uncc 49er rocketry team · uncc 49er rocketry team university of north carolina at charlotte the...

UNCC 49er Rocketry Team

University of North Carolina at Charlotte

The William States Lee College of Engineering

2017-2018 NASA Student Launch Initiative

Preliminary Design Review

University of North Carolina at CharlotteWilliam States Lee College of Engineering

9201 University City Blvd.Charlotte, North Carolina, 28223

November 3, 2017

Contents

1 Summary of PDR 81.1 Team Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81.2 Launch Vehicle Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 81.3 Payload Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8

2 Changes made since Proposal 92.1 Changes made to Vehicle Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 92.2 Changes made to Payload Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 92.3 Changes made to Project Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9

3 Vehicle Criteria 103.0.1 Mission Statement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 103.0.2 Mission Success Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10

3.1 Selection, Design, and Rationale of Launch Vehicle . . . . . . . . . . . . . . . . . . . . . . . . 103.1.1 Airframe Diameter . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 113.1.2 Airframe Materials . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 123.1.3 Fincan Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 143.1.4 Drag Modulation System (DMS) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 173.1.5 Nosecone Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 183.1.6 Motor Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 203.1.7 Leading Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 22

3.2 Recovery Subsystem . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 303.2.1 Parachute Layout . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 323.2.2 Recovery Procedure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 333.2.3 Avionics Placement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 343.2.4 Payload Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 353.2.5 Booster Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 353.2.6 Parachute Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 363.2.7 Parachute Sizing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 363.2.8 Payload Parachute Reefing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 373.2.9 Shock Cord Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 383.2.10 Separation Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 383.2.11 Recovery System Redundancy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 393.2.12 Bulkhead Design and Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 393.2.13 Tracking and Communications . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 41

3.3 Mission Performance Predictions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 423.3.1 Flight Profile Simulations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 423.3.2 Flight Stability . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 423.3.3 Kinetic Energy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 443.3.4 Drift . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 44

3.4 Technical Challenges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 46

4 Safety 474.1 Safety Officer . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 47

4.1.1 Impact of Delays . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 484.1.2 Personnel Hazard Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 484.1.3 Failure Modes and Effects Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 524.1.4 Environmental Hazard Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 614.1.5 Risks to Project Completion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 63

1

5 Payload Criteria 655.0.1 Mission Statement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 655.0.2 Mission Success Criteria . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 65

5.1 Selection, Design, and Rationale of Rover Payload . . . . . . . . . . . . . . . . . . . . . . . . 655.1.1 Rover Drive System . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 655.1.2 Continuous Track Assembly Method . . . . . . . . . . . . . . . . . . . . . . . . . . . . 695.1.3 Tread Sprocket Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 695.1.4 Tread Material Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 715.1.5 Rover Dimensioning . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 725.1.6 Chassis Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 725.1.7 Chassis Material Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 735.1.8 Solar Panel Placement and Deployment . . . . . . . . . . . . . . . . . . . . . . . . . . 745.1.9 Object Detection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 745.1.10 Rover Distance Measurement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 775.1.11 Rover Leading Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 825.1.12 Preliminary Interface Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 905.1.13 Technical Challenges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 90

5.2 Selection, Design, and Rationale of Housing Payload . . . . . . . . . . . . . . . . . . . . . . . 925.2.1 Housing Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 925.2.2 Material Selection . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 945.2.3 Rover Retention . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 945.2.4 Rover Deployment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 975.2.5 Deployment Direction . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 985.2.6 Rover Housing Leading Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1005.2.7 Preliminary Interface Description . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1045.2.8 Technical Challenges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 104

6 Testing Plan 1066.1 Rover . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 106

6.1.1 Terrain Negotiation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1066.1.2 Partial Tread Soil Submersion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1066.1.3 Independent Track Elevation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1066.1.4 Ascending Ramp Angles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1076.1.5 Descending Ramp Angles . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1076.1.6 Sideway Slopes . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1076.1.7 Rover Autonomy . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 108

6.2 Rover Housing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1086.2.1 Rover Deployment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1096.2.2 Orientation Control . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 109

6.3 Launch Vehicle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1096.3.1 Drag Mod . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1096.3.2 Aerodynamics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1096.3.3 Seperation Testing . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1106.3.4 Airframe/Altimeter Bay Shear Strength . . . . . . . . . . . . . . . . . . . . . . . . . . 1106.3.5 Bulkhead Failure . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 110

6.4 Testing Methods Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 111

2

7 Project Plan 1157.1 Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 115

7.1.1 NASA Student Launch General Requirements . . . . . . . . . . . . . . . . . . . . . . . 1157.1.2 Team Derived Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 125

7.2 Budgeting and Timeline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1307.2.1 Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1307.2.2 Funding Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 1347.2.3 Major Project Timeline . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 137

8 Appendix 139

3

List of Figures

1 Payload Airframe Diameter Comparison . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 112 Drag Force Percent Multiplier . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 193 CAD Render of the Nosecone . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 194 L1500T Thrust Curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 205 L2200G Thrust Curve . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 216 CAD Model of the Current Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 227 Current Vehicle Design Dimensions (all dimensions in inches) . . . . . . . . . . . . . . . . . . 228 Locations of Fasteners and Shear Pins . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 239 CAD Model of the Booster Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2310 Fin Can Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2411 CAD Model of the Boattail . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2412 Fin Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2513 CAD Model of the Drag Modulation Structure . . . . . . . . . . . . . . . . . . . . . . . . . . 2514 Drag Modulation System Mechanism Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . 2615 Drag Modulation System Mechanism . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2616 Drag Modulation System Block Diagram . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2717 CAD Model of the Recovery Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2818 CAD Model of the Payload Section . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2819 CAD Model of the Payload Section Avionics Bay . . . . . . . . . . . . . . . . . . . . . . . . . 2920 Parachute Placement . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3321 Flight Plan and Recovery (Not to Scale) . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3322 Separation Points of the Vehicle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3423 Avionic Bays Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3424 Jolly Logics . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3525 Recovery Bulkhead Design . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4026 Bulkhead Locations . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4027 Flight Simulation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4228 Center of Pressure and Center of Gravity . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4329 Center of Pressure and Center of Gravity After Burnout . . . . . . . . . . . . . . . . . . . . . 4330 Current Rover Design Dimensions (all dimensions in inches) . . . . . . . . . . . . . . . . . . . 8231 Current Rover Design Dimensions (all dimensions in inches) . . . . . . . . . . . . . . . . . . . 8332 Rover Drive Servo Mounting . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8333 Tread Link Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8434 Rover Sprocket Configuration . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8435 Rover Chassis Mounting Points . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8536 Solar Panel Unfolded . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8537 Rover Body . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8638 Rover Electrical Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8739 Distance Measurement Data Flow . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9040 Rover Housing Dimensions (all dimensions in inches) . . . . . . . . . . . . . . . . . . . . . . . 10041 Leadscrew Retention System Overview . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10142 Electronics Sled and Altimeter Bay with Rotating Housing. . . . . . . . . . . . . . . . . . . . 10243 Payload Section Assembly . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10244 Rover Housing Electrical Summary . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10345 Categorized Project Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134

4

List of Tables

1 Airframe Diameter Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 122 Airframe Material Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 133 Airframe Material Properties . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 134 Fincan Design Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 155 Plastic Materials Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 166 Plastic Material Properties . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 167 Motor Retention Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 178 DMS Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 189 Nosecone Material Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2010 Aerotech L1500T . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2011 Aerotech L2200G . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 2112 Recovery Configurations Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3113 Payload Recovery Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3514 Parachute Design Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3615 Section Mass Estimates . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3616 Kinetic Energy and Velocity for Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3717 Kinetic Energy and Velocity for Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3718 Reefing System Data . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3819 Shock Cords Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3820 Separation Methods Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 3921 Simulation Results . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4222 Kinetic Energy and Velocity for Recovery . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4423 Payload Drift Distances . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4524 Launch Vehicle Technical Challenges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4625 Risk Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4826 Risk Acceptance Levels and Approval Authority . . . . . . . . . . . . . . . . . . . . . . . . . 4827 Severity Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4928 Probability Definitions . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 4929 Personnel Hazard Analysis . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5030 Booster Section FMEA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5231 Drag Modulation System FMEA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5332 Recovery FMEA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5433 Rover FMEA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5634 Rover Housing FMEA . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 5835 Launch Operations and Support Equipment FMEA . . . . . . . . . . . . . . . . . . . . . . . . 6036 Environmental Effects on Launch Vehicle . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6137 Launch Vehicle Effects on Environment . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6238 Risks to Project Completion . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6339 Rover Drive System Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6740 Rover Drive System Decision Matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6841 Tread Construction Method Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . 6942 Tread Configuration Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7043 Tread Material Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7144 Tread Material Properties . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7245 Chassis Position Selection Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7346 Chassis Material Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7447 Chassis Material Properties . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7448 Object Detection Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 7649 Distance Measurement Methods Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . 80

5

50 Microcontroller Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8151 Servo Motor Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8152 Rover Battery Budget Estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 8853 Rover Technical Challenges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9154 Rover Housing Design Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9355 Housing Design Decision Matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9356 Housing Material Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9457 Rover Retention System Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9658 Retention System Decision Matrix . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9759 Deployment Method Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9860 Deployment Direction Pros and Cons . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 9961 Rover Housing Battery Budget Estimation . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10362 Housing Technical Challenges . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 10563 Testing Methods . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11264 General Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11665 Launch Vehicle Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 11966 Recovery System Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12167 Experiment Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12268 Safety Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12369 Team Derived General Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . 12670 Team Derived Vehicle Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . 12771 Team Derived Recovery Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 12772 Team Derived Payload Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . 12873 Team Derived Safety Requirements Verification . . . . . . . . . . . . . . . . . . . . . . . . . . 12974 Launch Vehicle Anticipated Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13075 Payload Anticipated Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13176 Testing Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13177 Outreach Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13178 Anticipated Travel Budget . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13279 Purchased Parts . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13380 Projected Funding Plan . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13481 NASA Requirements . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 13782 Anticipated Project Development . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . . 137

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Acronym Dictionary

ABS = Acrylonitrile Butadiene StyreneAGL = Above Ground LevelAPCP = Ammonium Perchlorate Composite Propel-lantCAD = Computer Aided DesignCAR = Canadian Association of RocketryCATA = Central Academy of Technology and ArtsCATO = Catastrophe at TakeoffCDR = Critical Design ReviewCFD = Computational Fluid DynamicsCFR = Combined Federal RegulationsCG = Center of GravityCP = Center of PressureCOE = College of EngineeringDMS = Drag Modulation SystemEITY = Electronic and Information TechnologyEM = ElectromagneticFAA = Federal Aviation AdministrationFEA = Finite Element AnalysisFN = Foreign NationalFRR = Flight Readiness ReviewGPS = Global Positioning SystemHEO = Human Exploration and OperationsIAW = In Accordance WithLCO = Launch Control OfficerLRR = Launch Readiness ReviewMCU = Micro Control UnitMSRL = Motorsports Research Lab

MSFC = Marshall Space Flight CenterNASA = National Aeronautics and Space Adminis-trationNAR = National Association of RocketryNFPA = National Fire Protection AgencyPCB = Printed Circuit BoardPDR = Preliminary Design ReviewPIV = Particle Image VelocimetryPLAR = Post Launch Assessment ReviewPPE = Personal Protection EquipmentRAC = Risk Assessment CodeROCC = Rocketry of Central CarolinaRP = Rapid PrototypeRSC = Rocketry South CarolinaRSO = Range Safety OfficerSDS = Safety Data SheetsSLI = Student Launch InitiativeSME = Subject Matter ExpertSTEM = Science, Technology, Engineering, andMathematicsSOW = Statement of WorkTRA = Tripoli Rocket AssociationUNCC = University of North Carolina CharlotteUSLI = University Student Launch InitiativeVBS = Vehicle Booster SystemVTC = Vehicle Technical ChallengesRTC = Rover Technical ChallengesHTC = Housing Technical Challenges

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1 Summary of PDR

1.1 Team Summary

Team Name: 49er Rocketry TeamMailing Address: 9201 University City Blvd. Charlotte, North Carolina, 28223

NAR/TRA Mentor: Charles OginoAddress: 3100 Windrow Lane Matthews, NC 28105Email: [email protected]: (704) 771-3665TRA L3 Certified: TRA #12083NAR L3 Certified: NAR #87917

1.2 Launch Vehicle Summary

The launch vehicle will be constructed using 5.13 in. and 6.16 in. diameter airframe separated by a 6in. transition. The transition and 5.13 in. diameter sections will be made of carbon fiber while the6.16 in. airframe will be made from fiberglass. The vehicle is designed with a length of 98.5 in. and anestimated loaded weight of 50 lbs. The vehicle will have a 98 mm diameter motor mount tube and launchon an Aerotech L1500T motor. The recovery system involves two separate sections landing with their ownindependent parachute systems. The lower airframe will employ a standard dual deployment system, whilethe upper section will use a single reefed parachute. The Milestone Review Flysheet for the vehicle can befound at: https://www.unccrocketry.org/nasa-documents/

1.3 Payload Summary

Primary Payload: Autonomous Rover

The purpose of the rover is to autonomously travel 5 ft. from the landing site and deploy a foldable solarpanel array. To do this it will employ camera vision and ultrasonic sensors to navigate around obstacles andinertial navigation to ensure the 5 ft. mark is reached. The goal of the experiment is to simulate autonomousrover deployment upon landing on an extra-terrestrial body; therefore, the rover will only employ systemsthat would be representative of this mission goal.

Secondary Payload: Rover Housing

The purpose of the rover housing is to protect the rover through launch, separation, and landing. Then,upon receiving a signal from the ground team, it will orient the rover to an upright position and deploy it.The retention and deployment will be done using a leadscrew system, and the orientation will be done usinga position controlled servo and a single accelerometer.

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2 Changes made since Proposal

2.1 Changes made to Vehicle Criteria

The lengths of several sections of the launch vehicle have changed since the proposal. The length of thebooster section has been reduced to 39.68 in. after reevaluating the required length for the drag modulationsystem and motor tube. The airframe section above the drag modulation system has been reduced to 15in. after reviewing the packing dimensions of the booster main parachute. The length of the upper tube onthe transition was reduced to 5 in. after reviewing the packing dimensions of the booster drogue parachute.And finally, the airframe of the payload section will be exactly 27 in., rather than the predicted 28.2 in. aswas stated in the proposal.

2.2 Changes made to Payload Criteria

Deployable Rover

The most significant change to the rover design was the retention system that will be used to secure therover through flight and landing. The proposal outlined a system that used a braided steel cable to securethe rover in the housing. The current design calls for a 5/16 in. diameter leadscrew to retain and deploy therover. This necessitated a change to the retention interface, which will now be a 5/16 in. flanged leadnutthat will be attached to the underside of the rover chassis. The leadscrew will also be used to deploy therover, which is a change from the rack and pinion system described in the proposal. A DC motor has beenadded to the housing electronics bay to drive the leadscrew.

The change in the retention method required the drive servos to be shifted out from the centerline of therover to allow for passage of the leadscrew. This outward shift led to the tread width being reduced from25.4 mm to 20 mm in order to minimize the increase in width of the rover.

The current design utilizes a combination of camera vision and ultrasonic rangefinding, rather than justcamera vision, to detect objects and autonomously navigate around them. The rover will no longer employencoders for the dead reckoning process. Instead, a single IMU will be used to calculate the distance travelledand will compare that distance to GPS coordinates.

Rover Housing

The most significant change made to the rover housing, aside from the retention system described above,was the increase in the thickness of the aluminum bushings that will allow for the housing to rotate in theairframe. The thickness was increased from 1/16 in. to 1/8 in. This was primarily to make it easier toadhere the bushings to the inside of the airframe. The added thickness adds rigidity to the airframe, makingit more robust to impact forces.

A cap was added to the housing design to protect the rover from drogue parachute deployment at apogee.The cap will thread onto the leadscrew and be secured with nylon shear pins to easily break free when therover is deployed.

2.3 Changes made to Project Plan

A set of team derived requirements for Vehicle, Payload, Recovery, Safety, and General requirements wascreated since the submission of the proposal. The team derived requirements are located in Section 7.1.The team has established additional milestones and minor changes were made to the project time line.The updated time line is tabulated in the Section 7.2 and the Gantt chart is referenced in the appendix.The budget has been updated with current purchases and more accurately displays the allocation of funds,detailed in Section 7.2.

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3 Vehicle Criteria

3.0.1 Mission Statement

The 2017-2018 UNCC Vehicle Team’s mission is to successfully build, test, and fly a reusable high-poweredrocket to 1 mile AGL while carrying a deployable autonomous rover. The objective of the vehicle is to safelydeposit the rover to the ground and allow for its deployment.

3.0.2 Mission Success Criteria

The mission will be considered successful if it meets all vehicle, recovery, and safety payload requirementsoutlined in the 2018 NASA Student Launch Handbook, and meets the following criteria:

MS1 - The launch vehicle reaches the 5,280 ft. target within 100 ft.

MS2 - The vehicle carries the rover safely through flight and landing

MS3 - The rover successfully deploys and travels 5 ft. from the landing site to deploy its solar panels

The team-defined mission success criteria for the launch vehicle are as follows:

VS1 - The vehicle must safely carry the rover payload through the entirety of the flight

VS2 - The airframe must withstand the forces of takeoff, ascent, and landing

VS3 - The vehicle must successfully deploy parachutes for landing

VS4 - The vehicle must not cause any significant safety risk to onlookers or property

VS5 - The team must be able to assemble the vehicle within 4 hours of reaching the launch site

The team-defined mission success criteria for the recovery system are as follows:

RS1 - The recovery system must be easily armed at the launchpad

RS2 - The payload parachute must deploy within 2 seconds of apogee separation

RS3 - All recovery systems must successfully deploy at set altitudes

RS4 - The landing energy of all sections must be below 75 lbf-ft

3.1 Selection, Design, and Rationale of Launch Vehicle

The mission success criteria outlined above served as a guide for the team while designing the vehicle. Fourmain guiding principles were extrapolated from the success criteria and taken into account for the design,in descending order of importance: (1) the safety of the onlookers, (2) the robustness of the vehicle, (3) thesuccessful protection and deployment of the payload, and (4) the ease of assembly on the launch field. Withthese four criteria in mind, a number of design alternatives were considered for each major section of thelaunch vehicle. Pros and cons of each alternative are identified and tabulated in the following sections.

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3.1.1 Airframe Diameter

The diameter of the airframe is restricted by the width of the rover payload. To successfully complete themission, the diameter of the upper airframe will need to be large enough for all of the rover and housingcomponents. Initial estimate of the rover width is approximately 4.8 in. to fit the servo motors and othernecessary components. Based on this width value and a tread height of 2 in., a simplified CAD model wascreated to determine which airframe diameter will best accommodate the rover and housing.

Figure 1: Payload Airframe Diameter Comparison

As shown in Figure 1, a 5 in. diameter airframe will not accommodate the rover and housing without adramatic redesign of the entire payload system. A 6 in. diameter will allow sufficient room for the currentpayload design, but will require a precision fit of the supporting components. A 7 in. diameter providesample room for the current payload design, but may add excess mass, vehicle drag, and unused interiorvolume.

The diameter of the lower sections are less restricted by internal components. A 5 in. diameter was decidedas the minimum allowable value due to the Drag Modulation System (DMS) that sits above the booster.The DMS could fit in a 4 in. diameter airframe; however, the reduced diameter would be difficult and timeconsuming to manufacture.

Another consideration is weight and surface area. Wider airframes are heavier and more expensive due toincreased material usage and raise drag forces due to their increased surface area. Also, more material wouldbe needed to fit the rover, housing, and other payloads in the vehicle thereby increasing the mass. Theincreased mass and drag would impede the vehicle’s ability to reach an altitude of 1 mile AGL.

While considering the use of a transition, the trade off is cost versus efficient use of space. A 5 in. diameterwould be sufficient for the booster and DMS design, but a 6 in. diameter will be required for the rover. A

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6 in. booster section would unnecessarily increase the mass and drag of the booster section. It was decidedthat, despite the cost, a transition will lead to a better overall vehicle design.

Table 1 shows the pros and cons associated with the various airframe configurations the team considered.The final decision for the airframe diameter was driven by the need to accommodate the rover payload, whileminimizing weight, drag, and cost. Based on the pros and cons of the different alternatives outlined in Table1, the current launch vehicle design consists of a 5 to 6 in. transition midway up the airframe.

Table 1: Airframe Diameter Pros and Cons

Alternatives Pros Cons

Full 5 in.Diameter

• Simplifies aerodynamic characteristics

• Minimizes volume, weight, and cost• Difficult to fit payload

Full 6 in.Diameter


• Easier to fit payload

• Decreases necessary length of the vehicle

• Potential for wasted volume

• Potential to be expensive

• Potential for excessive weight

Full 7 in.Diameter


• Easily fit payload

• Further decrease necessary length of the ve-hicle

• Higher potential for wasted volume

• Higher potential to be expensive

• Higher potential for excessive weight

5 in. to 6 in.DiameterTransition

• Less difficult to fit payload with efficient useof volume

• Complicates aerodynamic characteristics

• Potential to be expensive

6 in. to 7 in.DiameterTransition

• Easily fit payload while minimizing interiorvolume

• Complicates aerodynamic characteristics

• High potential to be expensive

• High potential for excessive weight

3.1.2 Airframe Materials

To minimize the weight of the vehicle, the airframe materials need to be lightweight, but must also be strongenough to withstand the rigors of flight. The team considered numerous materials including carbon fiber,fiberglass, and 7075 aluminum because they are known to be both strong and lightweight. Table 2 comparesthe pros and cons of the three materials, while Table 3 shows their individual material properties.

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Table 2: Airframe Material Pros and Cons


CarbonFiber

• Highest strength to weight ratio

• Easily customizable diameter

• Easy sanding and machining

• Conductive

• High cost

• Dust from machining is hazardous

• Fails catastrophically when it doesfail

• Does not show that it is structurallycompromised before failure

Fiberglass

• Moderate strength to weight ratio

• Lower cost

• Easily customizable diameter

• Easy sanding and machining

• Shows stress cracks before failure

• Non-conductive

• Weaker than carbon fiber

• Dust from machining is hazardous

• Fails catastrophically when it doesfail

7075Aluminum

• High strength to weight ratio

• Easily machinable

• Less hazardous to machine

• Inexpensive

• Readily available

• Deforms before failure

• Conductive

• Requires greater wall thickness forsame strength as carbon fiber orfiberglass

• Heavier for same performance

Table 3: Airframe Material Properties

Material Tensile Strength(psi)

CompressiveStrength (psi)

Density (lbm/in3) Cost

Carbon Fiber 87000 126000 0.0578 High

Fiberglass 42000 50800 0.055 Medium

Aluminum 7075 73000 83000 0.102 Low

Booster Airframe Material

Carbon fiber was selected for the booster section airframe. This is primarily due to its high strength to weightratio. The booster of the rocket is required to transfer all of the thrust of the motor to the vehicle, therefore

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it must be exceedingly strong to withstand these forces. Additionally, carbon fiber is easily machined andable to be rolled to custom specifications. As shown in Table 3, carbon fiber has a high tensile strengthwhich defines its capacity to withstand most forces during flight. For these reasons it was chosen as thematerial for the booster section airframe. Carbon fiber will also be used for the motor tube and centeringrings to ensure these parts have comparable strength to the airframe.

In an effort to show that the carbon fiber on the booster section can withstand the forces from the launch,Equation 1 was used to find the compressive Force on the airframe from the motor thrust.

σ =Force

Area(1)

σ is the variable for the compression force.

Equation 1 yielded the following calculations, where the maximum thrust from the two motors consideredwas used.

For the L1500T:

337.21π/4(5.132−52) = 326.03 psi

For the L2200G:

697.13π/4(5.132−52) = 674.02 psi

These values when compared to the compression strength of carbon fiber, which is shown in Table 3, provesthat the carbon fiber has sufficient strength to withstand flight forces.

Recovery Airframe Material

An integrated avionics bay is housed within the recovery section. These avionics include recovery altimetersthat must be shielded from interference from exterior EM signals. Carbon fiber, which is a conductive ma-terial, would help to shield the altimeters in this section. Carbon fiber is also easily molded into complexshapes, which would allow for this section to made entirely from carbon fiber. Due to the diameter transitionin this section, high strength is required to ensure the structural stability of the vehicle. For these reasons,carbon fiber is the material that was selected for this section.

Payload Airframe Material

The payload section houses the rover payload which must be autonomously deployed. To activate this de-ployment a wireless signal must reach the internal system. Because of this, a conductive material cannotbe used; however, the airframe must still maintain a high degree of structural integrity. Fiberglass, thoughnot as strong as carbon fiber, is non-conductive and has a good strength to weight ratio. For these reasonsfiberglass was selected to be the payload airframe material.

3.1.3 Fincan Design

Two main design alternatives were proposed for the fincan section of the vehicle: a fixed fin design and amodular fincan design. The fixed fin design is simple to design and construct, however the modular fincan

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design increases the reusability and maintainability of the vehicle. The fins of the vehicle are the most sus-ceptible to damage upon landing. Removable fins allow for quick and easy replacement of damaged fins whileon the field, meeting the requirement that the vehicle must be reusable following the competition launch.

Using removable fins offers the ability to easily change the fin size and shape on the vehicle in varyingweather conditions. For example, if there were higher than optimal winds at a test launch, the fin sizecould be decreased to minimize wind cocking. The fins could also be adjusted to compensate for unexpectedchanges in the vehicle design. For example, if the weight of the payload section varied from the predictedweight, the fin size could be adjusted to maintain the desired stability margin for the vehicle.

Table 4 shows the pros and cons associated with the fincan designs the team considered.

Table 4: Fincan Design Pros and Cons


Fixed FinDesign

• Simplifies design and construction

• No assembly on field

• Minimizes cost

• No ability to adjust fin shape or re-pair fins

• Must replace entire section if fins aredamaged, potentially increasing thecost

ModularFincanDesign

• Easy to replace fins and adjust finshape

• Simple design minimizes on assem-bly time on the field

• 3D printed manufacturing allows forconsistent dimensions across prints

• Complicates design and construc-tion

The trade off for the fincan of the vehicle was between two of the guiding principles previously defined: (2)the robustness of the vehicle, and (4) the ease of assembly on the field. A modular fincan design allows foreasy vehicle repair, making it more robust overall. To reduce the added assembly time, the design employsthe minimal number of components necessary, the four fins and boattail, which holds the fins in place throughflight.

Fincan Materials

The fincan consists of the booster airframe, motor tube, fins and boattail. As stated above, the airframeand motor tube will be made from carbon fiber. The remaining parts will be made from rapid prototypedplastic due to the easy customizability and manufacturing it provides.

The options for the fin materials are shown in Table 5, and their respective material properties are shownin Table 6.

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Table 5: Plastic Materials Pros and Cons


Polycarbonate

• High fatigue which allows forreusability

• Allows for good tolerances on parts

• High Strength

• High cost

• Heavy

ABS

• Medium cost

• Good Strength

• Lightweight

• High ductility may show warpingwhich decreases reusability

• High porosity

ULTEM 9085

• Very lightweight

• High heat resistance

• Low cost

• Most brittle

Table 6: Plastic Material Properties



Polycarbonate 8400 0.043 High

ABS 6600 0.0376 Medium

ULTEM 9085 9300 0.0484 Medium

Polycarbonate was chosen for the fins due to its high rigidity and high print tolerances. This results ina higher print quality for the fins, which is desirable due to the need for them to be as close to perfectlyconsistent in size and shape as possible. The high rigidity of polycarbonate also reduces the chance of finbreakage upon landing.

Motor Retention

A 98 mm diameter motor will be used in the launch vehicle. In order to ensure the motor does not exit thevehicle due to parachute shock forces, various retention system alternatives were considered. These includea standard screw cap retainer that would be epoxied to the motor tube, and an integrated boattail motorretainer that could double as a fin retention device. Table 7 shows the comparison between both motorretention systems.

