NASA CDRfinal

[Type text] I) Summary of CDR Report [Type text]
Critical Design
Review
2012-2013 NASA USLI
“Research is what I’m doing when I don’t know what I’m doing.” - Werner Von Braun

Tarleton State University
Critical Design Review
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Note to reader:
To facilitate the reading of the Critical Design Review, we have mirrored the Student
Launch Project Statement of Work. In the body of the CDR, you will find extensive detail
in the design of our SMD payload. The payload’s features are threefold; atmospheric
data gathering sensors, a self-leveling camera system, and a video camera. One of the
two major strengths of our payload design is the originality of our autonomous real-time
camera orientation system (ARTCOS). The other major strength can be found in the
originality of our self-designed Printed Circuit Board layouts. This feature alone
represents over 150 man hours of work. Along with space and power efficiencies, the
PCBs provide major enhancement of the signal integrity of the sensor data. For ease of
reading, you will find documents such as itemized final build budget and launch
procedures moved to the appendix along with Sensor and Material Safety Data sheets.
We have enjoyed the challenges presented in the writing of this document and submit it
for your review.

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Table of Contents
I) Summary of CDR Report ............................................................................................. 1
Team Summary ........................................................................................................... 1
Launch Vehicle Summary ............................................................................................ 1
Payload Summary........................................................................................................ 1
II) Changes Made since PDR.......................................................................................... 2
III) Vehicle Criteria........................................................................................................... 5
Design and Verification of Launch Vehicle................................................................... 5
Launch Vehicle Mission Statement........................................................................... 5
Mission Success Criteria .......................................................................................... 5
Review the design at a system level......................................................................... 9
Verification of System Level Functional Requirements........................................... 19
Approach to Workmanship ..................................................................................... 23
Additional Planned Component, Functional, or Static Testing ................................ 24
Status and Plans of remaining manufacturing and assembly ................................. 24
Discuss the integrity of design ................................................................................ 24
Safety and Failure Analysis .................................................................................... 40
Subscale Flight Results ............................................................................................. 40
Subscale Flight Results ............................................................................................. 40
Flight Data .............................................................................................................. 41
Impact on Design Summary ................................................................................... 79
Recovery Subsystem ................................................................................................. 81
Physical Components............................................................................................. 81
Electrical Components............................................................................................ 86
Kinetic Energy......................................................................................................... 96
Test Results............................................................................................................ 98
Safety and Failure Analysis .................................................................................. 102
Mission Performance Predictions............................................................................. 116
Mission Performance Criterion.............................................................................. 116
Payload Integration.................................................................................................. 121
Payload Integration Plan....................................................................................... 121
Payload Installation and Removal......................................................................... 124
Payload Interface Dimensions .............................................................................. 126

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Payload Element Compatibility ............................................................................. 128
Simplicity of Integration Procedure ....................................................................... 128
Launch Concerns and Operation Procedures.......................................................... 128
Launch procedures............................................................................................... 128
Pre-launch Checklists and Procedures:................................................................ 128
Safety Materials Checklist .................................................................................... 128
Structure Preparation:........................................................................................... 129
Recovery Procedures: .......................................................................................... 130
Motor Preparation:................................................................................................ 132
Launch Checklist and Procedures ........................................................................ 133
Troubleshooting:................................................................................................... 134
In-Flight Inspection ............................................................................................... 135
Post-Flight Inspection ........................................................................................... 135
Travel.................................................................................................................... 135
Safety and Environment........................................................................................... 137
Failure Modes....................................................................................................... 137
Hazard Analysis.................................................................................................... 142
Environment.......................................................................................................... 146
IV) Payload Criteria ..................................................................................................... 149
Testing and Design of Payload Experiment ............................................................. 149
Design Review at a System Level ........................................................................ 149
System Level Functional Requirements ............................................................... 161
Approach to Workmanship ................................................................................... 163
Test Plan of Components and Functionality ......................................................... 163
Status and Plans of Remaining Manufacturing and Assembly ............................. 189
Integration Plan..................................................................................................... 192
Precision of Instrumentation and Repeatability of Measurements ........................ 194
Safety and Failure Analysis .................................................................................. 197
Uniqueness and Significance ............................................................................... 201
Suitable Level of Challenge .................................................................................. 201
Science Value.......................................................................................................... 202
Experimental Logic, Approach, and Method of Investigation ................................ 203
Relevance of Expected Data and Accuracy/Error Analysis .................................. 204
Safety and Environment........................................................................................... 205

