critical design review - tarleton.educritical design review i note to reader: to facilitate the...

[Type text] I) Summary of CDR Report [Type text]

Critical Design Review

2012-2013 NASA USLI

“Research is what I’m doing when I don’t know what I’m doing.” - Werner Von Braun

Tarleton State University Critical Design Review

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Note to reader:

To facilitate the reading of the Critical Design Review, we have mirrored the Student

Launch Project Statement of Work. In the body of the CDR, you will find extensive detail

in the design of our SMD payload. The payload’s features are threefold; atmospheric

data gathering sensors, a self-leveling camera system, and a video camera. One of the

two major strengths of our payload design is the originality of our autonomous real-time

camera orientation system (ARTCOS). The other major strength can be found in the

originality of our self-designed Printed Circuit Board layouts. This feature alone

represents over 150 man hours of work. Along with space and power efficiencies, the

PCBs provide major enhancement of the signal integrity of the sensor data. For ease of

reading, you will find documents such as itemized final build budget and launch

procedures moved to the appendix along with Sensor and Material Safety Data sheets.

We have enjoyed the challenges presented in the writing of this document and submit it

for your review.


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Table of Contents I) Summary of CDR Report ............................................................................................. 1

Team Summary ........................................................................................................... 1

Launch Vehicle Summary ............................................................................................ 1

Payload Summary ........................................................................................................ 1

II) Changes Made since PDR .......................................................................................... 2

III) Vehicle Criteria ........................................................................................................... 5

Design and Verification of Launch Vehicle ................................................................... 5

Launch Vehicle Mission Statement ........................................................................... 5

Mission Success Criteria .......................................................................................... 5

Review the design at a system level ......................................................................... 9

Verification of System Level Functional Requirements ........................................... 19

Approach to Workmanship ..................................................................................... 23

Additional Planned Component, Functional, or Static Testing ................................ 24

Status and Plans of remaining manufacturing and assembly ................................. 24

Discuss the integrity of design ................................................................................ 24

Safety and Failure Analysis .................................................................................... 40

Subscale Flight Results ............................................................................................. 40

Subscale Flight Results ............................................................................................. 40

Flight Data .............................................................................................................. 41

Impact on Design Summary ................................................................................... 79

Recovery Subsystem ................................................................................................. 81

Physical Components ............................................................................................. 81

Electrical Components ............................................................................................ 86

Kinetic Energy......................................................................................................... 96

Test Results ............................................................................................................ 98

Safety and Failure Analysis .................................................................................. 102

Mission Performance Predictions ............................................................................. 116

Mission Performance Criterion.............................................................................. 116

Payload Integration .................................................................................................. 121

Payload Integration Plan ....................................................................................... 121

Payload Installation and Removal ......................................................................... 124

Payload Interface Dimensions .............................................................................. 126


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Payload Element Compatibility ............................................................................. 128

Simplicity of Integration Procedure ....................................................................... 128

Launch Concerns and Operation Procedures .......................................................... 128

Launch procedures ............................................................................................... 128

Pre-launch Checklists and Procedures: ................................................................ 128

Safety Materials Checklist .................................................................................... 128

Structure Preparation: ........................................................................................... 129

Recovery Procedures: .......................................................................................... 130

Motor Preparation: ................................................................................................ 132

Launch Checklist and Procedures ........................................................................ 133

Troubleshooting: ................................................................................................... 134

In-Flight Inspection ............................................................................................... 135

Post-Flight Inspection ........................................................................................... 135

Travel .................................................................................................................... 135

Safety and Environment ........................................................................................... 137

Failure Modes ....................................................................................................... 137

Hazard Analysis .................................................................................................... 142

Environment.......................................................................................................... 146

IV) Payload Criteria ..................................................................................................... 149

Testing and Design of Payload Experiment ............................................................. 149

Design Review at a System Level ........................................................................ 149

System Level Functional Requirements ............................................................... 161

Approach to Workmanship ................................................................................... 163

Test Plan of Components and Functionality ......................................................... 163

Status and Plans of Remaining Manufacturing and Assembly ............................. 189

Integration Plan..................................................................................................... 192

Precision of Instrumentation and Repeatability of Measurements ........................ 194

Safety and Failure Analysis .................................................................................. 197

Uniqueness and Significance ............................................................................... 201

Suitable Level of Challenge .................................................................................. 201

Science Value .......................................................................................................... 202

Experimental Logic, Approach, and Method of Investigation ................................ 203

Relevance of Expected Data and Accuracy/Error Analysis .................................. 204

Safety and Environment ........................................................................................... 205


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The Safety Officer ................................................................................................. 205

Failure Modes ....................................................................................................... 206

Hazard Analysis .................................................................................................... 209

Environment.......................................................................................................... 212

V) Project Plan ............................................................................................................ 213

Budget Summary ..................................................................................................... 213

Funding Plan......................................................................................................... 225

Timeline ................................................................................................................... 225

Testing Timeline ................................................................................................... 227

Outreach Timeline ................................................................................................ 228

Education plan ......................................................................................................... 228

Outreach Plan ....................................................................................................... 228

Accomplished Educational Outreach .................................................................... 231

VI) Conclusion ............................................................................................................. 243


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Table of Figures Figure 1: Launch Vehicle Specifications .......................................................................... 9 Figure 2: Upper body Airframe ...................................................................................... 10 Figure 3: Clear Payload Housing .................................................................................. 11 Figure 4: Booster Section .............................................................................................. 12 Figure 5: Epoxy Strength Testing .................................................................................. 13 Figure 6: Acrylic Compression Testing .......................................................................... 14 Figure 7: Fin Testing Set up .......................................................................................... 15 Figure 8: Fin Detachment from the Motor Tube ............................................................ 16 Figure 9: Cesaroni L1720 Motor Thrust Curve from ThrustCurve.org ........................... 17

Figure 10: L1720-WT Thrust Curve from Cesaroni ....................................................... 18 Figure 11: Tarleton Aeronautical Team's Generated Thrust Curve ............................... 19 Figure 12: Fin Dimensions ............................................................................................ 25

Figure 13: Booster Assembly Steps 1-4 ........................................................................ 27 Figure 14: Booster Assembly Steps 5-8 ........................................................................ 28 Figure 15: Booster Assembly Steps 9-12 ...................................................................... 29

Figure 16: Coupler Assembly Procedure ....................................................................... 30 Figure 17: Avionics Assembly Steps 1-3 ....................................................................... 31 Figure 18: Avionics Assembly Steps 4-6 ....................................................................... 32

Figure 19: Payload Assembly ........................................................................................ 33 Figure 20: Ballast System Assembly ............................................................................. 34

Figure 21: Positive Motor Retainer ................................................................................ 35 Figure 22: Launch Vehicle Illustration ........................................................................... 36 Figure 23: Test Flight One Vehicle ................................................................................ 43

Figure 24: Test Flight One Simulation ........................................................................... 44

Figure 25: Raven3 Flight Data ...................................................................................... 45 Figure 26: Test Flight Two Vehicle ................................................................................ 46 Figure 27: Simulated Flight Two Data ........................................................................... 47

Figure 28: Raven3 Flight Data ...................................................................................... 48 Figure 29: Test Flight Three Vehicle ............................................................................. 48

Figure 30: Simulated Test Flight Three ......................................................................... 50 Figure 31: Test Flight Four Vehicle ............................................................................... 51 Figure 32: Simulated Test Flight Four Data ................................................................... 52 Figure 33: Raven3 Flight Data ...................................................................................... 53

Figure 34: Test Flight Five Vehicle ................................................................................ 53 Figure 35: Raven3 Test Flight Five Data ....................................................................... 54 Figure 36: Test Flight Six Vehicle .................................................................................. 55 Figure 37: Simulated Flight Six Data ............................................................................. 56

Figure 38: Test Flight Seven Vehicle............................................................................. 57 Figure 39: Test Flight Seven Simulated Data ................................................................ 58 Figure 40: Raven3 Test Flight Seven Data ................................................................... 59

Figure 41: Test Flight Eight Vehicle .............................................................................. 60 Figure 42: Simulated Flight Eight Data .......................................................................... 61 Figure 43: Test Flight Nine Vehicle ............................................................................... 62 Figure 44: Simulated Flight Nine Data........................................................................... 63 Figure 45: Flight Nine GPS Data ................................................................................... 63


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Figure 46: Test Flight Ten Vehicle ................................................................................ 64

Figure 47: Simulated Flight Ten Data ............................................................................ 65 Figure 48: Flight Ten GPS Data .................................................................................... 66 Figure 49: Test Flight Eleven Vehicle ............................................................................ 67 Figure 50: Simulated Flight Eleven Data ....................................................................... 68 Figure 51: Test Flight Eleven GPS Data ....................................................................... 68

Figure 52: Test Flight Twelve Vehicle ........................................................................... 69 Figure 53: Simulated Test Flight Twelve ....................................................................... 70 Figure 54: Raven3 Test Flight Twelve Data .................................................................. 71 Figure 55: Test Flight Thirteen Vehicle.......................................................................... 71 Figure 56: Test Flight Thirteen Simulation ..................................................................... 72

Figure 57: Test Flight Thirteen Stratologger Data ......................................................... 73 Figure 58: Test Flight Fourteen Vehicle ........................................................................ 74

Figure 59: Test Flight Fourteen Simulation ................................................................... 75

Figure 60: Test Flight Fourteen Stratologger Data ........................................................ 76 Figure 61: Test Flight Fifteen Vehicle ............................................................................ 77 Figure 62: Test Flight Fifteen Simulation ....................................................................... 78 Figure 63: Test Flight Fifteen Stratologger Data ........................................................... 78

Figure 64: Ejection Canister .......................................................................................... 80 Figure 65: 3F Black Powder .......................................................................................... 80

Figure 66: Astro 320 GPS System ................................................................................ 80 Figure 67: SkyAngle XXLarge Deployment Freebag ..................................................... 82 Figure 68: Main Parachute Attachment Scheme ........................................................... 83

Figure 69: Attachment Scheme to Couplers .................................................................. 84 Figure 70: Drogue Parachute Attachment Scheme ....................................................... 85

Figure 71: Altimeter Electronics Schematics ................................................................. 87

Figure 72: Raven3 Software Flow Diagram ................................................................... 89

Figure 73: Stratologger Software Flow Diagram ........................................................... 91 Figure 74: Example Drogue/Main Avionics Bay ............................................................ 93

Figure 75: Drawing of Avionics Sleds ............................................................................ 94 Figure 76: GPS Software Flow Diagram ....................................................................... 95 Figure 77: Launch Vehicle Prototype ............................................................................ 96

Figure 78: Final Vehicle Simulation ............................................................................... 96 Figure 79: Final Vehicle Simulation ............................................................................. 117 Figure 80: Input Parameters for Final Simulation ........................................................ 118

Figure 81: L1720-WT Actual Thrust Curve .................................................................. 118 Figure 82: Rear Payload Bulkhead to Frame Connection ........................................... 122 Figure 83: Telemetry Verification GUI ......................................................................... 123

Figure 84: SMD Payload ............................................................................................. 124 Figure 85: SMD Payload attached with Avionic Bays. ................................................. 125 Figure 86: Aluminum Angle ......................................................................................... 127 Figure 87: Altimeter Wiring Diagrams .......................................................................... 131

Figure 88: Materials and Components (Image obtained from the Cesaroni Pro 75 mm Motor Assembly Kit Instructions) ................................................................................. 133 Figure 89: Payload ...................................................................................................... 149

Figure 90: Upper Payload Circuit Boards .................................................................... 150


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Figure 91: UV Sensor Mounting .................................................................................. 150

Figure 92: ARTCOS .................................................................................................... 151 Figure 93: Test Flight Data .......................................................................................... 153 Figure 94: Test Flight Humidity Data ........................................................................... 154 Figure 95: Launch Pad Humidity Data......................................................................... 155 Figure 96: Test Flight Temperature Data .................................................................... 156

Figure 97: Launch Pad Temperature Data .................................................................. 156 Figure 98: Correlation between Temperature and Humidity ........................................ 157 Figure 99: Test Flight Pressure Data........................................................................... 158 Figure 100: Test Flight Altitude Data ........................................................................... 158 Figure 101: Test Flight GPS Data ............................................................................... 159

Figure 102: Test Flight Solar Irradiance Data .............................................................. 160 Figure 103: ARTCOS Image ....................................................................................... 161

Figure 104: BMP 180 Pressure Sensor Wiring ............................................................ 164

Figure 105: BMP 180 Software Flowchart ................................................................... 165 Figure 106: TSL2561 Pyranometer Wiring .................................................................. 166 Figure 107: TSL2561 Pseudo Code ............................................................................ 167 Figure 108: TSL2561 Lux Conversion Factors ............................................................ 167

Figure 109: BMP 180 and TSL2561 Wiring ................................................................. 168 Figure 110: HIH4030 Humidity Sensor Wiring ............................................................. 168

Figure 111: HIH4030 Software .................................................................................... 169 Figure 112: HIH4030, BMP180, and TSL2561 Wiring ................................................. 169 Figure 113: HH10D Humidity Sensor Wiring ............................................................... 170

Figure 114: HH10D Humidity Calculation Algorithm .................................................... 170 Figure 115: HH10D, HIH4030, BMP180, and TSL2561 Wiring ................................... 171

Figure 116: SU100 Testing ......................................................................................... 172

Figure 117: SU100 UV Sensor Wiring ......................................................................... 172

Figure 118: SU100 Software ....................................................................................... 173 Figure 119: GPS Wiring .............................................................................................. 174

Figure 120: MicroSD Wiring ........................................................................................ 175 Figure 121: XBee Wireless Transmitter Wiring ........................................................... 176 Figure 122: Digi Technical Support Forum Post .......................................................... 177

Figure 123: De-Soldering LED from XBee Adapter ..................................................... 177 Figure 124: XBee Range Test ..................................................................................... 178 Figure 125: Ground Station GUI .................................................................................. 179

Figure 126: ADGS Wiring Schematic .......................................................................... 180 Figure 127: VC0706 Camera Wiring ........................................................................... 181 Figure 128: VC0706 Configuration GUI ....................................................................... 182

Figure 129: ARTCOS Mounting .................................................................................. 183 Figure 130: ARTCOS Mounting .................................................................................. 184 Figure 131: ARTCOS Orientation Algorithm ................................................................ 184 Figure 132: ARTCOS Wiring Schematic ..................................................................... 186

Figure 133: Payload Block Diagram ............................................................................ 188 Figure 134: Breakout Board Compatible PCB ............................................................. 190 Figure 135: Surface Mount PCB ................................................................................. 191

Figure 136: Bulkhead Aluminum Frame Interface ....................................................... 192


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Figure 137: Bulkhead Recessed Slot .......................................................................... 193

Figure 138: Telemetry Verification GUI ....................................................................... 193 Figure 139: SU-100 Spectral Response ...................................................................... 195 Figure 140: SP-110 Spectral Response ...................................................................... 196 Figure 141: Clean Room ............................................................................................. 197 Figure 142: ARTCOS Epoxy Mounting Failure ............................................................ 198

Figure 143: Post-Flight Payload .................................................................................. 198 Figure 144: GPS Mounting Failure .............................................................................. 199 Figure 145: PCB Board ............................................................................................... 200 Figure 146: Self-Leveling Camera System .................................................................. 201 Figure 147: Allocated Funds ....................................................................................... 213

Figure 148: Budget Status ........................................................................................... 214 Figure 149: Vehicle Budget Status .............................................................................. 215

Figure 150: Payload Budget Status ............................................................................. 215

Figure 151: Propulsion Budget Status ......................................................................... 216 Figure 152: Outreach Budget Status ........................................................................... 216 Figure 153: Early Funding ........................................................................................... 225 Figure 154: Project Timeline ....................................................................................... 226

Figure 155: Testing Gantt Chart .................................................................................. 227 Figure 156: Outreach Timeline .................................................................................... 228

Figure 157: Acton Middle School ................................................................................ 229 Figure 158: Team Members Educate and Entertain Acton Students .......................... 231 Figure 159: Subject Interest ........................................................................................ 232

Figure 160: Presentation Learning Outcomes ............................................................. 233 Figure 161: Favorite Part ............................................................................................. 234

Figure 162: Students won NASA stickers for answering questions ............................. 237

Figure 163: Interactive Physiics at Morgan Mill ........................................................... 238

Figure 164: Preparing to Launch at BluffDale ............................................................. 239 Figure 165: Students Learning at the Recovery Station at Dublin Middle School ....... 242

Figure 166: Students Enjoying the Art Station, Decorating Parachutes ...................... 242


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Index of Tables Table 1: Vehicle Size and Mass ...................................................................................... 1 Table 2: Experiment Summary ........................................................................................ 1 Table 3: Changes Made to Vehicle Criteria ..................................................................... 2 Table 4: Changes Made to Payload Criteria .................................................................... 3 Table 5: Project Milestones Continued ............................................................................ 8 Table 6: Fin Force Resistance ...................................................................................... 15 Table 7: Motor Specifications ........................................................................................ 17 Table 8: Vehicle Verification Table ................................................................................ 23

Table 9: Mass Summary ............................................................................................... 38 Table 10: Mass by Subsection ...................................................................................... 40 Table 11: Flight Data ..................................................................................................... 42

Table 12: Test Flight One Conditions ............................................................................ 43 Table 13: Test Flight Two Conditions ............................................................................ 46 Table 14: Test Flight Three Conditions ......................................................................... 49

Table 15: Test Flight Four Conditions ........................................................................... 51 Table 16: Test Flight Five Conditions ............................................................................ 54 Table 17: Test Flight Six Conditions .............................................................................. 56

Table 18: Test Flight Seven Conditions ......................................................................... 58 Table 19: Test Flight Eight Conditions........................................................................... 60

Table 20: Test Flight Nine Conditions ........................................................................... 62 Table 21: Test Flight Ten Conditions............................................................................. 65 Table 22: Test Flight Eleven Conditions ........................................................................ 67

Table 23: Test Flight Twelve Conditions ....................................................................... 69

Table 24: Test Flight Thirteen Conditions ...................................................................... 72 Table 25: Test Flight Fourteen Conditions .................................................................... 75 Table 26: Test Flight Fifteen Conditions ........................................................................ 77

Table 27: Kinetic Energy Summary ............................................................................... 97 Table 28: Static Tests .................................................................................................. 102

Table 29: Safety and Failure Analysis 11-30-12 .......................................................... 103 Table 30: Safety and Failure Analysis 12-5-12 ............................................................ 104 Table 31: Safety and Failure Analysis 12-7-12 ............................................................ 105 Table 32: Safety and Failure Analysis 12-8-12 ............................................................ 106

Table 33: Safety and Failure Analysis 12-14-12 .......................................................... 107 Table 34: Safety and Failure Analysis 12-15-12 .......................................................... 108 Table 35: Safety and Failure Analysis 12-15-12 .......................................................... 109 Table 36: Safety and Failure Analysis 12-19-12 .......................................................... 110

Table 37: Safety and Failure Analysis 12-19-12 .......................................................... 111 Table 38: Safety and Failure Analysis 12-21-12 .......................................................... 112 Table 40: Safety and Failure Analysis 1-5-13 .............................................................. 114

Table 41: Safety and Failure Analysis 1-6-13 .............................................................. 115 Table 42: Safety and Failure Analysis 1-7-13 .............................................................. 116 Table 43: Calculated versus Simulated CG and CP Measurements ........................... 121 Table 44: Payload Preparation Steps .......................................................................... 124 Table 45: Payload Integration Steps ........................................................................... 126


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Table 46: Payload Framework Dimensions ................................................................. 127

Table 47: Potential Failure Modes for Design of the Vehicle ....................................... 138 Table 48: Potential Failure Modes during Payload Integration .................................... 139 Table 50: Potential Hazards to Personnel ................................................................... 144 Table 51: Summary of Legal Risks ............................................................................. 146 Table 52: Effects of Materials used in Construction and Launch ................................. 147

Table 53: Environmental Factors ................................................................................ 148 Table 54: Payload Functional Requirements ............................................................... 163 Table 55: XBee XSC S3B Specifications .................................................................... 176 Table 56: Payload Components and Qualities ............................................................ 189 Table 57: Payload Preparation Steps .......................................................................... 194

Table 58: Payload Sensor Precision ........................................................................... 196 Table 60: Potential Failure Modes during Payload Integration .................................... 206

Table 61: Potential Failure Modes during Launch ....................................................... 209

Table 62: Potential Hazards to Personnel ................................................................... 211 Table 63: Preliminary Budget Summary ...................................................................... 213 Table 64: Structure/Propulsion System Budget ........................................................... 218 Table 65: Recovery System Budget ............................................................................ 219

Table 66: Payload Budget (Through-Hole PCB) ......................................................... 221 Table 67: Payload Budget (Surface Mount PCB) ........................................................ 224

Table 68: Accomplished Educational Outreach ........................................................... 232 Table 70: Favorite Part ................................................................................................ 234 Table 71: Educational Outreach Stations .................................................................... 237

Table 72: Educational Outreach Stations .................................................................... 241

I) Summary of CDR Report

I) Summary of CDR Report

Team Summary Tarleton Aeronautical Team Tarleton State University Box T-0470 Stephenville, Texas 76402 Team Mentor: Pat Gordzelik. Past and Present Credentials: Tripoli Amarillo #92 Board of Directors Member, Technical Advisor Panel Panhandle of Texas Rocketry Society Inc. – Founder, President, Prefect TRA 5746 L3 NAR 70807 L3CC Committee Chair Married to Lauretta Gordzelik, TRA 7217, L2.

Launch Vehicle Summary Size and Mass

Length 109.25 inches

Outer Diameter 5.525 inches

Mass 37.125 pounds

Motor

Selection Cesaroni L1720-WT-P

Recovery

Drogue 24” Silicone Coated Rip stop Nylon Parachute, Apogee Deployment

Main 120” Silicone Coated Rip stop Nylon Parachute, 500 foot AGL Deployment

Avionics Primary Featherweight Raven3 Altimeter, Backup PerfectFlite Stratologger Altimeter, and Garmin GPS Tracking

Rail Size

Rail 1010 Aluminum

Milestone Review Flysheet – see Appendix B

Table 1: Vehicle Size and Mass

Payload Summary

Title Experiment

Science Mission Directorate (SMD) Payload

Sponsored by NASA; Gather Atmospheric and GPS Data, Autonomously Orientate Photographic Camera, Capture Video for Public Outreach

Table 2: Experiment Summary


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II) Changes Made since PDR


Changes Made to Vehicle Criteria Structure Rationale

Drogue avionics relocated to rear coupler from booster section

Ease of construction and accessibility

Ballast system relocated from nose cone to upper body airframe

Ease of construction

Coupler port hole rings added Eliminates need of lining up port holes through body and coupler

Centering ring in fin tab relocated to front of fin tab

Ease of construction

Bulkhead at upper end of motor tube replaced with centering ring

Allows access to anchor point on motor housing for shock cord

Centering ring added to lower end of motor tube

Used to secure motor retaining ring

Nose cone length changed from 7.5 inch to 8.5 inch

Manufactured at this length

Payload bulkheads reduced to 1 inch thickness from 2 inch

Reduces weight without compromising integrity

Payload bulkheads epoxied to couplers Creates seals between avionics bays and payload compartment

Coupler bulkheads added to avionics bay lids

Reinforce bay lids in event of failed main chute deployment

Recovery Rationale

U-bolts changed to welded eyebolts Reduce weight without compromising integrity

Removed deployment bag for drogue chute Unnecessary

Changed to XL ejection canisters Reduce friction by allowing chute more room lengthways

Changed to 3F black powder from 4F Availability

Switched to 3 portholes in each bay(sizing in SFR)

Following recommendations of manufacturers

LEDs added to allow visible confirmation of altimeter activations

Eliminates need of audio confirmation of altimeters

GPS relocated from nose cone to drogue shock cord

Easy to secure

Changed from Big Red Bee GPS to Garmin Astro DC40/320 system

Easy to implement

Increased deployment bag size of main chute to XXL from XL

Increase ease of deployment

Avionics Bays changed to standard sled containing design

Modular and easy to access

Table 3: Changes Made to Vehicle Criteria


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Recovery Rationale Tubular Kevlar shock cords reduced to .25 inch from .5 inch

Weight reduction and increase space in upper body section

Swivel removed from drogue chute Unnecessary

Backup shock cord (.25 inch, 4.5 feet) epoxied along motor tube

Safety/Redundancy

Shock cord of main chute length changed to 40 feet from 20 feet

limit multiple section collision

Shock cord of drogue chute length changed to 20 feet from 25 feet

limit multiple section collision

Increased size of all ejection charges Necessary for proper separation

Secondary charges have .4 grams more black powder than primary

Simple Logic

Switched to cross-form rip stop nylon parachutes

Availability and durability

Separation now occurs between the upper body airframe and payload section instead of at the nose cone

Allows for easier transportation and preparation

Changes Made to Payload Criteria Payload Rationale

Rail changed to .5 inch x .5 inch x .0625 inch aluminum angle from .5 inch x .125 inch flat aluminum

Add rigidity and reduce weight

Payload centered in payload section Allows avionics bays to have uniform dimensions

Rear coupler removable to access payload Easier to access

Port holes changed to 5 evenly spaced .25 inch holes

Provide adequate ventilation

Reduced 9V battery count from 8 to 4 Unnecessary

MS5611 pressure sensor removed Availability

HH10D humidity sensor removed Simplify circuit

Video camera changed to Keyfob from VCC-003-MUVI-BLK

Availability and cost

Arduino Mini added to ARTCOS Dedicated for video processing

Moved HIH4030 to ARTCOS The reference voltage required by the SU100 and SP110

Changed to buck converters from linear regulators

Power efficiency

Mounted ARTCOS to fiberglass brackets More secure installation

BMP180 placed in between circuit boards Shields the sensor from light

Magnetic switch connected to relay to activate entire payload

Simplify and speed up launch preparation

Table 4: Changes Made to Payload Criteria


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The team made no significant changes to the project plan.


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III) Vehicle Criteria


Design and Verification of Launch Vehicle

Launch Vehicle Mission Statement The mission is to design, build, and launch a reusable vehicle capable of delivering a payload to 5,280 feet above ground level (AGL). The vehicle will carry a barometric altimeter for official scoring and the Science Mission Directorate (SMD) payload. The design of the vehicle ensures a subsonic flight and must be recoverable and reusable on the day of the official launch. The launch vehicle meets the customer prescribed requirements set forth in the Statement of Work (SOW) of the NASA 2012-2013 Student Launch Projects (SLP) handbook.

Launch Vehicle Requirements The vehicle adheres to the following primary requirements. The complete list of requirements is in the Vehicle Verification Table (Table 8).

Vehicle shall carry a scientific or engineering payload. (Requirement 1.1)

Vehicle shall reach an apogee altitude of one mile above ground level. (Requirement 1.1)

Vehicle shall carry one official scoring altimeter. (Requirement 1.2)

Vehicle must remain subsonic from launch until landing. (Requirement 1.3)

Vehicle must be recoverable within a 2500 foot radius from the launch pad and reusable on the day of the official launch. (Requirement 2.3)

Vehicle must use a commercially available APCP motor with no more than 5,120 Newton-seconds of impulse. (Requirement 1.11, 1.12)

Mission Success Criteria The project defines the mission as a vehicle flight with a payload onboard where both the vehicle and SMD payload are recovered and able to be reused on the day of the official launch. Moreover, the vehicle will not exceed 5,600 feet of altitude, and the official scoring altimeter will be intact, audible, and report altitude. The recovery system stages a deployment of the drogue parachute at apogee and deploys the main parachute at 700 feet. After apogee and descent, the entire vehicle lands within 2,500 feet of the launch pad. If the above conditions are met, the mission will be considered partially successful in that requirements have been met by the vehicle design. However, because the actual altitude of the vehicle at apogee is scored based on comparison to one mile above ground level, a successful mission would be warranted only if the aforementioned


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conditions are met and an apogee of exactly 5,280 feet is achieved, plus or minus 0.1% plus 1 foot due to precision of the scoring altimeter.

Major Milestone Schedule Significant milestones of the project from initiation to final launch day and announcement of contest winners are detailed in Table 5. Each date has a description as well as the completion status of each event up to the time that the CDR is submitted (Jan 14). Additionally, the type of event is specified as either being provided by the NASA USLI SOW, a test date, a deadline for verification, or a deadline for manufacturing/assembly of the vehicle.


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Date Milestone Description Status Type

8/31/12 Proposal Due Met USLI

9/27/12 Schools Notified Met USLI

10/4/12 Team Teleconference Met USLI

10/11/12 PDR Q&A Met USLI

10/22/12 Web Presence est. Met USLI

10/23/12 SMD Award 1 ($780) Met USLI

10/28/12 Subscale Dual Deployment Test Met Test Launch

10/29/12 PDR due Met USLI

11/7-16/12 PDR Presentations Met USLI

11/17/12 Dual Deployment Test Not Met Test Launch

11/30/12 Subscale Launch Met Test Launch

11/30/12 Lab Prototyping Met Verification

12/1/12 Low Altitude Flight Met Test Launch

12/3/12 Post Launch Failure Analysis Met Verification

12/3/12 Full Scale Prototype Assembly 1 Not Met Manufacturing

12/3/12 CDR Q&A Not Met USLI

12/18/12 Range Radio Testing Met Verification

12/20/12 E-match Testing Met Verification

12/22/12 Subscale Low Altitude Flight Met Test Launch

12/31/12 PCB Testing Not Met Verification

12/31/12 Programming Met Verification

1/5/13 Static Black Powder Testing Met Verification

1/5/13 Static Ejection Test Met Verification

1/5/13 Low Altitude, Full Scale Launch (w/o SMD)

Not Met Test Launch

1/5/13 Low Altitude, Full Scale Launch (w/ SMD)

Met Test Launch

1/6/13 Alternative Launch Day-Used Met Test Launch

1/6/13 Post Launch Failure Analysis Met Verification

1/7/13 Low Altitude, Full Scale Launch (w/o SMD)

Met Test Launch

1/7/13 Static Motor Test Met Verification

1/12/13 Alternative Launch Day TBD Test Launch

1/14/13 Post Launch Failure Analysis TBD Verification

1/14/13 CDR due Met USLI

1/14/13 Spring Semester Begins …. ….