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Table 7: Motor Retention Pros and Cons


Boattail

• Allows for modular fincan

• Rapid prototyping allows for com-plex and constant geometries be-tween prints

• Decreases aerodynamic drag forceson the vehicle

• Requires 4 extra fasteners forbooster section assembly

• Possibility of heat damage to thestructure

Screw-onRetainer

• Low cost

• Simple assembly

• Negates possibility of modular de-sign on the fincan

• Increases aerodynamic drag forces

The use of a boattail motor retention system allows for a simplified module fincan design and reducesaerodynamic drag on the vehicle. The boattail design also doubles as a fin retainer, therefore if this methodwas not selected another method would need to be devised to hold the fins in place during flight. For thesereasons the boattail retention system was chosen.

Boattail Material

The possibility of heat damage to the boattail risks catastrophic failure to the fincan structure. This isbecause the boattail is the securing element of the fincan assembly, and therefore the assembly would breakapart if the boat tail is compromised. To mitigate this, the boattail material was chosen to be UTLEM 9085due to its high heat resistance.

3.1.4 Drag Modulation System (DMS)

In order to more accurately reach the target altitude, a drag modulation system was considered during thedesign phase. A DMS allows for higher accuracy and precision in reaching specific altitudes compared totraditional methods such as adding ballast weight to the vehicle. The trade off however, is the added com-plexity it adds to the vehicle design.

Table 8 shows the pros and cons associated with the decision to use a drag modulation system for altitudecontrol.

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Table 8: DMS Pros and Cons


DragModulation

• Allows higher precision control ofmaximum altitude irrespective ofweather conditions

• No need for ballast weight

• Provides a testbed for control the-ory

• Increases complexity of vehicle de-sign and construction

• Increases cost and length of the ve-hicle

• Requires a larger motor to overshootthe required altitude

BallastingWeight

• Simplifies rocket design and con-struction

• Allows for a smaller overall vehiclesize

• No need to overshoot altitude

• Susceptible to variations in theweather, making the maximum al-titude difficult to control

• Ballast for dialing in altitudeadds wasted space and unnecessaryweight, decreasing efficiency of thevehicle

The drag modulation system was ultimately chosen to be used on the vehicle. This decision was due tothe ballast resulting in wasted space within the vehicle volume. Also, the addition of the drag modulationsystem allowed for more precise altitude attenuation during flight. This attenuation can also be achieved invarying weather conditions, as the drag modulation system can operate at full capacity regardless of weatherconditions.

DMS Material

Due to the complex geometries of the DMS, rapid prototyped plastics were heavily considered when choosinga material. The material had to have sufficient strength while also being lightweight and easily modifiable.Table 5 shows the comparison between all plastics considered. ABS was ultimately chosen as the DMSmaterial because it is both cheaper and lighter than polycarbonate. Polycarbonate is stronger than ABS,however the excess strength is not needed as the structure will not be experiencing excessive forces throughoutthe flight.

3.1.5 Nosecone Design

The geometric shape of the nosecone drastically influences the flight characteristics of the vehicle. Theprimary factor that would govern the characteristics of flight is the drag coefficient of the nosecone whichdefines the magnitude of the drag forces exerted on the vehicle. Figure 2 shows the percent increase to thedrag when compared to a flat faced shape.

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Figure 2: Drag Force Percent Multiplier

The comparison shown in Figure 2 compares flat face drag to nosecone shape at subsonic speeds. Theseresults thus show that a parabolic nosecone will minimize the drag forces at subsonic velocities. Because thevehicle is not designed to reach supersonic speeds during flight, a parabolic nosecone was chosen to be usedon the vehicle. Figure 3 shows a CAD render of the nosecone shape.

Figure 3: CAD Render of the Nosecone

Nosecone Material

To acquire the preferred nosecone shape, the nosecone will either be manufactured through means of rapidprototyping, or be purchased commercially. These options present various material options which are shown

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in Table 9.

Table 9: Nosecone Material Pros and Cons


G10Fiberglass

• Good strength to weight ratio

• Low cost

• Cheaper than Carbon fiber

• Low impact resistance

• Heavy

• Difficult to post process

CarbonFiber

• Very light weight

• Great strength to weight ratio

• High cost

• Difficult to post process

ABS

• Rapid prototyped manufacturing al-lows for desired shape

• Low cost

• Lightweight

• Simple to post process

• Low strength makes if vulnerable toimpact forces

These options were narrowed down to the nosecone being produced by rapid prototyping from ABS plasticbecause the 3D printing process allows for the desired shape to be obtained.

3.1.6 Motor Selection

The motors that were considered could not be above an impulse of 5120 N-sec (L class). Having this re-striction, the motors that were considered for the launch vehicle are shown in Tables 10 and 11. Thesespecifications are a combination of manufacturer data as well as flight simulation data.

Impulse 5056 N · secCost $ 280

Loaded Mass 50 lbm

Empty Mass 40 lbm

Apogee 5877 ft.

Velocity Off-Rod 65 ft/sec

Max Thrust 337.21 lbf

Table 10: Aerotech L1500T Figure 4: L1500T Thrust Curve

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Impulse 5104 N · secCost $ 220

Loaded Mass 50 lbm

Empty Mass 40 lbm

Apogee 6083 ft.

Velocity Off-Rod 94 ft/sec

Max Thrust 697.31 lbf

Table 11: Aerotech L2200G Figure 5: L2200G Thrust Curve

With these motors in mind, the L1500T was selected because the specifications show that it provides thehighest impulse to propel the vehicle to 1 mile AGL. The L952W does present a more economically viableoption, however, the motor is not strong enough to propel the vehicle to the minimum 52 fps off-rail velocity.These specifications can be seen on Tables 5 and 4. The L2200G will be used as a backup motor should theL1500T prove to be insufficient.

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3.1.7 Leading Design

Overview

The launch vehicle will be constructed using 5.13 in. and 6.16 in. diameter airframe. The vehicle is designedwith a length of 95.5 in. and an estimated mass of 50 lbm. A 98 mm diameter motor mount tube and launchon an Aerotech L1500T motor. The recovery system involves two separate sections landing with their ownindependent recovery systems. Figure 6 shows the CAD model of the current vehicle design, while Figure 7shows the lengths of the vehicle and its sections.

Figure 6: CAD Model of the Current Design

Figure 7: Current Vehicle Design Dimensions (all dimensions in inches)

The full dimensions of the vehicle are shown in Figure 7. This includes all individual section and airframelengths, the lengths of all components, and the outer diameters of each part. These measurements representthe vehicle fully assembled. Not shown are the couplers that bind each section together. The location ofthese couplers are shown in Figure 8.

Some of these couplers will keep sections together throughout flight, while others will keep sections joineduntil recovery is initiated. The only difference between them is that the permanent sections use fasteners forsecuring, while the temporary use shear pins. The placement of these couplers, alongside their attachmentmethods are shown in Figure 8.

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Figure 8: Locations of Fasteners and Shear Pins

Booster Section

The booster section is comprised of the rocket motor, its mounting and retention system, the removable finsystem, and the drag modulation system shown in Figure 9. The projected weight of the booster sectionand drag modulation system is 25.3 lbm. This mass includes the loaded motor mass.

Figure 9: CAD Model of the Booster Section

The lower airframe of the booster section will be constructed from rolled carbon fiber with four slots for thefins providing a strong but lightweight structure to withstand launch and impact forces. The mid airframe,above the drag modulation system, will be constructed with rolled fiberglass which has a slightly lowerstrength to weight ratio, but will allow the booster section to communicate with the ground station as itdoes not impede radio waves from transmitting through the material. Both airframe components will havea diameter of 5 in. and a minimum wall thickness of 1/16 in.

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Figure 10: Fin Can Assembly

Figure 11: CAD Model of the Boattail

The vehicle will incorporate an interchangeable fin system. This allows on-field adjustments to the launchvehicle in the presence of different environmental conditions as well as the ability to replace fins that maybe damaged during flight or transport. The fins will be rapid prototyped out of polycarbonate, for its highstrength to weight ratio and tight printing tolerances, with a symmetric airfoil cross-section. The fin di-mensions can be seen in Figure 12. The boattail, which is shown in Figure 11, will secure the fins on the fincan.

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Figure 12: Fin Design

Drag Modulation System (DMS)

The drag modulation system (DMS) will be comprised of four drag flaps that will deploy symmetricallyabout the airframe to alter the vehicle’s drag characteristics and control its velocity to 0 ft/sec at apogee.The flaps will be desgined to deploy to a maximum of 22.5 degress from the airframe, which would allowthe DMS to attenuate up to 15% of the vehicle’s maximum altitude. This can be determined from CFDsimulations and will be confirmed through full-scale test flights. The body will couple the lower airframe tothe mid airframe and house the actuator mechanism as well as the DMS electronics. The body and flapswill be rapid prototyped out of ABS for its good strength to weight ratio and lightweight properties toaccommodate every component that will be housed within the body. The DMS will use a single actuatorso that all flaps will deploy simultaneously, preventing the launch vehicle from altering its trajectory duringflight. Figure 13 shows the external DMS structure.

Figure 13: CAD Model of the Drag Modulation Structure

The DMS will be actuated by a single high torque servo motor (HS-7950) that will rotate a steel shaft alongthe vehicle’s center axis. A machined steel collar connected to the shaft will have four aluminum links thatwill be connected to steel pins such that when the center shaft rotates, the brass links will drive the steelpins against the inner surface of the flaps, deploying them symmetrically about the airframe. The steel pinswill slide through linear bearings to maintain a fixed orientation. The ABS drag flaps will be connected to

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the body of the DMS with a 1/8 in. steel pin through aluminum hinge to provide a strong anchor to rotateon. Steel extension springs will be attached from an internal anchor to each flap so that the inner surfaceof each flap will remain in contact with each deployment pin during flight. The servo motor was chosenbecause of its internal feedback controller that will ensure accurate positioning as well as its small compactfootprint. The HS-7950 will be powered by 7.4 volts, which creates a stall torque of 486 oz-in. This is muchgreater than the expected torque the motor will have to overcome due to dynamic pressure and the retainingsprings, which will ensure rapid actuations. Figures 14 and 15 shows this mechanism.

Figure 14: Drag Modulation System Mechanism Assembly

Figure 15: Drag Modulation System Mechanism

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As seen in Figures 14 and 15, the drag modulation mechanism is designed to execute a symmetrical extensionof all flaps. This mechanism is operated by a servo motor which sits above the mechanism assembly. As theservo motor turns, the mechanism rotates which then pushes a set of pins through linear bearings. Thesepins would then push out the flaps to their full extension. Should this system fail then the springs will beable to retract the fins back to their closed position.

Figure 16 shows the control system block diagram of the DMS. The velocity of the vehicle will be measuredusing a combination of an accelerometer and an atmospheric pressure sensor. The accelerometer outputwill be integrated to find velocity. The pressure sensor will be used to estimate altitude, the derivative ofwhich will then be taken to get velocity. The two velocity values will be combined in a Kalman estimator toaccount for integration drift and derivated noise from the signals. The system will attempt to match a setvelocity curve that will be based on simulated data. A PID controller will be used to rapidly respond to theerror signal and match the ideal velocity curve as closely as possible.

Figure 16: Drag Modulation System Block Diagram

Recovery Section

The recovery section will be comprised of the booster section drogue and main parachutes as well as thealtimeter bay. They will be housed in the airframe transition, constructed from rolled carbon fiber for itshigh strength to weight ratio and its ability to shield electromagnetic radiation that may interfere with therecovery equipment. The smaller diameter of the transition will be 5 in. and the larger diameter will be 6in. with a wall thickness of 1/16 in. The combined weight of this system is projected to be 6 lbm.

The recovery section is completely made from carbon fiber because of its high strength to weight ratio. Thisallows for a very rigid structure but it also shields the internal avionics from EM signals. This is becausethe conductive properties of carbon fiber. These altimeters will control the deployment sequences for thebooster section. The area above the altimeter bay will house a 3 ft. drogue parachute which will deploy atapogee, while the lower section will house the 10 ft. main parachute which will be deployed at 600 ft. Figure32 shows an isolated CAD model of the recovery section assembly.

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Figure 17: CAD Model of the Recovery Section

For the avionics bay assembly within the recovery section, four altimeters along with their batteries willbe held to the top bulkhead by means of fasteners. The airframe will include a hole that exposes a switchthat turns on the electronics. On the two bulkheads that surround the bay, 1/4 in. diameter eye bolts areattached and secured to a single threaded rod that goes through the entire bay. The drogue and the mainparachutes are then attached to these eye bolts.

Payload Section

The payload section is the top-most section of the vehicle, where the rover will be housed throughout theflight. This section will include its own dedicated recovery system which will be initiated 2 seconds afterreaching apogee. This is to avoid a parachute collision with the other section. The 10 ft. parachute will bedeployed from the nosecone so that the rover has an unobstructed pathway for its deployment. Figure 18shows an isolated CAD model of the payload section.

Figure 18: CAD Model of the Payload Section

The avionics within the payload section, as they are shown in Figure 18, are placed at the top of the roverhousing. Here, two altimeters along with their batteries are housed within a removable bulkhead which isattached to the internal airframe by means of fasteners. A 1/4 in. diameter eye bolt will attach to thisbulkhead through the use of nuts securing to the eye bolt threads. The main parachute for the payload willthen attach to this eye bolt. Figure 19 shows the external and internal view of the payload avionics bay.

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Figure 19: CAD Model of the Payload Section Avionics Bay

The bulkhead structure shown in Figure 19 will be constructed from rapid prototyped ABS. This is dueto the use of complex geometries, such as the charge wells and the electronic attachment points, on thebulkhead. Rapid prototyping would allow the complex geometries to be made with ease. In addition to theABS, the bulkhead will be reinforced with carbon fiber plates.

The nosecone will be rapid prototyped from ABS. This is primarily due to the rapid prototyping processresulting in the desired parabolic shape. Additionally, the rapid prototyping allows for a coupler to builtdirectly to the nosecone structure.

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3.2 Recovery Subsystem

A regular dual deployment system uses a drogue and main parachute. This system provides a simple re-covery solution that is low cost and safe. Using only two parachutes minimizes the risk of tangling andonly requires two separation events. Reducing the total number of parachutes and altimeters will reduce theoverall mass and lower the cost. Without separating the rocket into two separate systems, a large parachutewill be required to land the rocket with in the kinetic energy requirements. There will also be an increasedrisk of inference between the booster and payload section while descending. This system adds complexity todeploying the rover due to having parachutes deploy at both separation points.

Another possible recovery solution separates the rocket into the payload and booster section at apogee.After apogee both sections will deploy drogue parachutes and then deploy main parachutes. Separating thesections at apogee will distribute the mass of the rocket to each section thus reducing the parachute sizerequired to land within the kinetic energy requirement. This will also separate the two sections to help withreducing possible tangling. This system requires the most parachutes which will require more space in therocket and increase cost. Using more parachutes also requires more altimeters to separate the body sectionsand deploy the parachutes.

The leading recovery design uses a separate booster and payload recovery system: a reefed payload parachuteand a dual deployment system for the booster section. Separate recovery systems distribute the rocket’s massbetween two separate system allowing for smaller parachutes. The payload section will fall under a reefedsingle parachute until fully deploying at 300 ft. above ground. This system allows to the payload to fall ata different rate than the booster section, thus reducing the risk of tangling. This system also incorporates areefing ring which will restrict the parachute from violently opening and significantly reduce the shock seenby the payload system. Since the booster section can withstand larger shock forces, dual deployment wasselected. Dual deployment reduces the complexity of the system and reduces the amount of possible failuremodes.

The last recovery solution considered separates the rocket at apogee and uses a reefed main parachute forboth booster and payload. This system only requires two parachutes which will use the least volume in therocket as well as minimizes the number of separation events. Using reefing rings on the parachutes reducesthe shock force when inflating to full size. Using this system complicates the overall recovery system andwhile also increasing the possible failure modes.

Table 12 compares the various possibilities for the recovery configuration.

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Table 12: Recovery Configurations Pros and Cons


Vehicle alltied to single

main,standard

dual-deploy

• Simplifies configuration

• Reduces volume allotted for recov-ery

• Two parachutes reduces cost andweight

• Minimizes tangling risk

• Minimizes separation events

• Large parachutes for full vehicleweight to land within KE require-ments

• Risk of booster interfering with pay-load

• Increased shock to payload system

• Complicates rover deployment

• Minimum number of altimeters(4+1)

Separatebooster and

payload,double

standarddual-deploy

• Separate booster and payload, dou-ble standard dual-deploy

• Except for tangling, reduces risk ofbooster interfering with payload

• Maximum required volume will beused

• Four parachutes increases cost andweight

• Maximum tangling risk

• Three separation events

• Increased shock to payload system

• Increased altimeter count (6+1)

• Increased length of rocket for neces-sary packing volume

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Separatebooster and

payload,reefed

payload,dual-deploy

booster

• Smaller parachutes for split vehicleweight

• Reduces risk of tangling by usingreefed payload main (falls fasterthan drogue)

• Significantly reduces shock to pay-load system

• Complicates configuration

• Medium required volume

• Three parachutes increases cost andweight


• Increased length in rocket for neces-sary packing volume

• Risk of reefing failure

Separatebooster and

payload,reefed

payload,reefed

booster

• Smaller parachutes for split vehicleweight

• Two parachutes, rather than threeor four

• Decreased weight/volume/length ofvehicle

• Reduced risk of tangling

• Minimizes separation events

• Minimizes shock to payload andbooster

• Complicates configuration

• Second least required volume


• Increased risk of reefing failure

3.2.1 Parachute Layout

The recovery system layout was designed to easily integrate with the payload system. The recovery systemis broken into two main sections, payload and booster. The payload section uses a reefed single parachutesystem. After main separation from the booster section, the payload section will separate the nose conefrom the body of the section. When designing the recovery system, the payload section had to allow oneside of the body tube to be open for rover deployment. Since the nose cone must still be attached to thebody this removes one possible exit for the rover. Using a conventional dual deployment system requires twoparachutes which would restrict both possible exits for the rover. The payload parachute is located abovethe rover housing and below the nose cone.

The booster section utilizes a dual deployment system. This system was selected since the booster section hasmore available space due to the recovery section. The drogue parachute will be located between the boosterand the payload section, which is deployed when the sections separate at apogee. The main parachute islocated below the avionics bay and above the booster section. Figure 20 shows the location of all parachuteswithin the vehicle.

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Figure 20: Parachute Placement

3.2.2 Recovery Procedure

From these options, the separate booster and payload with a reefed payload was chosen. This configurationwas chosen primarily due to the mitigation of parachute tangling and the shock reduction to the payloadsection. The configuration of the recovery system will also dictate the order of recovery procedures. Figure21 shows the flight plan of the vehicle with the implemented recovery configuration.

Figure 21: Flight Plan and Recovery (Not to Scale)

As shown in Figure 21, there will be 7 main events following the launch of the vehicle:

1. At apogee the vehicle will separate the payload section from the booster and recovery sections.

2. The booster drogue parachute will be deployed during the separation at apogee.

3. The payload section will fall completely separate from the booster section and will deploy its reefed

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parachute following a two second delay. The delay will ensure the payload section is clear of the boostersection to minimize the risk of the parachutes getting tangled.

4. At 600 ft. the booster section main parachute will deploy.

5. At 300 ft. the Jolly Logic reefing system will be released forcing the reefing ring to slide down theshroud lines. This will allow the main parachute to unfurl to its maximum diameter.

6. The booster and payload sections will reach the ground within the kinetic energy requirements.

7. After becoming stationary, the payload will activate and the rover will be deployed.

Figure 22 shows the separation points of the vehicle for this recovery configuration.

Figure 22: Separation Points of the Vehicle

3.2.3 Avionics Placement

The vehicle will include two avionic bays. One will be in the recovery section, while the other will be in thepayload section. The recovery avionics bay sits at the center of the section, because both the drogue andthe main parachutes are connected to, and deployed from, this structure. Here, the avionics are directlyconnected to the black powder charges that sit just above and below the bay. This reduces the length ofwire connections required to establish this connection.

The avionics bay within the payload section is housed within a removable bulkhead that is secured abovethe rover housing system. This placement was due to the need for the rover to be unobstructed duringdeployment. The avionics bay is then located opposite the airframe opening where the rover will deployfrom. Here, the bay sits below the nosecone Additionally, the avionics package is directly connected to theblack powder charges.

Figure 23 shows the placement of the avionic bays.

Figure 23: Avionic Bays Locations

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The avionic bay construction and assembly is discussed for the bays of each individual section in subsections3.1.7 and 3.1.7, for the recovery and payload sections respectively.

3.2.4 Payload Recovery

The team considered implementing a Jolly Logic reefing system for the payload section’s 10 ft. mainparachute. Once the parachute is free from the deployment bag, the Jolly Logic opening device wold holdthe parachute closed at the base preventing inflation. Once the payload section reaches an altitude of 300ft. the Jolly Logics will open, and the parachute will inflate while a reefing ring will slow down deployment.This system is compared with a traditional drogue parachute in Table 13.

Table 13: Payload Recovery Pros and Cons


ReefingSystem

• Volume efficient

• Low cost• Higher risk of recovery malfunction

TraditionalDrogue

Parachute

• Robust system that has been provento work

• Requires one extra set of B.P.charges for main parachute deploy-ment

• Uses more volume

• High cost

The team decided that the payload recovery would implement a Jolly Logic reefing system because it requiresless volume to fit within the vehicle. Additionally, an extra set of black powder charges would be redundantas the main parachute would already be deployed in a reefed orientation. Although the reefing system doesincrease the risk of recovery failure, past team’s experience has resulted in a more robust system whichdecreases the risk of failure. Figure 24 show what the Jolly Logics.

Figure 24: Jolly Logics

3.2.5 Booster Recovery

At main separation a drogue parachute will be deployed which will bring the booster to a safe terminalvelocity prior to main parachute deployment at 600 ft. Dual deployment provides a more reliable and safedeployment method over the reefed system. The dual deployment system uses two altimeters while a reefed

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system would require four. Since the booster section can withstand larger shock forces, a reefing ring is notnecessary.

3.2.6 Parachute Design

When selecting parachute shapes the team considered elliptical and annular parachutes. When comparingdesigns, the main emphasis was laid on size and weight. Annular parachutes have a coefficient of drag of 2.2while elliptical parachutes coefficient of drag is 1.5. Since annular parachutes coefficient is much larger thisallowed for a smaller parachute to be used. When comparing same sized elliptical and annular parachutes,annular parachutes use less material resulting in smaller packing volumes and lighter overall weights. Table14 compares these two parachute types against each other.

Table 14: Parachute Design Pros and Cons


Elliptical • Medium cost

• 1.5 coefficient of drag

• Heavier weight

• Requires more volume for storage

Annular

• 2.2 coefficient of drag

• Space efficient

• Lightweight

• High cost

After comparing both designs, annular shaped parachutes were selected for recovery because of the higherdrag coefficient and volumetric efficiency. This shape is to be used on both the main parachutes and thedrogue parachute.

3.2.7 Parachute Sizing

The vehicle’s recovery systems were designed to meet the maximum restriction of 75 lbf · ft kinetic energyat landing. The mass of each separate section of the vehicle was estimated in order to predict the kineticenergy for each section at landing. These estimations are shown in Table 15.

Table 15: Section Mass Estimates

Payload 18 lbm

Booster 15.3 lbm

Recovery 6 lbm

The calculations done for the recovery systems accounted for these masses. Additionally, 10 ft. diameterparachutes were considered for the main parachutes on both independent sections. Equations 2 and 3 wereused to find the values for both the velocity and the kinetic energy respectively while taking into accountboth the mass and parachute size. Also, parachute manufacturer FruityChutes report a coefficient of dragof 2.2 for their parachute systems which is required for Equation 2.

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v =

√2 ·W

Cd · ρ ·Ap(2)

Ke =1

2·W · v2 (3)

Table 16 shows the definitions for the variables:

Table 16: Kinetic Energy and Velocity for Recovery

Variable Definition

W Weight of Section

Cd Coefficient of Drag

ρ Density of Air

Ap Area of Parachute

Ke Kinetic Energy

v Velocity

These equations yielded the values shown in Table 17. These values include both the individual and tetheredsections which share the same parachute. For the tethered sections, the nosecone and payload section aretethered together while the booster and the transition sections are tethered together. Sections that aretethered together are numbered the same under the section column in Table 17. Also, in an effort to reducethe kinetic energy upon impact the lighter parts will be tethered to the parachute with a longer shock cordso that the overall mass on the parachute is reduced as parts tethered to a longer shock cord will touchdownfirst.


Section Weight (lbm) Velocity (ft/sec) Kinetic Energy (lbs · ft)(1) Payload 17 1.6 21.76

(1) Nosecone 1 1.65 1.36

(2) Booster 15.3 1.52 17.67

(2) Transition 6 1.79 9.61

With these kinetic energy values, it’s possible to see that a 10 ft. diameter FruityChute parachute is idealto be used for the recovery systems.

3.2.8 Payload Parachute Reefing

The purpose of the reefing system is to dampen the opening forces upon parachute opening. Additionally, thereefed parachute acts as a drogue during freefall. This system was implemented with the goal of mitigatingthe risk of damage to the payload during recovery. Past team’s analysis of the reefing ring system provedthat it is capable of dampening the opening forces by significant amounts. This analysis consisted of droptesting a mass with both a normal and reefed parachute, while measuring the acceleration due to openingof the parachute. Table 18 shows the results from these tests.

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Table 18: Reefing System Data

Parachute System Acceleration Due to Opening (G)Annular 8

Annular (Reefed) 4

3.2.9 Shock Cord Selection

To absorb the residual kinetic energy from deployments, shock cords will be used to connect all sections totheir respective parachutes. The lengths of the shock cords and the tying positions of the parachutes will besuch that the sections do not collide during descent. The shock cords will be tied onto 1/4 in. diameter steeleye bolts that are bolted through the bulkheads of the vehicle. Swivels will be used to attach all parachutesto their shock cord to reduce swaying when descending. Table 19 shows the various shock cords considered.

Table 19: Shock Cords Pros and Cons


9/16 in.Nylon ratedto 2400 lbf

• Great strength

• Long lengths are available

• Medium cost

• Requires more volume

3/16 in.Spectra

Microlinerated to1500 lbf

• Does not require much volume

• Low cost

• Only short lengths available

• Medium strength

1/4 in.Kevlar ratedto 2200 lbf

• Good strength

• Medium lengths available

• May take up large volume

• Medium cost

From the shock cords considered, the 9/16 nylon was chosen because of it has the best strength of all optionsavailable. This shock cord will be used for the attachment of the parachutes to their respective sections, aswell as tether sections that share the same recovery systems.

3.2.10 Separation Methods

Two main separation methods were considered, Black Powder Charges and Compressed Gas Systems. Alarge emphasis was placed on safety and reliability because hazardous materials are heavily involved in theseparations. Table 20 compares the two methods against each other.

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Table 20: Separation Methods Pros and Cons


BlackPowderCharges

• Robust system

• Low cost

• Requires increased safety for han-dling

• Requires precise amounts to sepa-rate but not damage the vehicle

CompressedGas

Canisters

• Safer to operate

• Requires a higher use of volumewithin the vehicle

• Inconsistent operation

Black powder charges were selected due to their simplicity and cost. In order to maximize the team’ssafety when handling the black powder, the team mentor will handle the assembly of all charges. Chargeswill be used to separate the booster and vehicle sections, deploy vehicle reefed parachute systems, and themain/drogue parachutes for the booster section. At each separation point a robust deployment charge andaltimeter will be used. Each charge will be ignited using two igniters to ensure detonation. Shear pinswill be used to hold the various sections in place where parachute deployment occurs. Calculations will bemade to determine the correct charges to be used for separation of the shear pins. In order to verify thesecalculations, ground ejection testing will be performed.

