[Type text] I) Summary of CDR Report [Type text]
Critical Design
Review
2012-2013 NASA USLI
“Research is what I’m doing wh...
Tarleton State University
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Note to reader:
To facilitate the reading of the Critical Design Review...
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Table of Contents
I) Summary of CDR Report ..............................
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Payload Element Compatibility ..........................................
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The Safety Officer.......................................................
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Table of Figures
Figure 1: Launch Vehicle Specifications..................
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Figure 46: Test Flight Ten Vehicle ......................................
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Figure 91: UV Sensor Mounting ..........................................
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Figure 137: Bulkhead Recessed Slot ....................................
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Index of Tables
Table 1: Vehicle Size and Mass ..........................
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Table 46: Payload Framework Dimensions....................................
I) Summary of CDR Report
I) Summary of CDR Report
Team Summary
Tarleton Aeronautical Team
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Box T-...
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II) Changes Made since PDR
II) Changes Made since PDR
Changes Made to V...
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II) Changes Made since PDR
Recovery Rationale
Tubular Kevlar shock cord...
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II) Changes Made since PDR
The team made no significant changes to the ...
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III) Vehicle Criteria
III) Vehicle Criteria
Design and Verification of ...
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III) Vehicle Criteria
conditions are met and an apogee of exactly 5,280...
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III) Vehicle Criteria
Date Milestone Description Status Type
8/31/12 Pr...
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III) Vehicle Criteria
Date Milestone Description Status Type
SMD)
1/21/...
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III) Vehicle Criteria
Review the design at a system level
Final Drawing...
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III) Vehicle Criteria
The Upper Body Airframe is 28.0 inches long. Thi...
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III) Vehicle Criteria
The Acrylic Housing Structure is 36 inches long ...
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III) Vehicle Criteria
The Booster Section of the vehicle is 36 inches ...
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III) Vehicle Criteria
Final analysis and model results, anchored to te...
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III) Vehicle Criteria
Tube Crushes
Using a hydraulic press, the spare
...
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III) Vehicle Criteria
Angle of force on fin
(degrees)
45 60 75 90
Max ...
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III) Vehicle Criteria
Final motor selection
The selected motor is a Ce...
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III) Vehicle Criteria
Motor
Apogee
(ft.)
Velocity
Off Rail
(ft./s)
Tot...
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III) Vehicle Criteria
Figure 10 is the motor thrust curve provided on ...
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III) Vehicle Criteria
Verification of System Level Functional Requirem...
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III) Vehicle Criteria
Requirement
(SOW)
Vehicle Requirement
Satisfying...
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III) Vehicle Criteria
Requirement
(SOW)
Vehicle Requirement
Satisfying...
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III) Vehicle Criteria
Requirement
(SOW)
Vehicle Requirement
Satisfying...
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III) Vehicle Criteria
Requirement
(SOW)
Vehicle Requirement
Satisfying...
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III) Vehicle Criteria
degree of workmanship of the vehicle design. The...
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III) Vehicle Criteria
Root Chord: 12 inches
Tip Chord: 0 inches
Height...
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III) Vehicle Criteria
deployment.
The payload bulkheads are constructe...
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III) Vehicle Criteria
Figure 13: Booster Assembly Steps 1-4
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III) Vehicle Criteria
Figure 14: Booster Assembly Steps 5-8
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III) Vehicle Criteria
Figure 15: Booster Assembly Steps 9-12
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III) Vehicle Criteria
Figure 16: Coupler Assembly Procedure
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III) Vehicle Criteria
Figure 17: Avionics Assembly Steps 1-3
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III) Vehicle Criteria
Figure 18: Avionics Assembly Steps 4-6
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III) Vehicle Criteria
Figure 19: Payload Assembly
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III) Vehicle Criteria
Proper Attachment and Alignment of elements
The ...
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III) Vehicle Criteria
Fins are aligned using pre-measured slots in the...
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III) Vehicle Criteria
Status of verification
The verification table sh...
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III) Vehicle Criteria
Mass Statement
The mass summary of the vehicle i...
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III) Vehicle Criteria
Mass Summary
Subsection Mass (oz.) Mass (lb.)
Pa...
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III) Vehicle Criteria
Mass per Subsection
Payload
Component Quantity M...
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III) Vehicle Criteria
Component Quantity Mass (oz.) Total Mass (oz.)
N...
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III) Vehicle Criteria
Flight Data
The following table, Table 11, provi...
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III) Vehicle Criteria
Date Apogee Predicted
Apogee
Drift Video link
No...
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III) Vehicle Criteria
Sub-scale Test Flight One
On November 30, 2012 a...
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III) Vehicle Criteria
Simulated Flight
The simulation for this test la...
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III) Vehicle Criteria
Flight Analysis and Impact on Design
Dual deploy...