1/19/13 Low Altitude, Full Scale Launch (w/ TBD Test Launch


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Date Milestone Description Status Type

SMD)


1/22/13 Full Scale Prototype Assembly 2 TBD Manufacturing

1/26/13 High Altitude, Full Scale Launch (w/o SMD)

TBD Test Launch

1/27/13 High Altitude, Full Scale Launch (w/SMD)

TBD Test Launch


2/1/13 CDR Presentations TBD USLI

2/2/13 SMD Award 2 ($1400) TBD USLI


TBD Test Launch


2/11/13 FRR Q&A TBD USLI


TBD Test Launch



TBD Test Launch


TBD Test Launch


3/1/13 Final Vehicle Assembly TBD Manufacturing

3/2/13 Final Demonstration Flight TBD Verification

3/9/13 Final Demonstration Flight (alt) TBD Verification

3/16/13 Final Demonstration Flight (alt) TBD Verification

3/18/13 FRR due TBD USLI

3/25-4/3/13 FRR Presentations TBD USLI


4/17/13 LRR Begin TBD USLI

4/18-19/13 Welcome Day TBD USLI

4/20/13 Launch Day TBD USLI

4/21/13 Launch Rain Day TBD USLI

5/6/13 PLAR due TBD USLI


5/17/13 Winners Announced TBD USLI

Table 5: Project Milestones Continued


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Review the design at a system level

Final Drawings and Specifications

The overall launch vehicle, as shown in Figure 1, is 109.25 inches long. The fin span is 15.525 inches. This includes the 5.525 inch width of the airframe. Each section of the launch vehicle will be further specified below.

Figure 1: Launch Vehicle Specifications


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The Upper Body Airframe is 28.0 inches long. This includes an 8.5 inch elliptical nose cone as shown in Figure 2. Note that the ballast system is provided in the drawing as well.

Figure 2: Upper body Airframe


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The Acrylic Housing Structure is 36 inches long as illustrated in Figure 3. The couplers will remain attached throughout the entire flight. The couplers are 11.25 inches long and have two diameters to integrate the different inside diameters of the Acrylic Housing Structure and the fiberglass airframes. The diameters are 5.373 inches for the side coupling the fiberglass airframes and 5.178 inches for the side coupling the Acrylic Housing Structure.

Figure 3: Clear Payload Housing


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The Booster Section of the vehicle is 36 inches long. Mounted inside the Booster section is a 20 inch motor mount tube as pictured in Figure 4. The motor mount tube is three inches in diameter to accommodate a 75 mm motor. 0.125 inch wide slots are cut into the Booster Section starting 1.125 inches from the bottom of the airframe and extending 9.7 inches for the fin tabs.

Figure 4: Booster Section


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Final analysis and model results, anchored to test data

After analysis of the initial test launch, it was found that the first prototype launch

vehicle had severely outgrown the motor. At 37 pounds un-ballasted and without SMD payload, the first prototype vehicle was heavier than designed. Also, considering a tilted rail, the vehicle achieved 4,992 feet in altitude. Redesign of heavier components has reduced the second prototype launch vehicle weight to approximately 34.625 pounds without the SMD payload.

Through testing, it was found that the epoxy and both airframe materials could

withstand 1,500 pounds. During the first test launch, the rocket experienced a high velocity impact which the airframe survived. This result provides confidence that the strength of the materials far exceeds the expected loads on the airframe.

Test description and results

Epoxy Test

To test the Proline 4500 epoxy, a bulkhead was epoxied into an airframe, filleting one side to simulate how the bulkheads are incorporated in the full scale rocket. A 2x3 inch block was placed on the bulkhead to simulate the mounting hardware for the recovery system. Then, using the hydraulic press pictured in Figure 5, pressure was applied to the bulk head in increments of ~10 pounds. At every 100 pound increment, the press was released, and then reapplied to that weight instantly to represent shock force. The scale used to measure the force was an airplane scale with a maximum of 1,500 pounds. The force on the bulkhead reached 1,500 pounds and held this force for 60 seconds before it was released with no sign of wear or damage.

Fiberglass Bulkhead Strength

The epoxy test also shows that the .125 inch thick flat sheet of fiberglass can hold over 1,500 pounds. Using this number and dividing by the contact area, the flat sheet of fiberglass can hold over 250 pounds per square inch.

Figure 5: Epoxy Strength Testing


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Tube Crushes Using a hydraulic press, the spare fiberglass section and the acrylic section was subjected to forces simulating the expected loads during motor thrust. To do this, a steel plate was placed on top of and below the tube and pressed in the center as shown in Figure 6. The coupler, fiberglass airframe, and acrylic airframe were all tested and each withstood the maximum weight of 1,500 pounds from the scale with no signs of wear or damage.

Fin Testing

The fins for the final vehicle are twice as thick as the fiberglass bulkhead. Thus, the shear strength of the fins is greater than 250 pounds per square inch, the minimum tested strength of the

bulkhead. To test the mounting of the fins, the fins were mounted to a tube in the same manner as the prototype build. This will also replicate the fin mounting in the final build. The tube was then secured using clamps, and the hydraulic press was used to apply weight at the point of the fin furthest from the rocket as featured in Figure 7. These forces were increased in ~10 pound increments reapplying the weight in bursts simulating shock force. With 110 pounds of force applied at 5 inches from the airframe, the press provided enough torque to fracture the epoxy bond at the motor tube and the bond from the fins to the external airframe surface as shown in Figure 8. Using τ = F x d, a torque of 550 inch-pounds is the maximum force applicable before the epoxy is compromised. Table 6 shows the force each fin can withstand when applied from different angles. Calculations were found by using τ = r F sin(ϴ) and altering the angle at which the force is applied.

Figure 6: Acrylic Compression Testing


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Angle of force on fin (degrees)

45 60 75 90

Max Force (lbs.) 156 127 114 110

Table 6: Fin Force Resistance

Figure 7: Fin Testing Set up


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Final motor selection

The selected motor is a Cesaroni L1720-WT-P. The high initial thrust helps to stabilize the rocket as it departs from the launch rail. Through simulations that take into consideration the average conditions for the launch site and date, the Cesaroni L1720-WT-P is the best choice of motors available to achieve an apogee of just less than one mile AGL.

Figure 8: Fin Detachment from the Motor Tube


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Motor Apogee (ft.)

Velocity Off Rail (ft./s)

Total Impulse

Max. Velocity (ft./s)

Average Thrust

Burn Time (s)

Thrust to Weight Ratio

Cost

Cesaroni L1720-WT-P

4852 69.8 831 lbfs (3,696 Ns)

656 394.3lbf (1,754 N)

2.15 10.6 $170.96

The Cesaroni L1720 has a total impulse of 3,696 Newton-seconds, which does not exceed a total impulse maximum of 5,120 Newton-seconds as required. The motor’s corresponding thrust curve, as calculated by ThrustCurve.org, is represented in Figure 9. As shown in the thrust curve, the motor has a fairly neutral motor burn. Average thrust for this motor is 394.3lbf = 1,754N as shown in Table 7 and marked in Figure 9. Noting that the acceleration of gravity is approximately 9.8m/s², this motor’s thrust to weight ratio is achievable by 10.6:1, which exceeds the suggested ratio of 5:1.

Table 7: Motor Specifications

Figure 9: Cesaroni L1720 Motor Thrust Curve from ThrustCurve.org


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Figure 10 is the motor thrust curve provided on the Cesaroni website. The thrust curve shape and thrust values are very similar to that from ThrustCurve.org.

Figure 11 is the actual motor thrust curve found by static testing an L1720 WT using a thrust stand. Data is collected by a thrust sensor connected to a WinDAQ analog to digital converter. This test was performed at Pat Gordzelik’s motor testing facility at P&L Ranch. Although the curve shape is very similar, the actual thrust values vary from the previous figures. This is attributed to a calibration error of the thrust sensor. If the opportunity arises, the test will be conducted again using a properly calibrated sensor.

Figure 10: L1720-WT Thrust Curve from Cesaroni


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Verification of System Level Functional Requirements

The verification plan in effect reflects how each requirement to the vehicle and recovery system satisfies its function. Requirements from the SOW are paraphrased followed by the design feature that satisfies that requirement. Ultimately, each design feature undergoes verification to ensure that it actually meets its requirements. Testing, analysis, and inspection serve as the mode of verification for each feature. A detailed Gantt chart containing test dates is in Figure 157. Table 8 gives each vehicle requirement, coupled with how it will be satisfied by the vehicle design and verified.

Figure 11: Tarleton Aeronautical Team's Generated Thrust Curve


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Requirement (SOW)

Vehicle Requirement Satisfying Design Feature

Verification Method

1.1 Vehicle shall deliver payload to 5,280 feet AGL

Cessaroni L1720-WT Testing, Analysis

1.2 Vehicle shall carry one official scoring barometric altimeter

Adept A1E, included in the SMD payload

Inspection

1.2.1 Official scoring altimeter shall report the official competition altitude via a series of beeps

Adept A1E Functionality

Testing, Inspection

1.2.2 Teams may have additional altimeters

Four additional altimeters for redundancy will be used to stage separation events as required by Recovery System

Inspection

1.2.2.1

At Launch Readiness Review, a NASA official will mark the altimeter to be used for scoring

Adept A1E can be located easily through clear acrylic body section

Inspection

1.2.2.2

At launch field, a NASA official will obtain altitude by listening to beeps reported by altimeter

Adept A1E functionality

Testing, Inspection

1.2.2.3

At launch field, all audible electronics except for scoring altimeter shall be capable to turn off

Recovery altimeters can be disabled externally via magnetic arming switch

Testing


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Requirement (SOW)


Verification Method

1.2.3.1

Official, marked altimeter cannot be damaged; must report an altitude with a series of beeps

Successfully recovery system; sufficient mounting

Testing, Inspection

1.2.3.2

Team must report to NASA official designated to record altitude with official marked altimeter on launch day

This task will be assigned to an appropriate team member

Inspection

1.2.3.3 Altimeter must not report apogee altitude of over 5,600 feet

Cessaroni L1720-WT/ Vehicle Mass

Testing, Analysis

1.3 Launch vehicle remains subsonic from launch until landing

Cesaroni L1720-WT Testing, Analysis

1.4 Vehicle must be recoverable and reusable

Successful recovery system

Testing, Inspection, Analysis

1.5 Launch vehicle shall have a maximum of four independent sections

Vehicle is composed of 3 tethered sections

Inspection

1.6 Launch vehicle shall be prepared for flight at launch site within 2 hours

Launch operations and assembly procedures

Testing, Inspection

1.7

Launch vehicle will remain launch-ready for a minimum of one hour with critical functionality

System runtime capability of at least 2 hours

Testing, Inspection, Analysis


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Requirement (SOW)


Verification Method

1.8 Vehicle shall be compatible with 8 feet long 1 inch rail (1010)

1010 rail buttons attached to vehicle body

Inspection

1.9 Launch vehicle will be launched by a standard 12 volt DC firing system

Cesaroni L1720-WT Igniter

Testing

1.10 Launch vehicle shall require no external circuitry or special equipment to initiate launch

Motor ignition only requires the 12V DC firing system

Testing

1.11 Launch vehicle shall use a commercially available, certified APCP motor

Cesaroni L1720-WT Inspection

1.12 Total impulse provided by launch vehicle will not exceed 5,120 Newton-seconds

Motor total impulse of 3695.6 N-s

Inspection

1.15

The full scale vehicle, in final flight configuration, must be successfully launched and recovered prior to FRR

Test Launch Schedule Testing

1.15.1 Vehicle and recovery system function as intended

Testing Schedule Testing

1.15.2 Payload can, but does not have to be, flown during full-scale test flight.

Payload will be flight-ready for the full-scale test flight

Testing

1.15.2.1 If payload is not flown, mass simulators shall be used to simulate payload mass

Payload mass simulator will be available, if needed

Inspection


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Requirement (SOW)


Verification Method

1.15.2.1.1

Mass simulators shall be located in same location on vehicle as the missing payload mass

Payload mass simulator will be placed in appropriate location, if needed

Inspection

1.15.2.2

Any energy management system or external changes to the surface of the vehicle shall be active in full scale flight

No energy management system; No external changes to vehicle surfaces

Not Applicable

1.15.2.3

Unmanned aerial vehicles, and/or recovery systems that control flight path of vehicle, will fly as designed during full scale demonstration flight

No unmanned aerial vehicles/flight-altering recovery systems

Not Applicable

1.15.3 Full scale motor does not have to be flown during full scale test flight

Testing schedule includes launches with full scale motor

Testing

1.15.4 Vehicle shall be flown in fully ballasted configuration during full scale test flight

Nose cone ballast system

Inspection, Testing

1.15.5

Success of full scale demonstration flight shall be documented on flight certification form, by a Level 2 or Level 3 NAR/TRA observer, and documented in FRR package

Team mentor (Pat Gordzelik)

Inspection

1.16 Maximum amount teams may spend on vehicle and payload is $5000

Budget indicates that the total spent on the vehicle and payload is less than $5000

Inspection, Analysis

Approach to Workmanship

The mission success criterion provides key goals that must be met in order for the mission to be deemed successful. Completing these goals reflects directly upon the

Table 8: Vehicle Verification Table


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degree of workmanship of the vehicle design. The team approaches workmanship by understanding the crucial importance of building the vehicle as closely to the intended design as possible. The attachment, construction, fabrication, manufacture, and assembly of all structural elements dictate the overall robustness of the vehicle design. A primary concern is building the launch vehicle such that it has a safe and stable flight. It must possess an acceptable degree of survivability so that it may be reusable on the day of the official launch. The team understands that the vehicle is only as good as its construction. Proper care and attention must be taken in the construction of the vehicle. The team benefits from around the clock access to manufacturing facilities as well as a remote testing site. Recently, the team has been invited to produce precision components at the Polen facility in Granbury, TX. This facility provides aircraft quality precision tools including: analog calipers accurate to .001 inch, a manual lathe, a mill digitally controlled to within .0001 inches, a CNC mill, PTC Creo Parametric, Solid Works, a band saw, a grinding station, a hydraulic press, and an extensive assortment of hand tools. The precision manufacturing process is overseen by facility owner Richard Keyt, a former Air Force pilot and licensed aircraft mechanic/machinist who holds a Bachelor of Science degree in Aeronautical Engineering from the University of Minnesota. Keyt was also involved in the development and testing of the parachute design during the Apollo program.

Additional Planned Component, Functional, or Static Testing

At this time, the last remaining structure test is the welded eyebolt strength test. All other testing to the structural components and the launch vehicle itself is completed. The results reveal that the launch vehicle has an acceptable design for meeting all requirements.

Status and Plans of remaining manufacturing and assembly

The PVC bulkheads, fins, couplers, and avionics sleds for the final vehicle still need to be manufactured. Some of these parts will be manufactured at the Polen facility at Pecan Plantation Airpark in Granbury, Texas. The team still requires assembly for the second prototype, and the final rocket must be constructed.

Discuss the integrity of design

Suitability of shape and fin style for mission

The selected fin shape was chosen due to the predetermined rocket shape and extensive simulation. The four-fin design provides a more corrective moment force rendering better stability. In addition, weather cocking reduces the lateral landing radius. A thickness of .125 inch has been chosen for structural integrity as dimensioned in Figure 12.


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Root Chord: 12 inches Tip Chord: 0 inches Height: 5 inches Sweep Length: 9.8 inches Sweep Angle: 63 degrees

Proper use of materials in fins, bulkheads, and structural elements

Airframe/Motor Tube/Nosecone

The upper airframe, booster section, motor tube, and nose cone will be made of fiberglass. The durability of fiberglass improves the chances of the rocket being reusable (Requirement 1.4). Fiberglass is also readily available. This material was chosen for structural components due to strength and ease of manufacturing. The consistency in component materials allows for use of a single type of epoxy manufactured especially for fiberglass. The epoxy used for attaching structural components is Proline 4500 epoxy. This will provide a strong and uniform bond throughout the vehicle’s structural components.

Bulkheads/centering rings

The bulkheads and centering rings are made of fiberglass due to the material’s superior strength to mass ratio as well as adhesive qualities between components. A .2 inch thickness was chosen to provide a larger surface area to epoxy to the airframe and provide adequate strength of mounting hardware during parachute

Figure 12: Fin Dimensions


26


deployment. The payload bulkheads are constructed out of PVC. This material was chosen because of its strength at the desired one inch thickness in addition to its availability.

Fins

The fin material is fiberglass. This choice is justified for similar reasons to that of the bulkheads and centering rings. The fins can be properly secured with the adhesive. The strength of fiberglass also allows for a greater durability upon possible high impact with the ground.

Couplers

The couplers are handmade to fit the acrylic section to both the upper body airframe and the booster section. The acrylic section has a different thickness than that of the fiberglass airframe. This is due to a difference in inside diameter, 5.25 inches versus 5.375 inches respectively. Thus, two different couplers are overlapped to make one single coupler. The couplers are made of fiberglass for ease of integration and strength.

Proper assembly procedures, attachment and alignment of elements, solid connection points, and load paths

The assembly is carried out in a multistep process which allows for each piece to be epoxied and cured. This allows sufficient time to focus on proper construction for each individual component before moving on to the next step. The follow figures are assembly procedures for the major parts of the launch vehicle.


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Figure 13: Booster Assembly Steps 1-4


28




29




30


Figure 16: Coupler Assembly Procedure


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Figure 17: Avionics Assembly Steps 1-3


32


Figure 18: Avionics Assembly Steps 4-6


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Figure 19: Payload Assembly


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Proper Attachment and Alignment of elements

The first full-scale prototype utilized 3/8 inch U-bolts for their high tensile strength to with stand the force of the main parachute deployment. Due to their size and weight, testing has begun on smaller welded eyebolts in order to lighten the overall system. It was found that .25 inch welded eyebolts will withstand the parachute deployment force. Based on testing, the lighter eyebolt will be used. The eyebolts will be welded to increase the load strength to a 400 pound working load limit. The structure will be cut to precise measurements and fine adjustments are made by hand to assure solid connection points. These measurements are done with a digital caliper to an accuracy of 0.0001 inch.

Figure 20: Ballast System Assembly


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Fins are aligned using pre-measured slots in the airframe. They are secured in perpendicular angles using a fin jig. This device ensures proper alignment and installation of fins to the booster section. Various mathematical techniques can be used to calculate the internal stress levels for each of the components to include the analysis of the stiffness, strength, and tolerance before damage occurs. COSMOSXpress in SolidWorks will also analyze the load path of the vehicle. However, for simple structures, visual inspection and simple logic and testing is sufficient for establishing load transfer. The thrust force of the motor pushing against the rocket vehicle and drag creates the load paths. The load path is as follows; from the motor retainer, the load is directed to the motor tube and centering rings inside the booster section. The load is then directed to the external surface of the booster section via the epoxy bond. The booster section then distributes the load to the coupler fitting between the booster and acrylic housing sections. The coupler then directs the force vertically to the external surface of the acrylic housing section. From the acrylic section, loads are directed vertically, to the next coupled section where the acrylic section couples to the upper body airframe.

Motor mounting and retention

The motor tube is attached with four .2 inch thick fiberglass centering rings. Each of these is epoxied to the inner airframe and allows the motor tube to provide a secure motor mount. After testing, a single centering ring epoxied to the inner airframe can withstand over 1,500 pounds of constant force. By using four centering rings, the design is sufficient for loads expected by the motor. The motor retainer as shown in Figure 21 uses 12 bolts and threaded inserts in the rear centering ring that is epoxied to the inner airframe of the booster section.

Figure 21: Positive Motor Retainer


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Status of verification

The verification table shown previously in Figure 8 provides the detailed requirements of the launch vehicle. Each requirement has a corresponding design feature to meet that requirement along with the verification method. To date, all aspects of the launch vehicle design are verified to meet SOW requirements.

Final CAD Rendering of Launch Vehicle

The final layout of the launch vehicle is shown in Figure 22. All subsystems and major components are included. These consist of the nose cone and upper body airframe, acrylic payload housing with SMD installed, and the booster section. Avionics bays are located within the couplers to the upper body airframe and booster section. The main parachute is packed into the upper body airframe, and the drogue parachute is packed into the booster section. The ballast system is located in the upper body airframe.

Figure 22: Launch Vehicle Illustration


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Mass Statement

The mass summary of the vehicle is located in Table 9 (Mass Summary). Each subsection is broken down into its respective components in Table 10 (Mass Subsections). The mass calculations for the launch vehicle, subsections, and individual components were obtained by three main methods. First, the mass of components was retrieved from data sheets when available. The second method of obtaining mass involves components not exceeding 2.5 pounds. These components were measured on a digital scale to an accuracy of .0001 ounces. The third method involves obtaining mass estimates of components exceeding 2.5 pounds. Density of the materials in question and the volume of the structural components are used to find the mass. This allowed for a much higher level of accuracy than was obtainable for the PDR. The design of the final launch vehicle has a mass of 37.1 pounds on-the-rail, which is 3.6 pounds over the original mass estimate but still within the original expected mass growth of 2-5 pounds. Due to the construction of a full-scale prototype and the measuring of actual components, the mass is expected to be within one pound of the current estimation. Simulations in OpenRocket show the apogee of the vehicle is reduced by 100-150 feet for every additional pound. With an average thrust of 394.3 pounds-force from the Cesaroni L1720-WT-P, the rocket has a thrust to weight ratio of 10.6:1. This requires more than 393 pounds of additional mass to be added to prevent the vehicle from launching. Currently, the increased mass and redistributing of mass has led to the vehicle being unable to achieve the targeted one mile goal. Simulations are being performed to study these effects and potentially reduce the current mass by five to 10 percent to counter this.


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Mass Summary Subsection Mass (oz.) Mass (lb.)

Payload 39.52 2.47

Recovery 131.13 8.20

Structure 423.41 26.46

Total Mass (Launch) 594.05 37.13

Total Mass (Apogee) 531.95 33.25

Table 9: Mass Summary


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Mass per Subsection Payload Component Quantity Mass (oz.) Total Mass (oz.)

Battery - 9-volt 3 1.28 3.84

Circuit Boards 1 6 6

Miscellaneous Components 1 8.56 8.56

Payload Frame 1 6.15 6.15

Sensors/Electronics 1 13.1 13.1

Servo 2 0.67 1.34

Video Camera 1 0.529 0.529

Subtotal

39.519

Recovery Component Quantity Mass (oz.) Total Mass (oz.)

Attachment Hardware 4 2.86 11.44

Charges - Drogue 1 5.2 5.2

Charges - Main 1 11.2 11.2

Deployment Bag - Main 1 5 5

GPS 1 4.8 4.8

Parachute - Drogue 1 7 7

Parachute - Main 1 64 64

Recovery Electronics - Drogue 1 5.22 5.22

Recovery Electronics - Main 1 5.22 5.22

Shock Cord - Drogue 1 4.65 4.65

Shock Cord - Main 1 7.395 7.395

Subtotal

131.125

Structure Component Quantity Mass (oz.) Total Mass (oz.)

Acrylic Payload Section 1 52.3 52.3

Ballast1 1 1.44 1.4

Bulkhead 3 4.85 14.6

Bulkhead - Coupler 2 2.81 5.62

Bulkhead - Payload 2 16.4 32.8

Center Rings 4 3.21 12.84

Coupler 2 20.95 41.9

Engine Compartment 1 12.9 12.9

Body Tube - Upper 1 38.4 38.4

Body Tube - Rear 1 49.4 49.4

Fin 4 5.625 22.5

Motor2 1 118 118

Motor Retaining Ring 1 4.96 4.96


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Component Quantity Mass (oz.) Total Mass (oz.)

Nosecone 1 15.8 15.8

Subtotal

423.4

1Mass of Ballast varies with configuration.

2Mass listed is for launch. The empty mass (oz.) is: 55.9

Safety and Failure Analysis

After each test launch, the team follows procedures for post-launch analysis. All test launches have video data of assembly, launch, flight, and recovery. In addition, the landing site is undisturbed until pictures are taken and evidence is gathered. This evidence is triangulated with sensor data from payload and onboard altimeters. Failure analysis is conducted on the same day upon return to the rocket lab. Analysis of failures is conducted by sub-teams and presented for group discussion, and updates to the design and additional testing plans are prepared. Additionally, a safety analysis of events is used to update procedures and operations checklists. For example, the correct procedures for arming the deployment altimeters were established in this manner to reflect safety in handling live black powder charges. Each test launch and post-launch analysis allows the team to adequately educate each member on the proper procedures and precautions taken during a launch.

Subscale Flight Results

Subscale Flight Results Throughout the course of testing, the team conducted fifteen subscale launches with various vehicles, motors, and recovery system assemblies in order to learn about and improve the design proposed in the PDR. Eleven flights were conducted with 2.56 inch diameter Level 1 Arcus vehicles constructed during the 2012 Advanced Rocketry Workshop. These vehicles were modified for dual deployment and flew under various Cesaroni G and H motors. Four flights were conducted with the first full-scale prototype. Of the full-scale vehicles, two were launched with a Cesaroni L585 motor, one with a Cesaroni L1720 motor, and one with a custom L667. Of the prototype flights under a Cesaroni L585, one vehicle carried an active payload and the other carried no payload. The flight under the Cesaroni L1720 as well as the custom L667 carried no payload.

Table 10: Mass by Subsection


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Flight Data

The following table, Table 11, provides a visual summary of the available flight data from onboard altimeters and GPS units for each flight. A full description of each flight follows the table. Launch conditions for each flight including weather, elevation, launch coordinates, launch rail position relative to wind, wind speed, wind direction, and pre-flight screen-captures from the Featherweight Raven3 altimeters. Any space showing "N/A" indicates that the data from the altimeter for that piece of information is either unavailable or unreliable. The lack of reliability in the case of the Stratologger SL100 altimeter is the result of a pressure spike in the avionics bay in one of the subscale test vehicles. These pressure spikes were caused either by incorrect port hole sizing, debris in the port holes, poorly maintained port hole alignment throughout the flight, or ventilation between the avionics bay and the black powder charges.


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Date Apogee Predicted Apogee

Drift Video link

November 30, 2012

N/A 401 ft AGL N/A N/A

December 5, 2012

460ft AGL 450 ft AGL 168ft Video

December 7, 2012


December 8, 2012

519ft AGL 600 ft AGL N/A N/A

December 8, 2012

509ft AGL 690 ft AGL N/A N/A

December 14, 2012


December 15, 2012

666ft AGL 615 ft AGL 50ft N/A

December 15, 2012


December 19, 2012


December 19, 2012

1,061ft AGL 1848 ft AGL 342ft Video

December 21, 2012

629ft AGL 780 ft AGL N/A Video

December 21, 2012

4,992ft AGL 5227 ft AGL 2,118ft Video

January 5, 2013


January 6, 2013


January 7, 2013

2,402ft AGL 2908 ft AGL N/A N/A

Table 11: Flight Data

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Sub-scale Test Flight One

On November 30, 2012 at Hunewell Ranch, one launch occurred. Launch conditions were 75 degrees Fahrenheit, 30.05 inches of Mercury, ten mile per hour winds, and elevation of 1,309 feet MSL. The subscale vehicle, shown above in Figure 23, used a Cesaroni G185VMAX motor, with a seven second delay charge for redundant drogue ejection. The recovery avionics utilized a Featherweight Raven3 to control the drogue parachute ejection and main parachute ejection; using a 0.65 gram 3F black powder charge for the drogue parachute ejection and a 1.38 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 256 feet AGL on descent.

Date Location Coordinate Motor November 30,

1012 Hunewell (32.216114, -98.096019)

Cesaroni G185VMAX

Altimeter Drogue Charge Size

Main Charge Size

Main Deploy Altitude

Featherweight Raven3 0.65g 1.38g 256 AGL

Temperature Wind Pressure Elevation 75° F 10 mph 30.05 in Hg 1309 ft

Figure 23: Test Flight One Vehicle

Table 12: Test Flight One Conditions

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Simulated Flight

The simulation for this test launch is shown below in Figure 24. This simulation was conducted through OpenRocket.

Actual Flight

The actual data from the flight was acquired from the onboard Raven3 altimeter, and shown below in Figure 25.

Figure 24: Test Flight One Simulation


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Flight Analysis and Impact on Design

Dual deployment failed with no parachutes deployed. As a result the GPS sled shifted and became lodged in the nose cone, fracturing the nose cone. Post flight analysis revealed two issues; the altimeter never entered flight mode and the port holes were not properly aligned prior to launch. The discovery of the port holes misalignment ultimately led to a design change, which was later implemented on December 16, 2012. The design consisted of adding a static ring with pre-drilled port holes to the coupler, mounted at the center of the exterior. This eliminated the need to align the port holes. A second flight which took place December 8, 2012, before this design was implemented, is suspected of being caused by improper porting as well. Post flight inspection ruled out the possibility of improperly wired electronics. The battery connection is suspected to be a possible cause of failing during launch. The aggressive acceleration of the rocket might have temporarily disrupted the physical connection of the battery terminals. Participants at the December 15, 2012 launch in Asa, TX informed the team this was a common avionic failure. As a result zip ties are now used secure the battery terminal to the battery. The apogee was less than 200 feet, while the altitude predicted in the OpenRocket simulation of the flight was 557 feet. Post flight analysis lead to the discovery of hardware elements not added to the mass calculations properly. These elements were

Figure 25: Raven3 Flight Data


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weighed and added to the mass for the repeat launch on December 5, 2012 at Hunewell.

Sub-scale Test Flight Two

On December 5, 2012 at Hunewell Ranch, one launch occurred. Launch conditions were 70 degrees Fahrenheit, 31.29 inches of Mercury, five mile per hour winds, and elevation of 1,309 feet MSL. The subscale vehicle, shown above in Figure 26, used a Cesaroni G185VMAX motor, with a seven second delay charge for redundant drogue ejection. The recovery avionics utilized a Featherweight Raven3 to control the drogue parachute ejection and main parachute ejection; using a 1.38 gram 3F black powder charge for the drogue parachute ejection and a 1.0 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 256 feet AGL on descent.