3.2.11 Recovery System Redundancy

Redundancy is a key aspect of safety in the recovery section and will be incorporated into their design. At eachrecovery event that requires a black powder ejection charge and altimeter, a secondary fully independentcharge and altimeter will be installed. At apogee, the main separation charge will ignite, then a secondaltimeter will ignite at 1 second post apogee to ensure separation if the main charge fails. This redundantcharge will also include a safety factor larger than the main charge to ensure separation if the main chargefails to completely separate. Each black powder charge will use a dual igniter system to ensure ignition.Both the payload and booster sections will use the same redundant system on their respective separation andparachute deployment. The reefing system on the payload section’s main parachute will include a redundantsystem as well. The Jolly Logics will be attached to the parachute in such a way that if only one Jolly Logicopens, it will release the parachute and allow to fully inflate. The recovery system will make use of four maindeployment altimeters and 4 redundant altimeters to ensure every recovery system will have a redundantsystem.

3.2.12 Bulkhead Design and Testing

The vehicle will include two bulkhead types: permanent and removable. Removable bulkheads will includecharge wells that will house the black powder charges, along with eye bolt insertions which allows parachutesto be attached. In order to show that these bulkheads can withstand the force of detonation, ground testingwill be done with various charge sizes. In addition to this testing, removable bulkheads will be reinforcedwith carbon fiber plates. The removable bulkhead that is just above the payload will be made from carbonfiber.

The recovery electronics bay bulkheads will be constructed using rapid prototyped ABS thermoplastic forease of manufacturing. To take advantage of the complex geometries available in rapid prototyping, the

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bulkheads will serve to both secure the electronics as well as form the charge wells necessary for black pow-der housing. Additionally, carbon fiber plates, shown in Figure 25, will be secured to both the upper andlower bulkheads for RF shielding of the recovery electronics. Lateral shielding will be accomplished usingthe carbon fiber airframe of the recovery section.

Figure 25: Recovery Bulkhead Design

A permanent bulkhead will be placed in an airframe section above the drag module. This bulkhead willsupport an eye bolt that will serve as the attachment point foe the booster main parachute. Because thebooster is the heaviest section at 15.3 lbm (without motor mass), the parachute opening will subject thebulkhead to very high accelerations. This results in the bulkhead being subjected to high force magnitudes.By making the bulkhead from carbon fiber it is possible that the opening forces will not cause fracture tothe structure. This is because of the high strength to weight ratio of the carbon fiber.

Bulkhead strength as well as charge well strength will be thoroughly tested before flight. Similar ABSbulkhead designs will constructed and tested at a smaller scale. During these tests, a 4 in. diameter ABSbulkhead (thickness of 1/2 in.) will be fitted into a 1/4 in. aluminum housing. The housing would thenplaced in an Instron tensile test machine and pulled until failure. A failure strength will then be measuredand the results will dictate if changes should be made to the bulkhead structure.

The locations of the various bulkheads of the vehicle are shown in Figure 26.

Figure 26: Bulkhead Locations

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3.2.13 Tracking and Communications

The launch vehicle will contain two independent GPS tracking units, each consisting of an Adafruit UltimateGPS v3, Teensy 3.2 microcontroller (MCU) and an XBee PRO 900HP with an attached antenna. Each unitwill have a dedicated power supply to ensure vehicle tracking cannot be compromised by other electronicsystem failures. One of the GPS units will be fitted into the drag modulation electronics bay and the otherwill be contained in the payload electronics bay. The XBee’s transmit in the frequency band of 902-928MHz, feature a software selectable channel mask for interference immunity, and have a nominal range of 6miles. The signal will be picked up at the ground station by another XBee PRO 900HP which will feed thedata into a MATLab program that will be capable of plotting the data in real time over a map of the launcharea.

As a contingency plan for the GPS tracking units, long range RF transmitters will also be used on the vehicle.The Communications Specialists, Inc. RC-HP transmitters, which broadcast on the 222 - 225 MHz hamband, have an effective range of up to 10 miles in clear conditions and a battery life of 1 week. Typical rangesvary between 6 - 12 miles depending on the transmitters altitude, weather conditions, battery strength, etc.These ranges were determined through past use and testing. The RF transmitters will be housed with theGPS units and will be used in case the GPS information can no longer be received as a result of unforeseencircumstances.

41

3.3 Mission Performance Predictions

3.3.1 Flight Profile Simulations

By using Open Rocket’s software, several simulations were done to predict the flight profile of the vehicle.These simulations allowed the possibility to fine tune the vehicle’s design so that it meets the set require-ments. Figure 27 shows an example plot of the various flight parameters while Table 21 shows the resultsfrom these simulations.

Figure 27: Flight Simulation

Table 21: Simulation Results

Wind Speed (mph) 0 5 10 15 20

Velocity Off Rail (ft./sec) 68 68 68 68 68

Apogee (ft.) 5877 5860 5810 5726 5657

Max Velocity (ft./sec) 643 642 642 640 639

Max Acceleration (ft./sec2) 226 227 227 228 228

Time to Apogee (sec) 19.80 19.80 19.80 19.6 19.5

3.3.2 Flight Stability

The stability of the rocket was determined through the use of OpenRocket software simulations. The stabil-ity margin (S) is calculated by using the difference between the known lengths of both the center of gravity(CG) and the center of pressure (CP), which is then divided by the diameter of the airframe (D) seen in

42

Equation 4 below. The launch vehicle is required to have a minimum static stability margin of 2.0 and wasexceeded with a stability margin of 2.54. Flight stability will also be verified during the full scale test flightswith a GPS, three axis accelerometers and gyroscopes (IMU) which will record the flight path, velocity, andacceleration in three dimensions.

To carry out stability calculations, the center of gravity and center of pressure were determined on thevehicle. Figure 28 shows the location of both centers, where the red circle is the CP and blue is the CG.

Figure 28: Center of Pressure and Center of Gravity

The stability margin for the overall vehicle was calculated using Equation 4. This calculation led to a sta-bility value that was well within the set requirements.

Stability =CP − CG

Diameter(4)

Using this equation the value for the stability was calculated. This is shown below.

Stability = 72.15·in.−56.93·in.6·in. = 2.54 ·cal

The stability was also determined for once the motor was completely burned out. This was done becausethe motor fuel is a significant amount of mass that is subtracted from the overall vehicle mass. Figure 29shows the new location for the gravity and center pressures on the vehicle after burnout occurs.

Figure 29: Center of Pressure and Center of Gravity After Burnout

Once again, Equation 4 was used to determine the new stability margin. This calculation is shown below.

Stability = 72.15·in.−48.78·in.6·in. = 3.89 ·cal

This new stability margin shows that the vehicle will still be stable after the motor mass is gone.

43

3.3.3 Kinetic Energy

The calculations that were done with Equations 2 and 3 yielded the value shown in Table 22. These cal-culated values are well below the kinetic energy requirements. These calculations were made with the massestimates, also shown in Table 22. Section 3.2.7 explains the details of the process involved to gain thesevalues.


Section Weight Velocity Kinetic Energy

(1) Payload 17 lbm 1.6 ft/sec 21.76 lbs · ft(1) Nosecone 1 lbm 1.65 ft/sec 1.36 lbs · ft(2) Booster 15.3 lbm 1.52 ft/sec 17.67 lbs · ft

(2) Transition 6 lbm 1.79 ft/sec 9.61 lbs · ft

NOTE: Equal section numbers indicate that the sections are tethered together.

3.3.4 Drift

In an effort to achieve a successful recovery of the vehicle, wind drift must be accounted for in the de-sign. This was done by implementing two different recovery systems for the two independent sections of thelaunch vehicle. First, the booster section recovery procedures will be initiated with a 3 ft. diameter drogueparachute at apogee. This drogue parachute will stabilize the section during freefall; because of the rapidfall velocity, drift will be minimal. A 10 ft. main parachute will be deployed at a lower altitude to achievea successful recovery. For the payload section, the main parachute will be deployed but with the use of atwin Jolly Logic opening system, the parachute will be not be allowed to fully inflate. This deployed closedparachute will act as a drogue for the payload section, until a lower altitude is reached. Here the Jolly Logicswill open and the parachute will be allowed to fully inflate.

With this, calculations were done to identify the amount of drift that would be seen for each section undermain parachute deployment. By using the altitudes at which the main parachute will deploy, as well as thecalculated velocities that were obtained from Equation 2, the time that it would take each section to landwas obtained. This is shown below.

At a 300 ft. main parachute deployment for the payload section:

300 ft.

1.65 ft/sec= 182 sec

At a 600 ft. main parachute deployment for the booster section:

600 ft.

1.79 ft/sec= 336 sec

Using various wind speed estimates, the drift distances were calculated with Equation 5.

Wind Speed · Time to Landing = Drift Distance (5)

44

With this, Table 23 shows the drift distances for the both payload and booster sections. This also includesthe tethered sections.

Table 23: Payload Drift Distances

Wind Speed Payload Drift Distance Booster Drift Distance

0 mph 0 ft. 0 ft

5 mph 1334 ft. 2028 ft.

10 mph 2667 ft. 4056 ft.

15 mph 4000 ft. 6083 ft.

20 mph 5331 ft. 8111 ft.

45

3.4 Technical Challenges

The primary challenge the 49er Rocketry Team faces with launch vehicle design is proper integration of thetwo scientific payloads. The recovery systems and vehicle transition section were designed with the roverand rover housing in mind. Table 24 lists the anticipated technical challenges associated with the vehicle.

Table 24: Launch Vehicle Technical Challenges

Item Challenge Solution

VTC 1Ability to launch again on the same daywith minimal repairs.

Robust design, repackable chutes, andreplaceable fins on hand at launch sitewill ensure relaunch on same day.

VTC 2Bulkheads can withstand black powderdetonations and section weight whenunder parachute.

Bulkheads will be made from robustmaterials and will be tensile tested.

VTC 3The launch vehicle will be capable ofhousing a deployable rover.

A 6 in. airframe was selected for properrover fit.

VTC 4The launch vehicle will be capable ofbeing prepared for flight at launch sitewithin 4 hours.

Vehicle will be designed with assemblyin mind.

VTC 5Sections will separate at designatedheights, to allow parachutes will deploy.

Separation testing will be done to en-sure proper separation.

VTC 6The internal components will be able tosustain the rigors of the flight.

Electronics will be secured to holdingbays that are strong enough to with-stand flight forces.

VTC 7 Fitting of parachutes in recovery bays.Vehicle will be designed to provide am-ple room for parachutes

VTC 8Launch vehicle landing safely withoutdamage to the payload section.

Parachutes will be sized to ensure pay-load lands with minimal kinetic energy.

VTC 9 Ensure fin symmetry about fincan. A jig will be used when cutting fin slots.

VTC 10Ensure structural stability of transitionsection.

Structural stability of transition sectionwill be tested in flight.

VTC 11Ensure vehicle reaches exact altitude of5,280 ft.

Drag Mod system will be incorporatedinto vehicle in order to prevent over-shoot of target altitude.

VTC 12Ensure proper fit of rapid prototypingparts.

A design review process will be imple-mented in order to ensure parts arewithin proper tolerances.

VTC 13Successfully integrate the rover housinginto the payload airframe.

Rigorous testing of payload assembly toensure adequate fit.

VTC 14Ensure proper fit and seal of bulkheadsbetween vehicle sections.

Utilize precision machining and gas-kets.

46

4 Safety

Maintaining a high level of safety is the highest priority with the 49er Rocketry Team. The following sectionsoutline plans for maintaining a safe environment during construction, testing, and operation of the launchvehicle. This section includes, but is not limited to: plans to abide by all NAR/TRA procedures, machineshop safety procedures, rocket motor handling procedures and all hazardous material handling procedures.

4.1 Safety Officer

The 49er Rocketry Team's Safety Officer, Dallas Solomon, is tasked with ensuring that safety remainsparamount in the minds of each team member. The Safety Officer will work closely with the project andteam leads to cultivate a safety conscious mindset as the team moves forward with the design, construction,testing, and launching of the team's launch vehicle. Taylor Forte and Ashton Johnston will act as alternateSafety Officers.

Training

The Safety Officer is required to have detailed knowledge of the National Association of Rocketry (NAR)High Power Rocket Safety Code. Additionally, the Safety Officer will be knowledgeable of the requirementsset forth in the following: NFPA 1127, Code for High Power Rocketry, 14 CFR Subchapter F, Part 101,subpart C-Amateur Rockets, and all governing FAA regulations. The Safety Officer will be familiar with therules and regulations contained in 27 CFR Part 55- Commerce in Explosives. All team members will abideby the launch site rules and all rulings made by the Range Safety Officer. The Safety Officer will be familiarwith and maintain copies of the most current Safety Data Sheets (SDS) for the chemicals/ materials utilizedby the team.

Responsibilities

The Safety Officer is responsible for the following, emphasizing safety:

• Monitor design, construction, and assembly of launch vehicle and payload

– Design of vehicle/ payload

– Construction of vehicle/ payload

– Assembly of vehicle/ payload

– Ground testing of vehicle/ payload

– Sub- and full-scale launch test(s)

– Launch day activities

– Recovery activities

– Preparation and installation of recovery devices

– Educational engagement activities

• Implement team developed construction, assembly, launch, and recovery activities

• Manage/ maintain all copies of Safety Data Sheets, hazard and failure mode analysis, and procedures

• Assist in writing team hazard and failure modes analyses and procedures

• Remains vigilant to identify new hazards during design, construction, and assembly

• Works closely with NAR mentor to ensure proper handling and storage of motors, ignitors, electricmatches, and other pyrotechnics

47

4.1.1 Impact of Delays

Table 38, Risks to Project Completion, identifies circumstances that could potentially cause disruptions tothe project at large. These circumstances include risks to the budget, risks to the schedule, etc. The tableprovides mitigations for these circumstances plus quantification for each. Best estimates for quatification aregiven at this point as the project is still young. As the project moves forward, more realistic and accurateinformation will become available.

A loss of either the sub-scale or full-scale vehicle would cause the greatest delay to the project. A total lossof either launch vehicle has far reaching effects. Each flight of either vehicle increases the likelihood of aloss. The delays would come in the form of reordering parts damaged or destroyed. An event of this naturewould further result in an immediate increase in budget. This increase in budget could potentially resultin a budget overrun. To mitigate the chances of such a loss, all components will be individually tested andoperated prior to final assembly, critical components will highlighted in assembly checklists with two teammembers verifying correct assembly and operation.

4.1.2 Personnel Hazard Analysis

Table 25 through table 28 outline the risk definitions, acceptance levels, severity and probability definitionsused in the evaluation of personnel, environmental, etc. hazards anticipated by the 49er Rocketry Team.These tables are used to interpret the Risk Assessment Codes (RAC) assigned to each personnel hazard,environmental hazard, and identified failure mode.

Table 25: Risk Definitions

Severity

Probability1

Catastrophic2

Critical3

Marginal4

Negligible

A - Frequent 1A 2A 3A 4A

B - Probable 1B 2B 3B 4B

C - Occasional 1C 2C 3C 4C

D - Remote 1D 2D 3D 4D

E - Improbable 1E 2E 3E 4E

Table 26: Risk Acceptance Levels and Approval Authority

Severity/ Probability Acceptance Level/ Approval Authority

High RiskUnacceptable. Documented approval from faculty adviser, project lead,and safety officer. Documented approval from RSO andNASA Student Launch officials, if required. Mitigation required.

Moderate RiskUndesirable. Documented approval from safety officer and project lead required.Notify faculty adviser. Mitigation required.

Low RiskAcceptable. Documented approval from safety officer and project leadrecommended. Mitigation required.

Minimal RiskAcceptable. Review by safety officer recommended.Little to no mitigation required.

48

Table 27: Severity Definitions

DescriptionPersonnel

Health andSafety

Facilities andHardware

EnvironmentalMissionSuccess

1-Catastrophic

Loss of Life,Permanent,disablinginjury

Loss of facility/laband hardware.

Irreversible severeenvironmentaldamage violating lawsand regulations.

Categorized as havingdrastic effects onmission success.

2-Critical

Severe injuryMajor damage tofacility/lab, systems,or equipment

Reversible environmentaldamage causing a violationof law or regulation.

Categorized as havingsignificant effects onmission success.

3-Marginal

Minor injury,occupational-relatedillness

Minor damage tofacilities, systems,or equipment.

Environmental damagewithout violation of lawor regulationwhere restoration activitiescan be accomplished.

Categorized as havingmoderate effects onmission success.

4-Negligible

First aid injuryoccupational-relatedillness

Minimal damage tofacility, systems,or equipment.

Minimal environmentaldamage not violatinglaw or regulation.

Categorized as havingnegligible effects onmission success.

Table 28: Probability Definitions

Description Qualitative Definition Quantitative Definition

A - FrequentHigh likelihood to occur immediatelyor expected to be continuously experienced.

Expected to occur at or greaterthan 50% of the time.

B - ProbableLikely to occur to expected to occurfrequently within time.

Expected to occur between30% and 50% of the time.

C - OccasionalExpected to occur several times oroccasionally within time.


D - RemoteUnlikely to occur, but can be reasonablyexpected to occur at some point within time.


E - ImprobableVery unlikely to occur and an occurrenceis not expected to be experienced within time.

Expected to occur between0.01% and 1% of the time.

49

Table 29: Personnel Hazard Analysis

Hazard Cause EffectPre-RAC

Mitigation

Team memberinjured by motor

detonation (CATO,installation, etc.)

Withdrawaldistances not

observed duringlaunch, improperassembly of solid

rocket motor,ignitor installed

prematurely

Loss of life, limb,eyesight

1C

Safe withdrawal limits willbe observed at all times.

Motor only purchased fromreputable dealers or

assembled by qualified/certified personnel. Ignitorwill not be installed intomotor until rocket is on

launch rail.

Injury resultingfrom Recoverysystem failure

(Booster/ Payload)

Tangled risers,Jolly Logic failure,

or failure ofseparation charges

Injury topersonnel, possible

death from highvelocity impact

1C

Risers will inspected fortangling prior to recovery

assembly. Continuity checksof e-matches conducted prior

to separation chargeassembly. Jolly Logics will

be turned on IAW JollyLogics user’s manual

Team memberinjured fromignition of

separation charge

Stray voltage,E-match connected

to battery,altimeter powered

on

Severe burns tohands, eyes. Loss

of limb1C

Separation chargepreparation limited to Safety

Officer(s). Connection ofseparation charge leads will

be done with batteriesdisconnected and power off

Electrical Shock Exposed wiringSevere burns,death from

electrocution1D

Pre Use inspection on powercords. Report exposedwiring to Lab Manager.

Lithium Polymerbattery explodes

injuring teammember(s)

Overcharging ofbatteries/ batteries

get punctured/batteries

overheated due toimproper storage

Moderate to severeburns

2C

Battery charge times will beobserved at all times.

Batteries will remain clear ofall sharp objects. Batteries

will be charged/ stored IAWmanufacturer’s LiPoWarning and Usage

Disclaimer

Machine Shopinjury

Improper machineuse, failure follow

to shop rules

Loss of limb, severeinjury

2C

Team members must passshop safety test and be

trained on machines. UNCCissues an access badge

allowing use of machiningequipment

Team memberinjury from

inadvertent solidpropellant grains

ignition

Improperhandling/ storage,

motor assemblywithin 25 ft of heat

source

Serious bodilyinjury/ major

burns2D

Motor handlingaccomplished by NAR

certified team members/Safety Officer(s). Motor

grains will not be handled orstored within 25 ft of a heat

source

50


Mitigation

Team memberexposure to radio

frequencytransmission

RF transmissionswithin closeproximity ofpersonnel for

extended period oftime

RF transmissionsof substantialstrength for

extended periods oftime have beenknown to causesterility, cancer,

etc.

2D

RF testing/ operation willbe conducted by minimal

personnel for as shortamount of time as necessaryto adequately test system.

Separation distance betweenantennae and persons will be

>25cm

Team memberinjury resultingfrom unstable

launch

Vehicle stabilitymiscalculation,

high winds

moderate to severeinjury to personnel

2D

Vehicle designed withminimum static stability

margin of 2.0 IAW NASA SLHandbook Statement of

Work para. 2.16, launchesnot performed when windsexceed 20 mph IAW NARHigh Power Rocket Safety

Code, para. 9 ”FlightSafety”

Team memberinjured duringground testing

Launch vehiclecomponent striking

personnel

Bruising, minorcuts, scrapes

2D

Safety Officer will ensurepersonnel are cleared fromthe testing area prior to

conducting tests.

Injury fromE-Match ignition

Improperhandling/ stray

voltageMinor burns 3C

E-match installation/connection will be

accomplished with power offand batteries disconnected

Soldering burns Inattention to task Minor burns 3CPersonnel will focus on taskand exercise caution when

using soldering iron

Team member(s)inhale carbon

fiber/ fiberglassdust/ particles

Lack of orimproper use of

personal protectiveequipment

Irritation to lungsand mucusmembranes

3C

PPE will be worn whilemachining or cutting

material IAW 29 CFR Part1910 Subpart I para. 132(d)

Personnelinhalation/ contact

during chemicaluse

Improper/ lack ofuse of personal

protectiveequipment

Respiratory, skin,eye irritation

3DSDS will be reviewed priorto use. PPE will be used

during use.

Finger(s) caught inrover tracks

inattention to taskMinor cuts,

bruising4C

Rover will be placed onsuitable stand or surfaceduring testing of rover

tracks. Personnel will keepfingers clear of moving partswhile rover is in operation

51

4.1.3 Failure Modes and Effects Analysis

Table 30: Booster Section FMEA

Subsystem/Component

Failure Mode Cause EffectPre-RAC

Mitigation

Motor CATO

Fracture ofmotor casing orimproper grain

packing.

Catastrophicloss of vehicle

at take-off.1A

Ensure licensedvendor for purchasingof motors. Inspection

of motor assemblyprior to use.

Ultemstructural

failure.

Excessive heatnear boattail.

Loss of motorretention.

Catastrophicmission failure.

1D

Proper design suchthat boattail is

minimally exposed toexhaust. Inspect

boattail section beforeand after flight forstructural damage

Motor retention

boattailattachment

failure

Impropersecuring of

boattail to thevehicle

Loss of motorretention.


1DInspect boattail forsecurity and proper

installation

Fracture of finduring flight

Excessive shearforces acting on

the fin

Drasticallyaltered flight

profile and lossvehicle stability

1CInspect fins before and

after flight forstructural damage

Fin attachmentfailure

Impropersecuring of

boattail to thevehicle



1DInspect fin and

boattail assembly forproper installation

Fins

Fracture of finupon recovery

Excessive shearforces acting on

the fin



3A

Recovery systemdesigned to minimize

impact force uponlanding

52

Table 31: Drag Modulation System FMEA

Subsystem/Component


Mitigation

Fracture of flap

Excessiveshearing forcesacting on the

flap


profile and lossof drag

modulation.

1C

Design flaps toaccommodate loadswith safety factor.

Inspect flaps beforeand after flight forstructural damage.

Drag Flaps

Fracture ofhinge pin

Excessiveshearing forcesacting on the

pin

Loss of flapcontrol

1D

Design hinges toaccommodate loadswith safety factor.

Inspect flap hinges forstructural damage and

migration of pins.

Power loss.

Improperassembly.

Insecure powersupply.


profile and lossof altitude

control.

2B

Detailed assemblychecklist with dual

verification of powersupply connections.Implement securitymeasures for power

supply.

MicrocontrollerImproper

configuration orconnections.

Loss of altitudecontrol.

3D

Detailed checklistsinsuring proper

electronics assembly.Hardware-in-the-looptesting and tuning.

Control Electronics

Software failureImproperMATLAB

Simulink model.

Altered flightprofile.

3D

Tuning of Simulinkmodel. Ground

testing of code anddrag mod operation.

Steel Push Pins Binding

Improperassembly.Improper

tolerances whenmanufactured.


profile and lossof altitude

control.

1D

Fit check componentsupon manufacture.Operational test of

drag mod and controlsoftware upon

assembly.

Retraction SpringsInsufficient

force forretraction.

Damage tospring during

assembly.Required torque

to moveunpowered

servo motor toogreat.

Altered springconstant.

Possible loss ofsafety

mechanism.

1D

Inspection of springprior to assembly andfunctional test after

assembly.

Mechanical Control

Loss ofconnection

between servoand turning

puck

Improperattachment

between servoand mechanism

Loss of dragmodulation

3D

Inspect and functionalcheck drag mod for

proper assembly andoperation prior to

connection to launchvehicle

53

Table 32: Recovery FMEA

Subsystem/Component


Mitigation

Failed ejectioncharge.

Defectiveigniter.

Defectivealtimeter. Black

powder leak.


Parachute fails todeploy.

1B

Continuity test ofe-matches prior to

flight. Use of 2E-matches per ejection

charge.

E-matchdisconnection

from altimeters

Assembly ofaltimeter bay

removese-match leads

from altimeter.


Primary parachutedoes not deploy.

1B

Dual verification ofE-match connection to

altimeter. SecureE-match wiring toinner portion of

recovery electronicsbay.

ParachuteDamage from

Ejection Charge

Improperpacking ofrecovery

parachutes.

Altered dragcharacteristics ondescent. Possible

catastrophicmission failure.

1CAll parachutes will be

wrapped in fireretardant material.

PayloadReefing Ring

Tangling inshroud lines.

Primary parachutefails to fully

deploy. Possiblecatastrophic

mission failure.

1C

Parachute shroud lineswill be inspected toensure reefing ring

properly installed andnot tangled in shroud

lines.

Recovery

Tangling ofshroud lines


parachutes.


Possiblecatastrophic

mission failure.

1C

Parachute shroud linesinspected and all

tangles removed priorto packing.

Power loss toaltimeters

Power supplybecomes loose,

breaksconnection.



1C

Implement securitymeasures to altimeterpower supply. Dualverification of secure

connections.

Jolly Logicparachute

Release failure

Insufficientbattery charge.Wrong release

setting.


Possiblecatastrophic

mission failure.

1D

Charge Jolly Logicparachute release after

every flight. Useredundant Jolly Logic

parachute release.Dual verification ofrelease setting and

operation.

54

Subsystem/Component


Mitigation

Payload

Shearing ofbulkhead from

airframe

Excessive stresson bulkhead.

Catastrophicmission failure. Noform of controlled

descent.

1D

Ensure adequate setuptime for adhesive

securing bulkhead toairframe. Ground testsecurity of adhesive.

Recovery

GPS ReceiverPower loss toGPS receiver

Inability to locatepayload section via

GPS.3C

Implement securitymeasures to GPS

power supply. Dualverification of secure

connections.

E-matchdisconnects

from altimeters

Assembly ofaltimeter bay

removesE-match leadsfrom altimeter.



1B

Dual verification ofE-match connection to

altimeter. SecureE-match wiring toinner portion of

recovery electronicsbay.

Vehicle BoosterFailed ejection

charge.

Defectiveigniter.

Defectivealtimeter.


Parachute fails todeploy.

1B

Continuity test ofe-matches prior to

flight. Use of 2E-matches per ejection

charge. Altimeterstested for operation

prior to use.

Section Recovery

ParachuteDamaged from

Ejection Charge


parachutes.

Altered dragcharacteristics ondescent. Possible

catastrophicmission failure.

1C

Parachutes will bewrapped in fire

retardant material toprevent damage

55

Subsystem/Component


Mitigation

Tangling ofshroud lines


parachutes.


1C

parachute shroud linesinspected and all

tangles removed priorto packing.