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III) Vehicle Criteria
weighed and added to the mass for the repeat lau...
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III) Vehicle Criteria
Simulated Flight
The simulation for this test la...
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III) Vehicle Criteria
Flight Analysis and Impact on Design
Dual deploy...
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III) Vehicle Criteria
On December 7, 2012 at Hunewell Ranch, one launc...
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III) Vehicle Criteria
Actual Flight
No flight data was recovered or re...
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III) Vehicle Criteria
On December 8, 2012 at Hunewell Ranch, two launc...
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III) Vehicle Criteria
Actual Flight
The actual data from the flight wa...
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III) Vehicle Criteria
Flight Analysis and Impact on Design
The G78 lau...
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III) Vehicle Criteria
ejection and a 1.0 gram 3F black powder charge f...
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III) Vehicle Criteria
Flight Analysis and Impact on Design
Dual deploy...
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III) Vehicle Criteria
Date Location Coordinate Motor
December 14, 2012...
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III) Vehicle Criteria
Actual Flight
No flight data was recorded or rel...
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III) Vehicle Criteria
main parachute ejection; using a 1.38 gram 3F bl...
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III) Vehicle Criteria
Actual Flight
The actual data from the flight wa...
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III) Vehicle Criteria
Flight Eight
The subscale vehicle, shown above i...
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III) Vehicle Criteria
Actual Flight
No flight data was recovered or re...
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III) Vehicle Criteria
Flight Nine
The subscale vehicle, shown above in...
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NASA CDRfinal

Published on: Mar 3, 2016
Source: www.slideshare.net


Transcripts - NASA CDRfinal

  • 1. [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
  • 2. Tarleton State University Critical Design Review i 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.
  • 3. Tarleton State University Critical Design Review ii 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
  • 4. Tarleton State University Critical Design Review iii 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
  • 5. Tarleton State University Critical Design Review iv 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
  • 6. Tarleton State University Critical Design Review v 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
  • 7. Tarleton State University Critical Design Review vi 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
  • 8. Tarleton State University Critical Design Review vii 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
  • 9. Tarleton State University Critical Design Review viii 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
  • 10. Tarleton State University Critical Design Review ix 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
  • 11. Tarleton State University Critical Design Review x 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
  • 12. 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
  • 13. Tarleton State University Critical Design Review 2 II) Changes Made since PDR 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
  • 14. Tarleton State University Critical Design Review 3 II) Changes Made since PDR 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
  • 15. Tarleton State University Critical Design Review 4 II) Changes Made since PDR The team made no significant changes to the project plan.
  • 16. Tarleton State University Critical Design Review 5 III) Vehicle Criteria 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
  • 17. Tarleton State University Critical Design Review 6 III) Vehicle Criteria 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.
  • 18. Tarleton State University Critical Design Review 7 III) Vehicle Criteria 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
  • 19. Tarleton State University Critical Design Review 8 III) Vehicle Criteria Date Milestone Description Status Type SMD) 1/21/13 Post Launch Failure Analysis TBD Verification 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 1/28/13 Post Launch Failure Analysis TBD Verification 2/1/13 CDR Presentations TBD USLI 2/2/13 SMD Award 2 ($1400) TBD USLI 2/2/13 Low Altitude, Full Scale Launch (w/ SMD) TBD Test Launch 2/4/13 Post Launch Failure Analysis TBD Verification 2/11/13 FRR Q&A TBD USLI 2/16/13 Low Altitude, Full Scale Launch (w/ SMD) TBD Test Launch 2/18/13 Post Launch Failure Analysis TBD Verification 2/23/13 High Altitude, Full Scale Launch (w/SMD) TBD Test Launch 2/24/13 High Altitude, Full Scale Launch (w/SMD) TBD Test Launch 2/25/13 Post Launch Failure Analysis TBD Verification 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 4/4/13 SMD Award 3 ($400) 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 5/7/13 SMD Award 4 ($200) TBD USLI 5/17/13 Winners Announced TBD USLI Table 5: Project Milestones Continued
  • 20. Tarleton State University Critical Design Review 9 III) Vehicle Criteria 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
  • 21. Tarleton State University Critical Design Review 10 III) Vehicle Criteria 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
  • 22. Tarleton State University Critical Design Review 11 III) Vehicle Criteria 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
  • 23. Tarleton State University Critical Design Review 12 III) Vehicle Criteria 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
  • 24. Tarleton State University Critical Design Review 13 III) Vehicle Criteria 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
  • 25. Tarleton State University Critical Design Review 14 III) Vehicle Criteria 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
  • 26. Tarleton State University Critical Design Review 15 III) Vehicle Criteria 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
  • 27. Tarleton State University Critical Design Review 16 III) Vehicle Criteria 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
  • 28. Tarleton State University Critical Design Review 17 III) Vehicle Criteria 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
  • 29. Tarleton State University Critical Design Review 18 III) Vehicle Criteria 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
  • 30. Tarleton State University Critical Design Review 19 III) Vehicle Criteria 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
  • 31. Tarleton State University Critical Design Review 20 III) Vehicle Criteria 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
  • 32. Tarleton State University Critical Design Review 21 III) Vehicle Criteria Requirement (SOW) Vehicle Requirement Satisfying Design Feature 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
  • 33. Tarleton State University Critical Design Review 22 III) Vehicle Criteria Requirement (SOW) Vehicle Requirement Satisfying Design Feature 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
  • 34. Tarleton State University Critical Design Review 23 III) Vehicle Criteria Requirement (SOW) Vehicle Requirement Satisfying Design Feature 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
  • 35. Tarleton State University Critical Design Review 24 III) Vehicle Criteria 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.