Date Location Coordinate Motor

December 5, 2012 Hunewell (32.216114, -98.096019)

Cesaroni G185VMAX


Main Charge Size


Featherweight Raven3 1.0 g 1.38 g 256 ft AGL


Figure 26: Test Flight Two Vehicle

Table 13: Test Flight Two Conditions




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Simulated Flight


Actual Flight


Figure 27: Simulated Flight Two Data


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Dual deployment was not achieved due to the main parachute not ejecting. Post flight analysis revealed the e-match leads were the failure point. The leads were wired to the ground and main, rather than the being wired to the power and main. As a result the recovery procedures were modified to include inspection of the ejection canister connections to the altimeter output channel.

Sub-scale Test Flight Three


Figure 29: Test Flight Three Vehicle


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On December 7, 2012 at Hunewell Ranch, one launch occurred. Launch conditions were 49 degrees Fahrenheit, 28.43 inches of Mercury, six mile per hour winds, and elevation of 1,309 feet MSL. The subscale vehicle, shown above in Figure 29, used a Cesaroni G78 Blue Streak motor, with a six second delay charge for redundant drogue ejection. The recovery avionics utilized a Featherweight Raven3 to control the drogue parachute ejection and main parachute ejection; using a 1.38 gram 3F black powder charge for the drogue parachute ejection and a 1.0 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 256 feet AGL on descent.



Cesaroni G78 Blue Streak


Main Charge Size




Simulated Flight


Table 14: Test Flight Three Conditions




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Actual Flight

No flight data was recovered or reliable for this test launch.


The dual deployment was only partially successful. This was due to both the drogue parachute and the main parachute deploying at apogee. The conclusion of post flight analysis was the main parachute prematurely deployed because the upper body sections friction fit was not strong enough to withstand the force of the drogue parachute ejection charge. This resulted in a sheer pins being implemented to secure the upper body section.

Sub-scale Test Flights 4 and 5

Figure 30: Simulated Test Flight Three


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On December 8, 2012 at Hunewell Ranch, two launches occurred. Launch conditions were 54 degrees Fahrenheit, 29.96 inches of Mercury, 0 mile per hour winds, and elevation of 1,309 feet MSL.

Flight Four

The subscale vehicle, shown above in Figure 31, used a Cesaroni G78 Blue Streak motor, with a six second delay charge for redundant drogue ejection. The recovery avionics utilized a Featherweight Raven3 to control the drogue parachute ejection and main parachute ejection; using a 1.38 gram 3F black powder charge for the drogue parachute ejection and a 1.6 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 256 feet AGL on descent.

Date Location Coordinate Motor December 8, 12 Hunewell (32.216114, -98.096019) Cesaroni G78 Blue Streak


Main Charge Size




Simulated Flight


Figure 31: Test Flight Four Vehicle

Table 15: Test Flight Four Conditions



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Actual Flight


Figure 32: Simulated Test Flight Four Data


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The G78 launch was successful and dual deployment was achieved. While this did not directly impact the design, the experience gained in the test launch was valuable to all future launch operations.

Sub-scale Test Flights Five

The subscale vehicle, shown above in Figure 34, used a Cesaroni G115 White Thunder motor, with a four second delay charge for redundant drogue ejection. The recovery avionics utilized a Stratologger SL100to control the drogue parachute ejection and main parachute ejection; using a 1.38 gram 3F black powder charge for the drogue parachute


Figure 34: Test Flight Five Vehicle


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ejection and a 1.0 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 256 feet AGL on descent.



Cesaroni G115 White Thunder

Altimeter Drogue Charge Size Main Charge Size Main Deploy Altitude

Stratologger SL100 1.38 g 1.0 g 256 ft AGL

Temperature Wind Pressure Elevation

54° F 5 mph 29.96 in Hg 1309 ft

Simulated Flight


Actual Flight


Table 16: Test Flight Five Conditions

Figure 35: Raven3 Test Flight Five Data




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Dual deployment was not achieved due to the main parachute not deploying. Post-flight analysis did not reveal a conclusive reason for this failure. The team suspects a porting issue to be the failure point. The avionics bay had been modified several times for various flights, resulting in extra port holes. No flight data was retrieved from the Stratologger SL100 because a DT2U cable is required to access stored data, which was not available to the team at the time. Other suspicions include not painting the rocket after it had been christened with a successful flight!

Sub-scale Test Flight Six

On December 14, 2012 at Hunewell Ranch, one launch occurred. Launch conditions were 58 degrees Fahrenheit, 29.93 inches of Mercury, sixteen mile per hour winds, and elevation of 1,309 feet MSL. The subscale vehicle, shown above in Figure 36, used a Cesaroni G79 Smoky Sam motor, with a six second delay charge for redundant drogue ejection. The recovery avionics utilized two Featherweight Raven3 altimeters to control the drogue parachute ejection and main parachute ejection; using a 0.8 gram 3F black powder charge for the drogue parachute ejection and a 1.0 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 256 feet AGL on descent.

Figure 36: Test Flight Six Vehicle


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Date Location Coordinate Motor December 14, 2012 Hunewell (32.216114, -98.096019) Cesaroni G79 Smoky Sam


Main Charge Size


2x Featherweight Raven3 0.8 g 1.0 g 256 ft AGL


Simulated Flight


Table 17: Test Flight Six Conditions

Figure 37: Simulated Flight Six Data



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Actual Flight

No flight data was recorded or reliable for this test launch.


Dual deployment was unsuccessful. All recovery systems failed, resulting in a ballistic descent and a lawn dart. The nose cone, upper airframe, and coupler were all destroyed, though the rest of the rocket was deemed reusable. Post flight analysis revealed that both altimeters failed to enter flight mode, and no flight data was recovered. It is suspected the failure point was human error in preparing the altimeters for launch.

Sub-scale Test Flights Seven and Eight

On December 15, 2012 in Asa, two launches under the supervision of our team mentor, Pat Gordzelik, at a Hotroc launch event just outside of Waco. Launch conditions for the first flight were 67 degrees Fahrenheit, 29.93 inches of Mercury, 6 mile per hour winds, and elevation of 427 feet MSL. Launch conditions for the second flight were 74 degrees Fahrenheit, 29.74 inches of Mercury, 10 mile per hour winds, and elevation of 427 feet MSL. A ten foot 1010 launch rail was used for both flights.

Flight 7

The subscale vehicle, shown above in Figure 38, used a Cesaroni G78 Blue Streak motor, with a nine second delay charge for redundant drogue ejection. The recovery avionics utilized a Featherweight Raven3 to control the drogue parachute ejection and

Figure 38: Test Flight Seven Vehicle


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main parachute ejection; using a 1.38 gram 3F black powder charge for the drogue parachute ejection and a 5.0 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 128 feet AGL on descent. A main parachute ejection charge was made on location since the remaining Apogee ejection canisters available appeared to be defective after conducting continuity checks. A charge of 5.0 grams was constructed with the intent of testing the effect of a larger charge size; since the opportunity had not been present previously.

Date Location Coordinate Motor December 15, 2012 Asa (31.438403, -97.027519) Cesaroni G78 Blue Streak


Main Charge Size Main Deploy Altitude



Simulated Flight


Table 18: Test Flight Seven Conditions

Figure 39: Test Flight Seven Simulated Data

https://maps.google.com/maps?q=31.438403,+-97.027519&ie=UTF8&hnear=0x864f7e1fed74495b:0xcf82d39a721792b5,%2B31%C2%B0+26'+13.51%22,+-97%C2%B0+1'+39.65%22&gl=us&t=h&z=15


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Actual Flight


Flight Analysis and Impact on Design (Flight Seven)

Dual deployment was successful, but the shock cord between the upper airframe to the booster section failed upon drogue ejection. Upon apogee, the 5.0 gram black powder charge ignited, with a very loud report, and broke the shock cord tethering the booster section to the upper airframe. This resulted in two sections descending; a booster section with a drogue parachute and an upper airframe descending without decent control. At the pre-programed height of 128 feet AGL, the main parachute deployed and both sections were recovered. Post flight analysis revealed damage to the shock cord from heat. This damage and the 5.0 gram black powder charge were concluded to be the point of failure. As a result we implemented the use of tubular Kevlar shock chord.

Figure 40: Raven3 Test Flight Seven Data


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Flight Eight

The subscale vehicle, shown above in Figure 41, used a Cesaroni G129 Smoky Sam motor, with a ten second delay charge for redundant drogue ejection. The recovery avionics utilized a Stratologger SL100 to control the drogue parachute ejection and main parachute ejection; using a 1.38 gram 3F black powder charge for the drogue parachute ejection and a 2.5 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 200 feet AGL on descent.

Date Location Coordinate Motor December 15, 2012 Asa (31.438403, -97.027519) Cesaroni G129 Smoky Sam


Main Charge Size Main Deploy Altitude



Simulated Flight


Figure 41: Test Flight Eight Vehicle

Table 19: Test Flight Eight Conditions

https://maps.google.com/maps?q=31.438403,+-97.027519&ie=UTF8&hnear=0x864f7e1fed74495b:0xcf82d39a721792b5,%2B31%C2%B0+26'+13.51%22,+-97%C2%B0+1'+39.65%22&gl=us&t=h&z=15


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Actual Flight


Flight Analysis and Impact on Design (Flight Eight)

The flight was a partial success as both parachutes deployed, but both parachutes deployed at apogee. The rocket was recovered. Post flight analysis revealed both ejection charges being fired at once and it was concluded the failure was due to human error. No flight data was retrieved from the Stratologger SL100 because a DT2U cable was needed to access the stored flight data, which was not available to the team.

Sub-scale Test Flights Nine and Ten

On December 19, 2012 at Hunewell Ranch, two launches occurred. Launch conditions were 77 degrees Fahrenheit, 28.87 inches of Mercury, 19 mile per hour winds, and elevation of 1,309 feet MSL. A ten foot 1010 launch rail was used for this flight.

Figure 42: Simulated Flight Eight Data


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Flight Nine

The subscale vehicle, shown above in Figure 43, used a Cesaroni H125motor, with a twelve second delay charge for redundant drogue ejection. The recovery avionics utilized a Stratologger SL100 to control the drogue parachute ejection and main parachute ejection; using a 1.6 gram 3F black powder charge for the drogue parachute ejection and a 1.8 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 200 feet AGL on descent.

Date Location Coordinate Motor December 19, 2012 Hunewell (32.216114, -98.096019) Cesaroni H125


Main Charge Size



Temperature Wind Pressure Elevation 77° F 19 mph 28.87inHg 1309 ft

Simulated Flight


Figure 43: Test Flight Nine Vehicle

Table 20: Test Flight Nine Conditions



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Actual Flight

The actual data from the flight was compiled in Microsoft Excel from GPS readings during the flight and is shown below in Figure 45.

1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

-0.3 0.2 0.7 1.2 1.7 2.2 2.7 3.2 3.7 4.2

Figure 44: Simulated Flight Nine Data

Figure 45: Flight Nine GPS Data

Altitu

de (

Fee

t-M

SL

)

Altitude Versus Time

Time (Seconds)


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Flight Analysis and Impact on Design (Flight Nine)

Partial dual deployment was achieved as both parachutes deployed, but the main parachute deployed at apogee. Post flight analysis revealed both the ejection charges going on at the proper time. The conclusion was the sheer pins in the upper airframe were not installed prior to flight to be the point of failure. This revealed that pre-flight protocols were not followed. As this error was assessed in a previous test flight, the importance of proper implementation of protocols was established to overall mission success. No flight data was retrieved from the Stratologger SL100 because a DT2U cable was need to access the; which was not available to the team.

Flight Ten

The subscale vehicle, shown above in Figure 46, used a Cesaroni H125motor, with a twelve second delay charge for redundant drogue ejection. The recovery avionics utilized a Stratologger SL100 to control the drogue parachute ejection and main parachute ejection; using a 1.6 gram 3F black powder charge for the drogue parachute ejection and a 1.8 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 200 feet AGL on descent.

Figure 46: Test Flight Ten Vehicle


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Main Charge Size



Temperature Wind Pressure Elevation 77° F 19 mph 28.87inHg 1309 ft

Simulated Flight


Actual Flight


Table 21: Test Flight Ten Conditions

Figure 47: Simulated Flight Ten Data



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Flight Analysis and Impact on Design (Flight 10)

This flight was successful and dual deployment was achieved. The subscale vehicle was recovered after landing in a cactus plant, which was documented as an environmental concern for the test launch site. No flight data was retrieved from the Stratologger SL100 because a DT2U cable was needed to access the stored flight data, which was not available to the team.

Sub-scale Test Flights Eleven and Twelve

On December 21, 2012 at Hunewell Ranch, two launches occurred. Launch conditions of both flights were 61 degrees Fahrenheit, 29.32 inches of Mercury, 5 mile per hour winds, and elevation of 1,309 feet MSL. A ten foot 1010 launch rail was used for this flight.

Flight Eleven

1100

1200

1300

1400

1500

1600

1700

1800

-0.4 0.1 0.6 1.1 1.6 2.1 2.6 3.1 3.6 4.1

Figure 48: Flight Ten GPS Data


Altitu

de (

Fee

t-M

SL

)

Time (Seconds)


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The subscale vehicle, shown above in Figure 49, used a Cesaroni G185 VMAX motor, with a five second delay charge for redundant drogue ejection. The recovery avionics utilized a Stratologger SL100 to control the drogue parachute ejection and main parachute ejection; using a 1.4 gram 3F black powder charge for the drogue parachute ejection and a 1.4 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 200 feet AGL on descent.



Main Charge Size




Simulated Flight


Figure 49: Test Flight Eleven Vehicle

Table 22: Test Flight Eleven Conditions



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Actual Flight


1000

1100

1200

1300

1400

1500

1600

1700

1800

1900

2000

0 0.2 0.4 0.6 0.8 1 1.2 1.4 1.6 1.8 2

Figure 50: Simulated Flight Eleven Data

Figure 51: Test Flight Eleven GPS Data


Altitu

de (

Fee

t-M

SL

)

Time (Seconds)


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Flight Analysis and Impact on Design (Flight Eleven)

The flight was successful and dual deployment was achieved. Post flight analysis revealed unreliable data by the Stratologger SL100, but the apogee altitude was determined to be reliable. A pressure spike is suspected of causing the unreliable flight data.

Flight Twelve

The full scale vehicle, shown above in Figure 52, used a Cesaroni L1720 motor. The recovery avionics utilized a Featherweight Raven3 to control the drogue parachute ejection; using a 1.3 gram 3F black powder charge for the drogue parachute ejection. A Stratologger SL100 was used to control the main parachute ejection; using a 1.8 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 700 feet AGL on descent.

Date Location Coordinate Motor December 19, 2012 Hunewell (32.216114, -98.096019) Cesaroni L1720

Altimeters Drogue Charge Size

Main Charge Size


Featherweight Raven3 Stratologger SL100 1.3 g 1.8 g 700 ft AGL


Figure 52: Test Flight Twelve Vehicle

Table 23: Test Flight Twelve Conditions



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Simulated Flight


Actual Flight


Figure 53: Simulated Test Flight Twelve


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Flight Analysis and Impact on Design (Flight Twelve) The dual deployment was not successful due to the main parachute not deploying. The launch vehicle landed in a tree and sustained minor damage to one of the fins. Post flight analysis revealed the ejection charge ignited at the correct altitude. The conclusion was that the ejection charge did not contain enough black powder to deploy the packed main parachute.

Sub-scale Test Flight Thirteen

Figure 54: Raven3 Test Flight Twelve Data

Figure 55: Test Flight Thirteen Vehicle


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On January 5, 2013 at Hunewell Ranch, one launch occurred. Launch conditions were 56 degrees Fahrenheit, 30.00 inches of Mercury, 11 mile per hour winds, and elevation of 1,309 feet MSL. A ten foot 1010 launch rail was used for this flight. The full scale vehicle, shown above in Figure 55, used a Cesaroni L585 motor. The recovery avionics utilized a Featherweight Raven3 to control the drogue parachute ejection; using a primary 2.5 gram 3F black powder charge and a redundant 2.7 gram 3F black powder charge for the drogue parachute ejection. A Stratologger SL100 was used to control the main parachute ejection; using a 5.8 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 700 feet AGL on descent.

Date Location Coordinate Motor January 5, 2013 Hunewell (32.216114, -98.096019) Cesaroni L585


Main Charge Size


Featherweight Raven3 Stratologger SL100 2.5 g and 2.7 g 5.8 g 700 ft AGL


Simulated Flight


Table 24: Test Flight Thirteen Conditions

Figure 56: Test Flight Thirteen Simulation



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Actual Flight

The actual data from the flight was acquired from the onboard Stratologger altimeter, and shown below in Figure 57.


The dual deployment failed. Both parachutes deployed at apogee and the parachutes became entangled. The rocket was recovered, but sustained damage to the nose cone, payload housing section, and SMD payload. The nose cone was repaired and the payload housing section was modified to a 14.5 inch section that was used in two subsequent flights. The prototype SMD payload was deemed repairable, but ultimately retired. Post flight analysis revealed the Stratologger in the main parachute avionics to be incorrectly wired to fire at apogee. This was determined to be the first point of failure. The second point of failure was the drogue parachute becoming entangled with the main parachute. This was due to a design flaw with the deployment bag

Figure 57: Test Flight Thirteen Stratologger Data


Altitu

de (

Fee

t-A

GL

)

Time (Seconds)


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implementation. The packing procedures for the main parachute were documented as the cause of this failure. The scoring altimeter read an apogee altitude of 2,271 feet above ground level. The BMP-180 read an apogee altitude of 2,615 feet above ground level. The Garmin Astro 320 read an apogee altitude of 2,307 feet above ground level. The drogue Stratologger read an apogee altitude of 2,263 feet above ground level. The main Stratologger read an apogee altitude of 2,276 feet above ground level. Overall the launch vehicle achieved an apogee of approximately 1,000 feet less than the simulation in OpenRocket. A potential cause was a simulation error. As a result, it was decided to manually weigh each component prior to each test launch to ensure the actual weights were consistent with the simulation weights.

Sub-scale Test Flight Fourteen

On January 6, 2013 at Hunewell Ranch, one launch occurred. Launch conditions were 54 degrees Fahrenheit, 29.82 inches of Mercury, 11 mile per hour winds, and elevation of 1,309 feet MSL. A ten foot 1010 launch rail was used for this flight. The full scale vehicle, shown above in Figure 58, used a Cesaroni L585 motor. The recovery avionics utilized a Featherweight Raven3 and a Stratologger SL100 in each avionics bay; the drogue parachute avionics bay and the main parachute avionics bay. The drogue parachute avionics used a 2.6 gram 3F black powder charge and a redundant 2.8 gram 3F black powder charge for the drogue parachute ejection. The main parachute avionics used a 5.8 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 700 feet AGL on descent.

Figure 58: Test Flight Fourteen Vehicle


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January 5, 2013 Hunewell (32.216114, -98.096019) Cesaroni L585


Main Charge Size




Simulated Flight


Actual Flight


Table 25: Test Flight Fourteen Conditions

Figure 59: Test Flight Fourteen Simulation




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The dual deployment failed due to the main parachute becoming entangled. The only damage sustained was to the tracking GPS. The GPS antenna broke upon landing resulting in the loss of GPS transmission. Post flight analysis revealed a second design flaw in the deployment bags implementation to the vehicle. This resulted in another redesign for the deployment bag implementation. While the simulated apogee was 3,559 feet above ground level, the actual apogee reached by the launch vehicle was approximately 500 feet less according to on-board recovery electronics. The Garmin Astro DC40 read an apogee of 2,121 feet above ground level. The Stratologger in the main parachute avionics bay read an apogee of 2,920 feet above ground level, while that in the drogue parachute avionics bay read an apogee of 2,923 feet above ground level. The Raven3 in the drogue parachute avionics bay read an apogee of 2,731 feet above ground level.

Figure 60: Test Flight Fourteen Stratologger Data


Altitu

de (

Fee

t-A

GL

)

Time (Seconds)


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Sub-scale Test Flight Fifteen

On January 7, 2013 at P&L Ranch, one launch occurred. Launch conditions were 48 degrees Fahrenheit, 28.61 inches of Mercury, nine mile per hour winds, and elevation of 1,739 feet MSL. A ten foot 1010 launch rail was used for this flight. The modified full scale vehicle, shown above in Figure 61, used a modified L667 motor. The recovery avionics utilized a Featherweight Raven3 and a Stratologger SL100 to control the drogue parachute ejection; using a 3.5 gram 3F black powder charge for the drogue parachute ejection. A Stratologger SL100 was used to control the main parachute ejection; using a 5.6 gram 3F black powder charge for the main parachute ejection. Main parachute deployment was set to activate at 700 feet AGL on descent.

Date Location Coordinate Motor January 7, 2013 P&L Ranch (32.105039, -99.142556) Custom L667


Main Charge Size



Temperature Wind Pressure Elevation 48° F 9 mph 28.61 in Hg 1,738 ft

Simulated Flight


Figure 61: Test Flight Fifteen Vehicle

Table 26: Test Flight Fifteen Conditions

https://maps.google.com/maps?q=32.105039,+-99.142556&hl=en&sll=31.437086,-97.027681&sspn=0.030574,0.055747&t=h&gl=us&z=15


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Actual Flight


Figure 62: Test Flight Fifteen Simulation

Figure 63: Test Flight Fifteen Stratologger Data


Altitu

de (

Fee

t-A

GL

)

Time (Seconds)


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Dual deployment failed due to no main parachute deployment. The launch vehicle was recovered, sustaining damage only to the modified payload housing structure, which was destroyed on impact. Post flight analysis revealed both the main parachute ejection charge and the redundant drogue ejection charge failed to ignite. Failure to arm the main parachute avionics is suspected of being the failure point. Due to four different audible devices being activated at the same time, it was difficult to ensure all avionics devices were armed. The redundant drogue ejection charge was suspected of not igniting due being dislodged by the primary ejection charge. Although this redundant charge not firing does not result in critical failure for drogue ejection, it has been documented. As a result of this failure two changes were implemented. First, the recovery protocol was modified to standardize arming of the avionics bays. Secondly, design implementations have been proposed to add visual indicators of the altimeters in the form of L.E.D.’s mounted in the coupler, near the port holes. This will allow the altimeters audible beepers to be deactivated solving two issues; allowing the inspection of the altimeters visually rather than audibly and allowing the official scoring altimeter to be heard clearly without the need to deactivate the avionics. Of the four recovery electronics, only the Stratologger in the drogue parachute avionics bay logged recoverable data. According to this altimeter the time of flight was 39.85 seconds and the apogee of the flight was 2,402 feet above ground level. The simulated apogee with the custom L667 was 2,908 feet above ground level, but the difference between the actual recorded apogee and the predicted apogee cannot accurately be compared since the actual motor was different. Additionally, the thrust to weight ratio was approximately 4:1, resulting in a less than desirable rail exit velocity. This likely affected the apogee of the flight.

Impact on Design Summary

Sub-scale launches have shown that we need to drill three port holes. Sizing the port holes too small or too large affects the reliability of the altimeter systems. It was determined that the port holes for the drogue parachute altimeters should be 5/32 inches in diameter, while those for the main parachute altimeters should be 3/16 inches in diameter. It has been determined that LED lights will be implemented in place of audio output on all altimeters in the avionics bays. The use of the LED output enables distinction of the scoring altimeter in the payload section that beep the official apogee altitude. Further, it reduces noise when arming the altimeters at the launch site while providing visual verification that the magnetic switch in each avionics bay has been armed.


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Daveyfire e-matches are no longer going to be used, but Pratt Hobbies ejection canisters will be. On subscale tests, the team was packing 3F black powder with tissue wadding into large ejection canisters, which hold up to two grams. This method proved effective. However, based on the analysis of the full-scale flight conducted on December 21, 2012 at Hunewell Ranch, it was determined that extra-large canisters, which hold four grams, must be used in order to eject the main parachute system which is described in

more detail under Physical Components in the Recovery Subsystem section. One such canister is shown in Figure 64. The ejection canisters, while not reliably reusable, are more environmentally friendly and safer as ignition depends solely on a bridge between the ends of the exposed leads in the bottom of the canister rather than ignition of a chemical pyrogen. Based on the availability of non-synthetic gun powder, 3F black powder, as shown in Figure 65, implemented from, is going to be used rather than 4F. Subscale flights and static tests have been carried out successfully using 3F black powder, while no testing has been done with 4F. Synthetic smokeless powder was originally purchased by the team for testing with black powder charges, but it was determined that the burn time was too long and the ignition temperature too high to safely carry out the necessary testing. It has been determined that the most reliable way to cause separation while avoiding premature ejection is to make each parachute compartment a "cannon". That is, the ejection charges should be at the end farthest from the separation

point, with components packed down on top of the charges.

In order that each parachute compartment separates, the charge leads from each ejection charge to the corresponding altimeter will be precut to the length of the charge compartment plus four inches. A dog-tracking GPS system shall to be used in the final design rather than the Big Red Bee, due to the ease of interface and price. The chosen system is the Garmin Astro DC 40 transmitter unit, removed from the collar and mounted to the reverse side of the shielded altimeter

compartment. A ground receiver, the Astro 320, with an LCD interface that can track both transmitters at once

Figure 64: Ejection Canister

Figure 65: 3F Black Powder

Figure 66: Astro 320 GPS System

http://www.apogeerockets.com/Ejection_Systems/Ejection_Canisters/XL_Variable-Capacity_Ejection_Canister

http://www.seriousshooters.co.nz/content.asp?function=viewprd&prd_id=292199

http://www.seriousshooters.co.nz/content.asp?function=viewprd&prd_id=292199


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will be held by a team member and used to easily identify the location of the landed launch vehicle. This system is shown in Figure 66. Rather than being mounted in an avionics bay, the recovery system GPS transmitter, the Garmin Astro DC40, will be attached to the main parachute shock cord, as documented under Recovery Subsystem in the Physical Components section. The shock cords will be constructed of ¼ inch tubular Kevlar rather than ½ inch due to space constraints for packing, but still possess sufficient tensile strength. The cords will be cut such that those for the drogue parachute are half of the length of those for the main based on an experienced suggestion from team mentor Pat Gordzelik. Quick links rather than swivels will be used to attach the shock cord to the shroud lines for each parachute, since the number of shroud lines and their material is sufficient to prevent entanglement under normal flight conditions. The main parachute will be made of rip stop nylon due to a potentially high rate of descent upon deployment. Furthermore, based on the full-scale test performed on December 21, 2012 at Hunewell Ranch with a 14 foot main parachute, the main parachute will remain at 10 feet. This reduces the weight of the system and allows for looser packing of the main parachute into its deployment bag. The main parachute deployment bag is shielded from the firing of the ejection charge by “dog barf” wadding based on experience with scorching of our Nomex blankets in sub-scale flights. The attachment point of the main parachute deployment bag is at the eye bolt at the rear of the nose cone section rather than a knot near the main parachute avionics bay due to experience with entanglement. This entanglement was seen in the test flight conducted on January 6, 2013 described under Subscale Flight Results in the Predictions and Comparisons section. The primary drogue parachute will not be housed in a deployment bag due to the small packed size, but will also be shielded by “dog barf” wadding and wrapped in a Nomex blanket.

Recovery Subsystem

Physical Components

The drogue parachute remains at two feet in diameter as in the PDR. The drogue parachute is the Sky Angle CERT-3 24" Drogue. While this parachute has a manufacturer-tested load capacity of 1.0 to 2.2 pounds, it has been tested up to 36.660 pounds in-flight with no damage and safe recovery. Likewise, the main parachute remains 10 feet in diameter, and was chosen to be the Sky Angle CERT-3 XX Large, with a tested load capacity of 60 to 129 pounds. The chosen parachutes are cross-form and constructed of 1.9 ounce, silicon-coated, rip stop Nylon. Further, these parachutes have seams reinforced with nylon webbing to reduce the probability of shroud line disconnection as a result of ejection. With four shroud lines, each made of 5/8 inch woven tubular nylon for durability, these

http://www.autonavdirect.com/request-2-618.html


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parachutes have the fewest number of shroud lines to reduce the probability of entanglement. The length of the shroud lines on the main parachute is 10 feet. Similarly, the length of the shroud lines on the drogue parachutes is two feet. Manufacturer-tested drag coefficients for the main parachute and drogue parachute are 2.92 and 1.16, respectively. The main parachute has a surface area of 129 square feet. The drogue parachute has a surface area of 1.16 square feet.

The main parachute is neatly folded into its deployment bag, the SkyAngle XXLarge Deployment Freebag, with all shroud lines and shock cord fed through the loops in the front of the bag, as seen in Figure 67, as implemented from Apogee Rockets. The drogue parachute and its shroud lines are neatly folded into a Nomex cloth rather than a deployment bag due to its small size. Additional "dog barf" wadding is placed between the black powder charges and the parachutes to ensure minimal damage upon ejection. In order to enable deceleration of the two tethered sections on descent, .25 inch tubular Kevlar shock cords have been implemented based on strength and flame-proof construction.

The length of the shock cord for the drogue parachute is 20 feet, thus based on the advice of the team mentor; the main parachute’s shock cord is 40 feet long. The shock cord for each parachute is “z-folded” and bound with painter’s tape both for a reduced risk of entanglement and to save space in the parachute compartments.

The vehicle sections and descent control systems are tethered together using shock cords utilizing eyebolts and stainless steel, delta-shaped quick links.

The motor casing within the booster section has an eyebolt installed. This eyebolt is used to attach the booster section to the 20 foot shock cord. The other end of the shock cord is attached to an eyebolt installed on the exterior bulkhead of the drogue avionics bay (bottom of the payload housing structure); effectively tethering the booster section. The drogue parachute shroud lines are attached to a loop in the shroud line four feet below the drogue avionics bay.

The exterior bulkhead of the main avionics bay has an eye bolt installed. This eyebolt is used to attach the 40 foot shock cord to the payload housing section. The main parachute deployment bag is also anchored here.