Vehicle Booster

Power loss toaltimeters

Insecure powersupply.



1C

Implement securitymeasures to altimeterpower supply. Dualverification of secure

connections.

Section recoveryShearing of

bulkhead fromairframe

Excessive stresson bulkhead.

Catastrophicmission failure. Noform of controlled

descent.

1D

Ensure adequate setuptime for adhesive

securing bulkhead toairframe. Ground testsecurity of adhesive.

GPS ReceiverPower loss toGPS receiver

Inability to locatepayload section via

GPS.2C

Implement securitymeasures to GPS

power supply. Dualverification of secure

electrical connections.

Table 33: Rover FMEA

Subsystem/Component


Mitigation

Rover fails tofully disengage

from theleadscrew

Failure tounscrew fromchassis guide

nut

Rover fails todeploy

1A

Ground testing fromdifferent droppedheights. Full scale

flight test.Signal fails toactuate and

power on rover

Loss of signal.Failed MCU.

Rover fails todeploy

1AMultiple ground testsand full scale flight

test

One or bothdrive servos fail

to operate

Loss of powerdue to

electronicand/or battery

power

Drastically altereddisplacement.

Failure to completethe 5 ft. minimum

distancerequirement

1A

Include a high safetyfactor for torque.

Complete a checkoutof all wiring, batterypower, and electronic

signals.

Fails to operate

Electroniccomponent(s)

failure

Complete failure tomaintain motion

1A

Double check allwiring and electronicsources in accordance

to a check list.Pre-test full scale

rover.

Rover

upon deployment

Loss of tractionDrastically altered

displacement1B

Design tank tracks tobe able to stay in

optimum contact withthe ground. Test

rover.

56

Subsystem/Component


Mitigation

Rover binds inhousing

Housing warpsduring

parachutedeployment

and/or landing

Rover might notexit to complete

the mission1B

Test subscale and fullscale housings on

launches.

Rover tread(s)break

Sheared Pinand/or

fractured tread

Treads fall ofand/or bind

potentially causingthe rover to stopmotion or cause

MCU tomisinterpret overall

distance

1C

Design treads and pinsto take force of

objects. Pretest onsimilar terrain.

Insufficientpower in roverbattery pack

Miscalculationof needed power

Slowedperformance

and/or total roverfailure

1C

Include a high safetyfactor on power

consumption. Checkall batteries before

each flight.

Fails to travelBinded tracks

Drastic decrease inspeed and/or

complete failure tomove

1CDesign and test tracksfor proper tension and

operation.>5 ft.

Code ErrorMCU miscalculates

position or totaldisplacement

1DGround test all

coding. Pre-test rover.

Solar panels failto open

Servo failure orbinding

Panels do notdeploy

2CDouble check torque

on servo. Pre-testdeployment function.

Rover

Objectdetection arrayfails to operate

Unable tonegotiateobstacles

Collision withdebris

2D

Pre-test component toensure proper object

recognition. Useredundant GPS unit

57

Table 34: Rover Housing FMEA

Subsystem/Component


Mitigation

Housing willnot rotate

Hard impactcauses

out-of-roundcondition

Rover housingcannot orient rover

1ADesign housing towithstand impact

force

DC Motor failsto rotateleadscrew(binds)

Failure toprovide enoughtorque. Binding

within thebulkhead

and/or roverconnection

point

Rover does notdisengage from

leadscrew1A

Select motor withenough torque torotate leadscrew.

Ensure leadscrew andleadscrew nut share

idential threadpatterns.

Housing capfails to break

away

Does notunthread from

leadscrew

Rover fails todeploy from

housing1A

Design the cap andleadscrew to unthread

from each other.Ensure leadscrew and

housing cap shareidentical thread

patterns.Housing

electronics failto receive

deploy signal

Failure in code/loose wires


housing1A

Create correct codeoperation. Check

wires.

Rover Housing

Rover bindsduring

leadscrewoperation

leadscrew bindswithin the

leadscrew nutattached to thechassis of rover


housing1A

Dimension leadscrewand leadscrew nut tobe in the same x-aixs

orientation. Selectcorrect sizing forcomponents to

interaction

58

Subsystem/Component


Mitigation

DC Motor failsto operate

Battery failure/disconnection

leadscrew fails torotate, causing

rover not to deploy1A

Select appropriateaccelerometer. Orientthe accelerometer to

desired position.

Housing fails

Housingdeformation

upon parachutedeployment

and/or impact

Rover comes out inan undesired

direction or fails todeploy from

housing

1C

Design housing towithstand parachutedeployment force and

landing impact.Design recovery to

minimize impact forceupon landingto orient

Accelerometerfails to detect+z direction

Rover deploys inan undesired

direction, possiblycausing failure to

move

1C

Select appropriateaccelerometer. Orientthe accelerometer to

desired position.

Rover Housing

Leadscrew failsto fully

disengage fromrover

Leadscrew doesnot produce

enoughrotations to

fully disengagerover

Increaseddeployment time.Possible failure to

deploy

2A

Code the lead screwto rotate more than

necessary to disengagefrom rover

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Table 35: Launch Operations and Support Equipment FMEA

Subsystem/Component


Mitigation

Igniter fails to

Corrosion onlaunch module

firing leads

Rocket fails tolaunch

2C

Remove corrosionfrom launch modulefiring leads with wire

brush prior toconnection to igniter

leads

Launch

initiate launch

Faulty igniterMisfire during

launch sequence/failure to launch

2C

Verify continuitythrough igniter priorto installation into

rocket motor

OperationsLaunch vehicle

fails to fitlaunch rail

Rail buttonstoo large or too

smallScrubbed launch 2D

Verify types/ sizes oflaunch rails available.

Use rail buttonscompatible withlaunch site rails.

and Support

Failure to pass

Launch vehiclecomponents not

adequatelysecured.

Refusal by RSO tolaunch or delay of

planned launch3C

Safety Officer/ ProjectLead final preflight

inspectionaccomplished prior to

RSO inspection.

Equipment

RSO inspection

Failure toidentify apogeeevent to RSO

Refusal by RSO tolaunch or delay of

planned launch3D

Identify apogee eventto RSO. Document

settings duringassembly. Verify/validate altimetersettings in RSO’s

presence.

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4.1.4 Environmental Hazard Analysis

Table 36: Environmental Effects on Launch Vehicle


Mitigation

Wet/Damplaunch field

Rain. Recentwatering of field

Damage to electronicsystems and loss of

flight data.1C

Electronics kept in waterproofcontainers prior to installation

in launch vehicle.

Inclementweather, e.g.

rainStorm fronts

Damage to electronicsystems

1C

Suitable cover, i.e. pop-upcanopy, will be available at

every launch. Electronics keptin waterproof containers prior to

installation in launch vehicle.

TerrainLaunch fieldtopography

Failure to successfullydeploy payload

2BTesting and verification of

prototype(s) on multiple typesof terrain.

Excessive windspeeds

Storm frontsAltered flight profile/

recovery off launchfield

2C

Monitor wind speedsconsistently throughout day,

paying particular attention towinds speeds while launchvehicle is on launch rail.

Launches not performed whenwinds exceed 20 mph IAW NAR

High Power Rocket SafetyCode, para. 9 ”Flight Safety”

Low visibilityFog/ low cloud

ceiling

Inability to tracklaunch vehicle afterrecovery deployment

2C

Monitor weather throughoutday. No launch conducted withlow visibility IAW 14 CFR Part101 Subpart C para. 101.25(b)

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Table 37: Launch Vehicle Effects on Environment


Mitigation

Rocket impactsbystanders or

vehicles

Excessivewinds.

Injury to bystanders.Death. Damage to

vehicles orsurroundingequipment.

1C

Monitor wind speedsthroughout day. Launch will

not be performed if wind speedsexceed 20 mph IAW NAR HighPower Rocket Safety Code para.

9 ”Flight Safety”.

AsymmetricControl Surface

Deployment

Improperassembly.Improper

operation ofcontrol system.

Altered flighttrajectory and damage

to surroundingenvironment.

1C

Member(s) will follow allassembly procedures and

complete functional tests priorto flight.

Loss of launchvehicle

Failure ofrecovery system

Pollution to launchfield and surroundingwildlife due to high

velocity impact.

1B

Use of drogue system to reducedrift distance. Deployparachutes in reefed

configuration IAW NAR HighPower Rocket Safety Code para.

12 ”Recovery System”.

Inadvertentignition of blackpowder chargesor motor grains

Spark or flamenear blackpowder.

Prematureinsertion of

igniter wire intorocket motor

Damage to launchfield. Possible loss ofwildlife. Fire. Severe

burns. Damage toproperty

1C

Handling of black powder/rocket motors will be limited tosafety officer(s)/ project lead/

NAR mentor. All ignitionsources will be kept a minimum

of 25 ft. away from allflammable items IAW NAR

High Power Rocket Safety Codepara. 3 ”Motors”.

Epoxy orHardenerSpillage

Impropersecurement of

epoxies orhardeners.

Possible loss ofwildlife. Damage to

equipment or facilities.1C

Team members will secure anyopen containers when not in useor when leaving work area IAW

3M Epoxy safety data sheetSection 7, para. 7.2

FireRocket motor

ignitionDamage to launch

site. Spreading of fire.2C

Verify availability of fire controlequipment before launch. Clear

blast zone of all flammablematerial. Use of blast deflector.NAR High Power Rocket Safety

Code para. 7 ”Launcher”.

LithiumPolymer (LiPo)

ignites atlaunch field

Impropercharging/

dischargingcausing internal

damage tobattery

Explosion of battery,igniting fire to launchfield. Death to wildlife

2D

Battery maintenance limited toqualified team members. All

guidance set forth inmanufacturer’s Warning and

Usage Disclaimer will befollowed

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4.1.5 Risks to Project Completion

Table 38: Risks to Project Completion

Risk Impact Likelihood Mitigation Quantification

Loss of full- orsubscalevehicle

High Medium

All components are tested forproper operation/ assembly

prior to use on sub- or full scalevehicle

Loss of sub- or full-scale vehiclewill double our budget due to

having to rebuild and will resultin a several week setback in

construction and testing.

Long Leadtime for parts

High MediumParts with long lead times willbe ordered early to alleviate apotential schedule compression

A long lead time will cause anschedule setback equal to the

lead time and delay construction

Damage toelectronics

High MediumElectronics secured when not in

use and only used whenabsolutely necessary

Damaged electronics will resultin budget increases stemmingfrom their replacement, i.e.

Stratologger altimeter cost =$55, Jolly Logic parachute

release = $130

Loss/ lack offunding

High MediumCrowdfunding efforts, additional

fundraisers planned

Wrong partsreceived/ordered

High MediumParts orders are submitted toproject lead/ manager with

justification and intended use

Ordering wrong parts will incurincreased budgetary expenses

and cause schedule delayswaiting for the correct parts.

Payload failsto integratewith launch

vehicle

High Medium

vehicle and payload designedwith integration in mind.Extensive testing will beconducted to validate.

Integration failure wouldpotentially cause a complete

redesign of the payload resultingin a setback of several days to a

couple of weeks.

Custom madeparts don’tmeet design

specifications

High Low

Clear designs, drawing,specifications, etc. are made

available to the manufacturer ofany custom parts. Effective

communication to supplier iskey.

The schedule setback incurredwould be equal to the lead timerequired for the custom part to

be made a second time. Thebudgetary costs could result indouble what was budgeted for

the custom part.

Budgetoverrun

Medium Low

Only items required for thecompletion of testing and

construction are ordered. Onlyminimal quantities ordered.

All superfluous parts/components/ hardware orderedresults in a budgetary increase

possibly taking away frombudget of other critical

necessities

Deadlines notmet

Medium Low

Project plan is available for allteam members to stay on task.Weekly meeting held to discuss

all work: completed, inprogress, and planned. Cleardirection and time line givenand managed by project lead

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Risk Impact Likelihood Mitigation Quantification

Team memberremoved fromteam due to

injury

Medium Low

Team members are required tobe trained on and pass a safety

test to use machiningequipment.

Remaining team members wouldexperience an increase workloadwith the majority of the injuredperson’s workload concentrated

in their respective subteam

Parts/hardware lostin shipment

Medium Low

All parts ordered will be trackedusing supplier provided tracking

number. Will work withsupplier at first sign of possible

loss.

Lost shipment increases budgetthe cost of the reordered partsand their respective shipping

costs.

Conflicts inteam

members’schedules

Medium LowCommunication of potential

conflicts between Project leadand project team members.

UNCCprohibited

fromcontinuing

Medium Low

All NASA SL and UNCCguidelines will be followed by allteam members to ensure project

continuance

UNCCrestricts lab

useMedium Low

All members will follow allestablished lab rules. Team

members required to take/ passa lab safety test governing the

lab rules

Exceedingwaiver

approvedaltitude atlaunch site

(FAAviolation)

Medium Low

Only launch sites with approvedFAA waivers will be used forsub- and full scale launches.

Launches will not exceedmaximum FAA-approved

altitude: ROCC (Midland, NC)4000 ft. AGL; ICBM/ ROSCO

(Camden, SC) 9000 ft AGL

Potential exists to be bannedfrom launch sites for exceeding

FAA waiver resulting inschedule delays as much as a

week or more.

Excessivehardware in

designMedium Low

All sections designed to utilizeminimal fasteners while

maintaining high level of safety

Excessive hardware in designincreases budgetary costs bothin parts and in shipping of saidparts. Additional fund raisingmay be necessary resulting intime taken away from design,

construction, test, etc.

Naturaldisaster

Medium LowGood planning, i.e. back up

launch dates, timelyconstruction

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5 Payload Criteria

5.0.1 Mission Statement

The 2017-2018 UNCC Payload Team’s mission is to design, build, and test an autonomous rover that willbe retained in the vehicle airframe through launch and landing. The rover must be capable of navigating atleast 5 ft. from the landing site before deploying a foldable solar panel array. In addition to the rover, theteam will design, build, and test a rotating housing capable of retaining, orienting, and deploying the rover.

5.0.2 Mission Success Criteria

The mission will be considered successful if it meets all payload and safety requirements outlined in the 2018NASA Student Launch Handbook and meets the following criteria:

PS1 - The rover housing successfully protects the rover through launch, separation, and landing

PS2 - The ground team successfully transmits a signal to initiate deployment

PS3 - The rover housing orients the rover within ± 5° of a normal orientation

PS4 - The leadscrew properly deploys the rover from the rover housing

PS5 - The rover autonomously travels a minimum distance of 5 ft. from the vehicle

PS6 - The rover deploys its foldable solar panel array

5.1 Selection, Design, and Rationale of Rover Payload

The rover must accomplish three tasks: (1) survive launch and recovery, (2) travel a minimum of 5 ft. fromthe vehicle after landing, and (3) deploy folding solar panels after traveling the required distance. Survivalof flight is defined as all primary rover subsystems functioning to allow autonomous movement away fromthe airframe. Solar panels will deploy once the rover is outside of a 5 ft. radius from the landing site. Anumber of design alternatives were considered for each major section of the rover payload. Pros and cons ofeach alternative are identified and tabulated in the following sections.

5.1.1 Rover Drive System

In order to ensure mission success, the rover must be capable of moving over unknown terrain while stillbeing able to fit inside the airframe of the launch vehicle. The size of the rover depends on the chosen drivesystem, so the team began the design process by comparing various wheel and tread systems.

The first drive system considered was a continuous track, or tank tread type system. Due to their large con-tact area, tracks give unparalleled traction and stability in rugged, off-road environments. A larger contactarea also reduces ground pressure, though this was a minor consideration due to the low projected weight ofthe payload. With the expected landing site being a tilled field of an unknown soil composition, tank treadsprovide the best option for traversing the highly variable terrain; however, the drawbacks of a continuoustrack design are not insignificant. Tracks require a large number of moving parts and linkages, making themmore susceptible to failure. Tracks will also take a long time to assemble unless a one-piece molded tread isused. Continuous tracks will also add friction to the drive system and increase power requirements.

The next drive system the team considered was a more traditional four-wheeled design. Wheels are com-monly used for small robot designs due to their low cost, ease of design and assembly, and generally lowweight. They are also capable of achieving the same level of traction as tank treads if they are very wide orof a large diameter. This, however, would not be an option due to the requirement that the rover must fit

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in the vehicle airframe. Wheel compliance and/or a suspension system could improve stability in off-roadenvironments, but would also drastically complicate the rover design and assembly. An additional concernis steering, which could be accomplished without a steering mechanism if the wheels are sideways coupledand differentially driven (one side driven forward, and the other backward); however, the maneuverabilitybenefit of wheels would be negated with the requirement that the vehicle skid-steer. The addition of asteering system would further complicate a four-wheeled rover design.

Legged drive system designs were also considered. This type of drive system can be designed with a varietyof leg shapes and are very robust in off-road environments due to their ability to clear obstacles greaterthan half of the underbody clearance. Legged vehicles were found to demand significantly higher power thantracked or wheeled vehicles, and would be difficult to design for the tight confines of the vehicle airframe.

The last option the team considered was a large two-wheeled “axle rover”. This design would allow for largerwheels to be used by placing them coaxial to the housing, making their maximum diameter equal to theinterior diameter of the airframe. This design has several attractive benefits including increased clearanceheight, relatively simple design and manufacturing, and increased traction over smaller wheel designs. Thedrawback is the lack of available chassis space for the electronics necessary to make the rover autonomous.

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Table 39: Rover Drive System Pros and Cons


Tank Treads

• Maximum traction

• Inexpensive to replace individ-ual links and pins

• Best weight distribution

• More robust

• Clears uneven terrain

• Longer manufacturing leadtimes

• Does not fit easily in vehicle

• Track failure immobilizesrover

Wheels

• Easy to manufacture

• Low failure potential

• Maneuverable

• Agile

• Lower traction

Legs

• Larger effective drive diameterfor obstacle clearance

• Difficult to manufacture

• Potential for failure or entan-glement with obstacles

• Independent rotation requiresmultiple servos, adding weight

Axle Rover

• Large wheel diameter providessignificant ground clearance

• Fits easily in vehicle

• Difficult to deploy

The decision matrix used by the team to decide the leading design for the rover drive system is shown inTable 40. The criteria considered for drive system selection included ease of design, ease of assembly, rovertraction, robustness, timeline, cost, and rover weight.

Ease of design, ease of assembly, and rover traction were given the highest weighting, showing that the teamconsiders these the most important criterion, compared to 4 points for robustness, 3 points for timeline andcost, and 2 points for rover weight. Each rating is multiplied by the weight for that criterion and the scoresadded across each row. The design with the highest total score was selected as the leading design option.

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Table 40: Rover Drive System Decision Matrix

Criteria: Ease ofDesign

Ease ofAssem-bly

Traction Robust-ness

Timeline Cost RoverWeight

TotalScore

Multiplier: 5 5 5 4 3 3 2

ContinuousTrack

4 3 5 4 3 4 4 105

Wheels 4 5 2 3 4 4 5 101

Legs 2 4 5 3 3 2 3 88

LargeWheels/AxleRover

3 4 4 4 4 3 4 100

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5.1.2 Continuous Track Assembly Method

After deciding to use a continuous track system to drive the rover, the team explored how to assemble thedrive system. Continuous tracks can be constructed in a number of ways, but generally fall into one of twocategories: (1) a modular system of comprised of multiple plates that are linked together, or (2) a continuousband of molded rubber.

Modular links provide more options for tread material choice and construction methodology. The links andsprockets can be manufactured using RP plastics or traditional machining methods, both of which give re-peatability of the tread dimensions to ensure a precision fit. Modular tracks allow for replacement of failedlinks without having to replace the entire tread, and allow the team to quickly and easily adjust the lengthof the track should the need arise. The drawback of a modularly linked track system is that it extends theassembly time for the rover significantly over a one-piece molded track. Also, modular tracks inherentlycontain a high number of moving parts, increasing the risk of tread failure.

A molded track design would allow for quick assembly and a drastically reduced number of moving parts. Forthe track to be flexible, the material would be restricted to synthetic silicone rubbers given the equipmentavailable to the team. Also, the only feasible method for constructing molded tracks is the use of a sacrificialmold. This method would increase the cost and design time of the rover track system significantly, and doesnot allow for adjustments to the design to be easily implemented.

Table 41 summarizes the various pros and cons related to the tread construction methods the team considered.

Table 41: Tread Construction Method Pros and Cons


Modular Links

• Link failure allows replace-ment of only a portion of thetrack

• Adjustable length

• Many material and manufac-ture options

• Extends assembly time

• Numerous moving parts

Continuous One-PieceTread • Simplifies assembly

• Few moving parts

• Failure requires replacementof an entire track

• Fixed Length

• Material and manufacturinglimitations

5.1.3 Tread Sprocket Configuration

Driving the continuous track system requires at least one driven wheel and some combination of additionalrolling elements in order to maintain tension on the track and distribute rover weight over the terrain. Theteam considered three primary designs: (1) one driving sprocket and two idler sprockets, (2) one drivingsprocket and one idler sprocket, or (3) one drive sprocket, an idler sprocket, and multiple independent spring-loaded rollers.

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Adding an additional idler sprocket increases rover weight distribution over the deployment terrain. Twoidler sprockets also reduce the likelihood that lateral forces during skid steering will disengage the track fromthe sprocket teeth. The additional idler sprocket does not greatly increase design complexity; however, itdoes add weight and reduce flexibility in the track, diminishing the track’s ability to conform to terrain.

A single non-driven sprocket reduces rover complexity, assembly time, and weight. Two disadvantages to asingle non-drive sprocket are reduced load distribution over the length of the continuous track and a higherlikelihood of the track disengaging from the sprockets during skid-steering.

The final option the team considered was to add multiple independent rollers between the driven sprocketand a single idler sprocket. Each roller would be sprung to apply downward force on the tread that wouldtranslate to the deployment terrain, improving traction. Independent rollers would add significant weight,cost, and drastically increase the design complexity of the drive system.

The current design utilizes one driving gear and two idler sprockets. The pros and cons of each designalternative considered are shown in Table 42.

Table 42: Tread Configuration Pros and Cons


For each side: Onedriving gear, two idlerwheels

• More contact surface on theground

• Prevents tread disengage-ment from wheels duringskid-steering

• Better load distribution overtread area

• More moving parts, thus morefailure points

• Less flexibility in the tread

For each side: Onedriving gear, one idlerwheel

• Less manufacturing and as-sembly

• More flexibility in the tread

• Conforms better to terrain

• Less Load Distribution

• Higher likelihood of track dis-engagement leading to immo-bilization

Multiple IndependentSpring-Loaded Rollers • Increased traction due to

pushing against the ground atmultiple contact points

• Extra weight

• Drastically increases de-sign/manufacturing complex-ity

• More Expensive

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5.1.4 Tread Material Selection

The team considered a wide variety of materials when designing the treads. The primary considerationsincluded cost, robustness of the material, and ease of manufacturability.

Aluminum offers a low cost and high strength option for the treads; however, manufacturing lead times andoverall costs, due to machine time and labor, prevent it from being a viable option. Based on the currentdesign, each side of the rover will require approximately 32 tread links with tight tolerancing. Due to thesmall scale of the continuous track components, traditional machining would create significant manufacturinglead times.

ABS, nylon, and polycarbonate provide similar tensile strengths and are all available for use as RP materials.Due to their similar characteristics, any one of these materials could be a viable option for the tank treads.Based on cost alone, ABS is the leading material choice for the tread links.

Table 43 summarizes the various pros and cons associated with the materials the team considered, whileTable 44 contains the material properties of said materials.

Table 43: Tread Material Pros and Cons


ABS

• Lightweight

• Inexpensive

• Numerous Post-ProcessingOptions

• High Strength

• Rigid

• Can warp during production

Nylon

• Lightweight

• High Strength

• Flexible

• Expensive

• More limited post-processingtechniques than ABS

• Flexible

Polycarbonate

• Lightweight

• High Strength

• Expensive

• More limited post-processingTechniques than ABS

Aluminum

• High Strength • Heavy

• Difficult to Manufacture

• Long Manufacturing LeadTimes

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Table 44: Tread Material Properties



ABS 1890-9430 0.04 Low

Nylon 8100-14500 0.02-0.16 Low

Polycarbonate 8500-10200 0.04-0.05 Low

Aluminum 6061 40000 0.09 High

5.1.5 Rover Dimensioning

Overall rover dimensions were dictated by the components needed for the rover drive system and for theretention hardware. The width of the current rover design allows a 5/16 in. leadscrew to pass betweenall components mounted on the underside of the chassis. The width was also restricted by the airframediameter, but this was not as important in the decision making process. The decision was made early on tofirst minimize the rover width (based on the drive and retention system components) and then determine theappropriate airframe diameter to house the rover. This was done to reduce the airframe diameter requiredfor the payload section, which in turn keeps vehicle weight and drag forces to a minimum.

The length of the rover was selected to maximize the available space on the chassis for electronics and batterieswithout adversely impacting turning ability. Maneuverability was the main concern for this decision. Alonger tread length would result in a wider turning radius and higher stresses on the tread pins due to thelateral forces caused by skid steering. The increased turning radius could hinder the rover while navigatingthe terrain, and the increased stresses on the treads could overload the servos or disengage the tracks fromthe sprockets. According to the Second Edition Handbook of Industrial Robotics, the nominal length towidth ratio of a tracked robot is ideally between 1.0 and 1.8 for a balance of turning ability and traction.With an estimated length and width of 8 in. and 4.8 in. respectively, the current design has an aspect ratioof 1.7 which falls in this range.

5.1.6 Chassis Design

The team considered multiple design alternatives for the rover chassis, including placement of the chassisabove the track system and placement of the chassis between the tracks. When comparing chassis posi-tion options and general design of the chassis structure, the main considerations were fitting the rover inthe housing, component placement and mounting, and the overall mobility of the rover. The chassis wasrestricted to a generally rectangular shape by the use of the continuous track system. This allows for easydesign and assembly of the rover due to the simple geometry and maximized surface area. The placementof the chassis was the only remaining decision aside from the material used to construct it.

Mounting the chassis between the treads minimizes the height and lowers the center of gravity of the rover.The amount of material needed to manufacture in this configuration would be reduced, which would lowerweight and cost. The drawback to this configuration is the reduction in available area for mounting compo-nents.

The other mounting option the team considered was to place the chassis above the treads, maximizing theusable area for electronics and other components. This configuration would increase the height and raise thecenter of gravity of the rover, making it more likely to roll on slopes. The increased height would also makethe rover harder to fit in the payload housing. An additional concern was that small rocks or other debriscould lodge between the tracks and the chassis, potentially disabling the rover. This alternative would alsorequire additional components and material to align the axle and motor centerlines, increasing the weight ofthe rover and adding design complexity.

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The current design calls for the chassis to be located between the treads. Table 45 summarizes the pros andcons of the chassis position alternatives the team considered.

Table 45: Chassis Position Selection Pros and Cons


Between Treads

• Shorter than above treads;lower center of gravity

• Minimizes chance of tread in-terference

• Less material/lighter

• Less mounting area

Above Treads

• More surface area for electricalcomponents

• Taller than between treads;higher center of gravity

• Debris could lodge betweenunderside of chassis and treads

• Additional components or ma-terial to align axle and motorcenterlines

5.1.7 Chassis Material Selection

Material selection for the chassis was based on the need for low weight, high strength, and easy manufac-turability. The primary options included: (1) various RP plastics (ABS, nylon, or polycarbonate), (2) steel,or (3) aluminum.

RP plastics offer short manufacturing lead times, easy customization, and minimal post-processing, but havethe lowest strength of all the material options, making them more susceptible to breakage. RP plastics aregenerally less rigid than metals as well, meaning a chassis made from these materials would deflect more dueto driving and turning forces. An additional drawback to 3D printed plastics is the high cost of the materialsas compared to aluminum or steel.

