.04 

Travel 5,388.00 5,000.00 

Total 12,130.00 13,955.12 

Figure 144 shows the costs incurred over the project lifecycle. As expected, the costs incurred by 

the project were lowest in the conceptual phase and increased as the project evolved into the 

feasibility, detailed design and implementation phases. The costs in the implementation phase 

were the highest as this was the fabrication and test phase of the project. In the termination 

phase, the costs leveled out with the project closure, as expected. 

Cost Over the Project Lifecycle 

3,500.00 

3,000.00 

Conceptua 

l Phase 

Feasibility 

Phase 

Detailed 

Design Phase 

Implementation 

Phase 

Termination 

Phase 

2,500.00 

Cost ($) 

2,000.00 

1,500.00 

1,000.00 

500.00 

0.00 

Resources 

Used 

SEP 08 OCT 08 NOV 08 DEC 08 JAN 09 FEB 09 MAR 09 APR 09 

Cost 

Figure 144: Costs over Project Life Cycle 

14.5 Specialized Facilities and Resources 

14.5.1 RECUV Fabrication Lab 

As requested by our customer Dr. Brian Argrow, the CUDBF team will operate primarily out of 

the RECUV Fabrication Lab. All team members were given access to the lab in order to be able 

to work without restrictions. The lab offered a place for the team to meet as well as assemble 

and store the aircraft. The lab has three foam cutting wires of 28 inches, 40 inches, and 52 

inches in length as well as two power supplies to heat the wire. In the back room of the lab there 

are two belt/disc sanders, table sander, jig saw, band saw, and a drill press capable of satisfying 

most of the team’s wood and metal working needs. Finally, the lab has electronic capabilities 

159



required for assembling the aircraft propulsion system and battery chargers needed for the 

battery packs. 

14.5.2 Boulder Aeromodeling Society Airfield 

In order to test fly the aircraft safely under AMA requirements, the aircraft needs to be test flown 

at an AMA approved airfield. The BAS airfield was the closest and most accommodating 

airfield for the project. A requirement to use the airfield is to be a member of the AMA. Four 

team members were AMA members. Two members were also BAS members, which allowed 

the team to have copies of the key to the lock at the airfield, providing access at all times. 

14.5.3 AES Machine and Electronics Shop 

The team had access to the AES Machine Shop operated by Matt Rhode and the AES Electronics 

Shop operated by Trudy Schwartz. This provided the team access to specialized machines not 

available in the RECUV Fabrication Lab during regular business hours. 

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15.0 Lessons Learned 


15.1 Manufacturing Lessons Learned 

Since most components were built without the use of a CNC, the importance of precision and 

detail in the manufacturing process was a lesson learned among the team. The precision of 

several parts also improved between the three aircraft that were built. Another useful lesson 

learned was developing a naming convention when labeling SolidWorks files. Since several 

SolidWorks files were used to determine the dimensions to build parts, it became apparent that 

having a standard naming convention would make the process seamless, without requiring the 

presence of other team members to obtain dimensions. A manufacturing checklist would have 

kept the team up-to-date on the status of each component and it might have made the 

manufacturing process more efficient. Another important lesson learned during manufacturing 

was to improve knowledge redundancy across the team. For example, certain members 

developed expertise in building specific components. If more than one member had this 

expertise, it would have allowed for a more flexible schedule. Finally, a very important lesson 

learned was to be prepared for the worst during the process which include but are not limited to 

mishaps with gorilla glue, machining mistakes, and accidents. 

15.2 Testing Lessons Learned 

Since testing was a large part of assessing the system performance, the team learned that it is 

very important to be prepared before every flight test. A flight test checklist was created to 

ensure that all components were inspected before flying to avoid a crash. Another important 

lesson learned during testing was that any problems that arose had to be fixed rather than 

compensated for, and all assumptions and initial calculations that were performed needed to be 

re-visited. The most important lesson learned in testing was that flight tests needed to happen 

frequently because it allowed the team to assess the overall performance of the system. The test 

schedule must have a buffer to account for unpredictable weather. 

15.3 General Lessons Learned 

The major lessons learned for the project overall were to have a better schedule, improve team 

communication and to prepare for the worst. At the end of the project, there was a large 

discrepancy between the actual and predicted timelines in the project schedule. In order to 

prevent this in the future, it will be helpful to have a day-to-day schedule with each team 

member’s task. This will help evaluate each team member’s performance and will also 

demonstrate the work that each member is accountable for. It was also learned throughout this 

project that many unexpected events occur in the project life-cycle and it is extremely difficult to 

account for these events in an initial risk assessment. Therefore, it is always good to be prepared 

for the worst so that unexpected events can be handled appropriately. 

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16.0 Acknowledgements 

16.1 Professional Advisors 

The CUDBF team would like to thank Dr. Brian Argrow for being the team’s customer and a 

faculty advisor. The faculty advisors Dr. Donna Gerren and Kurt Maute, Senior Instructor Trudy 

Schwartz, and Facilities Coordinator Matt Rhode have given their valuable time and years of 

experience to help the team succeed this semester. Finally, the team would like to thank Scott 

Eichelberger and Kelvin Quarels for their advice throughout the semester. 

16.2 Graduate Advisors 

The team would like to thank graduate advisors Josh Fromm, Jason Roadman and Spencer Riggs 

for their valuable input throughout the design process. CUDBF would also like to thank Stefan 

Elsener and Oleg Usmanov, alumni of both CUDBF and the University of Colorado, for their 

assistance. 