Figure 67: SkyAngle XXLarge Deployment Freebag

http://www.apogeerockets.com/Buildi%20ng_Supplies/Parachutes_Recovery_Equipment/Parachute_Deployment_Bags/XXL_Parachute_Deployment_Bag

http://www.apogeerockets.com/Buildi%20ng_Supplies/Parachutes_Recovery_Equipment/Parachute_Deployment_Bags/XXL_Parachute_Deployment_Bag


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The upper body airframe has an eyebolt installed in a bulkhead just below the nose cone. This eyebolt is used to anchor a four foot Kevlar lead line. The lead provides an easily accessible quick link. This quick link is used to attach the shroud lines of the main parachute. The same quick link is also use to attach the free end of the 40 foot shock cord, effectively tethering the payload housing section to the upper body airframe. This attachment scheme is shown below in Figures 68, 69, and 70. Figure 68 shows the upper body airframe attachment to the upper acrylic housing section coupler. The main parachute is located inside its deployment bag.

Figure 69 shows the acrylic housing structure laid beside its couplers to demonstrate the attachment scheme further.

Figure 68: Main Parachute Attachment Scheme


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Figure 70 shows the attachment scheme from the lower acrylic housing structure coupler to the booster section. The drogue parachute is folded inside of its orange Nomex cloth.

Figure 69: Attachment Scheme to Couplers


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In order to ensure that separation in the body of the vehicle only occurs upon ejection, two removable nylon shear pins are spaced 180 degrees apart to temporarily secure the upper body airframe to the acrylic housing structure. When the ejection charge fires, the force of the coupler sliding past will snap the shear pins. However, other stresses under 25 pound force such as those caused by shifting mass, drag, or ejection from another compartment should not be strong enough to cause separation. Two sets of ejection charges, one set for main parachute ejection and one set for drogue parachute ejection, are constructed at each launch site. Each set of ejection charges includes one primary charge and one secondary charge. The secondary charge is sized larger than the primary to help ensure that separation occurs even if the primary charge is defective in some way. Packing Goex 3F black powder with wooden dowel rod and securing the pressure in each charge by packing tissue or “dog barf” wadding down on top of the packed black powder forms a reliable ejection charge. The ejection canisters used are Pratt Hobbies XL Ejection Canisters, which hold up to four grams of packed 3F black powder with the

Figure 70: Drogue Parachute Attachment Scheme


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foam packing included. Inspection shows that these canisters hold up to 6.2 grams packed 3F black powder with varying amounts of “dog barf” as wadding. The amount of black powder used in each ejection charge is affected by the diameter of the parachute chamber, the volume of the parachute chamber, the force against the separation point, and how well the compartment is sealed. Two suggestions from the team mentor were to not use more than four times the amount of black powder calculated for the charge, and to use more black powder in the secondary ejection charge than in the primary. As a general rule, a body diameter of between five and 5.5 inches requires one gram of 3F black powder per six inches of length. Since the drogue parachute compartment is 14.5 inches long, then the black powder charges for drogue parachute ejection must be at least 2.4 grams. Testing has shown that 2.4 grams for the primary and 2.8 grams for the secondary are sufficient amounts of 3F black powder packed into the ejection canister with “dog barf” wadding and covered with painter’s tape to eject all necessary components. Since the main parachute compartment is 23 inches long, the black powder charges for main parachute ejection must be at least 3.8 grams. Testing has revealed that with the deployment bag packed into the chamber, 5.4 grams for the primary and 5.8 grams for the secondary are sufficient amounts of 3F black powder packed into the ejection canister with “dog barf” wadding and covered with painter’s tape to eject all necessary components. As detailed under “Impact on Design” in the “Subscale Flight Results” section, three port holes have been drilled to vent each avionics bay to the exterior. This stabilizes chamber pressure and allows for accurate readings to be taken by the altimeters. The size of these port holes is 5/32 inches each in diameter for the drogue avionics bay, 3/16 inches each in diameter for the main avionics bay.

Electrical Components

The recovery system electronics consist of two avionics bays and one ground component. Each avionics bay includes a set of redundant altimeters maintained by a magnetic arming switch and a nine-volt battery. A GPS transmitting unit is secured to the main parachute shock cord. A member of the recovery team will operate a handheld GPS receiving unit on the ground to track the flight path and aid in physical recovery after landing.


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Altimeters

As shown in the above Figure 71, each set of altimeters consists of one Featherweight Raven3 and one Perfectflite Stratologger SL100. The Raven3 is capable of either accelerometer or barometer detection of altitude. Based on subscale flight results, it has been determined that accelerometer detection is more reliable than barometer detection; because, pressure spikes within the avionics bay can occur if debris is in the way of the port holes or if the chamber is not completely sealed from the black powder charges. Figure 72 details the flight cycle of a Raven3. Upon arming of the magnetic switch the Raven3 enters pre-flight mode, where it will remain until detecting launch. Launch is signified by an axial acceleration over three Gs integrating to a three mph upward

Figure 71: Altimeter Electronics Schematics


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velocity. Prior to entering flight mode the Raven3 must initialize an onboard data buffer with measurements ranging from 0.352 to 0.704 seconds prior to the launch event. While in flight mode a number of atmospheric measurements are gathered and stored at various frequencies. The axial acceleration is recorded every 2.5 milliseconds and the lateral acceleration is recorded every five milliseconds. The onboard processor integrates these values to determine the vertical, horizontal, and combined velocities. Every 25 milliseconds the output current at each terminal is recorded. Every 50 milliseconds the barometric pressure, barometrically-derived altitude, temperature, and output voltage at each terminal are recorded. When combined these measurements allow for detailed modeling of the rocket’s trajectory. Furthermore, the high sampling frequencies allow for an accurate real-time approximation of the flight. During flight the Raven3 tests for the criteria signaling the energizing of any of its four output terminals. The measurements of output voltage and current allow for exact verification of when any charges were ignited. The Raven3 continues to make atmospheric measurements until landing is detected. Landing is signified by a constant barometric pressure. After landing measurements cease, the flight data is finalized on the flash memory, and speakers begin to signal apogee via discretized beeps.


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Figure 72: Raven3 Software Flow Diagram


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The Raven3 in the drogue parachute avionics bay will be programmed to fire the primary drogue black powder charge upon detection of apogee with all other output channels disabled. Upon reaching apogee, the detection of decreasing pressure, signifying a decreasing altitude, will prompt the firing of the drogue ejection charge. The Raven3 in the main parachute avionics bay will be programmed for altitude-based detection of 704 feet above ground level following apogee detection to fire the primary main black powder charge with all other output channels disabled. (The Raven3 is only able to detect altitude increments of 32 feet above ground level, so the height comparator is set to 704 feet above ground level.) Figure 73 details the flight cycle of a Stratologger. Upon arming of the magnetic switch the Stratologger enters pre-flight mode, where it will remain until detecting launch. Launch is signified by an altitude above 160 ft. Prior to entering flight mode the Stratologger must initialize an onboard data buffer with 28 measurements corresponding to 1.4 seconds prior to the launch event. While in flight mode the altitude is recorded approximately every 50 milliseconds. This sampling frequency is limited by an analog-to-digital conversion. An analog barometer measures pressure, which is converted to a digital signal via a 24-bit delta-sigma analog-to-digital converter. Due to the large resolution of the conversion process the overall sampling frequency lowers tremendously. This analog pressure signal is used to approximate the altitude. The onboard computer compares these data points to determine velocities and accelerations. During flight the Stratologger tests for the criteria signaling the energizing of either of its two output terminals. The Stratologger continues to make atmospheric measurements until landing is detected. Landing is signified by a constant barometric pressure. After landing measurements cease, the flight data is finalized on the flash memory, and speakers begin to signal apogee via discretized beeps.


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Since the Stratologger only detects altitude barometrically, it will be the used as the secondary, or backup, altimeter. The Stratologger in the drogue parachute avionics bay will be programmed to fire the secondary black powder charge at apogee detection with the switch outputs grounded, the main outputs empty, and the drogue outputs wired to the ejection canister through the bulkhead. The Stratologger in the main parachute avionics bay will be programmed for detection of a height of 700 feet above ground level to fire the secondary black powder charge. The switch outputs will again be grounded,

Figure 73: Stratologger Software Flow Diagram


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the main outputs wired to the ejection canister through the bulkhead, and the drogue outputs empty. Figure 73 details the flight cycle of a Stratologger. Each set of altimeters is powered by a separate nine-volt battery and wired through a Featherweight Magnetic Arming Switch. A powerful handheld magnet is run over each switch from the exterior of the upper body airframe to arm each set of altimeters. Arming of the system is confirmed visually when a blue LED is seen through the porthole along with a series of red and blue LED flashes from the Raven3 in the compartment. Arming of the Raven3 is confirmed audibly by a series of beeps declaring the battery voltage. Arming of the Stratologger is confirmed audibly by a different series of beeps declaring the battery voltage, the apogee of the last recorded flight, and the height above ground level for which the main ejection-charge output has been programmed to fire. As previously discussed, the audible functionality of all recovery altimeters will be disabled. Confirmation of altimeter activation on the Launchpad will be performed by LED indicators. This design eliminates the need to disable the audible beeping of each altimeter at landing. Proper sizing of ejection charges along with the dual redundancy of the altimeter systems should ensure ejection and separation at the appropriate point during the flight. The avionics system near the booster section works to eject the drogue parachute at apogee to minimize drift and lessen descent rate. The avionics system near the nose cone works to eject the main parachute at a height of 700 feet above ground level on descent to greatly slow descent rate, as well as minimize the force of impact.

Avionics Bays

Each avionics bay is constructed around a vertical plywood sled 5.05 inches in width and 8.25 inches in length. An example setup of the avionics bays is shown below in Figure 74. One side of each bay is lined with 0.001575 inch thick aluminum flash shielding as well as the corresponding half of the bulkhead on either end. This shielding is intended to keep radio frequency interference from interacting with the altimeters and firing the deployment charges prematurely.


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Each sled is secured inside each avionics bay by two lengths of .25 inch all-thread three inches apart which have been padded with fuel line. Four contact points of 3/8 inch PVC piping, each one inch in length, are epoxied to the ends of the board as guides for the all-thread. A drawing of the avionics sled is shown in Figure 75.

Figure 74: Example Drogue/Main Avionics Bay


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Each set of altimeters is mounted with screws and held up on spacers to the shielded side in each avionics bay. Each battery is placed into a plastic mount that is screwed down to the shielded side and kept secure by a zip tie. All appropriate connections to and from the battery and altimeters are fed through a hole drilled through the board, soldered to the switch on the other side. The hole is then temporarily sealed with a bead of silicon. The switch is secured by a screw on the wall of the compartment for external accessibility.

Global Positioning System

The Garmin Astro DC40 transmits position every five seconds. The transmitter is securely fastened to the drogue parachute shock cord in the booster section. There are five options for transmitting frequency; 151.82 MHz, 151.88 MHz, 151.94 MHz, 154.57 MHz, and 154.80 MHz. The default setting, which can be changed at the launch site from the 320, is 151.82 MHz.

Figure 75: Drawing of Avionics Sleds


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The transmitted signal is received by the Garmin Astro 320 and used along with visual reference to track the path of the launch vehicle throughout the course of the flight. Lateral distance from the launch rail is easily determined upon landing using the readout on the screen of the Astro 320. A digital compass then aids in the physical recovery of the vehicle. The flight cycle of this system is detailed Figure 76.

Figure 76: GPS Software Flow Diagram


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Kinetic Energy

The kinetic energy at various phases throughout the mission is calculated using the launch vehicle shown above in Figure 77. The kinetic energy is calculated several times throughout the launch. The kinetic energy is calculated at rail exit, at apogee, and upon landing. The time, altitude, velocity, acceleration, and lateral drift provided in the table are based on an OpenRocket simulation for average launch conditions in Huntsville, Alabama in April under a 10 mile per hour winds. This simulated flight path is shown below in Figure 78. Note that the lateral distance from the launchpad of the vehicle at landing is approximately 1050 feet.

Figure 77: Launch Vehicle Prototype

Figure 78: Final Vehicle Simulation


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Note that while apogee is simulated to occur at a height of 4852 feet above ground level. Testing has revealed that actual launch conditions including weather, launch rail inclination, and ballast can greatly affect flight path. Kinetic energy is calculated by using the mass of each subsection of the vehicle, followed by the total contained weight, for three phases of the flight. These are the maximum kinetic energy from apogee to main parachute deployment, and main deployment until landing. The results for kinetic energy are summarized below in Table 27. Calculations follow the table.

Phase

Nose Cone and

Upper Body

Airframe

Payload Booster Total

Contained Vehicle

Apogee to Main (max)

965.6 ft·lbf 1280.2 ft·lbf 1225.3 ft·lbf 3730.3 ft·lbf

Main to Landing (max)

5.4 ft·lbf 17.7 ft·lbf 15.9 ft·lbf 51.6 ft·lbf

From apogee to main deployment:

From main deployment to landing:

Table 27: Kinetic Energy Summary


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The kinetic energy on landing of the contained launch vehicle is less than 75 foot-pounds of force. The final distance from the launch rod in a 10 mile per hour wind has been simulated and calculated to be less than 2,500 feet as well. Thus, the recovery system design for the vehicle is compliant with competition regulations.

Test Results

Static tests performed to date on the recovery system are outlined in the following Table 28. Dynamic tests are described in detail and summarized under “Predictions and Comparisons” in the “Subscale Flight Results” section. For each static test; the component being tested, the type of test, and the procedure used to conduct the test are identified. Detonation tests were used to determine the reliability of the ejection charge design. Separation tests were performed to determine reliability of ejection charge sizing. Burst tests were used to determine reliability of altimeters with ejection charges. A brief description of each test on the noted date is included. Finally, success or failure is determined.


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Component Test Type

Date Description Procedure Success

Drogue Ejection Charge

Burst 26 Nov 12 Raven3, drogue leads secured to +/- terminals of voltmeter

9V charge delivered to leads within interface program at simulated apogee based on a "descent" upon a 4G burst manually delivered to altimeter

Voltage drop registered; yes

Main Ejection Charge

Burst 26 Nov 12 Raven3, main leads secured to +/- terminals of voltmeter



Ejection Charge

Detonation

30 Nov 12 1.25g 3F black powder packed by tissue, 2g ejection canister

9V charge delivered manually to chamber by leads

Full blast; yes

Ejection Charge

Detonation

5 December 12

1.25g 3F black powder packed by tissue, 2g ejection canister


Full blast; yes


Separation, full-scale

14 December 12



Complete separation, no damage; yes


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Component Test Type



Burst 15 December 12

Raven3, drogue leads secured to +/- terminals of voltmeter














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Component Test Type







Ejection Charge

Detonation

20 December 12

1.2g 3F black powder, box tape envelope, Estes igniter


Full blast; yes

Ejection Charge

Detonation

20 December 12

2.2g 3F black powder, box tape envelope, medium Wildman igniter


Full blast; too long of a burn time, fire, not using

Ejection Charge

Detonation

20 December 12

2.2g 3F black powder, box tape envelope, large Wildman igniter


Full blast; too long of a burn time, fire, not using

Ejection Charge

Separation, subscale

21 December 12

1.8g 3F black powder, box tape envelope, Estes igniter


Full separation; yes


Burst 3 January 13

Stratologger, drogue leads secured to +/- terminals of voltmeter

9V charge delivered manually to leads within interface program



Burst 3 January 13





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Component Test Type



Burst 3 January 13





Burst 3 January 13

Stratologger, main leads secured to +/- terminals of voltmeter




Burst 3 January 13





Burst 3 January 13





Separation, full scale

5 January 13


9V charge delivered manually to leads

Full blast; yes


The following logs detail the safety and failure analysis of all test launches to date. A brief description of each launch precedes each table. The failure mode for each issue encountered is determined. Causal and effected components, along with potential hazards are identified. Finally, mitigations for each failure are presented to be implemented in future testing.

Table 28: Static Tests


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Test Date November 30, 2012 Brief Description of Test: Sub-scale dual deployment test launch at Hunewell Ranch. The altimeter used was Raven3, one black powder charge with 0.65 grams of packed 3F black powder. The main was programmed for 256 feet above ground level, backup main for 192 feet above ground level. Motor ejection was programmed for drogue ejection at seven seconds based on OpenRocket simulation with G185VMAX.

Test Date November 30, 2012 Failure Mode

Component(s) Responsible

Component(s) Affected

Potential Cause(s)

Potential Hazard(s)

Mitigation

Flight mode never entered

Port holes E-match firing, flight data, main deployment

Chamber not vented before accelerometer experienced 3G

Poor to no flight data, damage to upper body airframe

Port holes oversized, coupling secure

Nose cone fracture

GPS sled Nose cone structure, GPS housing, chamber pressure

Wooden dowel system not sturdy enough to hold GPS in place throughout flight

Poor to lost signal, loss of GPS unit, pressure reaches altimeter

Use metal pins to secure GPS sled

Open Vehicle simulation

Altimeter and GPS sleds

Incorrect motor burnout calculation for drogue ejection, simulated apogee not achieved

Altimeter and GPS sled masses not included in simulation

Inaccurate settings on altimeter, incorrect motor burnout

Include all component masses in simulation

Main Parachute did not deploy

Altimeter settings, motor burnout

Post-flight recovery

Altimeter counts motor burnout prior to ejection but motor burned out lower than programmed ejection

Ballistic impact, damage to upper body airframe

Simulation includes all masses so altimeter and motor ejection are prepared accurately

Table 29: Safety and Failure Analysis 11-30-12


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Test Date December 5, 2012 Brief Description of Test: Sub-scale dual deployment test launch at Hunewell Ranch. The altimeter used was Raven3 to charge one black powder well with one gram of packed 3F black powder for the main, which was programmed for 224 feet above ground level. Motor ignition for drogue ejection was shaved to three seconds based on OpenRocket simulation with a G185VMAX motor, which has a 1.38 gram 4F black powder charge.

Test Date December 5, 2012 Failure Mode



Potential Cause(s)

Potential Hazard(s)

Mitigation

Main parachute deployment

E-match wiring Ejection charge E-match wired incorrectly to ground and main rather than power and main


Properly wire e-match



105


Test Date December 7, 2012 Brief Description of Test: Sub-scale dual deployment test launch at Hunewell Ranch. The altimeter used was Raven3, having one black powder charge with one gram of packed 3F black powder. The main was programmed for 224 feet above ground level. Motor ignition for drogue ejection was shaved to three seconds based on OpenRocket simulation under G78 Blue Streak motor, which has a 1.38 gram 4F black powder charge.




Potential Cause(s)

Potential Hazard(s)

Mitigation

Premature separation of forward section from main body section

Coupling between forward section and main body section

Main parachute ejection

Friction fitting too weak to withstand force of drogue ejection at apogee

Shroud line, shock cord, and/or parachute entanglement, leading to ballistic impact and potential damage to upper body airframe

Employ nylon sheer pins with future launches

Altimeter data

Power supply Flight analysis Altimeter was left powered on for too long and went back into wait mode, so flight data was lost

Less data to support testing

Remove and power down altimeter directly upon recovery



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Test Date December 8, 2012 Brief Description of Test: Sub-scale dual deployment test launch at Hunewell Ranch. Altimeter used was Raven3, with one black powder charge with 1.6 grams of packed 3F black powder. The main output was programmed for 192 feet above ground level. The motor ignition was shaved to six seconds based on OpenRocket simulation with a G78 Blue Streak, which has a 1.38 gram 4F black powder ejection charge. Test Date December 8, 2012 Brief Description of Test: Sub-scale dual deployment test launch at Hunewell Ranch. Altimeter used was Stratologger, one black powder charge with one gram of packed 3F black powder. The main output was manually programmed for 100 feet above ground level deployment upon descent, with motor ignition for drogue ejection shaved to four seconds based on OpenRocket simulation with G115 White Thunder, having a 1.38 gram 4F black powder ejection charge.




Potential Cause(s)

Potential Hazard(s)

Mitigation


Altimeter Ejection charge Wiring error, chamber pressure


Check flight data with software once cable arrives to help determine cause



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Test Date December 14, 2012 Brief Description of Test: Sub-scale dual deployment test launch at Hunewell Ranch. Altimeters used were two Raven3 units. The main was equipped with a one gram of packed 3F black powder ejection charge, programmed for 162 feet above ground level deployment upon descent. The drogue was equipped with a 0.8 gram of packed 3F black powder ejection charge, programmed for accelerometer detection of apogee. Motor used was Cesaroni G79 Smoky Sam.




Potential Cause(s)

Potential Hazard(s)

Mitigation

Drogue altimeter flight mode

Drogue altimeter ejection charge firing

Drogue parachute ejection

Port hole alignment, poor friction fitting between booster and middle sections

Ballistic impact, damage to upper body airframe, no flight data

Test altimeter without reliance of ejection, use shear pins for points of separation


Impact of lower avionics bay with upper avionics bay

Main altimeter barometer and/or accelerometer components

Ballistic descent

Ballistic impact, damage to upper body airframe, damage to recovery components

Use shear pins for points of separation, switch ejection direction in next build



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Test Date December 15, 2012 Brief Description of Test: Sub-scale dual deployment test launch at HotRoc’s event in Asa. Altimeter used was a Raven3. The main parachute charge was a 2.5 gram 3F black powder ejection charge, programmed for 128 feet above ground level deployment upon descent. The drogue was equipped with a five gram 3F black powder ejection charge, programmed for accelerometer detection of apogee. Motor used was Cesaroni G78 Blue Streak, equipped with a 1.38 gram 4F black powder backup ejection charge for the drogue.




Potential Cause(s)

Potential Hazard(s)

Mitigation

Single recovery

Drogue parachute ejection

Tethering of booster section to middle body section

Drogue parachute black powder charge oversized

Ballistic impact of disconnected booster section, damage to upper body airframe or recovery electronics

Use fire retardant shock cord, use less black powder

Recovery wadding in booster section

Drogue parachute ejection charge firing

Reusability of drogue parachute

Drogue parachute black powder charge oversized

Failure of drogue parachute, high-speed ejection of main parachute

Wrap each parachute in fire retardant blanket or bag along with wadding



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Test Date December 15, 2012 Brief Description of Test: Sub-scale dual deployment test launch at HotRoc event in Asa. Altimeter used was one Stratologger SL100. The main parachute charge was 2.5 grams of 3F black powder, programmed for 200 feet above ground level deployment upon descent. The drogue was equipped with a 2.5 gram 3F black powder ejection charge, programmed for barometric detection of apogee. Motor used was Cesaroni G131 Smoky Sam, equipped with a 1.38 gram 4F black powder backup ejection charge for the drogue.




Potential Cause(s)

Potential Hazard(s)

Mitigation

Charge event staging

Timing of main parachute ejection

Dual event deployment

Main altimeter programmed too close to apogee

Parachute and/or shock cord entanglement and/or damage to upper body airframe

Program main ejection for lower altitude, make drogue shock cord half the length of main shock cord



110


Test Date December 19, 2012 Brief Description of Test: Sub-scale dual deployment test launch at Hunewell Ranch. Altimeter used was one Stratologger SL100. The main parachute charge was 1.8 grams of packed 3F black powder, programmed for 200 feet above ground level deployment upon descent. The drogue was equipped with a 1.6 gram packed 3F black powder ejection charge, wired for ignition upon barometric detection of apogee. Motor used was Cesaroni H125, equipped with a 1.38 gram 4F black powder backup ejection charge for the drogue.




Potential Cause(s)

Potential Hazard(s)

Mitigation

Dual deployment staging

Friction fitting between middle body section and nose cone section

Premature main parachute ejection

Lack of shear pins

Parachute and/or shock cord entanglement, high-speed impact

Implement shear pins on future flights



111


Test Date December 19, 2012 Brief Description of Test: Sub-scale dual deployment test launch at Hunewell Ranch; Altimeter used was one Stratologger SL100. The main parachute charge was 1.8 grams of packed 3F black powder, programmed for 200 feet above ground level deployment upon descent. The drogue was equipped with a 1.6 gram packed 3F black powder ejection charge, wired for ignition upon barometric detection of apogee. Motor used was Cesaroni H125, equipped with a 1.38 gram packed 4F black powder backup ejection charge for the drogue.




Potential Cause(s)

Potential Hazard(s)

Mitigation

Landing

Cactus plant Ease of recovery

Weather cocking sent vehicle on path toward greenery near edge of ranch

Cactus needles in vehicle components or personnel

Bring thick gloves to launch site in case of cactus recovery



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Test Date December 21, 2012 Brief Description of Test: Full-scale dual deployment test launch at Hunewell Ranch. Altimeters used were one Stratologger SL100 for both drogue and main parachute ejection. Both parachute charges were 1.4 grams of 3F black powder envelopes equipped with Estes igniters. The main parachute was set to deploy at 200 feet above ground level upon descent. The drogue parachute was set to deploy upon barometric detection of apogee. The motor used was Cesaroni G185VMAX, with a 1.38 gram 4F black powder charge for backup ejection of the drogue, ignition time shaved to five seconds after burnout.




Potential Cause(s)

Potential Hazard(s)

Mitigation

Ease of recovery

Tree line Landing Weather cocking guided vehicle back to tree line

Personnel injury, damage to exposed components

Position launch rod optimally based on current wind speed and direction



113


Test Date December 21, 2012 Brief Description of Test: Full-scale dual deployment test launch at Hunewell Ranch. Altimeters used were one Stratologger SL100 for main parachute ejection, and one Featherweight Raven3 for drogue parachute ejection. The main parachute charge was 1.8 grams of packed 3F black powder, programmed for 500 feet above ground level deployment upon descent. The drogue was equipped with a 1.5 gram packed 3F black powder ejection charge, programmed for ignition upon accelerometer detection of apogee. Motor used was Cesaroni L1720.




Potential Cause(s)

Potential Hazard(s)

Mitigation


Main parachute black powder ejection charge

Separation of main parachute compartment from body

Mass of main parachute compartment components too great, black powder charge undersized

Ballistic impact, damage to upper body airframe and/or on-board electronic devices

Over-size black powder charges for main parachute ejection

Apogee

Launch stability Apogee of one mile AGL or greater not achieved

Launch vehicle mass, launch rod angle

Loss of points on competition launch day

Mass entire launch vehicle before simulation for launch day, position launch rod optimally based on current wind speed and direction

Ease of recovery

Tree line Landing Weather cocking guided vehicle back to tree line

Personnel injury, damage to exposed component

Position launch rod optimally based on current wind speed and direction



114


Test Date January 5, 2013 Brief Description of Test: Full-scale dual deployment test launch with active payload at Hunewell Ranch. One Stratologger SL100 for each secondary drogue and main parachute ejection and one Featherweight Raven3 for primary drogue ejection were installed. The main ejection charge was 5.8 grams. The primary drogue charge was 2.5 grams, the secondary 2.7 grams. The main parachute was set to deploy at 700 feet above ground level on descent. The drogue parachute was set to deploy upon barometric detection of apogee, based on which altimeter detected apogee first. The motor used was a Cesaroni L585.

Test Date January 5, 2013 Failure Mode



Potential Cause(s)

Potential Hazard(s)

Mitigation


Main parachute avionics bay Stratologger

Rapidly increased descent rate

Incorrect ejection charge connection

Entanglement, pressure spike in avionics bays, fire, ballistic impact

Lead recovery system engineer personally checks each connection prior to flight

Clear acrylic

Force of impact

Clear acrylic payload housing section fractured and shattered in places

Ballistic descent

Damaged SMD payload, loss of or unreliable flight data, personnel injury

Size of main parachute deployment bag (mpbd) increased, position of mpbd made adjacent to nose cone

Predicted apogee not achieved

Inconsistency between actual and simulated launch

Reliability of simulated data

Weight and position of objects in launch vehicle did not match those in simulation

Unreliable simulation for competition launch

Physically weigh and position each component for each vehicle prior to flight simulation



115


Test Date January 6, 2013 Brief Description of Test: Full-scale dual deployment test launch with active payload at Hunewell Ranch. Altimeters used were one Stratologger SL100 for each secondary drogue and main parachute ejection, along with one Featherweight Raven3 for primary drogue ejection. The main ejection charge was 5.8 grams. The primary drogue charge was 2.6 grams, the secondary 2.8 grams. The main parachute was set to deploy at 700 feet above ground level on descent. The drogue parachute was set to deploy on accelerometer-based or barometric detection of apogee, based on which altimeter detected apogee first. The motor used was a Cesaroni L585.




Potential Cause(s)

Potential Hazard(s)

Mitigation

Main parachute envelopment

Attachment point of main parachute deployment bag to parachute shock cord, use of pulling drogue parachute

Streamer effect

Pulling drogue parachute became entangled

High-speed impact, damage to components

Position main parachute deployment near nose cone, do not use pulling drogue parachute

GPS transmission

GPS transmitter antenna crimped to fit in main parachute avionics bay

GPS transmitter antenna disconnected from base of device and rendered non-reusable

Force of impact

Loss of tracking signal, difficulty in recovering launch vehicle

Attach GPS transmitter to main parachute shock cord with free antenna



116


Test Date January 7, 2013 Brief Description of Test: Full-scale dual deployment test launch with no payload at Hunewell Ranch. Altimeters used were one Stratologger SL100 for each secondary drogue and main parachute ejection, along with one Featherweight Raven3 for primary drogue ejection. The main ejection charge was 5.6 grams. The primary drogue charge was 2.7 grams, the secondary 3.5 grams. The main parachute was set to deploy at 700 feet above ground level on descent. The drogue parachute was set to deploy on accelerometer-based or barometric detection of apogee, based on which altimeter detected apogee first. The motor used was a Cesaroni L585.




Potential Cause(s)

Potential Hazard(s)

Mitigation


Main parachute avionics bay magnetic switch not armed

Rapidly increased descent rate

Personnel error

Ballistic impact, damage to components

Protocol modified to standardize arming of altimeters

Mission Performance Predictions

Mission Performance Criterion

The launch vehicle carrying the SMD payload has direct requirements that determine how it must perform during the mission. The design is intended to result in a final product capable of performing these given requirements. The project defines a successful mission as a flight with payload, where the vehicle and SMD payload are recovered and able to be reused on the day of the official launch. Moreover, the vehicle will not exceed 5,600 feet, and the official scoring altimeter will be intact and report the official altitude. After apogee and descent, the entire vehicle lands within 2,500 feet of the launch pad. If the vehicle performs such that all requirements are met, it has met all performance criteria.