Aluminum and steel are more durable and lower cost options for the chassis, but require more time formanufacturing and post-processing than plastics. Between the two, aluminum is the lighter material andis much easier to machine. Aluminum also has a higher strength to weight ratio than steel, making it theleading material choice for the rover chassis.

Table 46 summarizes the various pros and cons associated with the chassis material options the team con-sidered. Material properties are listed in Table 47.

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Table 46: Chassis Material Pros and Cons


RP Plastics

• Short lead time

• Minimal post-processing (lessstudent labor)

• Easily customized

• Low strength

• Susceptible to failure

• High material cost

Aluminum

• Lower material cost than RP

• Durable

• Easy to machine

• Medium material cost

• High labor utilization

Steel

• Strong

• Low material cost

• Heavy

• Difficult to machine

• High labor utilization

Table 47: Chassis Material Properties



ABS 1890-9430 0.04 Low

Nylon 8100-14500 0.02-0.16 Low

Polycarbonate 8500-10200 0.04-0.05 Low

Aluminum 6061 40000 0.09 High

Steel 115000 0.28 Medium

5.1.8 Solar Panel Placement and Deployment

The current design calls for the solar panels to be located at the topmost point of the rover exterior. Thiswill allow the greatest amount of light to reach the panels after deployment. Two options were consideredfor solar panel deployment: (1) Panels would be spring loaded and mounted to a hinge. A solenoid wouldunlatch the sprung panel and allow it to open. (2) Panels would be opened with a servo to control theirdeployment rate. The leading design is to deploy the panels by servo due to the increased control this optionoffers.

5.1.9 Object Detection

For the rover payload to travel the required 5 ft. from the landing site, it may be necessary for the roverto be able to detect and avoid various objects that could impede its motion. A number of methods wereresearched when deciding how the team would approach this problem.

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The first option considered was the use of a camera to detect obstacles and determine a route around them.The main advantage of camera vision is its ability to identify color and contrast in its field of view. Throughmore advanced processing it’s also possible to identify specific objects and track their location in the image.Camera sensors also have a much less restricted range and could be used to identify objects at a distance.The drawbacks of camera vision though are the relatively high cost of the sensors, which are generally alsolarge and require more power than other sensor options. Cameras are also ineffective at very short distancesdepending on the angle of view. The team also has very limited experience with image processing, makingthis one of the more difficult object detection options we considered.

LiDar was another option the team considered due to its use on autonomous vehicles. Some benefits ofLiDar technology is its ability to work in any light conditions, whether it be daylight, cloudy, or dark out.There are also a variety of ranges available for LiDar sensors, however most are far too large to fit in thelaunch vehicle, let alone on the rover payload. Like camera sensors, LiDar tends to be a more expensiveand high power option for object detection. Unlike camera sensors though, LiDar does not detect color orcontrast, just distance. LiDar sensors are also subject to scattering due to particulate in the air, such asdust or fog. This is a distinct disadvantage in what we expect to be a dusty landing site in the middle of a field.

Due to the relatively large size of camera and LiDar sensors and the limited size of the rover payload, a num-ber of smaller sensor options were considered for object detection, namely: infrared and ultrasonic sensors.Both options offer distance measurement and object detection capabilities in small, inexpensive, lightweight,and low power options; however, there are definite drawbacks to either choice. Infrared sensors are generallyvery sensitive to light and weather variations. They are also subject to scattering if there is particulate inthe air. Ultrasonic on the other hand, performs fine in any light condition and is insensitive to particulate atshorter distances. However, the ultrasonic signal sent out by the sensor could be scattered by the numeroussmall objects in the environment, such as grass, corn stalks, or plant foliage.

When weighing the various options, the main considerations used for determining the best sensor choice forobject detection were the size and power consumption of the sensor, as well as the difficulty of implementingthe sensor in the environment. Due to the expected landing site being an open field, a large amount of dustis expected near ground level due to the wind and landing forces. Both infrared and LiDar are subject toscattering in dusty conditions, making them impractical solutions to the problem. Camera vision, thoughit is difficult to code, large, and power hungry in comparison to the other options, provides a wealth ofinformation about the environment and would be the ideal solution provided the image processing is donecorrectly. The drawbacks of camera vision though, are that it doesn’t work at short range and doesn’tactually give the distance to the object it is detecting.

With this in mind, the final alternative the team considered was the pairing of a camera sensor with a smallarray of ultrasonic sensors. The camera sensor would, at the very least, be capable of giving color andcontrast data to the control algorithm, while the ultrasonic sensors could measure the distance to nearbyobjects. By blending these datasets it would be possible to navigate a wide variety of environments. Due tothe unpredictability of the landing site, the team has chosen this option as our leading design.

Table 48 summarizes the pros and cons for the object detection alternatives the team considered.

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Table 48: Object Detection Pros and Cons


Camera Vision

• Could be used to identify andnavigate around obstructions

• Can detect color and contrast

• Long range functionality

• Limited experience with visionprocessing makes this optiondifficult to implement

• Requires high processingspeeds

• High power consumption

• Medium to large sensor size

• Relatively high cost

• Ineffective at short range

• Doesn’t give actual distance toobject

LiDar Rangefinding

• Works well in all light condi-tions

• Variety of ranges available

• Difficult to implement

• Subject to scattering due toparticulate (dust, fog, etc.)

• Large sensor size


• High cost

• Cannot detect color or con-trast

Infrared Distance Measure-ment • Small sensor size

• Inexpensive options readilyavailable

• Low power consumption

• Sensitive to variations in lightand weather conditions

• Short range

• Subject to scattering due toparticulate (dust, fog, etc.)

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Ultrasonic Distance Measure-ment • Works well in all light condi-

tions

• Insensitive to particulates atshorter ranges

• Small sensor size

• Inexpensive options readilyavailable

• Low power consumption

• Short range

• Cannot detect color or con-trast

• Subject to scattering from nu-merous objects (e.g. foliage,grass, stalks, etc.)

Camera Vision coupled withUltrasonic Distance Measure-ment

• Camera could supplycolor/contrast data fornavigation

• Ultrasonic rangefinding forsensing large solid objects

• Complicates design

• Difficult to implement


• Relatively high cost

5.1.10 Rover Distance Measurement

In order to better determine the distance travelled by the rover payload, and therefore minimize the amountof unnecessary travel distance, various methods of distance measurement were considered.

The first option considered was GPS due to its relatively low cost and ease of implementation, however,most low cost GPS receivers only give down to 5 meter accuracy. Even mid cost receivers using Wide AreaAugmentation System (WAAS) only get down to 3 meter accuracy. This is well beyond the 5 ft. distancethe rover is required to travel and could introduce significant error in determining the position of the roverrelative to its starting point. There are a few options available to improve the accuracy of this particularsystem. Commercially available receivers that give sub-meter accuracy, or the addition of a ground stationto help correct the GPS coordinates. The former was found to be extremely expensive while the latterwould drastically increase the complexity of the distance measurement system. Adding a ground station forGPS correction would also not be possible in an off-planet mission, like the USLI mission is supposed to besimulating.

The second option the team considered was inertial navigation, a process that uses accelerometers, gyro-scopes and magnetometers to continuously calculate a position, course, and velocity using dead reckoning.The simplest way to implement this type of system requires only a single IMU sensor be added to the roverelectronics, making this a lightweight option for distance measurement. However, this type of distance mea-surement is subject to significant integration drift, where small errors in the process get carried throughto the subsequent iterations. This is due to the double integration of the acceleration data in order to getposition data. Although the distance the rover needs to travel isn’t very far, the fear is that, while navigatingaround objects, the accumulated error will add up and give an inaccurate estimate of how far the rover hasactually travelled from the landing site.

When determining how well an inertial sensor would perform in an inertial navigation system, a first orderestimate is to assume that the errors in position will be caused only by the errors in the accelerometer bias.

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Errors in the position estimate will actually be caused by many factors including the gyro bias stability,accelerometer scale factor uncertainty, and many other parameters, however the largest contribution will bedue to uncorrected errors in the accelerometer bias. Any bias in the accelerometer horizontal axes will causea bias in the calculated orientation of the object in terms of pitch and roll. This is due to the amount ofacceleration of an object typically being much smaller than the acceleration due to gravity. The error inposition due to a bias in the accelerometer will have quadratic growth with time.

An estimation of horizontal position error can be found for a static case by simply using the accelerometerbias error given in the manufacturer’s datasheet. In the case of the Yost 3-Space Sensor, a high quality,medium cost miniature Attitude and Heading Reference System (AHRS) that is one of the team’s top picksfor the IMU, the accelerometer bias error is not explicitly given, but can be calculated using the voltagelevel of the chip and G range that would be measured. In the ±2G mode, the accelerometer bias error canbe calculated like so:

Accelerometer Bias Error (g) =V dd(V ) · Tolerance(%)

Sensitivity (V/g)(6)

From the datasheet: Vdd = 3.3V. This voltage is divided to allow for the sensing of positive and negativeaccelerations. In other words, 0G is found at 1.65V, and anything above that is a positive acceleration whileanything below 1.65V is a negative acceleration. Therefore, in ±2G mode, the sensitivity is:

Sensitivity(V/g) = 1.65(V )2(g) = 0.825(V/g)

A tolerance for Vdd is not given in the datasheet, however, if we assume a tolerance of 1% we find:

AccelerometerBiasError(g) = 3.3(V )·1(%)0.825(V/g) = 0.04g

To calculate the horizontal position error with the estimated accelerometer bias error we use the equation:

HorizontalPositionError(m) = 0.5 · 9.8(m/s2) ·BiasError · t2

Therefore, after 10 seconds the position error could be as much as:

HorizontalPositionError(m) = 0.5 · 9.8(m/s2) · 0.04 · 102 = 19.62m

If we assume a tighter tolerance of 0.01% we find the horizontal position error after 10 seconds is 1.96 m.

Another option for estimating the horizontal position error uses the error due to axis misalignment.

HorizontalPositionError(m) = 0.5 · 9.8(m/s2) · sin(Misalignment(◦)) · t2

If the chip is not properly calibrated and the measurements are not taken in a perfectly orthogonal coordinateplane, a significant amount of error can be introduced into the calculation. The data sheet states that theorientation accuracy of the chip is ±1° for dynamic conditions and all orientations. This is an averaged valuefound after factory calibration. Using this value in the above equation we get a horizontal position errorafter 10 seconds of:

HorizontalPositionError(m) = 0.5 · 9.8(m/s2) · sin(1◦) · 102 = 8.55m

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After 1 seconds however, the horizontal position accuracy (using the above equation) is a mere 0.0855 m, orabout 3 in.. These values seem to validate the tolerance assumptions from the earlier error calculation basedonly on the accelerometer bias error. The error is much smaller over shorter time periods due to the errorcalculation relying on the square of the elapsed time. Therefore, for short periods of time, this could stillbe a viable option for position determination. Furthermore, coupling this system with other sensors couldpotentially offset the error and still give an accurate measurement of the rover’s displacement.

One option for mitigating the integration error of the dead reckoning method is by adding another signalfor verification. Wheel encoders are simple and lightweight sensors that could be used to correct for theintegration drift. However, encoders are subject to wheel slippage and complicate the design of the chassis.The tank tread design of the rover relies on slip steering for turning and would introduce slippage whenclearing small or loose obstacles. This could introduce significant error in the distance estimation. One wayto mitigate this issue is by using clever coding. Slippage could be identified and rejected by comparing theaccelerations seen by the IMU to the encoder count. Also, encoder outputs could be ignored during turningassuming the rover is capable of a zero degree turn.

Another way to mitigate the integration error of the dead reckoning process would be to compare the deadreckoned distance to GPS data. This method is commonly found in high precision inertial navigation systemsand uses the GPS position to correct for the integration error. However, with the small distance the rover isrequired to travel, the low resolution GPS data may not accurately update the calculation and could poten-tially introduce more error. Significant testing would be needed to determine the efficacy of this techniquefor the rover design.

Another consideration for mitigating the integration error would be the addition of a Kalman filter to thedead reckoning system. Specifically, the filter would be used to blend the IMU and GPS derived positiondata and make a “best guess” prediction based on the two datasets. With appropriate tuning, a Kalmanfilter could effectively mitigate the inherent integration error of the dead reckoning method without the needfor additional sensor inputs.

The final alternative the team is considering for measuring the rover’s horizontal displacement is monocularvisual odometry. This method uses video data and feature detection and tracking algorithms to estimatevelocity and heading. This estimation can then be used to calculate the rover’s position in relation to itsstarting point. There are many concerns with using this method, such as: the potential lack of trackable fea-tures the rover will encounter in an open field, the processing capabilities of the image processing computer,and the minimal experience the team has with image processing. However, implementing this alternativewould not require adding more hardware to the rover system as the object detection camera could be usedfor navigating obstacles and estimating position.

With these things in mind, the team has decided on an inertial navigation system that will blend IMU andGPS data with a Kalman filter in order to better estimate the distance the rover has travelled.

Table 49 summarizes the pros and cons of the various methods the team considered for measuring thedistance the rover travels from the landing site.

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Table 49: Distance Measurement Methods Pros and Cons


GPS

• Simple to implement

• Several low cost options avail-able

• Introduces sync delay

• Mid cost (∼$60, WAAS) op-tions only give position accu-racy of 2-3 meters

• Sub-meter accuracy found tobe cost prohibitive

• Antenna must be integratedinto rover design

Inertial Navigation - IMUonly • Gives sub-meter accuracy

• No sync delay

• Significant integration drift(error ∼ t2)

Inertial Navigation - IMUwith wheel encoders for cor-rection

• Gives better indication of dis-tance travelled by using multi-ple data sources

• Gives sub-meter accuracy

• No sync delay

• Encoders are subject to slip-page, especially during turn-ing

• Increases code complexity

• Complicates mechanical de-sign

Inertial Navigation withKalman Filter - IMU withGPS for correction

• Gives better indication of dis-tance travelled by using multi-ple data sources


• No sync delay

• Introduces sync delay

• Drastically increases codecomplexity

• Complicates design and imple-mentation

Monocular Visual Odometry


• No sync delay

• No additional sensors

• More accurately simulatesdeep space mission

• Increases code complexity

• Potentially difficult to trackfeatures in natural environ-ment

• Increased processing capacityneeded from image processingcomputer

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Table 50: Microcontroller Pros and Cons


Raspberri PiZero

(Version 1.3)

• Low cost

• Small footprint

• Light weight

• RAM

• High processing speed

• Requires an SD card

• Not coded in C-based Arduino

• Limited pinout

Teensy 3.5

• Arduino compatible

• Small footprint

• Fast

• Many I/O pins

• Less RAM than Raspberri Pi Zero

• Slower processor

ArduinoZero

• Easy to code in

• Well documented

• Limited code size/ complexity

• Limited pinout

Table 51: Servo Motor Pros and Cons


HSR-2645CRH

• High torque output

• Already continuous motion

• Self-correcting

• High cost

• Large package size

HS-1425CR• Acceptable torque output

• Already continuous motion

• Lower torque for same package sizeas HSR-2645CRH

TowerPro-SG92R

• Low cost

• Light weight

• Small package size

• Low torque output

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5.1.11 Rover Leading Design

Overview

The leading rover design will use a tank tread system driven by two high torque continuous rotation servosto negotiate the terrain. The treads, sprockets, and body will be made from 3D printed ABS plastic. Thechassis, pillow blocks and servo hubs will be made from 6061 aluminum. The axles for the idler sprocketswill be made from stainless steel. Overall rover dimensions are shown in Figures 30 and 31.

In order for the rover to navigate the terrain, a camera and three ultrasonic sensors will be used for objectdetection and avoidance. This system will be coupled with a distance tracking system that will help tominimize the amount of time the rover will drive. This system will use a precision IMU and high accuracyGPS unit to track the rover’s motion. The solar panels will be mounted at the top-back of the rover bodyand will deployed using a small servo once the 5 ft. goal has been reached.

Figure 30: Current Rover Design Dimensions (all dimensions in inches)

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Figure 31: Current Rover Design Dimensions (all dimensions in inches)

Rover Mechanical Design

The continuous track drive system will use two HSR-2645CRH continuous rotation servos, mounted to thebottom-front of the chassis, to drive the rover. Turning will be done by differentially driving the tracks; onetread will rotate forward while the other rotates backward. Ideally, this will allow the rover to have a zerodegree turning radius in order to better navigate potential obstacles. The servos will be mounted under thechassis using custom machined aluminum mounts that will be bolted into the structure. The assembly ofthis is shown in Figure 32. Drive sprockets will be mounted to aluminum hubs that have splines matched tothe servo output shaft. These are shown in yellow in the image.

Figure 32: Rover Drive Servo Mounting

Modular tread links will be used to construct the track system. They will be linked using steel pins that willpress fit in place, as shown in Figure 33. Note, the interference fit is only on the outer portion of the treadlinkage. The inner portion is free to rotate to allow the treads to bend and contour to the terrain.

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Figure 33: Tread Link Assembly

The rover will utilize a total of three sprockets per side that will interface with the track in a way that issimilar to a chain drive system. The teeth on the sprockets will be matched to the profile of the tread linksto minimize binding and mitigate the chance of the tread disengaging. The layout of the sprockets can beseen in Figure 34

Figure 34: Rover Sprocket Configuration

The rover chassis will be machined from a single plate of 6061 aluminum and will be the mounting point forall rover subsystems. Mounting the drive system design to the chassis will be accomplished with machinedslots that will allow adjustment of individual components. This will allow the team to fine tune the alignmentof the drive system and adjust the rover width. Each idler sprocket will be supported by an individual axleusing two pillow block bearings per axle. The mounting holes and slots can be seen in Figure 35.

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Figure 35: Rover Chassis Mounting Points

The solar panels will be mounted on the top-back of the rover body and will be unfolded using a servo andhinge. A slot for the bottom panel will be printed as part of the rover body design, while the upper panelwill be in a separate retainer that will be hinged and attached to the servo horn. This system can be seenin Figure 36. The rover body is shown in Figure 37.

Figure 36: Solar Panel Unfolded

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Figure 37: Rover Body

Rover Electrical Design

In order for the rover to autonomously navigate the landing site terrain, travel 5 ft., then deploy its solarpanel array, it must be able to (1) identify and navigate around obstacles in its path, and (2) track thedistance that it has travelled from the payload airframe. Figure 38 shows a high level diagram of how therover navigation system will function.

To identify obstacles, the team will employ a camera sensor that will take static color images and an arrayof three ultrasonic sensors. The images will be processed on a dedicated image processing computer in orderto identify the location of objects that could impede the rover’s travel. The ultrasonic sensors will be usedas a close proximity backup for the imaging technique. Both sets of data will be passed to the centralmicrocontroller, which will then decide how to proceed.

To track the distance the rover has travelled, a high precision IMU and GPS will be used along with a deadreckoning algorithm to estimate the horizontal distance from the rover’s starting point. Once the 5 ft. markhas been passed with a high degree of certainty, the rover will cease movement and deploy its solar panel array.

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Figure 38: Rover Electrical Summary

Object Detection System

The object detection system was designed to give robust object detection capabilities in a highly variableenvironment. This was necessitated by the unknown landing position of the payload section after landing.Upon exiting the housing, the rover will need to determine the location of potential obstacles and hazards,then navigate around them to reach the 5 ft. goal. What follows is the current leading design the team willpursue for object detection.

Hardware

The primary device that will be used for object detection will be a Raspberry Pi Camera Module v2.1 thatwill interface directly with the CSI port on a Raspberry Pi Zero v1.3. The camera features an 8 MP nativeresolution sensor capable of taking 3280 x 2464 static images as well as video. It has a vertical field of viewof 40° and a horizontal field of view of 53°, which should easily capture the area in front of the rover. It willbe mounted on the top of the rover body, facing straight forward, with a clear line of sight unobstructed byother sensors and parts of the rover.

A Raspberry Pi Zero v1.3 will be used to perform all image processing functions and will be running OpenCVv3.0 on the Raspbian Jessie operating system. The Raspberry Pi will communicate with the main micro-controller, a Teensy 3.5, over UART. This connection will be used to control when images are taken, and tocommunicate obstacle locations to the controller.

In addition to the camera module, three XL-MaxSonar-EZ2 ultrasonic rangefinder sensors will be placed atthe front of the rover chassis to detect nearby objects. One sensor will be placed at the very center of thechassis, and the other two will be on the front corners facing out at a 35° to 45° angle. The ultrasonic sensorswill interface to the Teensy 3.5 microcontroller using a single GPIO pin each, and will provide close rangeobject detection capabilities for the rover to navigate with.

The final component in the object detection system will be the Teensy 3.5 microcontroller that will act asthe central control unit for the overall rover control system. A navigation program will dictate how the rovershould move based on whether an obstacle is detected in the rover’s path and the distance the rover hasalready travelled.

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Table 52: Rover Battery Budget Estimation

Battery Budget EstimationSystem Component Max Current Draw (mA) Amount Run Time (hr) Power Density (mAh)

Servos 650 2 1 1300Teensy 3.5 300 1 1 300GPS 100 1 1 100Antenna 14 1 1 14IMU 45 1 1 45Pi Zero v1.3 150 1 1 150Ultrasonic Sensors 3.4 3 1 10.2Camera 250 1 1 250

Rover

Solenoid 100 1 0.00278 0.277777778Total: 2169.47778

Software

Image processing will be handled by the Raspberry Pi Zero using OpenCV. OpenCV is an open sourcecomputer vision library that includes native functions with a variety of image processing capabilities. Uponreceiving an image from the camera, the image will be filtered using RGB mean shifting. This processsegments the image into areas that share a similar color and can be run for any number of iterations in orderto further delineate the image.

From there, an attempt will be made to characterize the color segments and identify them as obstaclesor passable terrain. Specific color ranges will searched for in the image in order to label individual colorsegments with a best guess of what those segments represent. This can be further refined using contextualinformation. For instance, a blue or white segment near the top of the image would most likely be the sky,and would not be labeled as an obstacle. However, if a gray, black, or white segment is found below theskyline, then there’s a high probability that this is a rock or large mound that the rover will need to navigatearound. This segment will then be labeled as an obstacle.

The final step will be to relay the position of the obstacle to the microcontroller. This will be done by sec-tioning the image into a grid and passing only the specific grid location of the object to the controller. Thecontroller will be programmed to adjust the heading of the rover based on the location of potential obstacles.For example, if an obstacle is on the left of an image, then the rover will steer right. Additional strategieswill be programmed and tested to ensure the decision making process is robust to both false positives andnegatives.

The final step in the object detection algorithm will use the ultrasonic sensor array described above. Thesensors were chosen to give accurate and noise-immune measurements to objects directly in the path of therover. With a minimum range of 20 cm (7.87 in.), the ultrasonic sensors will act as wireless bumpers for therover and allow it to steer clear of large objects that could impede its motion. The logic for these sensorswill be similarly simple. As an example, if a sensor reads an object on the right, the rover will steer leftuntil it no longer reads the object. Other strategies will be programmed and tested to ensure the navigationsystem can traverse the required 5 ft.

Inertial Navigation System

The inertial navigation system was designed to give an accurate estimation of the distance the rover hastravelled from its deployment location. The main objective of this system is to minimize the time anddistance the rover must travel to reach the 5 ft. goal with a high degree of certainty. This will in turnminimize the number of obstacles the rover needs to traverse, and will raise the chance of success of themission. What follows is the current leading design the team will pursue for the inertial navigation system.

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Hardware

The primary device that will be used for tracking the position of the rover will be the Yost Labs 3-SpaceSensor. The IMU features a small, lightweight, and robust package design with on chip Kalman filteringand correction of the individual sensor signals for improved accuracy, and a variety of data output formats.This will interface to the Teensy 3.5 over UART and will supply tri-axial gyroscope and acceleration datafor the Teensy to process. The IMU will be securely mounted as close to the center of the rover as possibleto ensure the readings are indicative of the rover’s movements.

The data gathered by the IMU, when processed, will accumulate a large amount of error very quickly dueto integration drift. This drift will be mitigated by comparing the calculated displacement to high accuracyGPS data. The GPS chip the team has selected is the Garmin GPS 15L due to its small package size, simpleinterface, and Wide Area Augmentation System (WAAS) capability which gives the chip an accuracy of lessthan 3 meters. The 15L is capable of position acquisition in under 5 minutes with no data known upon startup. This can be reduced to 45 seconds if a cold start is used and the initial position, time, and almanac areknown. The GPS is also compatible with a variety of active and passive antennas and uses a MCX connector,though adapters are available for antennas that terminate in a BNC male connector. It will interface withthe Teensy 3.5 through a RS-232 to UART adapter.

Both the IMU and GPS data will be passed to the Teensy 3.5 for processing and filtering. The data willbe used to determine the estimated distance the rover has travelled from the landing site and to guide therover in its navigation strategy.

Software

Due to the large amount of expected integration drift, a Kalman filter will be implemented on the controllerto blend the two data sets and make a best guess prediction of the actual displacement of the rover. Thereare a variety of ways to do this, but the current implementation scheme is to first parse the GPS data,then calculate the horizontal displacement since the last received fix. The IMU data will be integrated togive a position vector in the x and y directions and then the displacement from the last fix will also becalculated. These values will be averaged together, then run through a Kalman filter algorithm to accountfor the expected integration drift. If need be, the IMU data can be weighted differently before the filteringtakes place. The need for this will be determined through testing.

The Kalman filter will output a best guess of what the actual displacement of the rover was and that valuewill become the last known fix for the next iteration of the algorithm. The rover will be programmed toexceed the 5 ft. target by 1 to 2 ft., while the Kalman filter will be tuned to ensure a conservative estimateis made. The algorithm will be further tuned through testing.

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Figure 39: Distance Measurement Data Flow

5.1.12 Preliminary Interface Description

The rover will rest on a rapid prototyped platform that will serve as a linear guideway and orient the roverin the direction of deployment relative to the horizontal plane. The rover will be mounted with a flanged nutthat threads on a lead screw actuator which passes through the center of the rover housing. The flanged nutwill be mounted on an L-bracket which is bolted onto the rover chassis.The actuator acts as the deploymentmodule as well as the retention system for the rover. A cap will be threaded onto the leadscrew and setwithin the housing to protect the rover during descent. The cap will be forced off when the rover exits thepayload airframe.

5.1.13 Technical Challenges

The primary challenge the 49er Rocketry Team faces with the rover payload is successful terrain negotiation.The drive system and electronics have been designed to be robust and versatile. Table 53 lists the anticipatedtechnical challenges associated with the rover.

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Table 53: Rover Technical Challenges


RTC 1Verify structural integrity of all rovercomponents.

Simulation and material testing will beconducted.

RTC 2Rover binding in the housing before fulldeployment.

Leadscrew thread pitch and motorwill be selected to provide appropriatethrust force and verified through test-ing.

RTC 3Rover will be able to survive the rigorsof the flight.

Extensive material analysis will be per-formed as component dimensions are fi-nalized.

RTC 4Proper dimensioning of rover and pay-load components.

Proper tolerances will be carefully se-lected during design and manufactur-ing to ensure proper integration of rapidprototyped parts into the assembly.

RTC 5Inadequate strength due to layer orien-tation during rapid prototyping.

Proper orientation of the part duringrapid prototyping to ensure robust con-struction.

RTC 6Securing of the rover vehicle duringlaunch and landing.

A bronze leadnut will be fixed to theunderside of the rover chassis.

RTC 7 Object detection and avoidance.Image processing and object detectionalgorithms will be tested in various ter-rain conditions.

RTC 8Ultrasonic sensor height on rover bodyto obtain usable signals.

Appropriate height for object detectionwill be verified through testing.

RTC 9Camera image processing and colorrecognition.