16.3 Undergraduate Assistants 

CUDBF would like to thank our many underclass assistants. Their involvement is a project 

requirement and their contributions have been invaluable. Aaron Russert, Alex Wilkins, 

Alexander Granrud, Brandon Bosomworth, Caleb Bloodworth, Cameron Trussel, Cassie Clark, 

Colin Apke, Elliott Richerson, Emily Howard, Jacob Varhus, Josh Yeaton, Robert Rogers, Scott 

Brown, Tom Wormer, Vu Nguyen, Wences Shaw-Cortez, and Zach Dischner. 

16.4 Student Assistance 

The CUDBF team would like thank Jeff Mullen for his assistance in using PowerFLOW 

software. Nate Weigle also deserves thanks for his assistance for testing the nature of the foamcomposite 

material. Special thanks to David Berman for his assistance on the AutoIt software. 

16.5 Experienced RC Advisors 

This project would like to thank James Mack for his input and agreeing to pilot our aircraft. The 

team also thanks RC extraordinaire Frank Dilatush for his advice. 

16.6 Sponsors 

The CUDBF team would like to thank Lockheed Martin Corp. and EEF for their sponsorship on 

this project. This project would not be possible without their generous help. A special thanks to 

Neu Motors, Castle Creations, RC Hobbies, TruFire, and Cheap Battery Packs for discounts on 

their products. 

162



17.0 References 

1 "DBF Rules." AIAA Student Design/Build/Fly Competition. 13 Dec. 2008 

. 

2 "McMaster-Carr." McMaster-Carr. 13 Dec. 2008 . 

3 "AC Supply Wholesale Educational products." Welcome to AC Supply Co. 13 Dec. 2008 

. 

4 " TerraBreak.org: An independent Design/Build/Fly support site ." TerraBreak.org: An 

independent Design/Build/Fly support site . 13 Dec. 2008 . 

5 Drela, Mark, and Harold Youngren. Athena Vortex Lattice (AVL). Computer software. AVL. 4 

Aug. 2008. 13 Dec. 2008 . 

6 The MathWorks. Matlab. Vers. R2008b. Computer software. 

7 Bennett, Jonathan. "AutoIT V3."AutoIT. 13 Dec. 2008 

. 

8 Drela, Mark, and Harold Youngren. XFOIL. Vers. 6.97. Computer software. XFOIL. 7 Apr. 

2008. 13 Dec. 2008 . 

9 "T&J Models - Hughes H-1." T&J Models - homepage. 13 Dec. 2008 

. 

10 Gerren, Donna. "Aircraft Design." Aircraft Design. University of Colorado, Boulder, CO. 

11 Roskam, Jan. Airplane Design: Layout Design of Landing Gear & Systems. Lawremce, 

Kansas: Design Analysis & Research, 2000. 

12 National Instruments. LabView. Vers. 8.6. Computer software. 

13 "Eagle Tree Systems." 2005. 13 Dec. 2008 . 

14 "Neu Motors." Brantuas. 13 Dec. 2008 . 

1 5 "Servos." JR Radios. Horizon Hobbies. . 

16 Exa Corporation. PowerFLOW. Vers. 2008. Computer software. 

17 Exa Corporation. PowerCASE. Vers. 2008. Computer software. 

18 Exa Corporation. PowerVIZ. Vers. 2008. Computer software. 

19 "Terminal Velocity (gravity and drag)." Space Flight Systems Directorate / Glenn Research 

Center. 13 Dec. 2008 . 

163



20 Hoerner, Sighard F.. Aerodynamic Drag: Practical Data on Aerodynamic Drag. Brick Town, 

N.J.: Sighard F. Hoerner, 1951. 

21 "Strong Neodymium Magnets Rare Earth K&J Magnetics."Strong Neodymium Magnets Rare 

Earth K&J Magnetics. 13 Dec. 2008 . 

22 COSMOSWorks. Computer software. SolidWorks. 28 Oct. 2008 

. 

23 SLKelectronics. ElectriCalc. Vers. 2.2. Computer software. ElectriCalc. 26 Aug. 2006. 13 

Dec. 2008 . 

24 Maute, Kurt. Lecture. Advisor Meetings. University of Colorado, Boulder, CO. 

25 Ashby, Michael, David Cebon, and Hugh Shercliff. Materials: Engineering, Science, 

Processing and Design. St. Louis: Butterworth-Heinemann, 2007. 

26 "Young's Modulus - Density." Cambridge University Engineering Dept - Materials Group. 27 

Nov. 2008 . 

27 Dassault Systems SolidWorks Corporation. SolidWorks. Vers. 2008. Computer software. 

28 "IKEA | Built-in kitchens | AKURUM/RATIONELL system | INTEGRAL | Hinge." Welcome 

to IKEA.com. 28 Nov. 2008 . 

29 Advanced Aircraft Analysis 3.12. Released by Design, Analysis, Research Corporation 

(DARCorporation). 2008 

30 Vable, Madhukar. Mechanics of Materials. New York: Oxford University Press, USA, 2002. 

31 "Cooper Bussmann - Cooper Bussmann Home." Cooper Bussmann - Cooper Bussmann 

Home. 13 Dec. 2008 . 

32 "Spektrum RC." 2005. Horizon Hobby, Inc. 13 Dec. 2008 . 

33 "Hitec RCD." 2007. 13 Dec. 2008 . 

164

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