Flight Profile Simulations, Altitude Predictions, Weights, and Actual Motor



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Thrust Curve The flight simulation for the final vehicle design with recovery and payload subsystems is obtained from OpenRocket and shown in Figure 79. With the vehicle profile and the selected L1720-WT motor, the predicted altitude of the official flight is 4852 feet AGL. A summary of the input parameters for the simulation is given in Figure 80. The lateral distance from the launch pad was approximately 1,050 feet.

Figure 79: Final Vehicle Simulation


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The vehicle weight including the motor is 37.125 pounds. This weight must be as accurate as possible to validate the simulation. Weight can be adjusted for an increase of up to 10 percent by means of the ballast system. A static motor test was conducted to obtain an actual motor thrust curve that is provided in Figure 81. As discussed previously, the actual motor thrust curve closely reflects the theoretical thrust curve provided by the motor manufacturer in shape.

Figure 80: Input Parameters for Final Simulation

Figure 81: L1720-WT Actual Thrust Curve


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Thoroughness and Validity of Analysis, Drag Assessment, and Scale Modeling Results

As the team continues testing with full scale launches, the differences between actual performance and simulated performance of the vehicle and recovery system will become clearer. It is very important to understand and analyze these differences in order to make the simulation more accurate and valuable for altitude predictions. All parameters must be selected such that they most accurately represent the actual parameters of the vehicle flight. Drag assessment for the launch vehicle concerns two primary areas: vehicle profile characteristics and recovery system components. The vehicle experiences drag due to the shape and surface area of the nose cone as well as the surface area of the fins due to thickness. To date, no theoretical calculations of the drag profile of the launch vehicle have been completed. All predictions for the flight performance of the vehicle have been based on OpenRocket simulations. The recovery system creates drag from the drogue and main parachutes. Theoretical calculations have been completed to estimate landing radius of the vehicle due to parachute deployment events. These calculations are compared to the simulated landing radius from OpenRocket. Fortunately, weather cocking, which further reduces the actual landing radius of the vehicle, is taken into account by the OpenRocket simulation. This yields a smaller landing radius than theoretical calculations. Although multiple sub-scale test launches were conducted, these were primarily to gain experience and understanding pertaining to the successful dual deployment of parachutes. All scale modeling results are from OpenRocket and provided above in the “Subscale Flight Results” section. Upon conducting further test launches, the goal is to minimize the difference between simulated performance and actual performance. Although the primary tool utilized for this purpose is a flight simulator, further modeling may be necessary to identify inconsistencies between simulated flight parameters and actual flight parameters. As discussed below, quantified differences between simulated and actual CG/CP calculations have been developed.

Stability Margin and the Actual CP and CG Relationships and Locations

The simulation software that predicts the rocket’s flight path provides simulated measurements for center of gravity (CG) and center of pressure (CP). These simulated values are:


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An itemized mass budget was built to show the absolute locations, lengths, and weights of each component of the rocket. Software was developed to sum the itemized weights on either side of a predefined CG value. After several iterations of the software and refinements of the defined CG value, the program suggested a theoretical CG of 158.733 centimeters to balance the rocket. Using generalized equations from the 2012 NASA Advanced Rocketry Workshop (ARW) handbook these measurements were calculated by hand. According to the handbook, certain conditions must be satisfied to ensure the Barrowman equations adequately model the theoretical CP. These conditions are as follows:

The angle of attack (α) of the rocket is near zero (less than 10 degrees)

The speed of the rocket is much less than the speed of sound

The air flow over the rocket is smooth and does not change rapidly

The rocket is thin compared to its length

The nose of the rocket comes smoothly to a point

The rocket is an axially symmetrical rigid body

The fins are thin flat plates

Since each condition is met by the rocket, the Barrowman equations should accurately represent the theoretical CP. These calculations are as follows: Nose cone:

where L denotes length Airframe: for all portions of the airframe Fins:

(

)

√ (

)

( )

√ (

)

(

)


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( )

(

)

where “N” denotes the number of fins, “S” denotes the height of each fin, “D” denotes the diameter of the rocket airframe, “A” denotes the length of the fin’s root cord, “B” denotes the length of the fin’s tip cord, and “L” denotes the length of the fin’s half cord. Combined:

Table 43 compares the calculated and simulated values of CG and CP. The small error values demonstrate that the simulated values are quite accurate. Incorporating these simulated values in further analysis of the rocket will introduce negligible errors.

Measurement Calculated Simulated Error

CG 158.732939 cm 160 cm 0.7982 %

CP 203.751825 cm 205 cm 0.6126 %

Payload Integration

Payload Integration Plan The payload is designed to integrate into the acrylic payload housing structure of the launch vehicle with ease. The payload is constructed on a payload framework which consists of a forward payload bulkhead, a rear payload bulkhead, and a rectangular aluminum frame. The rear payload bulkhead will be epoxied into the rear coupler (booster to payload housing) which houses the drogue avionics bay. A rectangular frame constructed from four aluminum angle rails will also attach to this rear payload bulkhead with three screws.

Table 42: Calculated versus Simulated CG and CP Measurements


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The forward payload bulkhead is epoxied into the forward coupler (payload housing to upper body airframe) which houses the main parachute avionics bay. This forward payload bulkhead has a recessed slot where the open end of the aluminum frame will seat into once the payload is installed. Payload preparations involve the installation of two microSD cards. One card is used for the payload SMD sensor data, and a second card is used for storing camera photos. The Adafruit 254 microSD card reader features a card locking system, thus preventing the need to manually secure the microSD cards in place. Next, the power sources must be activated to conduct a payload functionality test. This will occur through the use of a magnetic single pole single throw (SPST) switch and three single pole single throw (SPST-NO) relays. The three relays will be used to activate three independent power supplies. After power is activated, wireless telemetry and L.E.D. visual inspection will verify the functionality of the S.M.D. payload. Additionally, the GPS will need time to find tracking satellites and save their position information onboard the SMD GPS.

Figure 82: Rear Payload Bulkhead to Frame Connection


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After payload functionality is verified and GPS satellite tracking is achieved, the power will be deactivated to preserve battery life. The onboard video camera in the payload must be activated manually at this time to start collecting video of the flight. Because the camera is a self-contained system there is no option for activating externally, or remotely. Power activation is achieved by pressing a “power on” button located on the video camera. At this point, the payload is ready to be integrated into the launch vehicle and ready to be activated on the launch pad when required.

Figure 83: Telemetry Verification GUI


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Step Component Action Taken 1 SMD Payload Sensor Data Card

Reader MicroSD card installed

2 Camera Card Reader MicroSD card installed

3 Payload Power Up Magnetic Switch Activation

4 Payload Functionality Verification Telemetry and L.E.D. Visual Inspection

5 Payload Power Down Magnetic Switch De-Activation

6 Video Camera Activation On-Board Power Activation

7 Payload Framework installation Payload Installation

Payload Installation and Removal

Payload installation is achieved by inserting the rear coupler to the payload housing structure with the payload framework inserted from the lower portion of the payload housing structure.

Table 43: Payload Preparation Steps

Figure 84: SMD Payload


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The upper portion of the aluminum rails will be aligned and seated into the slotted portion of the forward payload bulkhead near the top of the payload housing. Then the coupler will be secured with four screws; from the exterior of the payload housing, through the coupler, and into the rear payload bulkhead. At this point the physical integration of payload is complete. Payload removal is achieved by removing four screws securing the rear coupler to the payload housing structure and removing the coupler from the payload housing structure.

Figure 85: SMD Payload attached with Avionic Bays.


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Step Procedure Action taken 1 Inserting Payload

Framework Slide Payload Framework into Payload Housing

2 Aligning aluminum Framework with static bulkhead slot

Rotate Payload Framework Until Aluminum Frame is Aligned

3 Seating Aluminum Framework into static bulkhead slot

Press Payload Framework into Forward Payload Bulkhead Slot

4 Securing Payload Framework

Install Four Screws From the Exterior of the Vehicle into the Rear Payload Bulkhead

Payload Interface Dimensions The payload framework consists of six components; two bulkheads (rear payload bulkhead and forward payload bulkhead) and four aluminum angle rails. The rear payload bulkhead is a cylindrical piece of PVC which is 1.0 inches in height with a radius of 2.55 inches. This will provide a precision fit with the interior of the rear coupler which will be epoxied into place. The static bulkhead is identical with the exception of a 0.55 inch wide by 0.55 inch deep slot oriented on the center of the bulkhead and continuing across the width of the bulkhead surface. This will provide a 0.1 inch tolerance between the slot and aluminum framework; ensuring a precision fit, while allowing easy installation. The aluminum frame is constructed from four pieces of aluminum angle as seen in Figure 86. The aluminum Angle is 0.5 inches wide x 0.5 inches height x 0.125 inches thickness. Once constructed, the aluminum frame is a rectangular box 29.0 inches long x 4.85 inches wide. The Angles are joined with aluminum welds.

Table 44: Payload Integration Steps


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Item Dimensions Bulkheads 1.0 inches Height

2.55 inches Radius

Static Bulkhead Slot 0.55 inches Height 0.55 inches Depth 4.875 inches Width

Aluminum Angle 0.5 inches Height 0.5 inches Width 0.125 inches Thickness

Long Rails 29.0 inches

Short Rails 4.875 inches

Rail Separation Distance (Inner)

3.875 inches

Rail Separation Distance (Outer)

4.875 inches

Figure 86: Aluminum Angle

Table 45: Payload Framework Dimensions


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Payload Element Compatibility To ensure compatibility, the payload is designed to meet the specifications of the internal payload housing structure. All payload circuits mounted on the framework are oriented in a manner to provide sufficient clearances between the payload housing structure’s internal diameter and the components themselves. Port holes are installed to the payload housing structure to allow ambient atmospheric pressures to equalize within the payload housing structure to allow accurate readings of the BMP180 pressure sensors. Clear acrylic was chosen as the material for the payload housing structure. This allows for use of the camera, solar irradiance, and UV radiation sensors from within the launch vehicle. The acrylic is UV-T specification to allow ~85% of UV light to pass through the material without being filtered.

Simplicity of Integration Procedure Physical integration of the payload consists of preparing the payload for operation, installing the payload framework into the payload housing structure, and activating the payload. Due to the payload’s internal operations design, the need for payload ejection during flight is eliminated, significantly increasing simplicity.

Launch Concerns and Operation Procedures

Launch procedures

A successful launch is possible if the following procedures and checklists are implemented to ensure maximum safety for all those involved with the project. These procedures and checklists have been developed from the advice of the team mentor along with component operator’s manuals. The subsystem leads ensure that the proper preparation and implementation of the procedures and checklists are followed. The procedures and checklists are categorized into two sections: Pre-Launch and Launch.

Pre-launch Checklists and Procedures:

Safety Materials Checklist

1. MSDS Binder

2. Operators Manuals Binder

3. Launch Procedures Binder


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Complete this checklist at the launch pad under the supervision of the Safety Officer. The following checklists are to be completed and passed before the vehicle is cleared to launch.

Safety Checklist

This checklist ensures that the instructions from the awarded flight waivers have been followed and the team is in accordance with the law.

1. The Lockheed Martin Fort Worth Flight Service Station has been contacted

and a Notice to Airmen has been issued.

2. The Fort Worth ARTC Glen Rose Supervisor has been contacted the day of

the scheduled launch.

3. All applicable operating limitations of Federal Aviation Regulations (FAR’s),

Title 14, Part 101, expect for parts (e,f,g) have been checked.

4. Ensure that operations will be conducted in accordance with all applicable

state and local ordinances.

5. http://www.aviationweather.gov/adds/metars/ has been checked for horizontal

visibility of more than five miles and cloud coverage is less than five tenths of the intended altitude of the vehicle.

6. Aircraft spotters have been assigned and informed of their job.

7. Make sure that a safe launch radius has been obtained before launch (300

feet or greater)

8. A list of authorized personnel for launch operations has been made.

9. The team has been briefed on the following, expected altitude, horizontal

visibility, and cloud coverage height, who the aircraft spotters are, what members are allowed in the flight operation area, and what the simulated landing radius of the vehicle will be given the different wind speeds.

Structure Preparation:

Structure Checklist

1. Nose Cone

2. Airframe

3. Couplers

4. Avionics Bay

5. Bulkheads

6. Payload Housing (Clear Acrylic)

7. Fins

http://www.aviationweather.gov/adds/metars/


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8. Rail Buttons

9. Motor Retainer

Recovery Procedures:

The purpose of this checklist is to ensure that all the supplies have been gathered so that recovery system can be prepared.

Materials and Components Checklist

1. Black Powder (3F)

2. Ejection Canisters

3. Dowel Rod (for packing charges

4. Lead wire

5. Batteries (nine-volt)

6. Gorilla Tape

7. Altimeter Units(s)

8. Main parachute

9. Main parachute deployment bag

10. Drogue parachute

11. Flame proof blankets

12. Recovery wadding (Dog Barf)

13. Main Shock Cord

14. Drogue Shock Cord

15. Garmin GPS tracker

16. Garmin GPS unit(s)

17. Garmin GPS car charger

Recovery Checklist:

1. Open proper software

2. Provide power to altimeter

3. Connect altimeter via USB cable

4. Configure axes (if not previously done)

5. Check the battery voltage (replace if below 5V)

6. Disable any unused outputs

7. Program used output(s) for custom altitude or automatic based on test

description

8. For custom altitude, vary height above ground level by increments of 32


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9. Clear any unused banks

10. Take screen shots of altimeter configuration and parameter page

11. Each 9V battery voltage is checked with voltmeter

12. Raven3 altimeters are wired properly and connections are securely fastened

(Figure 87)

13. Stratologger altimeters are properly wired and connections securely fastened

(Figure 87)

14. Ensure that the both avionics bays are in the proper orientation by checking

the labeling on the couplers.

15. Astro DC40 (dog tracker GPS) is securely fastened to drogue parachute

shock cord

16. Check all port holes for obstructions.

17. Pack ejection charges

18. Run the ejection charge leads into the bottom of the booster section

19. Lightly pack recovery wadding into lower portion of booster section

Figure 87: Altimeter Wiring Diagrams


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20. Pack shock cord and then drogue parachute into booster section

21. Load avionics sleds

22. Run extension wires through avionics bay couplers

23. Secure avionics bay lids

24. Connect ejection charges to avionics bay lid leads

25. Secure drogue parachute shock cord to eye-bolt on avionics bay

26. Connect booster section to payload section via coupler

27. Run the electronic match into nose cone section

28. Lightly pack recovery wadding into nose cone section

29. Pack main parachute into deployment bag

30. Pack main parachute and main shock cord into nose cone section

31. Secure main parachute shock cord to eye-bolt

32. Connect ejection charges to avionics bay lid leads

33. Connect nose cone section to payload section

34. Insert sheer pins into designated locations

Motor Preparation:

The purpose of this checklist is to ensure that all the supplies have been gathered in order to begin motor preparation.

Materials and Components Checklist

1. 75 mm motor casing

2. Nozzle Holder

3. Tracking Smoke Element

4. Forward Closure

5. Nozzle

6. Case Liner (phenolic tube)

7. Forward Insulator Disk

8. Tracking Smoke Insulator

9. Retaining Ring

10. Nozzle O-ring

11. Fuel Grains

12. Igniter

13. Grain spacer O-rings


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Motor Preparation Checklist (Pro75® High-Power Reloadable Rocket Motor Systems Appendix C)

1. All proper materials have been gathered.

2. Assembly Instructions are read one at a time and completed by a trained

individual (Tripoli level two certified).

3. Give Igniter to designated vehicle preparation personnel.

Launch Checklist and Procedures

Setup on Launcher Preparation

1. Launch Rail – Inspect launch rail for excessive corrosion or snags that would

risk the vehicle jamming on the rail.

2. Slide the vehicle down onto the rail until it is against the rest.

3. Arm payload electronics

4. Arm the altimeters

5. Listen for the correct series of beeps (Raven3)

Figure 88: Materials and Components (Image obtained from the Cesaroni Pro 75 mm Motor Assembly Kit Instructions)


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a. 1st beep is for the drogue parachute if it is a high tone then the wire is connected.

b. 2nd beep is for the main parachute if it is a high tone then the wire is connected.

c. 3rd and 4th beeps are the alternate ejection charges and are low tones if disconnected.

6. Look for correct series of LED flashing (Stratologger)

d. Three quick bursts in a row means ready to launch.

7. The Launch Control System (LCS) box must be switched OFF.

8. Strip 1” – 2” of the wire’s sheath to expose both wire cores.

9. Insert igniter fully into the vehicle motor and install nozzle cover.

10. Secure the igniter to the launch rail.

11. Short LCS circuit by tapping both alligator clips together.

12. Connect one wire core to each alligator clip wrapping the excess wire around

the clip.

13. Before returning to the Range Safety Officer, switch the LCS box ON if you

are the last person leaving the area.

14. Visually inspect the launch pad area to ensure the area is clear.

15. Check with designated aircraft spotters to see if any planes have been

located.

16. Once the all clear has been given from the Range Safety Officer then

proceed with the countdown.

17. Arm switch

18. Initiate launch

Checklist is completed by: [Name] _____________________________[Date_________________________

Troubleshooting:

No audible/visible signal from altimeter?

1. Try rearming magnetic switch 2. Cancel/postpone launch 3. Disarm avionics 4. Remove vehicle 5. Disassemble avionics bay 6. Inspect avionics 7. Once problem is recognized and corrected prepare vehicle for launch

No motor ignition?


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Then proceed with the following: 1. Wait one minute 2. Disarm launch control system 3. Disarm avionics bays 4. Pull out and visually inspect the igniter 5. Check the continuity of igniter 6. Check the voltage on the Launch Control System with voltage meter 7. Use spare igniter

In-Flight Inspection

In-Flight Checklist

1. Listen for firing of ejection charges

2. Visually track vehicle

3. Monitor Garmin Astro 320 (handheld GPS receiver) for post-flight recovery

Post-Flight Inspection

Post-Flight Checklist

1. Recover vehicle

2. Take pictures for analysis

3. Disarm all electronics

4. Inspect upper airframe

5. Inspect payload section

6. Inspect lower airframe

7. Inspect shock cords

8. Inspect parachutes and shroud lines

9. Remove motor and store for later cleaning

Travel

Travel Checklists

The following checklists are to be completed and passed before the vehicle is cleared to travel. Structure/Propulsion checklist:

1. Fins are undamaged.


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2. Rail buttons are undamaged.

3. Motor casing and reloads are safely stored for travel.

4. Couplers, Nose cone, and Airframe are secured to ensure they remain

undamaged. Payload Inspection Checklist:

1. Inspect the camera

2. Inspect radio antenna

3. Inspect GPS unit

4. Inspect the UV Sensor

5. Inspect the Solar Irradiance

6. Inspect the pressure sensors

7. Inspect the humidity sensors

8. Inspect MicroSD card writer

9. Inspect the LCD screen

10. Inspect the video camera

11. Inspect the servos

12. Inspect the accelerometer

13. Power on the system to ensure that it is functioning properly and the LCD is

displaying the proper operating parameters.

14. Payload is properly assembled into the payload housing and is structurally

ready to travel. Avionics Checklist

1. Raven3 altimeter to ensure no damage has incurred to it and its wired

connections.

2. Stratologger altimeter to ensure no damage has incurred to it and its wired

connections.

3. Check Garmin dog tracking system

Recovery Checklist

1. Inspect Kevlar Shock Cord

2. Inspect Drogue Parachute

3. Inspect Drogue Parachute Protector

4. Inspect Main Parachute

5. Inspect Main Parachute Protector

6. Inspect E-matches

7. Inspect Flame Retardant Material

Checklist is completed by:


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[Name] _____________________________[Date]_________________________

Safety and Environment The Safety Officer The team safety officer, Blake, is level one certified with NAR. He has obtained an FAA flight waiver in his name for full scale launches. The responsibility of the safety officer is to design and implement safety plans that ensure all accidents are evaded. All hazards to people, the project, and the mission are determined so that mitigations can be enacted. The systematic identification of risks, failure modes, and personnel hazards allows the team to discover where single points of failure could occur throughout the project. The identification of single point failures allows for proactive design changes to counter these failures leading up to the CDR. The team has flight waivers from the FAA, allowing the team to launch at will. This gives the team a better opportunity to conduct a high number of full scale launches in preparation for the Flight Readiness Review. This allows practice and exercise with launch procedures, operations, and protocols. This reduces the risk of failure and promotes safety at the day of the official launch.

Failure Modes

A failure mode is the way in which a system could fail, causing an undesirable effect on some aspect of the project. The safety plan ensures development and implementation of mitigations for each failure mode. Each failure is listed with a resulting effect to the design. Proposed mitigations to prevent the failure are included along with the status of completion for that mitigation.

Launch Vehicle Design Failure Modes

Potential failure modes during the design of the vehicle are summarized in Table 47.


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Failure Effect Proposed Mitigation

Completed Mitigation

Vehicle Unstable Unpredictable Flight Path

Simulations Completed

Acrylics Does Not Withstand Forces Throughout Flight

Vehicle Not Reusable

Tensile Strength and Flight Testing

Completed

Fiberglass Does not Withstand Forces Throughout Flight

Vehicle Not Reusable

Tensile Strength and Flight Testing

Completed

Connection Between Acrylic Payload Housing and Upper Fiberglass Body Tube Becomes Detached

Unpredictable Flight Path, Damage to Vehicle Body

Research/ Design

Completed

Fins Cause To Much Drag

Expected Apogee Height Not Obtained

Simulations Completed

Thrust To Weight Ratio is Less Than 5:1

Unpredictable Flight Path

Simulations/ Calculations

Completed

Couplers Too Long or Too Short

Early or No Separation

Research/ Simulations/ Calculations

Completed

Table 46: Potential Failure Modes for Design of the Vehicle


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Payload Integration Failure Modes

Potential failure modes that could occur during payload integration are summarized in Table 48 provides an updated summary of.

Failure Effect Proposed Mitigation Completed Mitigation

Screw hole stripped out

Inadequately secured payload

Ensure the bulkhead is replaced when required

Completed

Incompatible hardware

Payload will not integrate properly

Ensure precision of fit during manufacturing

Completed

Components damaged during integration

Electronic malfunction

Be careful while inserting payload

Completed

Table 47: Potential Failure Modes during Payload Integration


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Launch Operations Failure Modes

Potential failure modes that could occur during launch operations are summarized in Table 49.

Subsystem Potential Failure Mode

Potential Effects of Failure

Proposed Mitigations

Completed Mitigations

Structure

Ejection charges damage air-frame/vehicle components

Critical systems become damaged

Calculations Completed

Proper testing Completed (12/21/2012)

Motor mount fails to properly retain motor

Damage to internal systems

Structural testing of the motor mount

Completed (10/27/2012

Rail button failure

Unpredictable flight path

Ensure rail buttons are properly installed and orientated

Completed (12/21/2012)

Insufficient component mounting

Potential system malfunction

Test mounting integrity,


Airframe stress failure

Loss of vehicle functionality, potential loss of vehicle

Structural testing of airframe


Fin Detachment Aerodynamic instability of vehicle

Ensure that the fins are properly epoxied




Proposed Mitigation


Recovery

Parachute shroud line fails

Uncontrollable descent

Research Completed

Verify parachute rating


Electronic matches do not fire

Parachute deployment does not occur

Redundant altimeter system

Completed

Eye bolt failure Uncontrollable descent

Verify eye bolt integrity



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Shock cord failure

Untethered vehicle components, violation of requirements

Properly fastened shock cord


Verify rating Completed (12/21/2012)

Premature black power ignition

Premature parachute ejection

Testing Completed (10/27/2012)

Recovery altimeter shielding

Completed

Recovery system ignites

Failure of recovery system, damage or loss to vehicle

Use Nomex cloth and fire retardant insulation


Main or drogue parachute comes untied from the swivel

Uncontrolled descent

Secure swivels along with quick links.


Main or drogue parachute shrouds become entangled

Uncontrolled descent rate

Ensure parachute shroud lines are attached to a swivel


Failed Separation

Un-functional recovery system, ballistic descent

Adequate separation testing


PerfectFlites power supply diminishes

failure of deployment of parachutes, Mission failure

Use new batteries before launch


Featherweight Power supply diminishes


Use new batteries before launch


PerfectFlite wired connections become damaged from handling


Use protected electrical components

Completed


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Featherweight wired connections become damaged from handling


Use protected electrical components

Completed

Scoring Altimeter failure

We lose all points associated with the altitude portion of the project

Redundant Systems

Completed



Proposed Mitigation


Propulsion

Igniter does not initiate the oxidation process for the propellant

The vehicle does not launch

Inspect igniter for concatenation

Completed

Always bring additional igniters for such an event

Completed

Propellant’s oxidation process does not commence

The vehicle does not launch

Use proper igniter, sue appropriate conditions when storing propellant

Completed

A pressure build-up occurs inside the motor

Explosion Inspect the motor

Completed

Hazard Analysis

Potential hazards to personnel through the course of the project are provided in Table 50. Personnel hazards refer to potential harm incurred by any individual. The development and implementation of the safety plan and protocols ensure that these hazards are appropriately mitigated.

Table 49: Potential Failure Modes during Launch


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Risk Sources Likelihood

Consequence

Mitigation Action

Laceration

Knives, routers, saws, file, Dremel tool

Medium Serious injury or death

Follow safety protocols, proper tool and equipment use, personal safety attire, refer to operators manual

Discontinue all operations, apply first aid, contact EMS

Burns

Chemicals (FFFg, fiberglass resin), welders, soldering Iron

Medium Minor to serious injury



Respiratory Damage

Chemicals (epoxy, solder), fumes, fiberglass

Low Brain damage or death

Follow safety protocols, proper tool and equipment use, personal safety attire, consult MSDS


Vision Damage

Welders, fiberglass, grinders, projectile debris

Low Partial to complete blindness

Use of goggles, force shields, consult MSDS, first aid kit available, refer to operators manual

Discontinue all operations, apply first aid, contact EMS, use eyewash


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Consequence

Mitigation Action

Allergic Reaction

Epoxy, chemicals, fiberglass

Low

Loss of respiration, inflammation (Internal & External)

Use of gloves, consult MSDS, first aid kit available

Discontinue all operations, apply first aid, contact EMS, administer antihistamine, safety shower

Hearing Damage

FFFg, Grinders, Ignition, Routers

Low Partial to complete deafness

Ear muffs, consult MSDS, first aid kit available, refer to operators manual


Dismemberment

Projectiles, Saws, Launches

Low Permanent injury or death

Make sure proper safety measures are taken, operators manual

Discontinue all operations, apply first aid, and contact EMS, tourniquet

The material safety data sheet (MSDS) that the manufacturer provides contains information about the material in consideration. It is comprised of 16 categories: identification, hazard(s) identification, composition/information on ingredients, first-aid measures, fire-fighting measures, accidental release measures, handling and storage, exposure controls/protection, physical and chemical properties, stability and reactivity, toxicological information, ecological information, disposal information, transport information, and regulatory information. MSDSs are referred to when a hazard occurs in order to enact the most effective mitigation. All team members shall be knowledgeable of the MSDS associated with each hazardous material. According to the safety plan, a binder containing all the MSDSs is always made available for personnel and brought to every launch. Operator manuals for each tool will be consistently referenced prior to each tool’s usage. This ensures each tool is used as intended. According to the safety plan, operator manuals for each component used during the project are kept in an operator manual binder. These documents will be made available by the safety officer at any

Table 48: Potential Hazards to Personnel


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location in which construction, testing, or launching of the vehicle could occur. It is important for all team members to be thoroughly briefed on the project risks, FAA laws and regulations regarding the use of airspace, and the NAR high-power safety code. The team is aware that the FAA must be notified of planned launch activities. For educational outreach events, notification to the closest airport within five miles of the launch site is required 72 hours prior to launch. For subscale launches, flight waivers are required to be obtained at least 45 days prior to the proposed activity. The team has obtained flight waivers for full scale launches. The flight waiver went into effect on December 15, 2012, and it lasts for one year. The conditions allow for CFR 101.25 (e) and 101.25 (f,g), which is located in Appendix D to be waived. Flight operations occur between sunrise and sunset within a controlled airspace from the surface to 5,280 feet above ground level, 6,569 feet mean sea level (MSL). Operations are coordinated prior to launch dates. The launch location is within a one nautical mile radius of 32.212954N/-0.98.092861W, on the Hunewell Ranch in Stephenville, Texas. To ensure that the conditions from the flight waivers are followed, a procedural checklist has been devised and implemented along with a pre-mission launch briefing which occurs prior to every launch. The flight waivers are located inside the launch procedures binder, which is brought to every launch. (Reference in Appendix E) The National Association of Rocketry and Tripoli are recognized as the primary rocketry associations of the United States. As such, their standards establish precedence throughout high powered model rocketry. Along with these standards, the team is cognizant of all federal, state, and local laws regarding unmanned vehicle launches and motor handling including the following regulations: CFR 101, Subchapter F, Subpart C: Amateur Rockets (Located in Appendix D) CFR Part 55: Commerce in Explosives (Located Appendix D) Handling and Use of Low-explosives Ammonium Perchlorate Rocket Motors (APCP) (Located in Appendix I.12) NAR Model Rocket Safety Code (Located in Appendix F) Hazardous Waste Management (Located in Appendix E) Fire Safety (Located in Appendix G) Lab Safety (Located in Appendix H) Launch Procedures and Checklists (Located in the Launch Procedures Section on Page XX) A summary of legal risks that could occur during the course of the project appears in Table 51.


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Risk Likelihood Severity Consequence Mitigation

FAA Violations

Low High Legal Repercussions

Adhering to Regulations

NAR/TRA Violations



Damage of Property


Insurance

OSHA Violations



Personal Injury


Redundant Calculations and Safety Preparedness

Environment

Environmental effects of the project

In the event of an unrecoverable or damaged vehicle, certain materials could be left exposed to the environment. The biodegradability of each material used effects the impact on the surrounding ecosystem. Much of the information concerning the hazards posed to the environment and ecology is available in the individual MSDSs. The effects of materials used in the construction and launch of the vehicle are summarized in the following Table 52.