Camera vision system and algorithmwill be extensively tested in varyinglight conditions.

RTC 10 Rover traction and terrain navigation.A test bed for soil and obstacle navi-gation will be built for controlled rovertesting.

RTC 11 Rover distance tracking.

GPS and IMU signals will be blendedand passed through a Kalman fil-ter, and accuracy of distance trackingtested.

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5.2 Selection, Design, and Rationale of Housing Payload

The rover housing must accomplish three tasks: (1) protect the rover through the entire launch sequence,(2) upon landing orient the rover to a within ±5° of a normal orientation, and (3) receive a wireless signal toinitialize the orientation and deployment of the rover. A number of design alternatives were considered foreach major section of the housing payload. Pros and cons of each alternative are identified and tabulated inthe following sections.

5.2.1 Housing Design

To ensure successful rover deployment, the rover housing must be able to properly orient the rover to thelanding terrain. The leading design for the rover payload requires the rover to be upright in order to drive;however, due to the unpredictable nature of an uncontrolled parachute descent, it is possible that the roverwill be upside down or otherwise incorrectly positioned upon landing. With this in mind, the team consid-ered several housing designs that could mitigate this condition.

The first housing alternative the team considered was a controlled landing module that would attempt toland vertically to ensure the proper orientation of the rover. If properly designed, this would allow the roverto drive out of the airframe after touching down with no post-landing modifications. With the expectedlanding site consisting of unknown topography, the attempt of a controlled vertical landing greatly increasedthe housing design complexity. Repeatability of such a system is low; therefore, the design was abandonedearly in the decision making process.

The next alternative considered was to eject the housing containing the rover from the airframe using com-pressed gas or black powder charges. Such a system would ensure the rover exited the airframe even if thedesignated exit was blocked. However, the orientation of the rover after ejection would be highly variablewithout an additional external system or ballast to weight the housing to the correct position. The team’sprimary concerns with a system using compressed gas or black powder charges are spectator safety anddamage to the airframe or rover.

The final alternative the team considered was a housing that would rotate in the interior of the airframe toorient the rover prior to deployment. This design would rotate a fiberglass tube on bushings fixed to theairframe interior. Rotation would be driven by a 360° position controlled servo using an accelerometer andhill-climbing algorithm. Correct orientation in this design is defined as the z-vector pointing upward. Thisgeneral design concept is the leading alternative being pursued by the team.

Table 54 summarizes the various pros and cons of the housing design alternatives the team considered.

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Table 54: Rover Housing Design Pros and Cons


Rotating Housing

• Gives greatest control of roverorientation upon landing

• Increases design complexity

• Possibility housing could jamafter separation, parachute de-ployment, and landing

Ejectable Housing

• Best chance of getting com-bined rover and housing clearof the airframe

• Black powder charges cause asafety risk

• Possibility of damaging rover,airframe, housing, electronicswithin the housing

Landing Module

• Controlled landing allows forthe rover to exit in an uprightposition

• Legs absorb some of the im-pact, causing less shock forceto the rover

• Difficult to design, code, man-ufacture, and assemble

• High probability of failure dueto unknown landing terrainwhich would lead to an un-known orientation of the roverfor deployment

The decision matrix used by the team to decide the leading design for the rover housing is shown in Table 55.The criteria considered for drive system selection included ease of design, ease of assembly, rover traction,robustness, timeline, cost, and rover weight.

Correct rover orientation, ease of assembly, and robustness were given the highest weighting, showing thatthe team considers these the most important criterion, compared to 4 points for ease of design and cost, and3 points for timeline and weight. Each rating is multiplied by the multiplier for that criterion and the scoresadded across each row. The design with the highest total score was selected as the leading design option.

Table 55: Housing Design Decision Matrix

Criteria: RoverOrienta-tion

Ease ofAssem-bly

Robust-ness

Ease ofDesign

Cost Timeline Weight TotalScore


RotatingHousing

4 3 3 3 3 4 3 95

EjectableHousing

1 4 2 3 4 4 3 84

ControlledLanding

4 2 2 1 2 2 2 64

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5.2.2 Material Selection

Material selection for the rover housing was based on the need for high strength, easy customizability, andthe ability to pass electromagnetic signals. The primary options included: (1) carbon fiber, (2) fiberglass,and (3) acrylic.

Carbon fiber offers high strength, low weight, and easy customizability. Based on the need to transmit awireless signal for rover deployment, carbon fiber’s signal attenuation characteristics would require an ex-ternal antenna for signal transmission. To realize the proposed rotating housing design a custom diametertube is likely necessary, so carbon fiber has the potential to be cost prohibitive.

Fiber glass and acrylic are lower cost alternatives to carbon fiber, and offer reduced signal attenuation;however, they are heavier. Custom diameters of acrylic are not easily obtained, and it is a more difficultmaterial to work with. Fiber glass can be easily customized, and offers easy post processing to correct anydefects from manufacturing. At present, fiber glass is the leading material choice for housing construction.

Table 54 contains pros and cons for each material alternative considered for the rover housing.

Table 56: Housing Material Pros and Cons


Carbon Fiber

• High Tensile and compressivestrength

• Low weight

• Easy Sanding/Machining

• Most expensive

• Attenuates electromagneticsignals

• Dust from machining is haz-ardous

Fiberglass

• Allows wireless signals to passthrough

• Easy Sanding/Machining

• Low cost

• Heavier than carbon fiber

• Dust from machining is haz-ardous

Acrylic

• Allows wireless signals to passthrough

• Medium cost

• Low impact strength

• Difficult to manufacture

5.2.3 Rover Retention

Rover retention is of critical importance during flight for safety reasons. During the separation event, lossof the rover would pose a hazard to spectators. Three possible solutions were considered including: (1) aleadscrew, (2) cable retention, or (3) a latch mechanism. Tethering of the rover to the housing is beingconsidered for redundancy.

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Maximum load is expected during parachute deployment. With a projected rover weight of 3.8 lb and as-sumed acceleration of 20 g during the recovery event, the retention system must be able to withstand anominal load of at least 76 lb.

A leadscrew would allow deployment of the rover with a large available thrust force. Mounting is simplifiedwith all necessary components attached to the electronics bay. A leadscrew nut would be the only addedcomponent to the rover itself. Installation of the rover into the housing would be accomplished by reversingthe motor, engaging the leadscrew nut, and threading the rover into position. Leadscrew and leadnut selec-tion can be based on static loading conditions. Stress concentrations are neglected. Because of the safetyhazard inherent to retention system failure, a safety factor of at least 4 is desired. Assuming a bronze leadnutis used, it is the most likely failure point. The rated load for a 5/16-12 round bronze leadnut is 490 lbs.which would provide a safety factor of 4.9 for the projected load. Deployment by leadscrew is the primarydesign being pursued.

The leadscrew can be threaded 8 in. or more through the leadnut located on the rover chassis. Givena leadscrew with 12 threads per inch, disengagement of the rover would require the rover or leadscrew torotate over 96 times during flight.

Another potential risk for disengagement exists at the motorshaft-to-leadscrew shaft coupler. In order todecouple the forces of recovery from an interface not intended to prevent axial movement (the shaft coupler),an additional bulkhead can be added to the rover housing. The leadscrew would be coupled to the motorthen pass through a threaded collar and shouldered bearing retained in the bulkhead to prevent axial motion.

Cable retention would add significant design complexity for system realization. Additional electronics wouldbe necessary to achieve secure and safe retention of the rover including but not limited to: a servo, solenoid,and limit switches. Assembly of the mechanism for flight would be difficult in the limited available airframevolume. The majority of braided steel cable available affordably and in a form factor appropriate for thedesign has a rated load of 100-200 lb; this provides a safety factor of 1-2. With the complexities of designingsuch a system and given the potential for injury and property damage associated with rover disengagement,cable retention is an unsuitable solution.

A latching mechanism has been considered to clip the rover to the housing. Latches would be spring loadedto the open position, requiring a pivot and additional attachment points in the housing. A solenoid wouldbe required to hold the latches in the closed position during launch, recover, and landing which would com-plicate rover installation into the housing.

Tethering of the rover to the housing is being considered for redundancy. A shock cord would retain therover in the event of primary retention system failure, or other catastrophic event. Addition of a tether hasthe potential to entangle in the rover drive system, and could result in mission failure. A tether may not beimplemented in the final design; however, the need for this safety feature is being explored.

Table 57 summarizes the pros and cons of the retention systems the team considered.

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Table 57: Rover Retention System Pros and Cons

Retention System Pros Cons

Leadscrew

• Stable and Reliable

• Low Friction

• Lowest risk for disengagement

• Doubles as deployment mech-anism; rover doesn’t drive outunder own power

• Heavy

• Leadscrew passes through therover; complicates design slow

Cable Retention

• Don’t have to pass leadscrewthrough rover

• Medium weight compared toleadscrew

• Drastically adds complexity

• Risk for entanglement withdrive system

• Risk of Disengagement

• Rover would drive out underits own power

Latch Mechanism

• Lightweight

• Secure

• Medium chance of disengage-ment

• Adds Complexity

• Large Volume Requirement

Tether

• Lightweight

• Increases Safety

• Added Redundancy

• Risk for Entanglement

• Not a standalone system(needs another form ofretention)

The decision matrix used by the team to decide how to execute rover retention is shown in Table 57. Thecriteria considered for system selection included ease of design, ease of assembly, robustness, safety, missionsuccess, cost, and weight.

Ease of design, ease of assembly, robustness, safety, mission success have the highest multiplier, showingthat the team considers these the most important criterion, compared to 3 points for cost, and 2 points forweight. Each criteria is scored for each design option, multiplied by the weighting, and the scores addedacross each row. The design with the highest total score was selected as the leading design option.

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Table 58: Retention System Decision Matrix

Criteria: Ease ofDesign

Ease ofAssem-bly

Robust-ness

Safety MissionSuccess

Cost Weight TotalScore


Leadscrew 5 5 5 5 5 3 2 138

Cable Re-tention

2 1 1 2 2 3 4 57

Latch 3 2 3 3 2 3 3 80

Tether 5 4 5 5 2 5 5 130

5.2.4 Rover Deployment

Primary design considerations for the rover deployment system include: (1) successful rover deployment aftersending the initiation signal, (2) thrust force which allows deployment under adverse circumstances (block-age, high friction, damage from landing, and with enough force to lift the entire housing), and (3) robustness.

The team first considered deployment using the rover’s drive system for thrust force.

Another design alternative would be to use black powder to eject the rover and housing from the airframe.Due to the uncertainty in rover orientation, potential for rover damage, and inherent safety issues associatedwith pyrotechnics, this system will not be pursued.

Based on the leading retention system option, the leadscrew can serve the dual purpose of rover deployment.

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Table 59: Deployment Method Pros and Cons

Deployment Method Pros Cons

Under its own power

• Extra materials, cost andweight not needed

• Could bind within the housing

• Not enough torque to getthrough obstacles at the en-trance (corn stalk, rocks, highgrass), or lift the airframe offthe rover

Leadscrew

• High thrust

• Allows the airframe to bepushed off the rover

• Double duty as retention sys-tem

• Higher weight than rover pow-ered deployment

Black Powder

• Dependable ejection of rover • Could cause damage to roverand/or airframe

• No position control

• Poses a safety hazard

5.2.5 Deployment Direction

Based on the location of the rover housing in the airframe, the deployment system can be designed for therover to come out towards the nose cone or towards the separation side of the payload section. Major con-siderations for deployment direction include ingress of black powder charge gases into the payload section,rover entanglement in parachute or shock cords, and vehicle stability.

If the rover deploys on the nose cone end of the payload section, the necessary hardware for recovery isreduced. A single altimeter bay could be housed in the transition section of the current leading design.The main booster parachute would be housed below the transition with a set of black powder charges; thepayload drogue and main parachutes would be housed above the transition with an additional set of separa-tion charges. All altimeters could be housed within the transition section. The primary drawback with thishousing orientation is the need to separate the nose cone after landing. Separation of the nosecone wouldneed to be accomplished by the rover deployment system; the nose cone would have the potential to impedesuccessful rover navigation.

Deploying the rover on the separation side of the payload section would place the payload recovery hardwareabove the rover housing. To avoid interference with the components necessary for the leading housing design,this configuration requires a separate altimeter bay from the main booster parachute. The rover itself wouldbe located further down in the airframe, thus lowering CG as a drawback. Additionally, the main boosterparachute could potentially cause damage to the housing or rover during separation. The benefit to thisconfiguration is that there would be less chance of vehicle interference with the rover after landing.

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Table 60: Deployment Direction Pros and Cons

Option Pros Cons

Deploy Rover on Nose ConeSide • Allows for all other sections of

rocket to be cleared

• Fewer necessary recovery sys-tem components

• Raises CG

• Would need the nose cone toopen opposite of the parachutecord

• Nosecone impedes rover de-ployment

Deploy Rover on SeparationSide • Already an open section for

the rover to come out• Lowers CG

• Separation could poten-tially cause damage tohousing/rover

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5.2.6 Rover Housing Leading Design

Overview

The leading rover housing design will utilize a rotating fiberglass tube that will sit inside the payload air-frame. The housing will contain all necessary electronics and rover retention hardware such that thesecomponents will rotate with the housing. Retention of the rover will use a leadscrew system that will drive aleadnut that is mounted to the rover chassis. The overall dimensions of the housing can be found in Figure 40.

The housing will rotate in order to orient the rover in an upright position before deployment. It will dothis using a single accelerometer and a hill finding algorithm that will maximize the z vector acceleration inthe downward direction. This process will be initiated by a wireless signal that will be transmitted by theground team.

Figure 40: Rover Housing Dimensions (all dimensions in inches)

Housing Mechanical Design

The housing will be constructed out of 1/16 in. fiberglass that will be rolled to a custom inner and outerdiameter. It will rotate on two 1 in. wide, 1/8 in. thick 6061 aluminum bushings that will be epoxied tothe inside of the airframe. The bushings will be dressed with a thin coat of graphite lubricant to minimizefriction. A shoulder will be cut in the housing that will rest on the upper bushing. This will act to retainthe housing from falling through the airframe during flight.

The rover retention/deployment system will consist of a 5/16 in. leadscrew that will be coupled to a DCmotor which will sit just behind the rover bay. The leadscrew will be retained in the housing using a threadedcollar that rests against a flanged bronze bushing. The bushing will sit in the bulkhead and keep the lead-screw from falling through during flight in case of motor coupler failure. The rover will be retained usinga 5/16 in. flanged bronze leadnut that will be mounted to the rover chassis. The housing portion of thissystem is shown in Figure 41.

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Figure 41: Leadscrew Retention System Overview

The rotation of the housing will be driven by a high torque 360° position controlled servo. The servo hornwill be mounted to the altimeter bay just above the housing with an aluminum hub. The case of the servowill be mounted to a custom electronics sled, which in turn will be fastened to the removable bulkhead. Thebulkhead will then be fastened to the walls of the fiberglass housing. These connection points will translatethe rotational force to the housing assembly. This interface is shown in Figure 42. The payload sectionassembly is shown in Figure 43.

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Figure 42: Electronics Sled and Altimeter Bay with Rotating Housing.

Figure 43: Payload Section Assembly

Housing Electrical Design

The purpose of the rover housing is to (1) retain the rover through flight and landing, (2) orient the roverupon receiving a wireless signal, and (3) ensure the rover is deployed in an upright position. Figure 44 showsa high level diagram of how the rover housing system will function after landing.

A long range XBee will be used to receive the deployment signal from the ground team. Upon receivingthe signal, the housing will begin the rotation procedure. To orient the rover, a hill climbing algorithmwill be used to maximize the z vector in the downward. Once the accelerometer reads that the housingis upright and stationary, the rover deployment will commence. If, upon deploying the rover, the housing

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shifts or rolls, the rover deployment will be stopped and the housing will rotate back to an upright orientation.

Figure 44: Rover Housing Electrical Summary

Hardware

An XBee PRO 900HP will be used to receive the deployment signal from the ground team. It has a nominalrange of up to 6 miles in ideal conditions and has been successfully flight tested numerous times. The XBeewill interface with a Teensy 3.5 microcontroller using UART and will signal the controller to initiate thedeployment procedure.

A 360° position controlled servo will be used to rotate the housing while the Teensy iteratively comparesthe z axis acceleration from the ADXL345 accelerometer. The accelerometer will be mounted in the housingelectronics bay in the same orientation as the rover. A hill-finding algorithm will be used to guide the housingto an upright position by maximizing the z vector in the downward direction. Once this is done and theaccelerometer registers that the housing is upright and stationary, the Teensy will activate the deploymentmotor, which will be a small, but robust, DC motor interfaced to the Teensy through an H-bridge circuit.

Table 61: Rover Housing Battery Budget Estimation

Battery Budget EstimationSystem Component Max Current Draw (mA) Amount Run Time (hr) Power Density (mAh)

Rotation Servo 650 1 0.33333 216.66667Deployment Motor 650 1 0.33330 216.64500Locking Solenoid 100 2 0.33333 66.66667Teensy 3.5 300 1 1 300ADXL345 1 1 1 1Teensy 3.2 250 1 2 500GPS 50 1 2 100

Rover Housing

Xbee 270 2 2 1080Total: 2480.978333

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Software

The hill finding algorithm that will be used to orient the rover will use a simple and fast comparative ap-proach to maximize the z vector acceleration. As the servo starts rotating, if the z axis acceleration decreases,then the controller will reverse the direction of rotation. Once the maximum value is reached, which shouldbe at an upright orientation, the z axis acceleration will once again start to decrease and the logic will forcethe servo to hold the upright position.

Deployment of the rover will use a PWM signal to drive the motor forward. If the accelerometer sensesa significant shift during deployment, the deployment motor will be stopped and the orientation will bechecked again to ensure the rover makes it out of the housing in an upright position.

5.2.7 Preliminary Interface Description

The rover housing will be located on two 1/8 in. aluminum bushings epoxied to the inner diameter of theairframe. The housing is free to rotate within the bushings and has a shoulder at the top. The minordiameter is dimensioned to allow rotation within the inner diameter of the aluminum bushings, acting asthe bearing journal, while the major diameter rests on the lip created by the aluminum bushings. Ejectionof the housing from the open end of the airframe is prevented by the shoulder. The payload section of theairframe will be attached to the transition section with a coupler and nylon shear pins.

5.2.8 Technical Challenges

The primary challenge the 49er Rocketry Team faces with the rover housing payload is successful roverorientation. Precision manufacturing will be required to successfully implement the rotational housing.Table 62 lists the anticipated technical challenges associated with the rover housing.

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Table 62: Housing Technical Challenges


HTC 1Retention system must be robustenough to secure the rover during therigors of flight.

Ensure structural stability of all reten-tion components through rigorous anal-ysis and testing.

HTC 2Uninhibited rotation of rover housingassembly within the airframe duringorientation sequence.

Ensure bushings are polished, dressedwith graphite lubricant, and allowsmooth rotation. Precision machiningof the bushing to allow accurate align-ment.

HTC 3Validate all wireless communicationwith rover housing electronics.

Range and signal attentuation testingof XBee transmitters.

HTC 4Wiring considerations during housingrotation.

No pass through wires between rotatingrover housing and fixed airframe.

HTC 5Sufficient space for electronics and me-chanical components within electronicsbay and rover housing.

Prototyping and design review of com-ponents to ensure proper fit within theelectronics bay and housing.

HTC 6

Prevent motion of the rover housingalong the longitudinal axis of the rocketbody during launch, recovery, and land-ing.

The rover housing will be manufacturedwith a shoulder to prevent movementlongitudinally and fixed to the electron-ics bay.

HTC 7Manufacture of 1/8 in. aluminum bush-ing.

Ensure slow enough feed rate to allowadequate chip evacuation, preventingpossible distortion of the workpiece.

HTC 8Ensure housing rotation algorithm re-sults in proper orientation.

Test housing rotation without roverpayload independent of launch and ona subscale test flight.

HTC 9Adhering bushings for housing rotationto the airframe.

Bearings will be installed with a jig andhigh strength epoxy.

HTC 10 Housing cap design

Precision machining will allow the capto thread onto the retention systemleadscrew and nylon shear pins will beimplemented for redundancy.

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6 Testing Plan

6.1 Rover

6.1.1 Terrain Negotiation

In order to ensure the rover can complete the mission requirement to travel a distance of 5 ft. from thepayload landing zone, it was decided to verify the rover could negotiate a variety of adverse terrain conditions.The rover’s tank tread design was considered the most robust system because of its versatility in a variety ofterrain conditions. In order test and validate this versatility the team has composed several testing methods.These testing methods include:

• Partial tread soil submersion

• Independent (or isolated) track elevation

• Ascending ramp angles

• Descending ramp angles

• Sideways slopes

6.1.2 Partial Tread Soil Submersion

This test is designed to test the ability of the rover to negotiate through damp or saturated soil. The testswill be conducted with several different soil types in order to ensure the soil texture does not impede on thetread performance. The test will be conducted in a controlled environment in a custom made test bed thatwill allow the team to control the saturation consistency of the soil utilized in each test. Each test will bepartial tread submersion that will not exceed submersion above 50% of the tread height.

This test is designed to not only evaluate whether or not the rover can travel the minimum distance of 5ft. in these adverse terrain conditions but also to assess two additional testing parameters. These testingparameters include the current draw of the motors in these conditions and the depth the treads vertically digwhile traversing through the terrain. These parameters enable the team to quantify and record the rover’sperformance during testing.

Because of the viscous conditions the rover will be maneuvering through, this can add significant strain onthe motors. If the stall torque is approached or reached the rover will slow or stop respectively. In order toquantify and record a more accurate depiction of the motor performance, current draw will be monitored asthe test is being conducted. In addition to monitoring the current output, the team will record the depththe rover digs vertically in the soil after the test is complete. The depth will be recorded whether the specifictest is a failure or success in order to evaluate the tread design and performance.

6.1.3 Independent Track Elevation

The variety of terrain conditions the rover could be exposed to include circumstances in which the rovertraverses an obstacle that is non-uniform or asymmetrical. This test specifically, is designed to elevate onetrack while the other remains on a flat surface. This test will evaluate the rover’s ability to continue tomaneuver as this track is elevated.

The test will be conducted in a controlled environment in a custom made test bed that will allow the teamto control and modify the angle the rover is tilted. The elevated tread will be subjected to heights rangingfrom 0.5 in. to 1.5 in. The obstacle itself will be placed in front of the rover’s line of travel to track the

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rover’s performance with an increasing tilt angle. This test will be performed on both sides of the rover inorder to evaluate whether the rover behaves differently.

The importance of this test is to determine if the elevated tread can still maneuver with the majority ofthe rover’s mass shifted to the opposing side and with the opposing tread having little contact surfaceleft to grip the soil. [TMS1] Because of this shifted mass, the rover has the possibility of rolling which willallow the team to determine if the rover’s center of mass needs to be adjusted to prevent the possibility of roll.

6.1.4 Ascending Ramp Angles

The tread design was designed to ensure that the rover could negotiate any obstacle that might be slopedupward; whether gradually sloped, steeply sloped, or curved. The test design was based on the worst caseconditions to which the rover might be subjected. The worst case scenario was based on the angle of reposeof a granular material. The worst case scenario for any soil composition is an angle of 45°.

The test will be conducted in a controlled environment in a custom made test bed that will allow the teamto control and modify the angle that the rover will be subjected to. Testing of the ramp angle will be inranges from 15° to 45° with different soil compositions. A linear testing distance of 3 ft. will be used as thestandardized testing method for each ramp angle.

This test not only evaluates the ability of the rover’s tread design to traverse the obstacle but also evaluatesan additional key parameter of torque output. The increased ramp angles can add additional stress on themotors and/ or possibly cause the motors to stall. In order to evaluate the motor performance the teamwill monitor the current draw of the motor while it negotiates the obstacle. This will enable the team toquantify and record the rover’s performance in varying environments.

6.1.5 Descending Ramp Angles

Most obstacles that the rover negotiates do not necessarily plateau which infers that there will be a descentthat follows an ascent when clearing obstacles. For this reason a testing method was developed to ensurethe rover can negotiate the apex of a steep slope and ascend the adjacent slope. This test will determinewhether the rover’s center of gravity needs to be shifted since this is the likely place the rover will transitionbetween slopes.

The test will be conducted in a controlled environment in a custom made test bed that will mate the as-cending and descending ramp angle tests. The test will include descending ramp angles that range from 15°to 45° paired with sharp transition zones at the apex and gently rolling transitions to evaluate how the rovernegotiates these with these conditions. This test will monitor if the rover experiences any teetering actionwhen transitioning between slopes and if the rover’s angle of descent changes with respect to the angle itascended the adjacent slope.

6.1.6 Sideway Slopes

The sideway slope test is designed to evaluate how the rover performs when both tracks are maneuvering ona slanted terrain feature. This test will determine at what angle the obstacle will cause the rover to beginto slide down the slope and if there is the possibility of roll. If the rover tread pattern provides sufficienttraction to prevent the rover from sliding, there is still the possibility of the rover rolling over. This willshow the limitations, if any, of the rover when traversing a slope.

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The test will be conducted in a controlled environment in a custom made testing apparatus that will allowthe team to control and modify the angle that the rover will be subjected to. The rover will be subjectedto traversing slopes that range from 15° to 45°. The distance of travel will be a minimum of 3 ft. or untilfailure. Failure for this test constitutes any slippage from which the rover can not maintain forward progressor if there is roll over.

6.1.7 Rover Autonomy

The mission requirements limit that the ground team to a single switch that will signal the orientation of therover followed by the deployment of the rover itself. Because of the absence of control, the team incorporatedsensors and a dead reckoning system that will enable the rover to navigate the terrain autonomously. Thesensors include both a camera and ultrasonic sensor that will enable the rover to detect potential objectsimpeding the rover’s path. This system will enable the rover to track its distance traveled as well as negotiateobstacles.

The tests will be conducted, initially, in a controlled environment with preset obstacles to ensure the sensorscan detect simple these simple objects. Once the object detection system is confirmed the rover will besubjected to tests in outdoor environments with random obstacles. The team will ensure that the outdoorenvironment has various non-uniform obstacles to ensure the rover object detection system is capable nego-tiating a multitude obstructions in order to accomplish the mission.

6.2 Rover Housing

The rover housing was designed such that the housing could rotate within the airframe in order to ensurethe rover deploys in an upright position. The housing rotates on polished aluminum bushing mounted insideof the airframe with the housing itself rotating via a servo motor. The rover sits inside of the housing andmounted on a leadscrew actuator. This actuator retains the rover securely inside of the housing and whensignaled the leadscrew rotates, forcing the rover out of the airframe. Two tests were designed in order toensure that the system functions as designed. These tests include:

• Rover Deployment

• Orientation Control

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6.2.1 Rover Deployment

The deployment of the rover is executed using a 5/16-12 leadscrew linear actuator mounted inside of therover housing. This system is designed to deploy the rover in various orientations in the event the payloadsection lands in an unfavorable condition. The test is designed to evaluate the ability of the rover to deploywhen placed in various unfavorable orientations.

The testing will be performed in a controlled environment placing the rover housing in several differentorientation angles ranging from 30° to 90° with the rover exit in the downward position. This test ensuresthat the rover can push the airframe off in the event it lands in a downward orientation.

6.2.2 Orientation Control

The deployment of the rover is dependent on its ability to exit the airframe undamaged and execute itsmission. In order to ensure that it is capable of deploying in an upright position when exiting the airframe,the team designed the rover housing so that it can rotate 360°. This design is paramount to the success ofthe rover because of the unlikelihood that the payload section of the airframe lands in a favorable orientationthat will enable to deploy the rover upright.