  • 36. Tarleton State University Critical Design Review 25 III) Vehicle Criteria 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
  • 37. Tarleton State University Critical Design Review 26 III) Vehicle Criteria 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.
  • 38. Tarleton State University Critical Design Review 27 III) Vehicle Criteria Figure 13: Booster Assembly Steps 1-4
  • 39. Tarleton State University Critical Design Review 28 III) Vehicle Criteria Figure 14: Booster Assembly Steps 5-8
  • 40. Tarleton State University Critical Design Review 29 III) Vehicle Criteria Figure 15: Booster Assembly Steps 9-12
  • 41. Tarleton State University Critical Design Review 30 III) Vehicle Criteria Figure 16: Coupler Assembly Procedure
  • 42. Tarleton State University Critical Design Review 31 III) Vehicle Criteria Figure 17: Avionics Assembly Steps 1-3
  • 43. Tarleton State University Critical Design Review 32 III) Vehicle Criteria Figure 18: Avionics Assembly Steps 4-6
  • 44. Tarleton State University Critical Design Review 33 III) Vehicle Criteria Figure 19: Payload Assembly
  • 45. Tarleton State University Critical Design Review 34 III) Vehicle Criteria 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
  • 46. Tarleton State University Critical Design Review 35 III) Vehicle Criteria 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
  • 47. Tarleton State University Critical Design Review 36 III) Vehicle Criteria 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
  • 48. Tarleton State University Critical Design Review 37 III) Vehicle Criteria 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.
  • 49. Tarleton State University Critical Design Review 38 III) Vehicle Criteria 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
  • 50. Tarleton State University Critical Design Review 39 III) Vehicle Criteria 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
  • 51. Tarleton State University Critical Design Review 40 III) Vehicle Criteria 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
  • 52. Tarleton State University Critical Design Review 41 III) Vehicle Criteria 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.
  • 53. Tarleton State University Critical Design Review 42 III) Vehicle Criteria 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 N/A 575 ft AGL N/A N/A 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 N/A 520 ft AGL N/A N/A December 15, 2012 666ft AGL 615 ft AGL 50ft N/A December 15, 2012 476ft AGL 530 ft AGL 50ft N/A December 19, 2012 929ft AGL 1848 ft AGL 833ft N/A 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 2,271ft AGL 3214 ft AGL 412ft Video January 6, 2013 2,920ft AGL 3559 ft AGL 693ft Video January 7, 2013 2,402ft AGL 2908 ft AGL N/A N/A Table 11: Flight Data
  • 54. Tarleton State University Critical Design Review 43 III) Vehicle Criteria 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
  • 55. Tarleton State University Critical Design Review 44 III) Vehicle Criteria 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
  • 56. Tarleton State University Critical Design Review 45 III) Vehicle Criteria 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
  • 57. Tarleton State University Critical Design Review 46 III) Vehicle Criteria 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 Altimeter Drogue Charge Size Main Charge Size Main Deploy Altitude Featherweight Raven3 1.0 g 1.38 g 256 ft AGL Temperature Wind Pressure Elevation 70° F 5 mph 31.29 in Hg 1309 ft Figure 26: Test Flight Two Vehicle Table 13: Test Flight Two Conditions
  • 58. Tarleton State University Critical Design Review 47 III) Vehicle Criteria Simulated Flight The simulation for this test launch is shown below in Figure 27. 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 28. Figure 27: Simulated Flight Two Data
  • 59. Tarleton State University Critical Design Review 48 III) Vehicle Criteria Flight Analysis and Impact on Design 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 28: Raven3 Flight Data Figure 29: Test Flight Three Vehicle
  • 60. Tarleton State University Critical Design Review 49 III) Vehicle Criteria 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. Date Location Coordinate Motor December 7, 12 Hunewell (32.216114, - 98.096019) Cesaroni G78 Blue Streak Altimeter Drogue Charge Size Main Charge Size Main Deploy Altitude Featherweight Raven3 1.0 g 1.38 g 256 ft AGL Temperature Wind Pressure Elevation 49° F 6 mph 28.43 in Hg 1309 ft Simulated Flight The simulation for this test launch is shown below in Figure 30. This simulation was conducted through OpenRocket. Table 14: Test Flight Three Conditions