Material Prevalence Mode of Biodegradability

Impact on Environment

Ammonium Perchlorate

Motor propellant Highly water soluble Iodization of local water table

Black Powder Ejection charges Remains solid No known impact

Epoxy 2M DP420

Fiberglass connections, sealed couplings

Decomposition begins within fifteen months

Leaching to local water table

Oil-Based Spray Paint:

External fiberglass structural components

Soluble Leaching to local water table

Clear Acrylic: Payload bay Soluble Leaching to local water table

Table 49: Summary of Legal Risks


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Material Prevalence Mode of Biodegradability

Impact on Environment

Fiberglass Structural components

Remains solid

No known environmental impact; may pose ecological hazard

Nomex Deployment bag Remains solid

No known environmental impact; may pose ecological hazard

Aluminum Motor tube, rivets, battery casing,

Highly water soluble Long-term degradation products

Cellulose Recovery wadding;

Remains solid Long-term degradation products

Steel Attachment hardware, ballast system;

Remains solid Leaching to local water table

Copper: Avionics bay lining, e-match lead wires

Highly reactive in air or moisture

Long-term degradation products

Sulfuric Acid Batteries Highly water soluble Long-term degradation products; may pose ecological hazard

Kevlar Shock harnesses Remains solid No known impact

Silicon Parachutes, batteries

Reactive in air or moisture

Irritating vapors form

Rip-Stop Nylon: Parachutes, shear pins

Remains solid No known impact

Environmental effect on the project

While some aspects of the project may adversely affect the surrounding environment, the environment can also have an impact upon the project. As primary test launches will take place in Texas during winter and spring, inclement weather will likely fall on test dates. In response to unforeseen issues in the weather, flight waivers have been obtained so that alternate test dates are easily rescheduled. Launch dates can be viewed in the testing timeline in Figure 157. Environmental factors such as surrounding flora, fauna, or sedimentary projections could cause the launch vehicle to become unrecoverable. These risks are outlined in Table 53.

Table 50: Effects of Materials used in Construction and Launch


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Risk Likelihood Severity Consequence Mitigation

Poor weather High High Delay in testing

Multiple test dates, Obtained Flight Waivers

Environment prevents recovery

Medium Medium Possible loss of Vehicle

Survey launch site, recovery tools

Burn ban in effect

Low High Delay in testing Multiple test locations

Table 51: Environmental Factors


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IV) Payload Criteria


Testing and Design of Payload Experiment

Design Review at a System Level

The payload design fulfills the requirements of the SMD payload. The payload consists of three main systems; the Atmospherics Data Gathering System (ADGS), the Autonomous Real-Time Camera Orientation System (ARTCOS), and the Video Capture System (VCS). The payload records measurements of pressure, temperature, relative humidity, solar irradiance and ultraviolet radiation; these measurements are stored onboard to a micro SD card. The payload does not eject from the vehicle, but rather takes all readings internally through a clear acrylic housing. The electronics are mounted on an aluminum rail framework. The camera is kept level throughout the descent phase of the flight by the ARTCOS, which includes a dual servo motor mechanism. Camera images are stored to a micro SD card. A 900 megahertz XBee S3B transmitter allows a 28 mile range of wireless transmission. A custom Printed Circuit Board (PCB) minimizes the space utilized by the electronics and improves the signal integrity between the components. Advanced Circuits in Aurora, Colorado has sponsored the PCB manufacturing. A video camera records flight footage for public outreach.

Drawings and Specifications Detailed schematics of the payload framework and electronics are contained in Figures 89-92 containing exact dimensions. The first prototype of the overall payload with installed components is seen in Figure 89.

Figure 89: Payload


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The first prototype ADGS consisting of a hygrometer, thermometer, pyranometer, barometric pressure sensor, GPS, lux sensor, UV sensor, wireless transmitter, and a micro SD card writer is viewable in Figure 90.

The first generation Autonomous Real-Time Camera Orientation System (ARTCOS) consisting of two servo motors, an accelerometer, and a camera is represented in Figure 92.

Figure 90: Upper Payload Circuit Boards

Figure 91: UV Sensor Mounting


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Analysis Results

Analysis of the payload’s preliminary design led to several changes. Upon ordering, several parts were deemed difficult to obtain and were removed from the design as a result including the secondary pressure sensor and the secondary UV sensor. The video camera chosen initially for the VCS was deemed difficult to obtain, so the group chose go with the Keyfob 808 Version 20. The ARTCOS requires a second microprocessor solely dedicated to controlling the servos, while the first microprocessor is dedicated to saving the pictures. When the camera is saving to the micro SD card, the microprocessor is not able to process other tasks, and it takes 15 seconds to save a picture.

Test Results

The HH10D relative humidity sensor was removed from the design as a result of testing. It outputs a digital frequency calculated through a clock signal provided by the Arduino Mega. The Arduino Mega has only one clock signal. The clock signal is also used for timing other purposes. In order to use the Arduino’s clock signal for other purposes, a new clock signal would have to be added to the circuit. The extra clock signal would add extra hardware and complexity to the circuit.

Full Scale Flight Test

On January 5, 2013, the first prototype of the payload design was launched in a full scale vehicle with a half scale L585 motor. The main parachute did not deploy correctly; therefore, the rough landing broke the payload housing and damaged several electrical components. While the landing was not ideal, the severity of the landing showed which components needed better mounting. In this test, the electronics were

Figure 92: ARTCOS


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wired in their perforated circuit boards configuration. The following paragraphs describe the successes and failures of the test launch and the changes made to the design. A major success of the test was the telemetry. The sensor values were transmitted from an onboard XBee to a ground station XBee. The ground station XBee used a high gain Yagi antenna. The vehicle remained on the launch pad for two hours, and the telemetry continued the entire time. The data was formatted into an ASCII encoded string of values. The ground station XBee, connected to a laptop, displayed the telemetry string on the default XBee software. The MATLAB ground station GUI is still being developed and will be used in the final implementation to the telemetry system. The string parsed the sensor values in the following order: humidity, lux, temperature, pressure, altitude in meters, GPS time, latitude, longitude, mean sea level (MSL) altitude in meters, number of satellites tracking the GPS, front ultraviolet radiation, back ultraviolet radiation, front solar irradiance, and back solar irradiance. Upon landing, the onboard XBee was dislodged from the XBee spacing adapter causing telemetry to stop. The solution to this problem is to secure the XBee with a zip-tie. Another major success of the test flight was storage of data to the micro SD card. Each telemetry string was saved to the micro SD card. Although the micro SD card also came dislodged upon landing, flight data was still obtained as seen in the screenshot of flight data from Figure 93.


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Figure 93: Test Flight Data


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To avoid using floating point numbers, several sensor values were multiplied by a factor of 10. The humidity, pressure, ultraviolet radiation, and solar irradiance values must be divided by 100 to achieve the actual value. The temperature and altitude must be divided by 10 to achieve the actual value. A few of the measurements were recorded in metric units; however, for the next prototype all the sensor measurements will be recorded in imperial units. The sensors transmitted their readings to the central microcontroller, Arduino Mega 2560, which then saved the readings to the micro SD card. The humidity sensor used was the Honeywell HIH4030. The sensor was ordered from Sparkfun pre-mounted on a breakout board. The humidity readings appear to drop throughout the flight as seen in Figure 94.

The vehicle was on the launch pad for two hours. The data gathered by the humidity sensor during this time shows an unexpected trend. According to the readings, the humidity inside the clear acrylic housing dropped about 35 percent within the first 30 minutes of sitting on the launch pad according to Figure 95.

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Figure 94: Test Flight Humidity Data


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The lux sensor used was the TAOS TSL2561. This sensor was ordered from Adafruit pre-mounted on a breakout board. The value of the sensor did not change throughout the flight; therefore, the sensor may have malfunctioned or the clear acrylic may have interfered with sensor readings. The sensor worked correctly when tested at the lab post-flight. More testing will be performed to determine the functionality of the TSL2561 inside the clear acrylic housing. The temperature was measured by a BOSCH BMP180. The raw sensor data saved to the micro SD card must be divided by 10 and then inserted into the following formula, the conversion factor between meters and feet, to achieve a correct value for pressure

in imperial units:

The adjusted temperature reading for the flight is

available in Figure 96. The temperature is much higher than the temperature for that day. A local weather station recorded the temperature for the time of the flight to be 55.4 degrees Fahrenheit, whereas the BMP recorded an average temperature of 95 degrees Fahrenheit.

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Figure 95: Launch Pad Humidity Data


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The temperature data for the time the vehicle was on the launch pad shows that the measured temperature inside the payload increased dramatically as seen in Figure 97.

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Launch Pad Temperature (oF)

Figure 96: Test Flight Temperature Data

Figure 97: Launch Pad Temperature Data


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The BMP180 is a light sensitive piezoelectric device. The hypothesis is that because the BMP180 was in direct sunlight for an extended period of time, the recorded measurements were adversely affected. The design change is to place the BMP180 in between the front and back electronic boards. The sensor will still be exposed to ambient pressure inside the payload, but it will no longer be exposed to direct sunlight. Humidity is directly related to temperature; therefore, the spike in temperature values may explain the drop in humidity. The correlation between temperature and humidity from the test flight data are displayed in Figure 98. As temperature increases, humidity decreases.

The pressure was measured by a BOSCH BMP180. The raw sensor data saved to the micro SD card must be divided by 100 and multiplied by .0293, the conversion factor between hectoPascals and inches of Mercury, to achieve a correct value for pressure in imperial units. The adjusted flight data for pressure in relation to time is available in Figure 99. As shown in the graph, the pressure begins to decrease at launch, and once the vehicle reaches apogee and begins descent, the pressure increases.

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Figure 98: Correlation between Temperature and Humidity

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The onboard flight computer software calculates altitude using the BMP180 pressure measurements as seen in the test flight data of Figure 100.

The GPS used in the test flight was a Locosys LS20031. The GPS recorded readings of latitude, longitude, and altitude. The GPS data is taken and formatted into a KML document, then opened in Google Earth as in Figure 101. The figure shows a three

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Figure 99: Test Flight Pressure Data

Figure 100: Test Flight Altitude Data


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dimensional graph of the flight on Google Earth. The flight took place at Honeywell ranch in Stephenville.

The UV sensor used was the Apogee Instruments SU-100. The measurements are recorded in the units of µmol m-2 s-1 and must be divided by 100 to achieve the actual value. There were two UV sensors, one mounted on the back of the payload circuit boards and one mounted on the front. Both UV sensors recorded basically no data. The hypothesis is that the clear acrylic blocked the UV radiation. This will be tested further to verify acrylic interference with UV radiation measurements by the SU-100. The final solution is to use a UV-T acrylic which does not block all UV radiation. The solar irradiance sensor used was the Apogee Instruments SP-110. The measurements are recorded in the units of W m-2 and must be divided by 100 to achieve the actual value. There was one SU-100 mounted on the back of the payload circuit boards and one mounted on the front. This positioning deviates from the final design in that there are four solar irradiance sensors mounted 90 degrees from each other. The front-to-back positioning is undesirable; because, when the payload is sideways relative to the sun (i.e. not facing the direction of the sun) neither sensor is taking accurate readings. The SP-110 accuracy is discussed in the precision of instrumentation section. The readings from the two solar irradiance sensors are available in Figure 102. The red line represents the front sensor and the blue line represents the back sensor. The analysis includes taking the maximum of the solar

Figure 101: Test Flight GPS Data


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irradiance sensors for each time step. With four sensors oriented facing every 90 degrees, the hypothesis is that the readings will be better correlated to altitude.

The ARTCOS was tested in the flight as well. The servos were operating, and the camera was taking pictures for about an hour and a half before flight; however, just before flight the epoxy holding the two servos together failed. The servo connected to the camera separated from the servo connected to the aluminum rails. In the next prototype a different type of epoxy will be tested. Another failure of the ARTCOS was the software for the camera was written such that the camera only saved the first one hundred pictures to the micro SD card. Due to the extended time the payload remained fully active on the launch pad coupled with the fact that the camera was not properly mounted, there were no images taken of the flight. A beneficial result of the test is that the pre-flight images show that the camera images were slightly tilted as seen in Figure 103. The hypothesis is that the ADXL was unintentionally mounted slightly offset. The accelerometer will be remounted, and the software will correct for any crooked mounting.

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Figure 102: Test Flight Solar Irradiance Data

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The video camera used in the test flight was the Hack HD 1080p Sparkfun video camera. There was video recorded pre-flight, but upon landing, the micro SD card on the video camera module came dislodged. The video of the flight did not save. Also, upon landing, the lens on the video camera broke, and the video camera no longer works. The video camera used costs approximately $180. After the flight, a different video camera has been chosen. The selected video camera is the Keyfob 808 video camera. It is self-contained, smaller, and much cheaper at $20.

System Level Functional Requirements

The functional requirements of the payload and the design feature that satisfies each requirement are compiled in Table 54.

Figure 103: ARTCOS Image


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SOW #

Requirement Satisfying Design Feature

Verification Method

3.1.3.1

UV Radiation SU-100

Lab Testing, Static and Flight Testing, Analysis, Demonstration

Solar Irradiance TSL2561, SP-110

Humidity HIH4030

Temperature BMP180

Pressure BMP180

3.1.3.2

0.2Hz Data During Descent

16MHz Arduino Mega 2560, Software

Lab Testing, Analysis

3.1.3.3

0.016Hz Data After Landing

16MHz Arduino Mega 2560, Software

Lab Testing, Analysis

3.1.3.4

Post-Landing Data Termination

Software Lab Testing, Flight Testing, Analysis

3.1.3.5

2 Descent Pictures

VC0706, Keyfob Video Camera Static and Flight Testing,

Analysis, Demonstration 3 Landing Pictures

VC0706, Keyfob Video Camera

3.1.3.6

Horizon Orientation

ARTCOS Lab Testing, Static and Flight Testing, Analysis

3.1.3.7

Onboard Data Storage

Micro SD Lab Testing, Static and Flight Testing, Analysis

Data Transmission

900MHz XBee Radios Range Testing, Flight Testing

3.1.3.8

Apogee Separation

No Separation, Clear Acrylic Housing

Inspection

3.1.3.9

GPS LS20031 Lab Testing, Static and Flight Testing, Analysis

3.1.3.10

2500ft Min. Separation


Inspection

3.2

Scientific Method Scheduling, Analysis, Testing, Documentation

Inspection

3.3

UAV Not Applicable, SMD Selection

Not Applicable


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SOW #

Requirement Satisfying Design Feature

Verification Method

3.4

Jettisoned Components


Inspection

3.5

Recoverable and Reusable

Aluminum Framework, No Separation

Lab Testing, Static and Flight Testing, Analysis

Approach to Workmanship

Construction of the payload requires coordination of several sub-teams. The structural team designs a framework to ensure compatibility between the payload and the launch vehicle. The Electronics Hardware Team designs a power system to power the payload and verifies the physical wiring and interfacing of the components. This involves the planning and designing of wiring schematics for connecting the various components of the payload as well as the power management system to ensure that all components have adequate power and proper voltages to operate. Additionally, the Electronics Hardware Team designs the board layouts and solders the electrical connections between components. The Electrical Software Team programs the payload and ensures functionality of the sensors and components. This includes compiling function libraries for all the sensors to ensure functionality. It also includes creating code to allow them all to be used as a system. Interface software includes the use of I2C, SPI, and USART.

Test Plan of Components and Functionality

Testing began on a breadboard. Each component was tested individually in order to determine the proper wiring and software. Next, the components were integrated together on the breadboard to determine the correct wiring for the components to operate together. Multimeters were used to determine if the components were receiving the correct voltage and amperage, and also to verify that the microcontrollers contained the correct voltage regulators. Verification of component functionality was measured by three parameters; hardware configuration, software configuration, and data interpretation. The following paragraphs specify each component’s hardware configuration, software, and data interpretation.

BMP180

The pressure and temperature sensor is a BOSCH BMP180. The BMP180 uses a piezoresistive sensor to detect applied pressure at a relative accuracy of plus or minus 0.017 pounds per square inch (psi) and plus or minus 1.8 degrees Fahrenheit. The piezoresistive sensor outputs an analog value, which is then converted to a 16 bit digital value via an analog to digital converter. The BMP180 communicates with the Arduino

Table 52: Payload Functional Requirements


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Mega 2560 over the I2C data bus and requires a 3.3 volt input voltage. The 3.3 volt power is provided by a nine-volt power regulated by a buck converter. The proper wiring of the sensor determined through breadboard testing is demonstrated in Figure 104.

The software used to read the BMP180 and convert the raw reading into usable pressure and temperature data is written in Arduino C. The data sheet for the BMP180 lists the conversion algorithms. A flow chart from the BMP180 datasheet that represents how to obtain pressure and temperature readings from the sensor are viewable in Figure 105. This flow chart has been determined to be accurate through testing.

Figure 104: BMP 180 Pressure Sensor Wiring


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Figure 105: BMP 180 Software Flowchart


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TSL2561 One pyranometer included in the payload design is the TAOS TSL2561. The TSL2561 measures the spectral range from 300 to 1100 nanometers. The sensor includes two photodiodes. The two photodiodes output an analog value which is then converted to a digital value by onboard ADCs; therefore, the TSL2561 outputs a 16 bit digital value. As stated in the TSL2561 datasheet, a digital output minimizes noise interference. While the sensor’s formal use is an ambient light sensor, as stated in the datasheet, the raw output of the sensor is irradiance. The sensor communicates with the Arduino Mega 2560 over the I2C data interface bus and requires a 3.3 volt power supply. The 3.3 volt power supply is provided by a nine-volt battery regulated by a buck converter. Through the use of the I2C interface, each individual photodiode is read separately. The correct wiring as determined through testing as shown in Figure 106.

The software for the TSL2561 was determined through analysis of the datasheet. The datasheet presents pseudo code for reading the TSL2561 registers. The pseudo code from the datasheet is in Figure 107.

Figure 106: TSL2561 Pyranometer Wiring


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The raw data from theTSL2561 registers must be converted in order for them to represent actual lux values. The datasheet presents the conversion factors used to calculate lux from the raw data. The conversion factors from the TSL2561 datasheet are in Figure 108. CH1 and CH0 are register values.

The TSL2561 and the BMP180 were then integrated together. Information found at the I2C website showed that 4.7 kilo-ohm pull-up resistors needed to be added to the Serial Clock and Serial Data lines of the bus when multiple devices are on the bus. The proper wiring for the TSL2561 and BMP180 on the same circuit is shown in Figure 109.

Figure 107: TSL2561 Pseudo Code

Figure 108: TSL2561 Lux Conversion Factors

http://www.i2c-bus.org/


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HIH4030 The humidity sensor is the Honeywell HIH4030. The HIH4030 is an analog sensor. The HIH4030 sensor uses a thermoset polymer capacitive sensing element to measure relative humidity to an accuracy of plus or minus 3.6 percent. For prototyping purposes, Sparkfun retails a sensor preinstalled on a breakout board that meets the necessary specifications. The breakout board allowed through-hole testing as opposed to surface-mount testing. The HIH4030 requires a five-volt power source to operate, which is provided by a nine-volt battery regulated by a buck converter. The HIH4030 outputs an analog signal between zero and five volts. This output signal wire is connected to an analog in pin of the Arduino Mega. The proper wiring of the HIH4030 as determined through testing is represented in Figure 110.

The software for the HIH4030 takes the analog input value and uses a conversion factor to convert from the raw reading to the actual relative humidity value. This conversion factor is found in the datasheet. The conversion factor also utilizes the temperature reading from the BMP180 to acquire an even more accurate relative humidity reading. After analyzing the datasheet, the following function was formulated to calculate the relative humidity from the HIH4030. The function is written in Arduino C. The code includes the conversion factor.

Figure 109: BMP 180 and TSL2561 Wiring

Figure 110: HIH4030 Humidity Sensor Wiring


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Integrating the HIH4030 with the BMP180 and TSL2561 requires no extra hardware. The devices do not utilize the same data busses. The proper wiring after integrating the HIH4030 with the other two sensors is shown in Figure 112.

HH10D The HH10D is a relative humidity sensor; it has been removed from the design, but testing still occurred and is described here. The sensor outputs a digital frequency signal. There are also two calibration values stored in the EEPROM which must be read from the I2C data bus. The digital frequency signal is simply read by the Arduino Mega through a digital I/O pin. A schematic showing the proper wiring of the HH10D alone on a circuit appears in Figure 113.

Figure 111: HIH4030 Software

Figure 112: HIH4030, BMP180, and TSL2561 Wiring


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In order to get a relative humidity reading from the HH10D, the two calibration values must be read from the EEPROM at the start of the program. Then, the frequency signal must be polled. The frequency value along with the calibration values are used in a formula specified in the datasheet. The calculation algorithm and EEPROM values are in Figure 114.

Integrating the HH10D into the circuit with the HIH4030, BMP180, and TSL2561 simply involves connecting the I2C lines and the digital frequency line to the Arduino Mega. To simplify the connections even further, since the calibration values do not change, the I2C connections can be disconnected after reading the calibration values the first time. A schematic showing the wiring of the four sensors is in Figure 115.

Figure 113: HH10D Humidity Sensor Wiring

Figure 114: HH10D Humidity Calculation Algorithm


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As mentioned previously, the HH10D has been removed from the design; because, the extra clock signal needed to measure the frequency made the circuitry overly complicated.

SU-100 The UV sensor is the Apogee Instruments SU-100. The SU-100 measures the light spectrum from 250 nanometers to 400 nanometers. A protective dome houses the sensor. The SU-100 requires no voltage source, but it accumulates power from the sun. The sensor outputs a voltage which represents the level of ultraviolet radiation. This output voltage is read into the Arduino Mega through an analog pin. There are multiple ways in which to wire the SU-100. The sensor has three wires: negative, positive, and ground. One way to use the SU-100 is to subtract the negative voltage from the positive voltage using the difference in an algorithm; this is known as a differential measurement. The second way to use the SU-100 is to connect the negative to the ground and read the output from the positive signal; this is known as a single-ended measurement. Testing was performed with a voltmeter as shown in Figure 116.

Figure 115: HH10D, HIH4030, BMP180, and TSL2561 Wiring


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Through testing, the single-ended measurement gives a more accurate reading. Therefore, this is the wiring used in the design. The wiring of the SU-100 is shown in Figure 117.

When forming the software, the first issue is the reference voltage of the analog pins of the Arduino Mega. The default reference voltage is five volts, and the analog to digital converter on the Arduino Mega is 10 bits. This means that the five volts is split into 1024 increments. Therefore, with a reference voltage of five volts, the software can only detect differences of about four millivolts. The problem with this is that the SU-100 outputs a voltage between zero and 27 millivolts. A difference of four millivolts will represent a large change in ultraviolet radiation. In order to making the readings more accurate, the reference voltage of the Arduino Mega is changed to 1.1 volts. This way the software can detect changes of about one millivolt. The datasheet provides an

Figure 116: SU100 Testing

Figure 117: SU100 UV Sensor Wiring


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algorithm for calculating the ultraviolet radiation from the voltage reading. This algorithm is used in the software. The following code is the code used in the software for the SU-100. The function returns a value accurate to two decimal places.

Figure 118: SU100 Software


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SP-110

The first pyranometer selection is the Apogee Instruments SP-110. The SP-110 measures the spectral light range from 380 to 1120 nanometers. The sensor outputs an analog value from zero to 350 millivolts. An increase of one millivolt corresponds to a radiation increase of 0.456 watts per square foot. The analog output connects to an analog input pin of the Arduino Mega 2560. The functionality of the SP-110 is extremely similar to the SU-100. The difference is that the SP-110 reads solar irradiance and outputs a higher voltage. In all other regards, the functionality of the two sensors is the same.

LS20031

The LS20031 is a GPS module manufactured by Locosys. The LS20031 contains a GPS antenna, a MC-1513 GPS module, and a transistor-transistor logic (TTL) data interface. The GPS has a refresh rate range of one Hertz to 10 Hertz. The optimal refresh rate, determined through testing, is five Hertz. The GPS is accurate to within nine feet. The GPS receives a NMEA string. There are six possibilities for NMEA output message selection. The software takes readings of latitude, longitude, and altitude from the GGA NMEA statement. The GPS transmit pin is connected to the USART receive pin of the Arduino Mega and the GPS receive pin is connected to USART transmit pin of the Arduino Mega. The LS20031 requires a 3.3 volt power supply which comes from a nine-volt battery regulated by a buck converter. The proper wiring of the LS20031 is shown in Figure 119 below.

The NMEA statement from the GPS is sent out of the transmit line character by character. The software looks at the statement and decides if it is the correct NMEA statement. Then, the software parses the latitude, longitude, and altitude from the statement. The GPS code can be found in Appendix L.

Figure 119: GPS Wiring


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Micro SD The payload saves all sensor data to a 16 gigabyte high capacity micro secure digital (SD) card. The micro SD card will organize data into a 32 bit file allocation table (FAT32) file system. A micro SD card allows for quick retrieval of data. The micro SD card utilizes the SPI bus of the Arduino Mega for data transfer. Digital pins 50 through 53 are the SPI pins of the Arduino Mega. From the advice of technicians at Parallax, there is a one Kilo-ohm resistor on each of the SPI data lines. The correct wiring of the micro SD card is shown in Figure 120.

Arduino has a pre-made library for micro SD cards. This library is what the software uses to store data to the micro SD card.

XBEE The wireless transmitter that relays atmospheric sensor readings and GPS data is the XBee-PRO XSC S3B. The XBee-PRO XSC S3B is capable of transmitting telemetry to a line of sight range of 28 miles when coupled with a high gain antenna. The transmitter operates at the 900 Megahertz frequency band. The 900 Megahertz band is organized into 12 different channels to help prevent interference from other devices transmitting at this frequency. The specifications of the XBee are detailed in Table 55.

Figure 120: MicroSD Wiring


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Specification Quantity

Processor ADF7023 transceiver, Cortex-M3 EFM32G230 @ 28 MHz

Frequency Band 902 MHz to 928 MHz

RF Data Rate 10 Kbps or 20 Kbps

Outdoor/Line-Of-Sight Range Up to 28 mi (45 km) w/ high-gain antenna

Transmit Power Up to 24 dBm (250 mW)

Receiver Sensitivity -109 dBm at 9600 baud -107 dBm at 19200 baud

Spread Spectrum FHSS

Operating Temperature -40° C to +85° C

Supply Voltage 2.4 to 3.6 VDC

Transmit Current 215 mA

Receive Current 26 mA

The XBee-PRO XSC S3B requires a low voltage of only 2.4 volts to 3.6 volts. Therefore, the XBee is connected to the 3.3 volt buck converter regulating a nine-volt battery. The data from the sensors is sent to the Arduino Mega and then transmitted to the XBee through one of the USART serial data busses. The transmit line of the Arduino Mega is connected to the receive line of the XBee. The wiring of the XBee is shown in Figure 122.

The XBee is located away from all deployment altimeters. The onboard XBee uses a 3.5 inch antenna which faces upward. The data is transmitted from the payload XBee to a corresponding XBee connected to a high gain Yagi antenna at the ground station. The ground station XBee uses a Sparkfun adapter with a USB connection to communicate with the ground station computer. When testing the adapter, a problem was realized with the communication link. Through the advice of Digi Technical Support, the adapter had to be modified. The following advice was found on Sparkfun’s forums:

Table 53: XBee XSC S3B Specifications

Figure 121: XBee Wireless Transmitter Wiring

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The LED was de-soldered from the XBee adapter as shown in Figure 124.

Range testing has been performed in order to verify the functionality of the wireless transmission. To specify how range testing was performed, a GPS was used to determine the starting location and final location where the XBee’s lost connection. Range testing shows that the XBee telemetry system can transmit over two miles with line of sight. Figure 125 shows the starting and stopping GPS coordinates along with the distance between them.

Figure 122: Digi Technical Support Forum Post

Figure 123: De-Soldering LED from XBee Adapter


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The ground station display will be through a MATLAB GUI. All sensor values will be graphed in real time and saved to a text file on the laptop. The current version of the ground station software is available in Figure 126. The format should not change, but the current version does not display the entirety of the sensor data.

Figure 124: XBee Range Test


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Liquid Crystal Display Sparkfun distributes the selected LCD, part number 11062, which mounts on a breakout board. The LCD has a color display with a resolution of 132 pixels by 132 pixels. The Arduino Mega 2560 microcontroller controls the LCD. The LCD screen utilizes the SPI data interface bus. One desirable feature of the chosen LCD screen is the low voltage requirements. At a minimum, the LCD screen requires 3.3 volts.

Arduino Mega 2560 The atmospheric sensors send readings to an Arduino Mega 2560-R3 microcontroller. The microcontroller includes an ATmega2560 microprocessor which has 16 analog input ports, 14 digital pulse width modulation enabled output ports, and 54 general purpose digital input and output ports. The ATmega2560 microprocessor is capable of supporting at least four serial data interface devices, communicating with Inter-Integrated Circuit (I2C) data interface devices, and containing a Serial Peripheral Interface (SPI) bus for communicating with SPI data interface devices. The microcontroller has built-in 3.3 volt and five-volt voltage regulators. The preceding characteristics are some of the determining factors to the microcontroller selection. Through testing, the Arduino Mega 2560 microcontroller has been determined to be the most suitable choice.

Figure 125: Ground Station GUI


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Electrical Schematic The wiring schematic of the overall ADGS is shown in Figure #127 and the code for the software is found in Appendix M.

Figure 126: ADGS Wiring Schematic


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Batteries The ADGS is powered by a nine-volt battery. The ARTCOS is powered by two nine-volt batteries, one for the servos and one for the image processing. The nine-volt battery is the Ultralife U9VLBP which has a capacity of 1.2 amp-hours. This capacity exceeds the capacity requirements of the payload. As mentioned in the “Test Results” section, the batteries allowed the payload to operate for over two hours. A runtime test will be performed with the payload running at full power in order to time the approximate maximum runtime. A magnetic switch activates the three payload system power supplies by the use of relays.

Buck Converters Linear voltage regulators are inefficient (up to 45% power loss) and buck converters are extremely efficient (up to 97% efficient); therefore, the payload design utilizes buck converters to regulate the nine-volt batteries. The buck converters are distributed by Adafruit. They are manufactured by Traco Power: product number TSR 1-2450. Testing has not begun on the buck converters.