The rover housing design incorporates a motor, accelerometer, and microcontroller to rotate upon commandfrom the ground crew. The initial test will be conducted on the ground in a controlled environment to ensurethe integration of the system components work. Proceeding test will be conducted that will include boththe full and subscale test launches. The subscale will not incorporate the rover in it’s testing because of thelimited size of the airframe itself. The purpose of the subscale test is to ensure that the housing orientationsystem can survive the rigors of launch, flight and landing while still being capable of rotating in the cor-rect orientation when given the signal by the ground crew. The full scale test launches will incorporate allcomponents of the payload orientation system. The goal of the test is to validate that the rover and housingorientation system can survive the rigors of launch, flight, and landing while still capable of orienting therover upright, and deploying the rover undamaged.

6.3 Launch Vehicle

6.3.1 Drag Mod

The drag mod system was designed in order to ensure that the launch vehicle reaches the target altitudeof 5,280 ft. The system operates by opening and closing its control surfaces to increase and decrease dragrespectively. Two different system aspects are tested in order to ensure the drag mod operates as designed.

• Control Surface Aerodynamics

• Electronic Control System

The drag mod system will be ground tested to ensure that the control surfaces actuate as designed beforea test flight is conducted. This ensures that the overall electronic control system within the drag mod isfunctioning properly with repeatable results. The test flight will determine the aerodynamics of the con-trol system by confirming the drag coefficient calculations and that the control surfaces actuated during flight.

6.3.2 Aerodynamics

The aerodynamics of the launch vehicle will be confirmed through test flights of the full and subscale rocket.These test flights are conducted in order to justify the launch vehicle design and to determine if any modi-fications to the nosecone, airframe, or fincan are needed to reach the target altitude. The coefficient of drag

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was estimated through calculations and the stability margin was determined through Rocket Sim. Thesevalues will be verified through both the subscale and full scale test flights.

6.3.3 Seperation Testing

In order to ensure that the black powder charges generate enough force to separate the airframe prior toparachute deployment, a ground separation test will be conducted. The ground separation test will confirmthat the charges are adequate shear pin separation and ensure that the impulse load experienced by therocket do not affect the other subsystems.

The test will be conducted using a gantry style crane in order to reduce the amount of friction with solidsurfaces (i.e. if the test was conducted on the ground or in a cradle). The rocket will be suspended from thegantry and fixed with a remote system that will detonate the black powder charges. The detonation will beinitiated outdoors when all team members and all other personnel are clear. Post detonation examinationwill determine if the shear pins have successfully broken and all other subsystem components are unaffected.

6.3.4 Airframe/Altimeter Bay Shear Strength

The carbon fiber altimeter bay plays a key role in that it serves in protecting the altimeters from physicaldamage and unwanted external signals that could trigger premature parachute deployment. In addition, thealtimeter bay bears the load experienced during parachute deployment. For this reason a tensile test willbe conducted in order to determine the necessary fasteners needed to withstand the force experienced whenthe parachute deploys.

This test will be conducted with an Instron Testing machine in the UNC Charlotte Material Science Lab-oratory. The test will utilize a carbon fiber sidewall and fiberglass sidewall mated together with a 1/16in thickness single fastener fixed to an apparatus. The apparatus end and sidewall end will are secured inthe Instron Testing Machine and pulled until failure. Failure criteria for this test will be classified as anystress fracture in the composites or bending in the fastener. This test will enable the team to determine thenumber and size of fasteners to withstand the force experienced at the parachute deployment.

6.3.5 Bulkhead Failure

In order to ensure the payload parachute bay can withstand the forces experienced during parachute deploy-ment a bulkhead failure test will be conducted. This test will utilize the Instron Tensile Testing Machine topull the bulkhead fixed within a section of airframe until failure. Failure criteria for this test will be classifiedas any stress fracture in the bulkhead or airframe and bending of the bulkhead.

The test will take a small section of airframe with two bulkheads secured on either end with epoxy. Eachbulkhead will have an eyebolt mounted in the center of both ends that will be secured in the jaws of theInstron Testing Machine. The section will then be pulled until failure. This test is designed to verify thatthe bulkhead thickness and the type of epoxy utilized are capable of withstanding the forces experiencedduring parachute deployment.

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6.4 Testing Methods Summary

The testing methods and objectives which will be used by the 49er Rocketry team are designed to testcritical subsystems of the payload and vehicle. These tests are intended to mirror the expected environmentalconditions each system will experience or to verify manufacturer part specifications. The system and systemaspect to be tested and the associated testing methods and objectives are shown in Table 63.

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Table 63: Testing Methods

System System Aspect Testing Methods Objectives ResourcesFeatures and

FunctionsRisk Factors

Rover TreadSinkage

Partial submersion ofrover in damp soil

Ensure tread design andmotors are adequate in

adverse terrain conditions

Test bed,microcon-

troller

Motors havesufficient output,

robust tread design

Rover electronicsdamaged in wet terrain

Grouser Designand RoverTraction

Steep terrainEnsure rover can climb

steep slopesTest bed

Grouser profileprovides optimum

traction

Rollover causing damageto rover

Rover

Roverautonomy

Obstacle Course,outdoor environment

Ensure rover canautonomously navigate

terrain

Naturalobstacles

(e.g. shrubs,rocks, ect.),

camera,ultrasonic

sensors

Object recognitionand avoidance

Fails to identify objects

Dead ReckoningDrive Rover knowndistance on various

terrains

Verify rover properlytracks distance travelled

IMU, GPS,test bed,outdoor

environment

Distance TrackingMission failure due to

rover stoppingprematurely

RoverRover

DeploymentTest deployment at

various anglesEnsure successful rover

deploymentRover,housing

leadscrew providessufficient force

Binding, failure to deploy

Housing OrientationControl

Full and subscalehousing orientation

Determine functionalityof algorithm, efficacy of

design

servo motor,accelerome-

ter,microcon-

troller,bushings

Housing rotatesrover to upright

position

Binding due to friction,inadequate torque

Control SurfaceAerodynamics

Full scale controlsurface

characterization

Determine the dragcoefficient range for

control surfacesTest Flights

Increasing anddecreasing of drag

Failure to achieve targetaltitude

Drag Mod ElectronicControl System

Hardware in the looptesting

Attenuate altitude to apredetermined value

Groundsimulations,test flights

Apogee altitudecontrol

Failure to achieve targetaltitude

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Aerodynamics Test FlightExperimentally derive

coefficient of drag, verifystability

Airframewith fincan

andnoseconeassembled

Stability, center ofpressure,

aerodynamics ofvehicle

Instability, failure toreach target altitude

StructuralIntegrity

Test Flight

Demonstrate airframestructural robustness

through test flights andground testing

Airframewith fincan

andnoseconeassembled

Structuralcapability of fincan

during flight

Instability at launch, lossof fin(s) during flight,

zippering of airframe atparachute deployment

ParachuteDeployment

Launch vehicle sectionmain and drogue.

Payload section mainas drogue

Validate the mechanics ofusing the main parachute

as a drogue, anddetermine coefficients of

drag

Altimeters,data-loggingaccelerome-

ter,microcon-

troller,Elevatedplatform.

Test Flights

Kinetic energy,drop velocity,

impact velocity.Diameter, material,

and shape ofparachute selection

Parachute deploymentfailure, catastrophic

impact

Vehicle

Radio LocatorRadio frequency

tracker and receiverunit

Determine range andability to locate launch

vehicle

Radiofrequency

tracker andreceiver unit

Radio frequencylocater, receiver

unit

Losing radio frequencytracker

VehicleSeparation

Ground separationtesting

Determine force requiredto break shear pins and

impulse loads experiencedby rocket sections, verify

calculations

Shear pins,black

powdercharges,

data-loggingaccelerome-

ters,remote-

detonationsystem,rocketsectiongantry

Proper shear pinselection to ensure

separation

Failure to separate anddeploy parachutes,

damaging electroniccomponents

Launch VehicleMotor

Determine impulsesand thrust curves

Determine the maximumaccelerations and thrust

during launch. Determinerail-exit velocity

Rocketmotor teststand, load

cells

Total impulse ofselected motor andmaximum thrust

Not achieving minimumrail-exit velocity andfailure to reach target

altitude

113



Airframe/Altimeter BayShear Strength

Instron Tensile Test

Determine the forcerequired to cause the

airframe sidewall,altimeter sidewall or

fastener to fail

InstronTesting

Machine,carbon fiber

sidewall,fiberglasssidewall,fastener,testing

apparatus

Proper fastenerselection type and

quantityMaterial failure

Vehicle BulkheadTensile Strength

Instron Tensile Test

Detrmine the forcerequired to cause thebulkhead, eyebolt, or

epoxy to fail

InstronTesting

Machine,airframe

section withepoxied

bulkhead ,testing

apparatus

Proper bulkheadthickness and

epoxy selectionMaterial failure

114

7 Project Plan

7.1 Requirements Verification

7.1.1 NASA Student Launch General Requirements

The NASA Student Launch has created a list of requirements that need to be satisfied to ensure the team,the launch vehicle, the design, and procedures are aligned with the guidelines and regulations for a safeand prosperous flight. Compliance with these requirements, which are separated into five separate systemcategories, is necessary to participate in the 2017/18 USLI competition. The categories are as follows:

1. General Requirements

2. Launch Vehicle Requirements

3. Recovery System Requirements

4. Experiment Requirements

5. Safety Requirements

Each sub-category for compliance will be allocated one (or more) of the four fundamental methods of require-ment verification: Inspection, Demonstration, Test, and/or Analysis. These methods verify the requirementsof the system in a hierarchical order, increasing in time and rigor:

1. Inspection:

• Inspection is the non-intrusive and nondestructive examination of the system or product using thefive senses and could include physical manipulation or measuring of the system, or observation ofthe results of the system.

2. Demonstration:

• Demonstration is the manipulation of the system by running the system as it is intended to beused to verify that the results match the expected outcome.

3. Test:

• Testing is verifying the system through using a predefined series of inputs to ensure the systemproduces a predefined output that are specified in the requirements.

4. Analysis:

• Analysis is the verification of a system through utilizing models, calculations and testing equip-ment. By using confirmed test results, predictive statements and assumptions can be made aboutthe performance of the system. This will most often be used to predict the failure point of thesystem through nondestructive testing.

The break-down of the requirements compliance for each category can be found in their respective sub-sections throughout the document. The tables will include the NASA Student Launch requirements, themethod of verification that will be used for compliance, and a brief description of the verification plan.

115

Table 64: General Requirements Verification


Item Requirement Verification VerificationMethod Plan

1.1

Students on the team will do 100% of theproject, including design, construction,

written reports, presentations, and flightpreparation with the exception of assemblingthe motors and handling black powder or anyvariant of ejection charges, or preparing andinstalling electric matches (to be done by the

team’s mentor).

Inspection

The 49er Rocketry Team iscompletely student-lead. An

inspection will be conducted toverify all work is performed by

student members, with theexception of operations thatteam mentors are required to

handle (i.e. assembling motors,handling ejection charges, orinstalling/preparing electric

matches)

1.2

The team will provide and maintain aproject plan to include, but not limited to

the following items: project milestones,budget and community support, checklists,personnel assigned, educational engagement

events, and risks and mitigations.

Inspection

The 49er Rocketry Team utilizesan online-database and weeklymeetings to communicate eventdates and project milestones.

1.3

Foreign National (FN) team members mustbe identified by the Preliminary Design

Review (PDR) and may or may not haveaccess to certain activities during launch

week due to security restrictions. In addition,FN’s may be separated from their team

during these activities.

N/A There is one FN on our team.

1.4

The team must identify all team membersattending launch week activities by theCritical Design Review (CDR). Team

members will include:

• 1.4.1 Students actively engaged in theproject throughout the entire year.

• 1.4.2 One mentor (see requirement 1.14)

• 1.4.3 No more than two adult educators.

Inspection

Team members, mentors, andeducators will express interestin attending the launch beforethe CDR is due. Team Leadswill verify that all personnelwho are interested meet thenecessary requirements to

attend Launch Week.

1.5

The team will engage a minimum of 200participants in educational, hands-on science,

technology, engineering, and mathematics(STEM) activities, as defined in the

Educational Engagement Activity Report, byFRR. An educational engagement activityreport will be completed and submittedwithin two weeks after completion of an

event. A sample of the educationalengagement activity report can be found on

page 31 of the handbook. To satisfy thisrequirement, all events must occur betweenproject acceptance and the FRR due date.

Inspection

Educational EngagementActivity Reports will give anaccurate description of the

team’s community engagementand meet all requirements

specifications needed.

116

1. General Requirements Continued


1.6The team will develop and host a Web site

for project documentationInspection

The team’s website,www.unccrocketry.org, will be

updated with the mostup-to-date project

documentation to verify theteam’s development and

progress throughout the entiretyof the competition season.

1.7

Teams will post, and make available fordownload, the required deliverables to the

team Web site by the due dates specified inthe project timeline.

Inspection

The designated siteadministrator will verify that alldeliverables will be uploaded to

the team website and madeavailable in PDF Format before

the required due dates.Additional members will verifythat all documents are available

for viewing after they areuploaded to the website.

1.8 All deliverables must be in PDF format. Inspection

All documentation will be donethrough the web-based

platform, ShareLaTex, and willbe compiled into PDF format.

1.9In every report, teams will provide a table ofcontents including major sections and their

respective sub-sections.Inspection

All documentation will be donethrough ShareLaTex, which uses

code tags to create a table ofcontents with sections andsub-sections that update

continuously as the document iscompiled so it remains accurate

and current with allinformation. Members in charge

of documentation will verifythat the sections are labeled

appropriately.

1.10In every report, the team will include thepage number at the bottom of the page.

Inspection

Using ShareLaTex, documentswill be written to include pagenumbers at the bottom of each

page. Members in charge ofdocumentation will verify thatpage numbers are present and

accurate.

1.11

The team will provide any computerequipment necessary to perform a video

teleconference with the review panel. Thisincludes, but is not limited to, a computer

system, video camera, speaker telephone, anda broadband Internet connection. Cellular

phones can be used for speakerphonecapability only as a last resort.

Inspection

The 49er Rocketry Team willcontact the College of

Engineering 2 weeks prior to allteleconferences to ensure a room

is available and equipped forvideo teleconferences with the

review panel.

117

1. General Requirements Continued


1.12

All teams will be required to use the launchpads provided by Student Launch’s launchservice provider. No custom pads will be

permitted on the launch field. Launchservices will have 8 ft. 1010 rails, and 8 and

12 ft. 1515 rails available for use.

Demonstration

The launch vehicle will bedesigned to launch from the

required launch pads and usingthe specified rails available from

the launch service providers.

1.13

Teams must implement the Architecturaland Transportation Barriers Compliance

Board Electronic and InformationTechnology (EIT) Accessibility Standards(36 CFR Part 1194) Subpart B-TechnicalStandards (http://www.section508.gov): §

1194.21 Software applications and operatingsystems. § 1194.22 Web-based intranet and

Internet information and applications.

Inspection

All members will beknowledgeable on the

Architectural andTransportation Barriers

Compliance Board Electronicand Information TechnologyAccessibility Standards and

designated team members willverify all standards are

implemented

1.14

Each team must identify a “mentor.” Amentor is defined as an adult who is

included as a team member, who will besupporting the team (or multiple teams)throughout the project year, and may or

may not be affiliated with the school,institution, or organization. The mentor

must maintain a current certification, andbe in good standing, through the NationalAssociation of Rocketry (NAR) or TripoliRocketry Association (TRA) for the motor

impulse of the launch vehicle and musthave flown and successfully recovered(using electronic, staged recovery) a

minimum of 2 flights in this or a higherimpulse class, prior to PDR. The mentor isdesignated as the individual owner of the

rocket for liability purposes and must travelwith the team to launch week. One travel

stipend will be provided per mentorregardless of the number of teams he or shesupports. The stipend will only be providedif the team passes FRR and the team and

mentor attends launch week in April.

Inspection

The 49er Rocketry Team hascreated an organization chartoutlining the hierarchy of theteam’s members, including all

mentor involvement. The TeamLead will ensure all ‘mentors’

are in compliance with thementor requirements.

118

Table 65: Launch Vehicle Requirements Verification

2. Launch Vehicle Requirements


2.1The vehicle will deliver the scientific payload

to an apogee altitude of 5,280 ft. aboveground level.

Analysis

The launch vehicle will reachthe target altitude through a

combination of motor selection,vehicle aerodynamics, dragmodulation, and the overall

mass.

2.2

The vehicle will carry one commerciallyavailable, barometric altimeter for recordingthe official altitude used in determining the

altitude award winner.

AnalysisOne altimeter will be designatedas the official scoring altimeter.

2.3

Each altimeter will be armed by a dedicatedarming switch that is accessible from theexterior of the rocket airframe when the

rocket is in the launch configuration on thelaunch pad.

InspectionHoles will be placed in the

airframe that reveal the armingswitches for the altimeters.

2.4Each altimeter will have a dedicated power

supply.Inspection

Each altimeter will have adedicated 9V battery supply.

2.5Each arming switch will be capable of being

locked in the ON position for launch (i.e.cannot be disarmed due to flight forces).

InspectionArming switches will be keylock

switches.

2.6

The launch vehicle will be designed to berecoverable and reusable. Reusable is defined

as being able to launch again on the sameday without repairs or modifications.

Inspection

Robust design, repackableparachutes, and replaceable fins

on hand at launch site willensure relaunch on same day.

2.7

The launch vehicle will have a maximum offour (4) independent sections. An

independent section is defined as a sectionthat is either tethered to the main vehicle oris recovered separately from the main vehicle

using its own parachute.

Inspection

The launch vehicle will becomprised of three main

sections, two of which will sharethe same parachute system.

2.8The launch vehicle will be limited to a single

stage.Inspection

Only one motor will be used forthe launch vehicle.

2.9

The launch vehicle will be capable of beingprepared for flight at the launch site within 3

hours of the time of Federal AviationAdministration flight waiver opens.

AnalysisVehicle will be designed with

assembly in mind.

2.10

The launch vehicle will be capable ofremaining in launch-ready configuration atthe pad for a minimum of 1 hour without

losing the functionality of any criticalon-board components.

Analysis

Battery budget will be done toensure vehicle can remain on

launch rail for upwards of threehours.

2.11

The launch vehicle will be capable of beinglaunched by a standard 12-volt direct current

firing system. The firing system will beprovided by the NASA-designated Range

Services Provider.

AnalysisCommercially available igniters

will be used.

119

2. Launch Vehicle Requirements Continued


2.12

The launch vehicle will require no externalcircuitry or special ground support

equipment to initiate launch (other thanwhat is provided by Range Services).

Inspection

All electrical components of thelaunch vehicle will be housed

internally, and each section willhave its own dedicated

electronics bay.

2.13

The launch vehicle will use a commerciallyavailable solid motor propulsion system usingammonium perchlorate composite propellant(APCP) which is approved and certified by

the National Association of Rocketry (NAR),Tripoli Rocketry Association (TRA), and/or

the Canadian Association of Rocketry(CAR).

AnalysisThe motor will be purchased

commercially through licensedvendors.

2.14Pressure vessels on the vehicle will be

approved by the RSO and will meet thecriteria laid out in the handbook.

InspectionThere will be no pressure vessels

on the vehicle.

2.15The total impulse provided by a collegeand/or university launch vehicle will notexceed 5,120 Newton-seconds (L-class).

InspectionThe motor selected will be

L-class or below.

2.16

The launch vehicle will have a minimumstatic stability margin of 2.0 at the point ofthe rail exit. Rail exit is defined as at thepoint where the forward rail button loses

contact with the rail.

AnalysisSimulation software will beutilized to verify stability

margin at rail exit.

2.17The launch vehicle will accelerate at a

minimum of 52 fps at rail exit.Analysis

Simulation software will beutilized to verify velocity at rail

exit.

2.20Any structural protuberance on the rocketwill be located aft of the burnout center of

gravity.Analysis

Drag modulation system will belocated aft of the center of

gravity.

2.21 Vehicle Prohibitions. AnalysisAll vehicle prohibitions will be

adhered to.

120

Table 66: Recovery System Requirements Verification

3. Recovery System Requirements


3.1

The launch vehicle will stage the deploymentof its recovery devices, where a drogue

parachute is deployed at apogee and a mainparachute is deployed at a lower altitude.

Tumble or streamer recovery from apogee tomain parachute deployment is also

permissible, provided that kinetic energyduring drogue-stage descent is reasonable, as

deemed by the RSO.

Analysis

Booster and Recovery sectionwill descend under a 3 ft.

drogue parachute until maindeployment at 600 ft. The

payload section will use a reefedmain parachute to act as a

drogue until opening fully at300 ft.

3.2

Each team must perform a successful groundejection test for both the drogue and mainparachutes. This must be done prior to the

initial subscale and full-scale launches.

TestGround testing systems will be

set up to test parachute ejectionprior to launches.

3.3At landing, each independent sections of thelaunch vehicle will have a maximum kinetic

energy of 75 lbf-ft.Test

Recovery and booster sectionwill land with a kinetic energy

of 20.96 lbf-ft and payloadsection will land with a kinetic

energy of 27.38 lbf-ft.

3.4The recovery system electrical circuits will be

completely independent of any payloadelectrical circuits.

Demonstration

All recovery systems will beindependently powered and

placed in electromagneticallyshielded sections in the rocket.

3.5All recovery electronics will be powered by

commercially available batteries.Demonstration

All altimeters will be poweredby independent commercially

available batteries.

3.6

The launch vehicle will be designed to berecoverable and reusable. Reusable is defined

as being able to launch again on the sameday without repairs or modifications.

Test & Analysis

The recovery system usesPerfectflite Stratologgers for

deployment. A secondaryaltimeter on independent power

is used as redundancy.

3.7Motor ejection is not a permissible form of

primary or secondary deployment.Inspection

The motor will be locked intothe booster section.

3.8Removable shear pins will be used for boththe main parachute compartment and the

drogue parachute compartment.Inspection

Shear pins will be purchasedcommercially to use for recovery

sections.

3.9Recovery area will be limited to a 2500 ft.

radius from the launch pads.Demonstration

& Analysis

Drogue parachutes and delayedopening systems will be used tolimit drift from wind currents.

3.10

An electronic tracking device will be installedin the launch vehicle and will transmit the

position of the tethered vehicle or anyindependent section to a ground receiver.

InspectionElectronic tracking devices will

be installed to the electronicbays.

3.11

The recovery system electronics will not beadversely affected by any other on-board

electronic devices during flight (from launchuntil landing).

DemonstrationElectronic systems used for

recovery will be isolated fromother major electronic bays.

121

3. Recovery System Requirements Continued


3.11.1

The recovery system altimeters will bephysically located in a separate compartment

within the vehicle from any other radiofrequency transmitting device and/or

magnetic wave producing device.

Inspection

Recovery system electronics willbe in a section surrounded bycarbon fiber which will shield

the electronics from radiotransmissions.

3.11.2

The recovery system electronics will beshielded from all onboard transmitting

devices, to avoid inadvertent excitation ofthe recovery system electronics.

Inspection &Demonstra-

tion



3.11.3

The recovery system electronics will beshielded from all onboard transmitting

devices which may generate magnetic waves(such as generators, solenoid valves, and

Tesla coils) to avoid inadvertent excitation ofthe recovery system.


tion



3.11.4

The recovery system electronics will beshielded from any other onboard deviceswhich may adversely affect the proper

operation of the recovery system electronics.


tion



Table 67: Experiment Requirements Verification



4.5.1Teams will design a custom rover that willdeploy from the internal structure of the

launch vehicle.Test

A wireless signal will be sentfrom the ground station to

deploy the rover.

4.5.2At landing, the team will remotely activate atrigger to deploy the rover from the rocket.

Test

The rover retention system willbe released and the rover will

turn on upon activation oftrigger.

4.5.3After deployment, the rover will

autonomously move at least 5 ft. (in anydirection) from the launch vehicle.

Test

Sufficient testing of the deadreckoning algorithm will be

conducted to ensure the 5 ft.mark is reached.

4.5.4Once the rover has reached its final

destination, it will deploy a set of foldablesolar cell panels.

TestNumerous tests will be run ondeployment mechanism of solar

cell panels.

122

Table 68: Safety Requirements Verification



5.1

Each team will use a launch and safetychecklist. The final checklists will beincluded in the FRR report and usedduring the Launch Readiness Review(LRR) and any launch day operations

Inspection,Demonstration

Safety checklists will begenerated for all assembly,

launch, and recovery activities.Final checklists will be used

during sub- and full-scale testsand included in the FRR report.All applicable checklists will beused during the LRR and any

launch day activities.

5.2Each team must identify a student safety

officer who will be responsible for allitems in section 5.3.

Demonstration

Safety officer has beenidentified: Dallas Solomon.

Alternate safety officers havebeen identified: Taylor Forte,

Ashton Johnston.

5.3

The role and responsibilities of each safetyofficer will include, but not limited to:

5.3.1 Monitor team activities with anemphasis on Safety during:

• 5.3.1.1 Design of vehicle and payload

• 5.3.1.2 Construction of vehicle andpayload

• 5.3.1.3 Assembly of vehicle and pay-load

• 5.3.1.4 Ground testing of vehicle andpayload

• 5.3.1.5 Sub-scale launch test(s)

• 5.3.1.6 Full-scale launch test(s)

• 5.3.1.7 Launch day

• 5.3.1.8 Recovery activities

• 5.3.1.9 Educational Engagement Ac-tivities


Safety officer will be present atall activities discussed in para.

5.3.1.1 through 5.3.1.9.Checklists will be used at all

applicable activities.

5.3.2Implement procedures developed by theteam for construction, assembly, launch,

and recovery activities


Detailed checklists will begenerated for the all assembly,

test, launch, and recoveryactivities. Use of the checklists

is mandatory

123

5. Safety Requirements Continued


5.3.3

Manage and maintain current revisions ofthe team’s hazard analyses, failure modesanalyses, procedures, and MSDS/chemical

inventory data


A UNCC Safety Binder will bemaintained. The binder willcontain the team’s hazard

analyses, FMEAs, assembly,test,and launch procedures, and all

applicable safetydocumentation.

5.3.4Assist in the writing and development ofthe team’s hazard analyses, failure modes

analyses, and procedures.

Analysis,Demonstration,

Inspection

Safety officer will work closelywith subteams to develop

team’s hazards analyses, FailureMode and Effects Analyses, and

assembly, test, and launchprocedures.

5.4

During test flights, teams will abide bythe rules and guidance of the local

rocketry club’s RSO. The allowance ofcertain vehicle configurations and/or

payloads at the NASA Student LaunchInitiative does not give explicit or implicit

authority for teams to fly those certainvehicle configurations and/or payloads at

other club launches. Teams shouldcommunicate their intentions to the local

club’s President or Prefect and RSObefore attending any NAR or TRA

launch.

Demonstration

The Safety officer will conduct aprelaunch safety briefing no

earlier than three days prior totraveling to the launch field.The briefing will include thespecific launch field rules and

emphasize that the RangeSafety Officer is the ultimate

launch authority.

5.5Teams will abide by all rules set forth by

the FAA.Inspection,

Demonstration

The 49er Rocketry Team willabide by all rules and

regulations set forth in 14 CFRCh. I Subchaper F Part 101

Subpart C - Amateur Rockets.This includes ensuring thelaunch vehicle meets the

requirements of §101.23 and101.25

124

7.1.2 Team Derived Requirements

In addition to the NASA Launch requirements presented in the handbook, team derived requirements werecreated that extend beyond the boundaries of the competition expectations to create further accountabilitywithin the team. The 49er Rocketry Team has generated a series of team requirements, broken down intothe five categories listed below:


2. Vehicle Requirements

3. Recovery Requirements



The verification process for the team derived requirements aligns with the process used to verify the manda-tory NASA Launch Requirements. Each requirement is itemized, assigned a verification method and verifi-cation plan, located below:

125

Table 69: Team Derived General Requirements Verification

1. Team Derived General Requirements


1.1

The day and time of weekly mentormeetings will be Thursdays at 6:30pm.