  • 61. Tarleton State University Critical Design Review 50 III) Vehicle Criteria Actual Flight No flight data was recovered or reliable for this test launch. Flight Analysis and Impact on Design 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
  • 62. Tarleton State University Critical Design Review 51 III) Vehicle Criteria 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 Altimeter Drogue Charge Size Main Charge Size Main Deploy Altitude Featherweight Raven3 1.38 g 1.6 g 256 ft AGL Temperature Wind Pressure Elevation 54° F 5 mph 29.96 in Hg 1309 ft Simulated Flight The simulation for this test launch is shown below in Figure 32. This simulation was conducted through OpenRocket. Figure 31: Test Flight Four Vehicle Table 15: Test Flight Four Conditions
  • 63. Tarleton State University Critical Design Review 52 III) Vehicle Criteria Actual Flight The actual data from the flight was acquired from the onboard Raven3 altimeter, and shown below in Figure 33. Figure 32: Simulated Test Flight Four Data
  • 64. Tarleton State University Critical Design Review 53 III) Vehicle Criteria Flight Analysis and Impact on Design 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 33: Raven3 Flight Data Figure 34: Test Flight Five Vehicle
  • 65. Tarleton State University Critical Design Review 54 III) Vehicle Criteria 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 8, 2012 Hunewell (32.216114, - 98.096019) 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 The simulation for this test launch is shown below in Figure 35. This simulation was conducted through OpenRocket. Actual Flight No flight data was recovered or reliable for this test launch. Table 16: Test Flight Five Conditions Figure 35: Raven3 Test Flight Five Data
  • 66. Tarleton State University Critical Design Review 55 III) Vehicle Criteria Flight Analysis and Impact on Design 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
  • 67. Tarleton State University Critical Design Review 56 III) Vehicle Criteria Date Location Coordinate Motor December 14, 2012 Hunewell (32.216114, -98.096019) Cesaroni G79 Smoky Sam Altimeter Drogue Charge Size Main Charge Size Main Deploy Altitude 2x Featherweight Raven3 0.8 g 1.0 g 256 ft AGL Temperature Wind Pressure Elevation 58° F 16 mph 29.93 in Hg 1309 ft Simulated Flight The simulation for this test launch is shown below in Figure 37. This simulation was conducted through OpenRocket. Table 17: Test Flight Six Conditions Figure 37: Simulated Flight Six Data
  • 68. Tarleton State University Critical Design Review 57 III) Vehicle Criteria Actual Flight No flight data was recorded or reliable for this test launch. Flight Analysis and Impact on Design 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
  • 69. Tarleton State University Critical Design Review 58 III) Vehicle Criteria 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 Altimeter Drogue Charge Size Main Charge Size Main Deploy Altitude Featherweight Raven3 1.38 g 5.0 g 128 ft AGL Temperature Wind Pressure Elevation 67° F 6 mph 29.93 in Hg 427 ft Simulated Flight The simulation for this test launch is shown below in Figure 39. This simulation was conducted through OpenRocket. Table 18: Test Flight Seven Conditions Figure 39: Test Flight Seven Simulated Data
  • 70. Tarleton State University Critical Design Review 59 III) Vehicle Criteria Actual Flight The actual data from the flight was acquired from the onboard Raven3 altimeter, and shown below in Figure 40. 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
  • 71. Tarleton State University Critical Design Review 60 III) Vehicle Criteria 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 Altimeter Drogue Charge Size Main Charge Size Main Deploy Altitude Stratologger SL100 1.38 g 2.5 g 200 ft AGL Temperature Wind Pressure Elevation 74° F 10 mph 29.74 in Hg 427 ft Simulated Flight The simulation for this test launch is shown below in Figure 42. This simulation was conducted through OpenRocket. Figure 41: Test Flight Eight Vehicle Table 19: Test Flight Eight Conditions
  • 72. Tarleton State University Critical Design Review 61 III) Vehicle Criteria Actual Flight No flight data was recovered or reliable for this test launch. 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
  • 73. Tarleton State University Critical Design Review 62 III) Vehicle Criteria 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 Altimeter Drogue Charge Size Main Charge Size Main Deploy Altitude Stratologger SL100 1.6 g 1.8 g 200 ft AGL Temperature Wind Pressure Elevation 77° F 19 mph 28.87inHg 1309 ft Simulated Flight The simulation for this test launch is shown below in Figure 44. This simulation was conducted through OpenRocket. Figure 43: Test Flight Nine Vehicle Table 20: Test Flight Nine Conditions

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