ARTCOS The photographic camera is the VC0706. Adafruit distributes a breakout board with a preinstalled module. The camera sends each photograph to an Arduino Pro Mini. The VC0706 communicates with the Arduino using the serial data interface bus. The photographs save to the micro SD card in Joint Photographic Experts Group (JPEG) format. The camera mounts from the servo motors onto the breakout board using the prefabricated mounting holes. The software dictates when the camera is to take pictures. Figure 128 shows the correct wiring of the VC0706.

The previously mentioned wiring configuration allows the use of the VC0706 Comm Tool software. This software comes free with the device. Using this tool, the functionality of the VC0706 camera module is verified. Figure 129 is a screenshot of the verification software.

Figure 127: VC0706 Camera Wiring


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The software for the camera does not use the hardware serial UART pins; instead it uses two digital pins and converts them to serial lines. A problem with the VC0706 is the amount of time the module takes to save each picture. Currently, the image processing configuration is only able to save two pictures each minute. The descent phase in which the main chute is deployed will be the most optimal time to record images. This phase will only last approximately 50 seconds. The delay time to save the pictures is being researched, and several solutions have been considered. More testing will need to be done in order to decide if the chosen camera will function appropriately. Servos control the orientation of the camera. One servo mounts to the aluminum rail on the side of payload. The second servo mounts to the first servo’s rotational arm. The camera mounts to the second servo’s rotational arm. This setup allows for camera orientation in the horizontal and vertical directions. The optimal servo for the payload has characteristics of being small, lightweight, and must have sufficient torque for orienting the ARTCOS properly. Both servo motors are HS-85BB+ Mighty Micros. The Mighty Micro is small, lightweight, and strong. A pulse width modulated output signal from the Arduino Pro Mini 328 controls the servo. The gears within the Mighty Micro are nylon. The Mighty Micro also contains a ball bearing to make the motor move smoother. The servos are connected to each other through the use of a mounting bracket. The first prototype connected the servos together with epoxy; however, the flight test showed

Figure 128: VC0706 Configuration GUI


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this type of epoxy is not sufficient and a different type of epoxy will be used in the second prototype. The mount attaching the servo to the aluminum rail is made of fiberglass, and the mount that will be manufactured for the servo connection and camera mounting will be made of fiberglass. The ARTCOS mounting can be viewed in Figure 130 and 131.

An accelerometer measures acceleration in multiple planes. Because gravity is essentially constant, it is possible to accurately measure the tilt by relative acceleration. The chosen accelerometer is an ADXL345. The ADXL345 measures acceleration in the X, Y, and Z planes. The ADXL345 utilizes the I2C data interface bus. The X and Y measurements of the accelerometer represent the tilt and yaw of the payload. Each reading stores to the micro SD card for post-flight analysis. The readings from the tilt sensor dictate the angle of the servo motors. An algorithm has been developed for the software. The following code is implemented in the software for servo control. The algorithm used is linear and was devised by observing the change in tilt in relation to the change in servo angle.

Figure 129: ARTCOS Mounting


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Figure 130: ARTCOS Mounting

Figure 131: ARTCOS Orientation Algorithm


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The ADXL345 transmits readings to the Arduino Pro Mini through the use of the I2C data bus. The servos operate by the use of digital output ports on the Arduino Pro Mini. The SU-100 and SP-110 require a reference voltage of 1.1 volts and the HIH4030 requires a reference voltage of five volts. For this reason, the HIH4030 is connected to an ARTCOS Arduino Pro Mini which transmits the readings to the Arduino Mega via the USART bus. The ARTCOS software is found in Appendix N. The overall wiring schematic of the ARTCOS is shown in Figure 133.


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Figure 132: ARTCOS Wiring Schematic


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Video Camera The HackHD - 1080p Camera Module was the camera used for the initial prototype. For the next prototype, the Keyfob video camera will be used. Testing has not begun; because, the Keyfob video camera has not arrived from the distributor. The video camera has a dedicated battery and an on/off switch. The video camera stores video to a 16GB SD card and is capable of recording multiple hours of video. The video is used for educational outreach and redundancy for the ARTCOS.

Official Scoring Altimeter The official scoring altimeter beeps the apogee height upon landing. The height records in feet in accordance with the altimeter’s data. The chosen official scoring altimeter is the Adept A1E. The altimeter wiring is completely independent of the other payload electronics; therefore, the altimeter utilizes a dedicated power supply. The required power supply is a 12 volt battery which comes with the altimeter upon the receipt of the order. The battery that ships with the altimeter is the GP-23A Alkaline Lighter Battery. This power supply allows the altimeter to function for up to 10 hours. The A1E was launched in the flight test and recorded an altitude of 2,271 feet.

Block Diagram A block diagram of the entire payload electrical system is presented in Figure 134.


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Payload Components Table 56 lists the payload components and their qualities.

Figure 133: Payload Block Diagram


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Purpose

Breakout Board Distributor

Part Number

Interface

Dimensions (L x W x H)

Input Voltage

Current Draw

Flight Computer

Sparkfun Arduino 2560-R3

N/A 2.125 x 4.3125” 7 – 12V 20 - 200mA

Flight Computer

Sparkfun Arduino Pro Mini 328

N/A 0.7 x 1.3” 5 – 12V 150mA

Hygrometer Sparkfun HIH4030 Analog 0.75 x 0.30” 4 – 5.8V 200μA

Thermometer

DSS Circuits BMP180 I2C 0.625 x 0.5” 1.8 - 3.6V 3 – 32μA

Pressure Sensor

DSS Circuits BMP180 I2C 0.625 x 0.5” 1.8 - 3.6V 3 – 32μA

UV Sensor Apogee SU-100 Analog 0.925 x 0.925 x 1.08”

0V 0A

Pyranometer

Adafruit TSL2561 I2C 0.75 x 0.75” 2.7 – 3.6V 0.5mA

Pyranometer

Apogee SP-110 Analog 0.925 x 0.925 x 1.11”

0V 0A

Wireless Transmitter / Receiver

Digi XBee-PRO XSC S3B

Serial 1 x 1” 2.4 – 3.6V 215mA

Spacer Parallax 32403 N/A 1.16 x 1.0 x 0.58”

0V 0A

GPS Locosys LS20031 Serial 1.18 x 1.18” 3 – 4.2V 29mA

Camera Adafruit VC0706 Serial 1.26 x 1.26” 5V 75mA

Servo Servo City HS55BB+ Mighty Micro

Pulse Width Modulation

1.14" x 0.51" x 1.18"

5V 240mA

Tilt Sensor Sparkfun ADXL345 I2C 1 x 0.5” 3.3V 40μA

Micro SD card

Wal-Mart 3FMUSD16FB-R

N/A 2.2 x 0.3 x 3.4” N/A N/A

Micro-SD Adapter

Adafruit 254 SPI 1.25 x 1 x 0.15” 3 – 5V 150mA

LCD Sparkfun 11062 SPI 1.5 x 2.5” 3.3 – 6V 108 – 324mA

Voltage Leveler

Adafruit Buck Converter 1065

N/A 0.25 x 0.25” 6.5 – 32V -

Battery Ultralife U9VLBP N/A 1.81x1.04x0.69” 5.4 – 9.9V Max 120 mA

Status and Plans of Remaining Manufacturing and Assembly The second prototype will include all of the solutions to the problems discovered through testing the first prototype. The ARTCOS will be mounted with a fiberglass mount and a different type of epoxy will be used to connect the servos to each other.

Table 54: Payload Components and Qualities


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The clear acrylic will be of UV-T specification in order to avoid blocking UV radiation. The sensors will be mounted more securely. The payload framework will be welded instead of using rivets to increase durability and strength. The PCB has been ordered, and testing will begin on January 14, 2013. The atmospheric sensors will be included in the PCB. The PCB design has been designed based on the prototyping research. PCB boards of two mounting types have been ordered. One allows the use of the commercially distributed breakout boards with through-hole mounts. The second PCB has fully redesigned each breakout board with surface mount pads. A schematic of the simplified PCB which is compatible with the breakout boards is in Figure 135.

Figure 134: Breakout Board Compatible PCB


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The surface mount PCB is much more intricate than the prototyping PCB. Each breakout board has been reduced to its operating components. The original PCB design has been changed to reflect design changes for the payload circuitry. Also, placement of components and routing of trace wires have been optimized in the PCB revisions. The current PCB design is shown below in Figure 136.

The team is fortunate to be sponsored by Advanced Circuits, who has donated their services to manufacture all PCB designs for the SMD payload.

Figure 135: Surface Mount PCB


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Integration Plan

The payload is designed to integrate simply into the acrylic payload housing structure of the launch vehicle. The payload is constructed on a payload framework which consists of a forward payload bulkhead, a rear payload bulkhead, and a rectangular aluminum frame. The rear payload bulkhead will be epoxied into the rear coupler (booster to payload housing) which houses the drogue avionics bay. A rectangular frame constructed from four aluminum angle rails, will also attach to this rear payload bulkhead with three screws.

The forward payload bulkhead is epoxied into the forward coupler (payload housing to upper body airframe) which houses the main parachute avionics bay. This forward payload bulkhead has a recessed slot where the open end of the aluminum frame will sit once the payload is installed.

Figure 136: Bulkhead Aluminum Frame Interface


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Payload preparations involve the installation of two microSD cards. One card is used for the payload SMD sensor data, and a second card is used for storing camera photos. The Adafruit 254 microSD card reader features a card locking system, thus preventing the need to manually secure the microSD cards in place. Next, the power sources must be activated to conduct a payload functionality test. This will occur through the use of a magnetic single pole single throw (SPST) switch and three single pole single throw (SPST-NO) relays. The three relays will be used to activate three independent power supplies. After power is activated, wireless telemetry and L.E.D. visual inspection will verify the functionality of the S.M.D. payload. Additionally, the GPS will need time to find tracking satellites and save their position information onboard the SMD GPS.

Figure 137: Bulkhead Recessed Slot

Figure 138: Telemetry Verification GUI


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After payload functionality is verified and GPS satellite tracking is achieved, the power will be deactivated to preserve battery life. The onboard video camera in the payload must be activated manually at this time to start collecting video of the flight. Because the camera is a self-contained system, there is no option for activating externally or remotely. Power activation is achieved by pressing a “power on” button located on the video camera. At this point, the payload is ready to be integrated into the launch vehicle and ready to be activated on the launch pad when required.

Step Component Action Taken 1 SMD Payload Sensor Data Card

Reader MicroSD card installed

2 Camera Card Reader MicroSD card installed

3 Payload Power Up Magnetic Switch Activation

4 Payload Functionality Verification Telemetry and L.E.D. Visual Inspection

5 Payload Power Down Magnetic Switch De-Activation

6 Video Camera Activation On-Board Power Activation

7 Payload Framework installation Payload Installation

Precision of Instrumentation and Repeatability of Measurements Ultraviolet Radiation The SU-100 datasheet specifies the range of the light spectrum that the sensor is sensitive to. The sensor is most responsive to UVA (320 – 400 nanometers) and UVB (280 – 320 nanometers). The SU-100 datasheet provides the sensor’s spectral response in Figure 140. The datasheet lists the absolute accuracy at 10% and the repeatability at 1%.

Table 55: Payload Preparation Steps


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Solar Irradiance The SU-110 measures solar irradiance. The sensor is responsive to the range of the light spectrum from 300 nanometers to 1100 nanometers, and it is most sensitive at approximately 975 nanometers. The SP-110 datasheet shows the sensor’s spectral response in Figure 141. The datasheet lists the absolute accuracy to be five percent and the repeatability to be one percent.

Figure 139: SU-100 Spectral Response


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The accuracy of each of the payload sensors is listed in Table 58.

Purpose Product Precision Barometric Pressure BMP180 ±0.017psi

Temperature BMP180 ±1.8° F

Humidity HIH4030 ±3.6% RH

Solar Irradiance SP-110 ±5%

Solar Irradiance TSL2561 ±5%

Ultraviolent Radiation SU-100 ±10%

GPS LS20031 ±9.84ft

Accelerometer ADXL345 ±4.3mg

Official Altimeter Adept A1E ±1ft

Figure 140: SP-110 Spectral Response

Table 56: Payload Sensor Precision


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Safety

When testing components, safety concerns are a major issue. To ensure the safety of the components, all lab testing takes place inside a clean room as seen in Figure 143. The table in the clean room is covered with antistatic pads. Before a worker tests a component, antistatic lotion is applied to their hands. There are lights mounted to the table to ensure that the worker can see adequately. A circuit board vice is on the clean room table, and it is used to hold the circuit board when soldering; the vice allows safe and accurate soldering.

A resulting feature of the payload design is safety; since the payload does not eject from the payload housing, the likelihood of jettisoned components decreases.

Failures The first flight consisted of several failures. The first failure occurred pre-flight. The ARTCOS servo mounting epoxy failed. The servo connected to the camera came disconnected from the servo mounted to the rail. The disconnected servos are documented in Figure 144. The solution to this failure is to use the proper type of epoxy in the second prototype; the current epoxy consideration is Loctite Plastic Epoxy.

Figure 141: Clean Room


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Even with this pre-flight failure, the payload was still launched safely. The housing keeps all components from jettisoning from the payload even when they become disconnected. When the main parachute did not eject properly, the clear acrylic housing broke and many components were damaged as shown in Figure 145.

The XBee was placed in a 0.1 inch circuit board through-hole spacing adapter and came dislodged upon landing. This failure shows that the XBee needs to be securely mounted to the circuit board. The solution is to use a zip-tie to secure the XBee to the spacing adapter. One of the SP-110s dismounted from the payload upon landing. The failure occurred due to the nylon screw holding the SP-110 in place breaking. The solution to this failure

Figure 142: ARTCOS Epoxy Mounting Failure

Figure 143: Post-Flight Payload


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is to use metal screws for the second prototype. The Hack HD 1080p video camera used in the first prototype broke upon landing. The solution to this problem is to use Keyfob 808 video camera, which is cheaper and more durable. The 3.7 volt battery required for the Hack HD video camera also came dismounted upon landing; the Keyfob 808 battery is internal to the device and will not need to be mounted. The micro SD card came dislodged from the micro SD card writer upon landing. The solution to this failure is to secure the micro SD card with tape. The GPS also came disconnected upon landing. The pre-flight and post-flight circuit board are shown in Figure 146. The solution to this failure is to use a zip-tie to secure the GPS.

The temperature recorded by the BMP180 while the launch vehicle was on the launch pad increased from 55 degrees Fahrenheit to 100 degrees Fahrenheit. The solution to this failure is to mount the BMP180 between the two circuit boards in order to shield it from light. The humidity reading decreased from 45 percent to 10 percent. It is hypothesized that this is due to the increase in temperature recorded by the BMP180, since the HIH4030 calibration function uses the temperature reading from the BMP180. The SU-100 recorded no ultraviolet measurements during the test flight or while the vehicle was on the launch pad. The solution to this failure is to use the proper type of Acrylic which shields UV radiation; this type of Acrylic is known as UV-T acrylic. The faulty, unchanging readings of the TSL2561 can also be solved through this solution.

Figure 144: GPS Mounting Failure


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Each of the aforementioned solutions are being researched and tested.

Payload Concept Features and Definitions

Creativity and Originality A clear acrylic section of the vehicle body houses the payload. This allows visual inspection of all sensors and indicator LEDs along with verification readouts from the LCD during assembly and pre-launch operations. This also verifies that the ARTCOS is operating properly after the payload is integrated and secured within the launch vehicle. Given the payload integration design, the payload is easily and quickly removed. All electrical components are placed on modular circuit boards so that they can easily be removed and replaced from the payload framework. This allows any repair, replacement, or upgrading of the payload to occur without altering other aspects of the payload. Given the modular design of the power circuits, disposable batteries are replaceable when necessary. Although testing and prototyping begins with breakout boards, the PCB design will comprise the majority of the electronics in the payload. This is advantageous not only for efficiency and signal integrity, but also for reducing the size requirements of the entire SMD payload. These completely original PCBs will be realized as a final SMD product for the official launch day. A silkscreen image with a team (or university) emblem on the PCB board in Tarleton purple will be included as well for aesthetic appeal. Figure 147 is an image of the front and back sides of the PCB boards.

The self-leveling camera system ensures that photographs are taken in proper orientation as required. The tilt sensor detects any changes in orientation, where two servo motors correct movement. This provides a lightweight and efficient solution, eliminating the need to design a control system for payload orientation during descent. The LCD displays appropriate data and relevant readouts to ensure and verify payload functionality. The payload includes a video camera for flight documentation and is used

Figure 145: PCB Board


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as an educational outreach tool.

Uniqueness and Significance The scientific payload is significant in that it meets the SMD requirements set forth in the NASA SLP SOW. All design features and component selections reflect compliance with the customer prescribed specifications. The ability of the payload to gather atmospheric data is significant for the analysis of changing conditions as the vehicle varies in altitude. All data gathering is useful for finding correlations to altitude such that atmospheric conditions from ground level to vehicle apogee can be modeled and tested. The clear acrylic airframe houses the payload. This is unique; because, it allows an external visual inspection of all components of the payload. Once preparation and integration of the payload takes place in the vehicle, a final verification ensures flight readiness relating to the payload. An LCD screen displays relevant checks and indicators from the flight software. The means to maintain proper orientation of the camera is unique as well; because, the orientation of the payload is not guaranteed upon descent, a unique corrective measure takes place. The chosen solution creates and implements a self-leveling camera. Any changes to the nominal orientation of the payload can be measured via the accelerometer, and the servos can then correct these offsets. This allows the camera to maintain proper orientation throughout the flight. A video camera in the payload captures video of the vehicle flights. Although this is a standard practice in rocketry, video recordings provide flight documentation that can be useful in educational engagement and community outreach.

Suitable Level of Challenge

The design complexity and implementation of a functional SMD payload is an extremely taxing endeavor. Not only must the appropriate components reside in the vehicle, but they must also function in a manner that achieves a useful atmospheric measuring

Figure 146: Self-Leveling Camera System


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instrument. This involves proper interfacing, sufficient power supply, and adequate programming logic of all constituent pieces of the payload. Although very challenging, design of the electrical circuits seek PCB implementation. This is a significant advantage over breakout board and perforated board mounting. The effective use of space and overall efficiency of the PCB design merits it as one of the best all-around payload components in the design. Despite the level of challenge and time necessary to complete this design, it is worthwhile in view of the integrity of the payload system as a whole.

Science Value Payload Objectives The main objectives of the payload are to record and store atmospheric and GPS data, transmit these measurements to a ground station, and capture photos during the rocket's flight. The NASA USLI SLP states additional objectives which are addressed in the following section. Payload Success Criteria The payload is deemed successful if it obtains and transmits valid atmospheric data and flight imagery. Validity of the measurements is based not on their accuracy, but rather on whether they lie within the established confidence intervals for their particular sensors. Should any data lie outside this acceptable range, it can be considered either a mechanical error of the sensor or a computational error within the software and consequently, invalid. Validity of the imagery is based on whether the orientation of the image meets the criterion of proper orientation, with the ground of the scenery at the bottom of the frame and the sky at the top of the frame. Successful operation of the ARTCOS system will determine whether captured images meet this standard. A minimum number of images ought to be captured. If the payload does not record two photos during the rocket's descent and three photos after the rocket's landing the imagery system is deemed unsuccessful. Validity of the data transmission is based on whether the information received at the ground station corresponds with the recorded measurements. Should any values differ, it can be assumed that bits were lost during the transmission and the values received are consequently invalid. These transmissions should occur at a specified frequency of 200 megahertz. If the ground station does not receive information every five seconds, the communication system is deemed unsuccessful. Table 59 summarizes the payload objectives, their science value and success criteria.


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Payload Objective Science Value Success Criteria

Gather Atmospheric and GPS Data

Representing Change in Atmospheric Variables Depending on Altitude

Collected Valid Data

Store and Transmit Atmospheric and GPS Data

Modeling Rocket Flight

Collected at Required Frequency, Transmitted Throughout Flight

Camera Orientation Test Multi-Servo Orientation Device

Two Pictures During Descent and Three after Landing, Correct Orientation

Experimental Logic, Approach, and Method of Investigation The SMD payload gathers data from approximately 5,280 feet above ground level (AGL) down to the landing site. Data gathered includes varied data for five atmospheric variables: pressure, temperature, relative humidity, solar irradiance, and ultraviolet radiation. This data determines the accuracy of the payload sensors and the statistical correlations between each of the various variables. These two calculations aid in the development of a regression model for each variable. By creating a model to represent these correlation effects, a new and comprehensive formula could demonstrate causal relationships between these five variables or any derivative subset.

Regression Model With a large number of samples ranging across the various test and demonstration flights, the data plots can be analyzed statistically to determine a model that accurately represents relationships between the different variables. Using R-based software a regression model of the form Y = Xβ + ε is computed. In these models, a chosen response variable (Y) is modeled against a column matrix of the other variables (X) multiplied by their derived coefficients (β) and a random error term (ε). Assuming that the response variables and error terms meet requirements of constancy and normality in their variance, these models can be tested for their validity. A t-test will suggest if any column in the variable matrix X should be removed. A stepwise model can suggest the optimal subset of variables to be included in the model. Each model that produced the R2 statistic will show how well the predictive values of the created model matches with the experimental data. Furthermore, the residual sum of squares of the error terms demonstrates how far the model deviates from the actual data. These tests, among others, can be used to create, refine, and validate any models.

Table 59: Payload Objectives Summary


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Correlations between SMD Sensor Readings The distributions of each variable aid in finding covariance between any two variables; therefore, the correlation between those variables can be attained. After determining the correlation, the model can evolve as necessary to provide more accuracy. Using R-based software, any data set can be analyzed by individual variables. Variance matrices and standard deviations from each variable can be combined in well-defined formulae to find covariance and correlation.

Accuracy of Sensors Though the datasheets for each sensor propose a certain level of accuracy, this cannot guarantee the sensors will perform to this level within the payload circuitry. By comparing collected data against atmospheric measurements from national databases, any discrepancy can establish itself. Furthermore, using both data sets can establish confidence intervals to ensure new data are within a particular range of the presumably absolute readings from these databases. This gives a measure of accuracy as it pertains specifically to the payload circuitry. Test and Measurement, Variables, and Controls Measurements will be taken by the payload sensor circuitry. Although some measurements might be taken at the ground station, these will be used for predicting the rocket's trajectory. The only data considered in the formal analysis will come from the onboard payload circuitry. The variables included in the measurement process are altitude, temperature, atmospheric pressure, relative humidity, solar irradiance, UV radiation and random electrical noise in the payload circuitry. The controls in consideration are the measurements from national agencies such as the National Oceanic and Atmospheric Administration (NOAA), whose measurement devices have presumably negligible errors. The testing process occurs post-flight in the form of a statistical analysis of the acquired data.

Relevance of Expected Data and Accuracy/Error Analysis The findings compare against documented formulae for atmospheric measurements to determine their validity. Furthermore, the models could potentially represent undocumented relations between these variables. Using these models, a select few measurements made on the ground level could accurately predict weather conditions at


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the launch site. This would enable amateur rocket enthusiasts to more precisely predict the exact flight path of their rocket. The accuracy and error analysis are an inherent part of the entire process. All errors factor into the final regression model. Small differences from the expectant values will demonstrate as random noise in the model. Preliminary Experiment Process Procedures A detailed preliminary experiment has been conducted to find any initial models or relationships against which to test. Using the same processes mentioned above, over five gigabytes of atmospheric data from NOAA was analyzed. This data was limited to a subset corresponding to Stephenville, Texas to eliminate any confounding variables relating to geographical differences. The results are as follows: CC = 478.5 – 1.028 ILW – 0.3985 PW – 0.003413 P – 614.1 SH ILW = 330.3 – 0.5912 CC – 0.2285 PW – 0.002283 P PW = -402.4 – 0.243 CC – 0.2496 ILW – 0.01109 ISW + 0.004428 P – 0.0595 RH + 1349 SH RH = 0.5537 PW + 12910 SH where “CC” denotes cloud coverage, “ILW” denotes long-wave solar irradiation, “ISW” denotes short-wave solar irradiation, “P” denotes atmospheric pressure, “PW” denotes precipitate water, “RH” denotes relative humidity, and “SH” denotes specific humidity. After establishing these models, a subset of the NOAA data corresponding to Huntsville, Alabama was used to validate the model. These calculations, among others, suggested these models to be valid. Concerning accuracy, atmospheric variables outside the payload's domain of measurement were considered. These extraneous variables were precipitate water, cloud coverage, ground roughness, specific humidity, geopotential height, and wind velocity measurements. As expected, many of these variables had nominal effects on the variables considered by the payload. Although the variables of cloud coverage and precipitate water demonstrated significant impact on the variables measured by the payload, their values can be either substituted or accurately calculated using versions of the formulae listed above.

Safety and Environment

The Safety Officer

The team safety officer, Blake, is level one certified with NAR, and has obtained an FAA flight waiver in his name for full scale launches. The responsibility of the safety officer is to design and implement safety plans that ensure all accidents are evaded. All hazards


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to people, the project, and the mission are determined so that mitigations can be enacted. The systematic identification of risks, failure modes, and personnel hazards allows the team to designate where single points of failure could occur throughout the course of the project. The identification of single point failures allows for proactive design changes to be made to counter these failures leading up to the CDR.

Failure Modes

A failure mode is the way a system could fail, causing an undesirable effect on some aspect of the project. The safety plan ensures development and implementation of mitigations for each failure mode.

Payload Integration Failure Modes

Through testing, many design flaws were discovered. Changes were made after each test upon failure analysis. The changes implemented in the final design were those found to be the most robust through testing. The failure analysis can be found in the Recovery Subsystem under “Safety” and “Failure Analysis”. An updated summary of potential failure modes that could occur during payload integration is detailed in Table 60.

Failure Effect Proposed Mitigation Completed Mitigation

Screw hole stripped out

Inadequately secured payload

Ensure the bulkhead is replaced when required


Incompatible hardware

Payload will not integrate properly

Ensure precision of fit during manufacturing


Components damaged during integration

Electronic malfunction

Be careful while inserting payload


Launch Operations Failure Modes

Failure analysis of test launches revealed both procedural missteps and simplifications. Checklists were amended to mitigate the procedures. The changes implemented in the final design were those found to be the most robust through testing. The failure analysis can be found in the Recovery Subsystem under “Safety” and “Failure Analysis.”

Table 57: Potential Failure Modes during Payload Integration


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An updated summary of potential failure modes that could occur to the payload during launch operations is in Table 61.


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Subsystem

Potential Failure Mode


Proposed Mitigation


Payload

Battery disconnects

Partial or complete failure of the SMD payload

Secure battery terminals


Port holes improperly sized

Inaccurate sensor readings

Test components within the payload section with portholes installed


Radio dislodges from breakout board

Telemetry lost and possible damage to SMD Payload

Secure the radio to the radio mount


BMP180 exposed to excessive sunlight

Inaccurate readings Mount sensor in a covered location


ARTCOS mounts break

Camera will not be capable of orienting properly and possible damage to SMD Payload

Ensure all components have adequate mounting and thoroughly inspected


Wired connections disconnect

Component and/or System Failure

Use of braided wire in non-static components and thorough inspection



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Subsystem

Potential Failure Mode


Proposed Mitigation


Component breaks away from board

Partial or complete failure of the SMD payload and possible damage to SMD Payload

Ensure all components have adequate mounting and are thoroughly inspected


Battery power fails

Partial or complete failure of the SMD payload

Check battery voltages prior to every flight and replacing batteries regularly


Circuits become un-mounted

Partial or complete failure of the SMD payload and possible damage to SMD Payload

Ensure all circuits have adequate mounting and are thoroughly inspected


Hazard Analysis

An updated overview of potential hazards to personnel through the course of the project can be viewed in Table 62. Personnel hazards refer to potential harm incurred by any individual. The development and implementation of the safety plan and protocols ensure that these hazards are appropriately mitigated.

Table 58: Potential Failure Modes during Launch


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Consequence

Mitigation Action

Laceration

Knives, routers, saws, file, Dremel tool

Medium Serious injury or death



Burns

Chemicals (FFFFg, fiberglass resin), welders, soldering Iron

Medium Minor to serious injury



Respiratory Damage

Chemicals (epoxy, solder), fumes, fiberglass

Low Brain damage or death

Follow safety protocols, proper tool and equipment use, personal safety attire, consult MSDS


Vision Damage

Welders, fiberglass, grinders, projectile debris

Low Partial to complete blindness

Use of goggles, force shields, consult MSDS, first aid kit available, refer to operators manual

Discontinue all operations, apply first aid, contact EMS, use eyewash


211



Consequence

Mitigation Action

Allergic Reaction Epoxy,

chemicals, fiberglass

Low

Loss of respiration, inflammation (Internal & External)

Use of gloves, consult MSDS, first aid kit available

Discontinue all operations, apply first aid, contact EMS, administer antihistamines, safety shower

Hearing Damage

FFFFg, Grinders, Ignition, Routers

Low Partial to complete deafness

Ear muffs, consult MSDS, first aid kit available, refer to operators manual


Dismemberment

Projectiles, Saws, Launches

Low Permanent injury or death

Make sure proper safety measures are taken, operators manual

Discontinue all operations, apply first aid, and contact EMS, tourniquet

The Material Safety Data Sheet (MSDS) that the manufacturer provides contains pertinent information about the material in consideration. It is comprised of 16 categories: identification, hazard(s) identification, composition/information on ingredients, first-aid measures, fire-fighting measures, accidental release measures, handling and storage, exposure controls/protection, physical and chemical properties, stability and reactivity, toxicological information, ecological information, disposal information, transport information, and regulatory information. MSDSs are referred to when a hazard occurs in order to enact the most effective mitigation. All team members shall be knowledgeable of the MSDS associated with each hazardous material. According to the safety plan, a binder containing all the MSDSs is always made available for personnel and brought to every launch.

Operator manuals for each tool will be consistently referenced prior to each tool’s usage. This ensures each tool is used as intended. According to the safety plan, operator manuals for each component used during the project our kept in an operator

Table 59: Potential Hazards to Personnel


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manual binder. These documents will be made available by the safety officer at any location in which construction, testing, or launching of the vehicle could occur.

It is important for all team members to be thoroughly briefed on the project risks, FAA laws and regulations regarding the use of airspace, and the NAR high-power safety code.