Additional metings will take placethroughout the week if necessary.

Inspection

Team members have signed acontract accepting responsibility

for attending all requiredmeetings. If additional meetingsare coordinated, the time andlocation will be announced viaemail with a minimum of 24

hours notice.

1.2

Members agree to attend all mandatorymeetings and be on time to all meetings,

with a tardiness of no more than 5minutes being tolerated.

Inspection

Team members have signed acontract accepting

accountability to be punctual toall mandatory team events and

meetings.

1.3Members agree that team communicationwill be professional and responses will betimely, within 24 hours of initial contact.

Inspection

Team utilizes groupme andemail, both of which will havenotifications ’on’ for all project

related discussion to ensurenecessary communication is

available at all times.

1.4

Members agree to take ownership of theirrespective projects and duties, includingdocumentation, drawings, calculations,

etc.

Inspection

Team members have signed acontract accepting ownershipand responsibility to completeprojects without substantial

oversight by project leadership.

1.5Members agree to complete assignments

by agreed upon end date.Inspection

Team members willcommunicate deadlines and time

conflicts to ensure all projectscan be done in a timely manner.

1.6Violations of team contracts will be

documented with the offending member’ssignature and signature of project lead.

Inspection

Contract will be reviewed withteam member, and will meet

with Project Lead (Ashton) todiscuss further consequences if

behavior continues further.

126

Table 70: Team Derived Vehicle Requirements Verification

2. Team Derived Vehicle Requirements


2.1 Weight Max: 60 lbm Analysis

Use lightweight materials, anddo not have any empty space.

Any onboard components mustbe absolutely required and must

have a purpose. If it is notneeded then it will not be used.Also, no excess material will be

used.

2.2 Size Max: 9 ft. Analysis

Absolutely no empty space. Allsections will be designed to bespace efficient and not waste

vehicle volume. No excessmaterial and no extra

components will be used.

2.3Do not overshoot required apogee altitude

by 10%Analysis

The motor will be powerfulenough to propel a vehicle thatweights 60 lbm. This will allowthe vehicle to reach the required

altitude but also be able toovershoot the mile.

2.4Allow for EM to pass through necessary

materials on Launch VehicleAnalysis

Purchase, and properly handle,fiberglass material.

2.5Do not allow for EM to pass throughcertain materials on Launch Vehicle

AnalysisPurchase, and properly handle,

carbon fiber and metalcomponents.

Table 71: Team Derived Recovery Requirements

3. Team Derived Recovery Requirements


3.1Parachute will not exceed a maximum

size of 10 ft.Analysis

Have the vehicle volumes andmasses be appropriate enoughto fit various 10 ft. parachutes.

3.2Black Powder quantity will not exceed a

maximum of 10 gramsAnalysis

Designed to have vehicleseparation be as simple as

possible with minimal use ofblack powder for safety

127

Table 72: Team Derived Payload Requirements Verification

4. Team Derived Payload Requirements


4.1 Weight will be ∼10lbs AnalysisCAD design for estimated mass.

Measure the weight using ascale for exact mass

4.2 Rover dimensions of 4.5” x 4.5” x 8” AnalysisCAD analysis. Placement of full

scale rover into housing

4.3Expected leadscrew retention will holdRover in place during flight, parachute

deployment, and landingTest Ground testing, full scale testing

4.4Rover Housing must be covered during

descent, the cap must stay on the housingTest

Ground testing, sub scaletesting, full scale testing

4.5Housing must be able to attenuate

wireless signals given by the ground teamAnalysis and

TestResearch, ground testing, sub

scale and full scale testing

128

Table 73: Team Derived Safety Requirements Verification

5. Team Derived Safety Requirements


5.1All team members will wear closed toeshoes in the Motorsports Research Lab.


Self Inspection. Safety officerwill ensure all team membersare wearing closed toe shoes

during construction, assembly,test, launch/ recovery activities.

5.2Safety Data Sheets will be reviewed byteam members prior to using hazardous

materials.


Self Inspection. Safety DataSheets will be centrally locatedfor review prior to hazardous

material use.

5.3All Ematches and igniters will be checked

for continuity prior to use. All failureswill be discarded.

Inspection, Test

Digital multimeter will be usedto verify continuity through the

ematch/ igniter leads uponarrival. Ematches/ igniters that

check ’bad will be discarded.Safety officer will record date oftesting and who conducted test.

5.4Safety glasses will be worn at all timeswhen working in Motorsports Research

Lab.


Self Inspection. Safety officerwill ensure all team members

are wearing safety glassesduring construction, assembly,

test, launch/ recovery activities.

5.5

Two personnel are required to visuallyverify critical steps in checklists, i.e.

altimeter settings, Jolly Logic settings,etc.)

Inspection

Checklists will be designed forboth safety officer and project

lead to initial critical steps.Altimeter/ parachute releasesettings will be annotated in

checklists

5.6All recovery parachutes will be wrapped

in fire retardant material to preventdamage from separation charges.


Checklists will include a criticalstep indicating this requirement.

This procedural step will bepreceded by a ”Caution”

statement

5.7Hearing protection will be worn whileoperating machinery with noise levels

exceeding 85dB.Inspection

Self Inspection. Safety officerwill verify that personnel using

machinery with noise levelsexceeding 85dB, are wearing

hearing protection IAW 29 CFRPart 1910 Subpart G para.

1910.95

129

7.2 Budgeting and Timeline

7.2.1 Budget

Table 74: Launch Vehicle Anticipated Budget

System Part Price Quantity Subtotal36” Drogue - FruityChutes $89.00 1 $89.00120” Mains - FruityChutes $355.00 2 $710.00Perfect Flight Stratologger $79.95 6 $479.702200 lb. Kevlar Shock Cord 1/4 in. $3.30 60 $198.004000 lb. Nylon Shock Cord 1 in. $2.50 30 $75.0013” Nomex Blanket $16.00 1 $16.00E-matches (Case of 40) $43.00 1 $43.00

Recovery

Black Powder 1 lb. $24.99 1 $24.99Total: $1,635.69

5” x 36” CF AF (Full scale) $225.00 1 $225.004” x 24” CF AF (Sub scale) $105.00 2 $210.005” CF Coupler (Full scale) $80.00 2 $160.004” CF Coupler (Sub scale) $60.00 2 $120.006” x 36” FG AF (Full scale) $200.00 1 $200.005” x 30” FB AF (Sub scale) $85.00 1 $85.00Rail Buttons $10.00 1 $10.00ABS RP $150.00 1 $150.00Polycarbonate RP $500.00 1 $500.00Nosecone $140.00 2 $280.00Fasteners $100.00 1 $100.00

Airframe

Miscellaneous $200.00 1 $200.00Total: $2,240.00

Aerotek K1100 $112.99 3 $338.97Aerotek L1420 $249.99 3 $749.9754 mm Motor Casing $200.00 1 $200.00

Motors

98mm Motor Casing $500.00 1 $500.00Total: $1,788.94

GPS $30.00 1 $30.00Xbee $130.00 1 $130.00MCU $30.00 2 $60.00Accelerometers $40.00 1 $40.00Pressure Sensor $40.00 1 $40.00Batteries $100.00 1 $100.00

Electronics

PCB $50.00 1 $50.00Total: $450.00

Grand Total: $6,726.09

130

Table 75: Payload Anticipated Budget

System Part Price Quantity SubtotalMechanical

Bearings $35.00 1 $35.00Chassis $50.00 1 $50.00Pillow Blocks $50.00 1 $50.00ABS RP $150.00 1 $150.00

ElectronicsDrive Servos $35.00 4 $140.00Solar Array Locking Solenoid $20.00 2 $40.00Accelerometer $40.00 2 $80.00MCU $30.00 2 $60.00PCB $50.00 4 $200.00Solar Cells $28.00 4 $112.00Object Detection Sensor $50.00 2 $100.00

Rover

Battery $50.00 2 $100.00Total: $1,117.00

Mechanical5 7/8” x 11” FG Housing Tube $400.00 1 $400.00ABS RP $150.00 1 $150.00Bearings $35.00 1 $35.00Bulkheads $50.00 2 $100.00

ElectronicsRotation Servo $35.00 1 $35.00Locking Solenoid $20.00 2 $40.00Retention Servo $50.00 2 $100.00MCU $30.00 3 $90.00Accelerometer $40.00 1 $40.00PCB $50.00 1 $50.00Batteries $50.00 2 $100.00GPS $30.00 1 $30.00

Rover Housing

Xbee $130.00 2 $260.00Total: $1,430.00

Table 76: Testing Budget

Item Description AmountRover Drive Testing $250.00Rover Deployment $250.00Wireless Communications $250.00Recovery Testing $250.00Total $1,000.00

Table 77: Outreach Budget

Item Description AmountLow Power Rocket Kits $150.00Low Power Motors $50.00NAR Lvl. 1 Kits $300.00Total $500.00

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Table 78: Anticipated Travel Budget

Item Description AmountFood $2,500.00Lodging $4,000.00Trailer Rental $200.00Gas $1,500.00Total: $8,200.00

The tables above represent the anticipated distribution of our budget for Launch Vehicle Fabrication, PayloadFabrication, Testing, Outreach and Travel. These values are subject to minor variation as fabricationcontinues.

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Table 79: Purchased Parts

System Part Part Price Quantity Total Price Company18-8 Stainless Steel Hex Drive RoundedHead Screw 6-32 Thread Size, 3/4” Long

$4.41 1 $4.41 McMaster-Carr

6061 Aluminum Rod-6” Diameter, 6” Long $104.72 1 $104.72 McMaster-Carr18-8 Stainless Steel Hex Drive RoundedHead Screw 6-32 Thread Size, 1/2” Long


Plastic Syringe with Taper Tip,0.2 oz. Capacity


Heat-Set Inserts for Plastics 6-32 ThreadSize, 0.150” Installed Length


Heat-Set Inserts for Plastics 6-32 ThreadSize, 0.250” Installed Length


Steel Eyebolt with Shoulder - for Lifting3/8”-24 Thread Size, 1-1/4” Thread Length


Nylon Rounded Head Screws Slotted DriveStyle, 6-32 Thread Size, 1/4” Long


Nylon Rounded Head Screws Slotted DriveStyle, 6-32 Thread Size, 1/2” Long


1117 Carbon Steel 1/4” Diameter 1 ft. Length $2.61 1 $2.61 McMaster-CarrAeropoxy Structural Adhesive - ES6209 $53.75 2 $107.50 Aircraft SpruceAERO PACK 54MM Retainer $31.03 1 $31.03 Apoge Components

Sub-Scale Vehicle

Black-Oxide Alloy Steel Socket HeadScrew 6-32 Thread Size, 1/4” Long


Subtotal: $366.94ACombo: 3pcs of Tenergy 30C 7.4 V 2200mAhLiPO Battery for Syma X8C/X8W/X8G

$39.99 1 $39.99 All-Battery.com

12 CPR Magnetic Encoder Pair Kit for Micro Metal $8.95 2 $17.90 Robot Shop900MHz Duck Antenna RP-SMA $7.95 1 $7.95 Robot ShopDigi International XBP9B-DMST-002 $39.00 1 $39.00 DigiKey

Electronics

Adafruit Ultimate GPS Breakout - 66channel w/10 Hz updates - Version 3

$39.95 1 $39.95 Adafruit

Subtotal: $144.79Alloy Steel Dowel Pin-3mm Diameter,18mm Long


6061 Aluminum Rod-6” Diameter, 6” Long $104.72 1 $104.72 McMaster-Carr6061 Aluminum Sheet- 0.19” Thick, 4” x 24” $26.96 1 $26.96 McMaster-CarrABS Filament for 3D Printers- 1.75mm Diameter $46.00 2 $92.00 Adafruit1/4” Stainless Steel D-Shafting 12” Length $4.69 2 $9.38 Servocity1515 Rail Button - Pair of 2 $2.00 1 $2.00 FruityChutes1010 Rail Buton - Pair of 2 $1.90 1 $1.90 FruityChutes

Airframe

AERO PACK 54MM RETAINER(Fiberglass Motor Tubes)

$29.00 1 $29.00 Madcow Rocketry

Subtotal: $291.10Shock Cord 1500lb Spectra Microline (3 yds) $30.00 1 $30.00 FruityChutesShock Cord 1500lb Spectra Microline (7 yds) $38.00 1 $38.00 FruityChutes3000 lb Stainless Steel Swivel $9.00 1 $9.00 FruityChutesIris 72” Ultra Light Parachute - 16bs @ 15fps $265.00 1 $265.00 Fruitychutes4” Deployment Bag (6” length) $39.00 1 $39.00 FruityChutes

Recovery

5/16” Stainless Steel Quick Link, 2100lb $8.00 4 $32.00 FruityChutesSubtotal: $413.00

Total: $1215.83

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Figure 45: Categorized Project Budget

Launch Vehicle Fabrication

38%

Payload Fabrication

8%Testing5%

Outreach3%

Travel46%

PROJECT BUDGET

From the categorized budget, the largest section is Travel which surpasses the budget of all Fabrication. TheTravel section includes traveling with the team and launch vehicle, housing and food for the competitionlaunch in Huntsville, Al.

7.2.2 Funding Plan

Contingent on the proposed budget above, the estimated funding necessary for the project is $19,227.79.To successfully subsidize this project, the 49er Rocketry Team is pursuing several avenues for funding. Theprojected funding plan is itemized in the table below.

Table 80: Projected Funding Plan

Funding Source AmountNC Space Grant $8,000.00Eclipse Glasses Sales $1,500.00UNCC Senior Design $3,000.00Crowdfunding/Sponsors $10,000.00Dr. Conrad NC Space Grant $1,500.00Department Donations $2,000.00Bridge Tournament $1,000.00Funds Carried Forward $2,800.00Ideal Total $29,800.00

William States Lee College of Engineering

The College of Engineering (COE) at UNC Charlotte makes $3,000 available to Senior Design teams. Addi-tional funding can be requested from individual departments in the COE, but the availability of additionalfunding is under the discretion of the individual departments.

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North Carolina Space Grant

The NC Space Grant has a financial grant available for student competitions with an upper limit of $8,000.An additional grant has been made available to the 49er Rocketry Team through the NC Space Grantadministrator on UNC Charlotte’s campus for $1,500. The 49er Rocketry Team will submit proposals forboth grants as soon as submissions are available.

Crowdfunding Campaign

The team is hosting a 30 day crowdfunding campaign through the university with the goal of raising $10,000and a stretch goal of $12,000. The crowdfunding campaign provides an avenue for the team to reach outto industry and community partners to raise awareness for our team and the NASA USLI Competition, aswell as significantly increase the team’s funding capabilities. The campaign will run from October 18th toNovember 17th.

Last year (2016-2017) the 49er Rocketry Team did a similar crowdfunding campaign and surpassed theirinitial goal of $10,000, raising $10,715 through crowdfunding alone. We hope to exceed our goal this year aswell by following a vigorous email and social media marketing strategy, detailed below:

Quiet Phase - 9/6/17

• 9/8/17 - Gather Data

– Emails, sorted from top 25% most likely to give down to 25% most likely to not give

• 9/18/17 - Create Content and Content Calendar

• 9/18/17 - Marketing Strategy in place

• 10/2/17 - Email #1

– Top 25% most likely to give

– Introduce Campaign/Project

• 10/10/17 - Email #2

– Next group of 25%

– Make a difference today (educational value of project)

• 10/16/17 - Drafts for alumni and UNCC donor list emails

Public Phase - 9/18/17 (30 days long)

• 10/18/17 - Social media push, announce campaign publicly

• 10/18/17 - Email #3

– Next group of 25%

– We need YOU (straight-forward ask - mention deadline)

• 10/20/17 - Update #1

– Updating donors and prospects on how fundraising is going so far for the group

• 10/25/17 - Email #4

– Last group of 25

135

– Update on the project so far

• 10/26/17 - Update #2

– Thanking donors and providing general update on funding status

• 11/2/17 - Email #5

– Alumni and UNCC donor list

– Push project and include deadline

• 11/6/17 - Update #3

– Update on project’s progress

• 11/13/17 - Email #6

– Alumni and UNCC donor list

– Push project and include deadline

• 11/14/17 - Email #7

– Entire email list

– Final push

• 11/16/17 - Public Phase ends

Stewardship Phase - 11/16/17

• 11/20/16 - Update #4

– Update on project and thank current donors

• 2/16/18 - Update #5

– Final thank you (include video if possible)

• 5/1/17 - Update #6

– Final project update following competition flight

Additional Funding Options

Additional funding for the project can be raised through volunteer efforts at events around Charlotte andpartnerships with local companies. In the past, the 49er Rocketry Team has volunteered at an internationalbridge tournament, working for an hourly rate of $10 that is donated directly to the project fund. We havealso worked at the local Renaissance Festival at an hourly rate of $7. The team is also considering partneringwith local restaurants to host fundraising nights where a portion of the sales from the night will be donatedto the project.

136

7.2.3 Major Project Timeline

The complete project plan is located in the Appendix which contains the NASA specified timeline andteam derived task completion dates, categorized by sub-components. The tables below highlights the majorproject milestones and their respective due dates.

Table 81: NASA Requirements

Task Completion DateProposal Due 9/20/2017Awarded Propsals Announced 10/6/2017Kickoff 10/12/2017Web presence Established 11/3/2017PDR Documentation Due 11/3/2017PDR Teleconference 11/29/2017CDR Q&A 12/6/2017CDR Documentation Due 1/12/2018CDR Teleconference 1/31/2018FRR Q&A 2/7/2018FRR Documentation Due 3/5/2018FRR Teleconferences 3/22/2018LRR 4/4/2018Launch Day 4/7/2018Post-Launch Assessment 4/27/2018

Table 82: Anticipated Project Development

Payload TimelineTask Start Date Completion DatePayload Design Decision 9/29/2017 9/29/2017Payload Design Testing 10/6/2017 10/27/2017Full Scale Design of Payload 11/23/2018 12/19/2018Final Assembly of Payload 1/5/2018 1/30/2018Full Scale Payload Testing 1/31/2018 2/20/2018

Vehicle TimelineTask Start Date Completion DateVehicle Design Decision 9/2/2017 9/30/2017Subscale Design of Vehicle 10/2/2017 10/15/2017Subscale Vehicle Launch 11/11/2017 11/12/2017Full Scale Vehicle Assembly 1/5/2018 1/30/2018Full Scale Vehicle Testing 1/31/2018 2/20/2018Full Scale Launch 2/24/2018 2/25/2018NASA Launch 4/7/2018 4/7/2018

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8 Appendix

ID Task Mode

Task Name Duration Start Finish Predecessors

1 NASA Student Launch Timeline

169 days Wed 8/23/17 Mon 4/16/18

2 RFP Available 0 days Wed 8/23/17 Wed 8/23/17

3 Proposal/ Statement of Work

20 days Thu 8/24/17 Wed 9/20/17

4 Research Proposal Requirements

3 days Thu 8/24/17 Mon 8/28/17

5 Begin Writing Proposal

17 days Sun 8/27/17 Mon 9/18/17

6 Review/ Edit Proposal

1 day Tue 9/19/17 Tue 9/19/17

7 Submit Proposal

1 day Wed 9/20/17 Wed 9/20/17

8 Awarded Proposals Announced

0 days Fri 10/6/17 Fri 10/6/17

9 Preliminary Design Review

35 days Thu 10/12/17 Wed 11/29/17

10 Kickoff and PDRQ&A

0 days Thu 10/12/17 Thu 10/12/17

11 Research PDR Requirements

2 days Thu 10/12/17 Fri 10/13/17 10

12 Begin Writing PDR

10 days Mon 10/16/17 Fri 10/27/17 11,8

13 Build PDR Presentation

10 days Mon 10/16/17 Fri 10/27/17 11

14 Review and EditPDR Report andPresentation

5 days Mon 10/30/17 Fri 11/3/17 12

15 Update/ Establish Web Presence

5 days Thu 10/26/17 Wed 11/1/17

16 Send Website URL to NASA

1 day Thu 11/2/17 Thu 11/2/17 15

17 Upload PDR Report, Presentation and Fly Sheet toWebsite

0 days Mon 11/6/17 Mon 11/6/17 15,14

18 PDR Video Teleconferences

18 days Mon 11/6/17 Wed 11/29/17 17

19 Critical Design Review

42 days Wed 12/6/17 Thu 2/1/18

20 CDR Q&A 1 day Wed 12/6/17 Wed 12/6/17

21 Research CDR Requirements

2 days Thu 12/7/17 Fri 12/8/17 20

8/23

10/6

10/12

11/6

7/16 7/30 8/13 8/27 9/10 9/24 10/8 10/22 11/5 11/19 12/3 12/17 12/31 1/14 1/28 2/11 2/25 3/11 3/25 4/8 4/22August 1 September 1 October 1 November 1 December 1 January 1 February 1 March 1 April 1 May

Task

Split

Milestone

Summary

Project Summary

Inactive Task

Inactive Milestone

Inactive Summary

Manual Task

Duration-only

Manual Summary Rollup

Manual Summary

Start-only

Finish-only

External Tasks

External Milestone

Deadline

Progress

Manual Progress

Project Plan

Page 1

Project: USLI_COMP3_ProjectPlDate: Tue 9/19/17

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ID Task Mode


22 Begin Writing CDR

22 days Mon 12/11/17 Tue 1/9/18 21

23 Begin CDR Presentation

22 days Mon 12/11/17 Tue 1/9/18 21

24 Review and EditCDR Report andPresentation

5 days Wed 1/10/18 Tue 1/16/18 23,22

25 CDR Report, Presentation and Flysheet onWebsite

0 days Wed 1/17/18 Wed 1/17/18 24

26 CDR Video Teleconferences

12 days Wed 1/17/18 Thu 2/1/18 25

27 Flight Readiness Review

49 days Wed 2/7/18 Mon 4/16/18

28 FRR Q&A 1 day Wed 2/7/18 Wed 2/7/18

29 Research FRR Requirements

2 days Thu 2/8/18 Fri 2/9/18 28

30 Begin Writing FRR

15 days Mon 2/12/18 Fri 3/2/18 29

31 Begin FRR Presentation

14 days Mon 3/5/18 Thu 3/22/18 30

32 Review and EditFRR Report and Presentation

4 days Fri 3/23/18 Wed 3/28/18 30,31

33 Upload FRR Report, Presentation and Flysheet onWebsite

0 days Thu 3/29/18 Thu 3/29/18 32

34 FRR Video Teleconferences

13 days Thu 3/29/18 Mon 4/16/18 33

35 Competition Launch

2 days Wed 4/4/18 Fri 4/6/18

36 Team Travels toHuntsville, AL

1 day Wed 4/4/18 Wed 4/4/18

37 Launch Readiness Reviews

1 day Thu 4/5/18 Thu 4/5/18 36

38 Official Launch Week Kickoff, LRRs, Launch Week Activities

0 days Thu 4/5/18 Thu 4/5/18 36

39 Launch Week Activities

1 day Thu 4/5/18 Thu 4/5/18 36

1/17

3/29

4/5


Task

Split

Milestone

Summary

Project Summary

Inactive Task

Inactive Milestone

Inactive Summary

Manual Task

Duration-only


Manual Summary

Start-only

Finish-only

External Tasks

External Milestone

Deadline

Progress

Manual Progress

Project Plan

Page 2


140

ID Task Mode


40 Launch Day 1 day Thu 4/5/18 Thu 4/5/18 36

41 Banquet 1 day Thu 4/5/18 Thu 4/5/18 36

42 Backup Launch Day

1 day Thu 4/5/18 Thu 4/5/18 36

43 Post‐Launch Assessment Review uploaded to website

0 days Fri 4/6/18 Fri 4/6/18 36,40

44 Payload Timeline 163 days Wed 8/23/17 Sat 4/7/18

45 Payload Design 36 days Wed 8/23/17 Wed 10/11/17

46 Conceptual Design

21 days Wed 8/23/17 Wed 9/20/17

47 Research Payload Design Alternatives

6 days Thu 9/21/17 Thu 9/28/17 46

48 Payload Design Decision

1 day Fri 9/29/17 Fri 9/29/17 47

49 Generate Preliminary CADModel

7 days Mon 10/2/17 Tue 10/10/17 46,47,48

50 Conceptual Design Review

1 day Wed 10/11/17 Wed 10/11/17 46,47,48,49

51 Payload Build 137 days Thu 9/21/17 Fri 3/30/18

52 Payload Design Analysis

32 days Thu 9/21/17 Fri 11/3/17

53 Initial Testing 32 days Mon 11/6/17 Tue 12/19/17 52

54 Rover Prototype Design

27 days Wed 12/20/17 Thu 1/25/18 53

55 Full Scale Design of Payload/Rover

41 days Thu 1/25/18 Thu 3/22/18 54

56 Provide Detailed CAD Model

7 days Thu 3/22/18 Fri 3/30/18 53,54,55

57 Final Assembly of Payload/Rover

6 days Thu 3/22/18 Thu 3/29/18 55,56

58 Payload Testing/Launch

56 days Fri 1/19/18 Sat 4/7/18

59 Prototype Payload/Rover Testing

5 days Fri 1/19/18 Thu 1/25/18 57

4/6


Task

Split

Milestone

Summary

Project Summary

Inactive Task

Inactive Milestone

Inactive Summary

Manual Task

Duration-only


Manual Summary

Start-only

Finish-only

External Tasks

External Milestone

Deadline

Progress

Manual Progress

Project Plan

Page 3


141

ID Task Mode


60 Full Scale Payload/Rover Testing

6 days Thu 3/22/18 Thu 3/29/18 59

61 Launch Day 1 day Sat 4/7/18 Sat 4/7/18 60

62 Vehicle Timeline 163 days Wed 8/23/17 Sat 4/7/18

63 Vehicle Design 30 days Wed 8/23/17 Tue 10/3/17

64 Conceptual Design

21 days Wed 8/23/17 Wed 9/20/17

65 Research Carbon Fiber Machining Methods

7 days Thu 9/21/17 Fri 9/29/17 46

66 Research 3D Printed Material Properties

7 days Thu 9/21/17 Fri 9/29/17 46

67 Vehicle Design Decision

1 day Mon 10/2/17 Mon 10/2/17 47,65

68 Conceptual Design Review

1 day Tue 10/3/17 Tue 10/3/17 46,65,66,67

69 Vehicle Build 131 days Thu 9/21/17 Thu 3/22/18

70 Design Analysis 32 days Thu 9/21/17 Fri 11/3/17

71 Initial Testing 32 days Mon 11/6/17 Tue 12/19/17 70

72 Vehicle Design Analysis

32 days Wed 12/20/17 Thu 2/1/18 71

73 Subscale Design 10 days Fri 2/2/18 Thu 2/15/18 72

74 Full Scale Design

16 days Thu 2/15/18 Thu 3/8/18 73

75 Full Scale Assembly

11 days Thu 3/8/18 Thu 3/22/18 73,72

76 Vehicle Testing/Launch

42 days Thu 2/8/18 Sat 4/7/18

77 Subscale Testing

21 days Thu 2/8/18 Thu 3/8/18 73

78 Full Scale Testing

6 days Thu 3/22/18 Thu 3/29/18 77

79 Full Scale Launch

1 day Sat 4/7/18 Sat 4/7/18 78


Task

Split

Milestone

Summary

Project Summary

Inactive Task

Inactive Milestone

Inactive Summary

Manual Task

Duration-only


Manual Summary

Start-only

Finish-only

External Tasks

External Milestone

Deadline

Progress

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Project Plan

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