The team is aware that the FAA must be notified of planned launch activities. For educational outreach events, notification to the closest airport within five miles of the launch site is required 72 hours prior to launch. For subscale launches, flight waivers are must be obtained at least 45 days prior to the proposed activity. For full-scale launches, flight waivers have been obtained.

Environment

There are no environmental concerns for the payload that are not also inherent to the vehicle. Please refer to the “Safety and Environment” section of the Vehicle Criteria.


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V) Project Plan

V) Project Plan

Budget Summary The following chart includes the projected budget for completing the project. The task of completing the NASA USLI is a complex interdisciplinary endeavor that tests the team’s knowledge and skills, including management of a budget. The first step in managing a budget is devising a budget that is sufficient in meeting all costs necessary to complete the mission. Table 63 breaks down the known project costs. The team has allocated funds to several areas of the project as charted in Figure 149, and a detailed budget is available for review in Appendix A.

Element Est. Cost Testing/Prototyping $13,972.23

Outreach $3,669.77

Final Build $4,202.33

Travel to Competition $8,200

Total $29,922.13

The current budget deviates slightly from the preliminary budget. These deviations will be explained throughout the budget section of this document. Actual budget

12% 14%

26%

9% 12%

27%

Allocated Funds

Vehicle Recovery Payoad Propulsion Outreach Travel

Table 60: Preliminary Budget Summary

Figure 147: Allocated Funds


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V) Project Plan

expenditures are shown in Figure 150. It is separated into the following seven categories: Vehicle, Recovery, Payload, Propulsion, Outreach, Travel, and Unspent. Each category has been labeled with a percentage of total budget spent.

The only budget that has been slightly exceeded is the Vehicle budget. The vehicle budget overage seen below in figure 151 is due to unforeseen hardware needs and necessary equipment need to assemble the vehicle (epoxy applicators, cutting blades, and changes to the original design). The overage totals to approximately $465.00. Furthermore, the vehicle purchase is complete; the parts for all future vehicles including the final competition vehicle have been purchased.

13% 8%

14%

6%

7% 0%

52%

Budget Status

Vehicle Recovery Payoad Propulsion

Outreach Travel Unspent

Figure 148: Budget Status


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V) Project Plan

The preliminary payload budget has been sufficient in providing the necessary components for the SMD payload. The spent section in Figure 152 below encompasses all of the testing of the SMD, it does not include the final build or PCB cost which have been estimated to be roughly the same as the original prototyping costs.

Propulsion remains under budget. The stock of motor reloads is sufficient for the remaining tests as well as our competition launch. The spent and unspent funds that have been allocated toward propulsion follows in figure 153.

spent 100%

Vehicle Budget Status

spent 55%

unspent 45%

Payload Budget Status

Figure 149: Vehicle Budget Status

Figure 150: Payload Budget Status


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V) Project Plan

Due to some unexpected contributions and carpooling, the cost of the educational outreach program has been kept under budget thus far. There are several outreach events scheduled after the CDR report and the current prediction is that our current budget will prove adequate. The chart below gives a visualization of our current outreach budget status in Figure 154.

The budget provides $8,200 dollars for travel to the competition. At this time, the travel funds have not been touched.

spent 67%

unspent 33%

Propulsion Budget Status

spent 56%

unspent 44%

Outreach Budget Status

Figure 151: Propulsion Budget Status

Figure 152: Outreach Budget Status


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In the following charts, itemized budgets are compiled to illustrate an “on the pad” cost for the current final build design. The data from Tables 64-67 are broken down between three categories. The first is Structure and Propulsion, the second is Recovery, and the last is SMD Payload, which has been subdivided between PCB and Non PCB budgets.


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Body Part Number Price Per

Unit Quantity Total

Nose Cone FNC5.5EL $47.50 1 $47.50

Fiberglass Body Tube

G12-5.5-60 $38.84 5.3 $233.04

Acrylic Body Tube

ACRCAT5.500ODX.250 $39.05 1 $39.05

Sheet for Fins 500SHT0.125X48X96 $208.00 0.2 $41.60

Motor Tube G12-3.0-48 $71.06 1 $71.06

Bulk Plate (Payload)

PVCGRAY2.00LAM12x24 $147.82 0.25 $36.96

Bulk Plate Standard

FBP5.5 $7.60 3 $38.00

Couplers G12CT-5.5 $47.03 1 $47.03

Centering Rings

FCR5.5-3.0 $8.55 4 $34.20

Epoxy 4500Q $69.00 0.25 $17.25

Motor Retainer

RA75 $52.00 1 $52.00

Casing and Hardware

Cesaroni 3 Grain Case and closures

260.96 1 260.96

Motor Reload Cesaroni L-1720 170.96 1 170.96

Miscellaneous Hardware $100.00 1 $100.00

Total $1189.61

Projected

Total $1199.64

Table 61: Structure/Propulsion System Budget


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V) Project Plan

Proposed Selection

Item Number Unit

Cost

Quantity Typical

Unit

Cost

Main Altimeters Raven3 $155.00 2 $310.00

Backup Altimeters StratoLogger $79.95 2 $159.90

Electric Matches XL Variable-Capacity Ejection Canister

$2.75 4 $11.00

FFFg Black Powder Goex 3F Black Powder $3 1 $3

Main Shock Cord Tubular Kevlar $37.99 1 $37.99

Drogue Shock Cord Tubular Kevlar $31.49 1 $31.49

Main Parachute 10ft Cert 3 $239.00 1 $239.00

Flameproof Main Parachute Deployment Bag

DB8 $40.00 1 $40.00

Drogue Parachute 2ft $27.50 1 $27.50

Flame-Proof Drogue Parachute Nomex

18x18 $10.95 1 $10.95

Eye bolts 0.25in. (compact) $2.00 2 $4.00

Quick Links 0.25in. Stainless Steel Delta Quick Link

$2.99 6 $17.94

Shear Pins 2-56 Nylon Shear-Pin (10 pack)

$1.00 1 $1.00

Arming Switches Featherweight Magnetic Switch

$25.00 2 $50.00

GPS Garmin DC-40 $144.50 1 $144.50

Battery Ultralife U9VLBP $6.65 2 $13.30

Total $1101.57

Projected $1383.35

Table 62: Recovery System Budget


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V) Project Plan

Price Price Quantity Total

Arduino 2560-R3 $58.95 1 $58.95

Arduino Pro Mini 328 - 5V $18.95 2 $37.90

Micro SDHC Card $9.99 3 $29.97

Adafruit 254 – Micro SD Adapter $15.00 2 $30.00

BMP180 – Pressure/Temperature Sensor $15.00 1 $15.00

HIH4030 – Humidity Sensor $16.95 1 $16.95

SP-110 – Pyranometer $169.00 2 $338.00

TSL2561 – Lux Sensor $12.50 1 $12.50

SU-100 – UV Sensor $159.00 2 $318.00

HS-85BB+ Mighty Micro Servo $19.99 2 $39.98

Adafruit 397 – Camera $42.00 1 $42.00

HackHD - 1080p Camera Module $159.95 1 $159.95

Xbee-PRO XSC S3B $42.00 1 $42.00

XBee 0.1” Through Hole Spacing Adapter $2.99 1 $2.99

3.5” RPSMA 900MHz Antenna $14.64 1 $14.64

LS20031 GPS $60.00 1 $60.00

9V Battery Holder $2.95 3 $5.85

LCD – Sparkfun 11062 $34.95 1 $34.95

ADXL345 – Accelerometer $27.95 1 $27.95

Tenergy Li-Ion 3.7V Battery $9.90 1 $9.90

Ultralife U9VLBP - 9V Battery $6.65 3 $53.20

Adafruit Buck Converter $14.95 2 $28.90

Adept A1E Altimeter $29.95 1 $29.95


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V) Project Plan

Price Price Quantity Total

Total $1409.53

Projected

Total $1558.24

Table 63: Payload Budget (Through-Hole PCB)


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V) Project Plan

Product Price Quantity Total

ATMega2560 $17.97 1 $17.97

ATMega16U2 $3.71 1 $3.71

Arduino Pro Mini 328 - 5V $18.95 2 $37.90

MicroSD Card Connector $3.08 1 $3.08

Adafruit 254 – Micro SD Adapter $15.00 1 $15.00

BMP180 (Chip Only) – Pressure/Temperature Sensor $2.88 1 $2.88

HIH4030 (Chip Only) – Humidity Sensor $13.68 1 $13.68

SP-110 – Pyranometer $169.00 2 $338.00

TSL2561 (Chip Only) – Lux Sensor $2.84 1 $2.84

SU-100 – UV Sensor $159.00 2 $318.00

HS-85BB+ Mighty Micro Servo $19.99 2 $39.98

Adafruit 397 – Camera $42.00 1 $42.00

HackHD - 1080p Camera Module $159.95 1 $159.95

Xbee-PRO XSC S3B $42.00 1 $42.00

XBee 0.1” Through Hole Spacing Adapter $2.99 1 $2.99

3.5” RPSMA 900MHz Antenna $14.64 1 $14.64

LS20031 GPS $60.00 1 $60.00

9V Battery Holder $2.95 3 $5.85

LCD – Sparkfun 11062 $34.95 1 $34.95

ADXL345 – Accelerometer $27.95 1 $27.95

Tenergy Li-Ion 3.7V Battery $9.90 1 $9.50

Ultralife U9VLBP - 9V Battery $6.65 3 $53.20


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V) Project Plan


Adafruit Buck Converter $14.95 2 $28.90

Adept A1E Altimeter $29.95 1 $29.95

47 µF Capacitor $2.25 2 $4.50

100 nF Capacitor $0.19 12 $2.28

22 pF Capacitor $0.35 3 $1.05

1 µF Capacitor $0.29 3 $0.87

100 nF Polarized Capacitor $0.77 2 $1.54

10 nF Capacitor $0.25 2 $0.50

10 µF Capacitor $0.42 1 $0.42

2.2 µF Capacitor $0.96 1 $0.96

10 µF Polarized Capacitor $0.43 2 $0.86

Switching Diode $0.06 1 $0.06

Resettable Fuse $0.46 1 $0.46

EMI Filter Bead $0.40 1 $0.40

215 nH Inductor $1.90 1 $1.90

Green LED $0.14 1 $0.14

Red LED $0.08 1 $0.08

White LED $0.17 2 $0.34

Yellow LED $0.08 1 $0.08

Blue LED $0.11 1 $0.11

10 kΩ Resistor $0.25 3 $0.75

1 kΩ Resistor $0.81 9 $7.29

1 MΩ Resistor $0.20 2 $0.40

22 Ω Resistor $0.19 2 $0.38


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V) Project Plan


330 Ω Resistor $0.20 3 $0.60

68 kΩ Resistor $0.20 1 $0.20

4.7 kΩ Resistor $0.19 2 $0.38

500 kΩ Resistor $0.55 1 $0.55

50 kΩ Resistor $1.40 1 $1.40

25 Ω Resistor $0.58 4 $2.32

5 kΩ Resistor $0.08 4 $0.32

250 Ω Resistor $0.45 4 $1.80

Plastic Spacers $0.19 4 $0.76

P-Channel MOSFET $0.43 1 $0.43

16 MHz Crystal $0.96 1 $0.96

16 MHz Crystal $1.16 1 $1.16

ESD Suppressor Diode $0.25 2 $0.50

Operational Amplifier $0.83 1 $0.83

Operational Amplifier $0.67 1 $0.67

Line Driver $0.38 1 $0.38

Digital Potentiometer $2.06 1 $2.06

USB Connector $1.07 1 $1.07

Non-PCB $1346.68

PCB $85.92

Total 1432.6

Projected

Total 1558.24

Table 64: Payload Budget (Surface Mount PCB)


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V) Project Plan

Items in red were used on the original design. The purchase of new items is shown in the bottom row. The bulk of the cost in building the PCB design comes from the costly elements that cannot be reproduced (SP-110, SU-100, etc.).

Funding Plan A significant portion of the funding necessary for this project derives from a wide range of University organizations and other community support functions. Thus far, $11,500 in donations from the Tarleton President’s Circle (as seen in Figure 155), the Provost’s Office, the Dean of the College of Science, and the Tarleton Foundation fund the project. The Office of Student Research has provisions for the project amounting to $17,000. USLI Science Mission Directorate (SMD) funding also stems from NASA in the amount of $2,780. The total allocation for the project currently amounts to $31,280.

Timeline The Tarleton Aeronautical Team understands that a project of this magnitude requires a great deal of time and dedication. The following schedule to meet the requirements of the project serves as evidence in Figure 156. Gantt Charts detailing the project timeline follow in Figure 157. The following chart gives a visual representation of major project deliverables.

Figure 153: Early Funding


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V) Project Plan

Figure 154: Project Timeline


227

V) Project Plan

Testing Timeline

The Gantt chart below includes preliminary testing dates. Time is in the schedule to allow for lab prototyping and testing. The team plans to conduct multiple test launches including low altitude as well as high altitude test launches when possible. Using data gathered from these test launches, the team performs a failure analysis after each launch. These analyses are useful to optimize successive launches and overall design.

Figure 155: Testing Gantt Chart


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V) Project Plan

Outreach Timeline The chart below delineates the dates the team plans for educational outreach. The star parties each involve a simple vehicle demonstration and a presentation about the basics of rocketry. During class trips, team members travel to area middle schools to actively engage students in safe, basic rocketry. The Tarleton Science Olympiad consists of area middle school and high school students convening at Tarleton to compete in science related activities. During the Science Olympiad, the team demonstrates a static motor test in addition to giving presentations explaining basic rocketry. For more comprehensive information on the educational outreach component of the project, please refer to the following outreach section of this document.

Education plan

Outreach Plan

Vehicle design, creation, and implementation are important components of this competition. Conjunctively, the educational engagement portion of this project is crucial, as its main goal is to promote enthusiasm for the necessary subjects that relate to rocketry and other important STEM fields. The team’s plan is to host several events for diverse audiences. All events aim to promote the global necessity of math, science, engineering, and technology. Furthermore, the team includes a vehicle launch with each

Figure 156: Outreach Timeline


229

V) Project Plan

event to provide a real world experience to reinforce the addressed STEM concepts in the lesson portion of each event. The team aims to encourage interest in the relevant subjects with the intent of increasing the number of people that choose to pursue STEM related careers.

Educational Outreach

Educational outreach targets three main audiences through a variety of events; students, teachers, and the community as a whole. The team is currently establishing contacts and scheduling dates to visit the local middle schools. Several schools plan to participate in the educational outreach events already. Outreach to students in the schools occurs through a classroom lesson or an assembly style presentation. During the classroom sessions, small groups from the team present an original interdisciplinary lesson over rocketry with an emphasis on math and science. The goal of these lessons is to demonstrate to students the importance of the STEM subjects and their role in a variety of topics such as engineering and rocketry. The necessity of these careers with companies such as NASA is a primary focus. By working in a setting which allows for a smaller student to presenter ratio, students are receive an opportunity to work closely with the team members on a lesson which reinforces concepts learned previously in a novel manner. Beyond the classroom lesson, the assembly format allows the team to communicate the same information to students on a larger scale. This portion of the outreach began at the request of some of the local middle schools. The assemblies take place toward the end of the school day and involve multiple classes and grade levels. To conclude each school presentation, students join our team outside for a vehicle launch. The rocket launch adds to the lesson by giving the students a visualization of

Figure 157: Acton Middle School


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V) Project Plan

what they just learned. This increases students’ retention. In a classroom setting, interactive, hands-on lessons encourage learning.

Educator Outreach

By aligning the lessons created for the classroom presentations with state and national curriculum, these repeatable lessons remain relevant for reuse. The team is working to create a live webcast of a vehicle launch for teachers to access in their classes. This webcast allows teachers to use this online content as a real-life application in their classroom. Furthermore, the team is communicating with teachers throughout the state to distribute lesson plans. We hope to raise interest in STEM fields by having teachers join our group. Lesson plans are available from the Tarleton Aeronautical Team discussing rocketry and the importance of NASA.

Community Outreach

Star Party

Outreach beyond schools allows for students, teachers, and community members to join in learning about rocketry and STEM concepts. The team coordinates with Tarleton State University to co-host their Star Party event which occurs in both the fall and spring semesters. The event includes a discussion about the program, rocketry, the need for growth in the STEM fields, and a vehicle launch. The Star Party is an open invitation event; the team reaches audiences with a range from children to adults from local and surrounding areas.

Tarleton Regional Science Olympiad

The team will be at the eighth annual Tarleton Regional Science Olympiad on February 23, 2013. Students participating in this event along with their sponsors and family join the team for several vehicle launches including a static launch demonstration. A presentation and question and answer session follow. The day concludes with an awards ceremony. Again, the focus of this presentation is to promote the STEM fields and reiterate their importance pertaining to the nation’s progress.

Participation Goal

The team expects to involve approximately 2,500 students in total. Comprehensive feedback from teachers and students will be gathered through surveys. This feedback helps the team alter presentations to ensure the quality of each event. Outreach efforts


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V) Project Plan

for boy scouts, girl scouts, and after school programs are being developed. The team members have a passion for the STEM fields, thus outreach is an important goal.

Accomplished Educational Outreach

October 5, 2012

On October 5, 2012, members of the Tarleton Aeronautical Team traveled to Granbury. Students at Acton Middle School and Granbury Middle School participated in basic rocketry presentations. The team’s presentation explained STEM fields and related careers. As part of Career Day at Acton Middle School, the team specifically discussed careers available at NASA. Bert led presentations featuring 7 Minutes of Terror, a NASA video highlighting interviews with NASA engineers on the Curiosity Rover project. While gaining exposure to career options with NASA, the students also learned the importance of safety protocol. To emphasize the message of the video, the students were given the opportunity to experience rocketry in a safe environment. Before the team launched their rocket, each class was given the opportunity to launch two-liter water bottle rockets.

Figure 158: Team Members Educate and Entertain Acton Students


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V) Project Plan

23%

20%

21%

19%

8% 9%

Subject Interest

Science Math Engineering Technology Rocketry None

The surveys reflected that the majority of the students were delighted with their experience. In the survey, the students were asked to report whether they felt a greater interest in Science, Mathematics, Engineering, or Technology, three things they learned from the presentation, and what their favorite parts of the presentation were. The data is given in Table 68 and illustrated in Figure 161.

Table 65: Accomplished Educational Outreach

Figure 159: Subject Interest

Areas of learning include force concepts, propulsion, rocket construction, rocket design, launch procedures, failure modes, qualifications for building rockets, careers in rocketry, competitions in amateur rocketry, and information about NASA. The categories with the greatest percentage of student learning were rocket design, qualifications for building rockets, launch procedures, and propulsion. This indicates that more time should be spent in future presentations on the other learning categories, but further sampling is required. The data is given in Table 69 and illustrated in Figure 162.

Subject Interest Count Science 39

Math 35

Engineering 36

Technology 33

Rocketry 14

None 15

Presentation Learning Outcomes Count Forces 5

Propulsion 22


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V) Project Plan

3% 13%

5%

19%

13% 8%

15%

9% 3%

7% 5%

Presentation Learning Outcomes

Forces Propulsion Construction Design

Procedures Failure Modes Qualifications Careers

Competitions NASA None

Table 69: Presentation Learning Outcomes

Figure 160: Presentation Learning Outcomes

When asked what their favorite part of the presentation was, the greatest number of students responded in favor of the water bottle rocket activity. The data is given in Table 70 and illustrated in Figure 161.

Construction 8

Design 31

Procedures 21

Failure Modes 13

Qualifications 24

Careers 14

Competitions 5

NASA 11

None 8


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V) Project Plan

13%

4%

6%

12%

32%

25%

8%

Favorite Part

Launch Launch Failure Flight

Outside Water Rockets Big Rocket

None

Table 66: Favorite Part

Figure 161: Favorite Part

The surveys conducted at the two middle schools on October 5, 2012 were free response. The Granbury events were the first conducted by the team. They provided a wide variety of student responses concerning the presentation and demonstration. This feedback will ultimately be used to formulate a comprehensive and unbiased multiple-choice survey to be conducted at subsequent events. This will boost the quality of questions posed at future presentations. November 9, 2012 The Tarleton Aeronautical Team co-hosted the Tarleton Star Party this semester. Before the sun set, the team presented information concerning the STEM fields and the necessity for growth and interest in those fields to the community and grade school students. The team discussed the importance of doing well in all levels of school as the knowledge scaffolds as students move from grade level to grade level. The NASA USLI competition and our involvement were also heavily discussed. Event participants were

Favorite Part Count Launch 11

Launch Failure 3

Flight 5

Outside 10

Water Rockets 27

Big Rocket 21

None 7


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V) Project Plan

given time for a question and answering session as we watched team members launch off a series of rockets before the sun set. Although the setting didn’t allow for a thorough educational evaluation, the team did receive many oral critiques and feedback, and the event was said to deliver a very positive and informative message. A few parents in the group mentioned that they better understood the importance of the STEM classes and they would be working more diligently with their children to encourage growth and understanding in their students’ science and math classes. Other participants wanted more information on the competition and what was expected of our group. Two of the children present expressed the desire to have our group visit their school and one of the adults present said she came to our event after watching our presentation to the junior high earlier that day. November 12, 2012 The Tarleton Aeronautical Team invited Glen Rose High School to come to Tarleton State University in Stephenville, TX to attend an educational outreach. Students were divided up into three random groups upon arriving. Identical materials were distributed to each team and they were given a twenty minute time limit to create a water bottle rocket that would be flight ready. The group did a debriefing session following the exercise. During the debriefing, the team discussed important aspects that were called upon during the exercise. These included communication skills, team work, and the ability to recall prior STEM topics. A discussion about STEM fields occurred. This was followed by a question and answering session, where the students were able to question our team about STEM, the competition, college, and future careers. An evaluation of the educational engagement proved that the outreach was highly valuable and influential. Students gave positive feedback about the event. It was stated that it was very informative and interesting. Their teacher appreciated the impact of the impromptu rocket building session as an introduction to the importance of the STEM fields. During the session, the team paid close attention to the classes the students were taking, in order to make the information relevant to what the students were already learning. This correlation helped the students understand the importance and relevance of STEM to their lives and their education. Furthermore, one of the students decided after visiting with us and our school to pursue a degree in computer science when he graduates high school.


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V) Project Plan

December 17, 2012 On December 17, 2012, members of the Tarleton Aeronautical Team held an outreach at Morgan Mill ISD in Morgan Mill, TX. The team hosted a “rocket fair” for the students. Students in grades 4-8 attended this portion of the event. The students rotated through a series of six stations where Tarleton Aeronautical Team members used a variety of formats to discuss STEM topics. The stations are discribed in Table 71 below.


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V) Project Plan

Station Description

History The history of space exploration & NASA were investigated at this table.

Vocabulary Word searches, matching, and an introduction to rocketry related words. A small rocket was used as a visual for this station.

Math

Discovering and calculating speed & velocity. Students were able to walk and find the distance, time and direction of their travel. They then learned how to calculate speed & velocity. Discussions about the correlation to rocketry occurred.

Technology Video presentations of rocket launches. Discussion about STEM fields and their importance. Discussion about the USLI competition

Problem Solving

Wind tunnel discussion. Students cut paper cups to experiment with the concepts of stability and lift. To test their designs, students placed their creations in a team created wind tunnel. Results were discussed as well as STEM implications.

Art

Students were provided coffee filter parachutes so that they could learn to fold a parachute as our team does before each launch. A parachute from our sub-scale rocket is used for the demonstration. Students were then able to create their own design for their parachute. Students learned about the recovery systems of our competition rocket. Duel deployment was discussed as well as how the STEM fields were used in conjunction with choosing our parachutes for the competition.

Table 67: Educational Outreach Stations

Figure 162: Students won NASA stickers for answering questions


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V) Project Plan

The team conducted a pre-test before the students participated in the stations and a post-test consisting of six questions. Each question related to the content discussed during the six various stations in the rocket fair. The difference in the student’s responses from pre-test to post-test was quite exciting. The group, as a whole, showed growth in knowledge gained. Questions that were commonly left unanswered on the pre-test were correctly answered following the rocket fair.

On the pre-test, students were prompted with these questions, “What does NASA stand for?”, and “What does STEM stand for?” The students’ responses were all a variation on the team’s favorites, “National Asteroid and Space Association” and “Space Team Economic Mission”. When pretesting, most students incorrectly answered these

Figure 163: Interactive Physiics at Morgan Mill


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V) Project Plan

questions or chose to leave them blank. However, on the post-tests nearly every student could accurately identify what STEM stands for, and over half could identify the entire NASA acronym. Though some students did not correctly answer what NASA stands for, many did list a variety of facts they learned about NASA and thus showed a growth in knowledge. Beyond paper testing, a group discussion was conducted concerning the topics presented at the stations. Once started, the discussions quickly became student-led and had to be stopped and redirected by our team in order to cover all the topics in the allotted time. The Tarleton Aeronautical Team truly felt this was a success because it demonstrated the amount of information the students had received and the interest they found in the topics!

The verbal feedback from the teachers and school principle was highly positive and there was interest expressed in the team returning to the school again. After the rocket fair, the entire school (grades K-8) joined the team outside for a rocket launch. Discussions about the STEM fields continued, and students were given the

Figure 164: Preparing to Launch at BluffDale


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V) Project Plan

opportunity to ask questions related to our visit. Students were also given a demonstration about how to properly prepare a rocket for launch. Following our event, the team was invited to stay for lunch and recess. Members of our group were able to talk in a less formal environment with students in grades 4-8 as they enjoyed their lunch and recess break. Photos of the event are available on the team’s Facebook page under the “Morgan Mill ISD Outreach” album. December 18, 2012 The Tarleton Aeronautical Team held an educational outreach event at Bluff Dale ISD in Bluff Dale, TX on December 18, 2012. The team hosted a “rocket fair” for the students. Initially it was planned for grades four through eight, but upon arriving at the school grades 2-4 were added to the group. Students rotated through a series of seven stations where Tarleton Aeronautical Team members discussed STEM topics in a variety of ways. The stations are described in Table 72.

https://www.facebook.com/TarletonAeronauticalTeam/photos_albums


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V) Project Plan

Students in grades five through eight were given a pre-test and post-test at the outreach session. Students showed a large amount of improvement in their knowledge concerning NASA and the STEM fields. The pre-test and post-test asked the same questions and some students who could not answer any questions prior to the outreach, answered all questions correctly on the post-test. The final question on the test asks the student to list words that come to mind when they think of rockets. The most popular responses on the pretests are “fire,” “space,” and “loud” as a response. Tests returned

Station Description

History The history of space exploration & NASA were investigated at this table.

Vocabulary Word searches, matching, and an introduction to rocketry related words. A small rocket was used as a visual for this station.

Math

Discovering and calculating speed & velocity. Students were able to walk and find the distance, time and direction of their travel. They then learned how to calculate speed & velocity. Discussions about the correlation to rocketry occurred.

Technology Video presentations of rocket launches. Discussion about STEM fields and their importance. Discussion about the USLI competition

Problem Solving

Wind tunnel discussion. Students cut paper cups to experiment with the concepts of stability and lift. To test their designs, students placed their creations in a team created wind tunnel. Results were discussed as well as STEM implications.

Art Students were able to create their own design for their parachute. Further discussion concerning STEM fields and the NASA competition were conducted.

Recovery

(This table was split from the art station into a new station due to the large number of students attending the event.) Students learned about the recovery systems of our competition rocket. Duel deployment was discussed as well as how the STEM fields were used in conjunction with choosing our parachutes for the competition. Students were provided coffee filter parachutes so that they could learn to fold a parachute as our team does before each launch. A parachute from our sub-scale rocket is used for demonstration.

Table 68: Educational Outreach Stations


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V) Project Plan

after the rocket fair included words such as “launch,” “airframe,” “apogee,” and “Tarleton.”

The students in grades 2-4 were verbally evaluated, and they too demonstrated a growth in understanding of the topics presented during the team’s visit. For example, many students did not know what NASA or STEM stood for, but could explain them after attending the rocket fair. Verbal feedback from students indicated excitement about rocketry and NASA, as well as an interest in science. The most mentioned station was the wind tunnel, the parachute folding, and the parachute design tables.

After the rocket fair, the entire school (grades K-8) joined the team outside for a rocket launch. Discussions about the STEM fields continued and students were given the opportunity to ask questions related to our visit. Photos of the event are available on the team’s Facebook page. Visit in the album entitled “Bluff Dale ISD Outreach.”

Figure 166: Students Enjoying the Art Station, Decorating Parachutes

Figure 165: Students Learning at the Recovery Station at Dublin

Middle School

https://www.facebook.com/TarletonAeronauticalTeam/photos_albums


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VI) Conclusion

VI) Conclusion The team is eager to continue the design process and see the final product come to fruition. Up to this point, the project is both challenging and rewarding. Creative, legitimate solutions have been posed and implemented, defining the performance characteristics of the vehicle and payload systems as a whole. The vehicle design features reflect standard practices in amateur rocketry. All components of the vehicle aim to achieve the goal of delivering the SMD payload to one mile above ground level. The SMD payload criterion calls for an atmospheric data gathering instrument that meets all requirements as stated in the SOW. Controlling the sensors to make measurements is among a long list of complicated problems creating a fully functional payload. Other taxing problems include the orientation of the image camera and the placement of the UV and solar irradiance sensors. The self-leveling camera design aims to take proper images by correcting offsets in the payload orientation during descent. Pyranometers and UV sensors sit in an opposing manner to effectively increase the field of view for solar irradiance and ultraviolet measurement. Optical properties of the clear acrylic payload housing must be addressed, and sensor performance testing must occur to verify this housing structure selection. The payload will ultimately be inserted in a PCB due to advantages in size, efficiency, repeatability and overall fidelity in design. The educational outreach portion of the project is going extremely well. The minimum requirement for the number of students to be reached was exceeded on first day of this competition. The team continues to go above and beyond this minimum requirement. The intention of the team is to expose as many students to the STEM fields as possible. With the support of the community and University, the team is building upon the momentum spawned by the success in the 2012 CanSat competition. The team is eager to progress through the design life cycle. Ultimately, flying the final vehicle and payload on launch day will illustrate the team’s achievements. It is a true testament to the abilities and ingenuity of the team members, and it provides an invaluable exercise in creating real-world engineering experience.

critical design review - tarleton.educritical design review i note to reader: to facilitate the...

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