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The Safety Officer................................................................................................. 205
Failure Modes....................................................................................................... 206
Hazard Analysis.................................................................................................... 209
Environment.......................................................................................................... 212
V) Project Plan ............................................................................................................ 213
Budget Summary ..................................................................................................... 213
Funding Plan......................................................................................................... 225
Timeline ................................................................................................................... 225
Testing Timeline ................................................................................................... 227
Outreach Timeline ................................................................................................ 228
Education plan ......................................................................................................... 228
Outreach Plan....................................................................................................... 228
Accomplished Educational Outreach.................................................................... 231
VI) Conclusion............................................................................................................. 243

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Table of Figures
Figure 1: Launch Vehicle Specifications.......................................................................... 9
Figure 2: Upper body Airframe...................................................................................... 10
Figure 3: Clear Payload Housing .................................................................................. 11
Figure 4: Booster Section.............................................................................................. 12
Figure 5: Epoxy Strength Testing.................................................................................. 13
Figure 6: Acrylic Compression Testing.......................................................................... 14
Figure 7: Fin Testing Set up.......................................................................................... 15
Figure 8: Fin Detachment from the Motor Tube ............................................................ 16
Figure 9: Cesaroni L1720 Motor Thrust Curve from ThrustCurve.org ........................... 17
Figure 10: L1720-WT Thrust Curve from Cesaroni ....................................................... 18
Figure 11: Tarleton Aeronautical Team's Generated Thrust Curve ............................... 19
Figure 12: Fin Dimensions ............................................................................................ 25
Figure 13: Booster Assembly Steps 1-4........................................................................ 27
Figure 14: Booster Assembly Steps 5-8........................................................................ 28
Figure 15: Booster Assembly Steps 9-12...................................................................... 29
Figure 16: Coupler Assembly Procedure....................................................................... 30
Figure 17: Avionics Assembly Steps 1-3....................................................................... 31
Figure 18: Avionics Assembly Steps 4-6....................................................................... 32
Figure 19: Payload Assembly........................................................................................ 33
Figure 20: Ballast System Assembly............................................................................. 34
Figure 21: Positive Motor Retainer................................................................................ 35
Figure 22: Launch Vehicle Illustration ........................................................................... 36
Figure 23: Test Flight One Vehicle................................................................................ 43
Figure 24: Test Flight One Simulation........................................................................... 44
Figure 25: Raven3 Flight Data ...................................................................................... 45
Figure 26: Test Flight Two Vehicle................................................................................ 46
Figure 27: Simulated Flight Two Data ........................................................................... 47
Figure 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 34: Test Flight Five Vehicle................................................................................ 53
Figure 35: Raven3 Test Flight Five Data....................................................................... 54
Figure 36: Test Flight Six Vehicle.................................................................................. 55
Figure 37: Simulated Flight Six Data............................................................................. 56
Figure 38: Test Flight Seven Vehicle............................................................................. 57
Figure 39: Test Flight Seven Simulated Data................................................................ 58
Figure 40: Raven3 Test Flight Seven Data ................................................................... 59
Figure 41: Test Flight Eight Vehicle .............................................................................. 60
Figure 42: Simulated Flight Eight Data.......................................................................... 61
Figure 43: Test Flight Nine Vehicle ............................................................................... 62
Figure 44: Simulated Flight Nine Data........................................................................... 63
Figure 45: Flight Nine GPS Data................................................................................... 63

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Figure 46: Test Flight Ten Vehicle ................................................................................ 64
Figure 47: Simulated Flight Ten Data............................................................................ 65
Figure 48: Flight Ten GPS Data .................................................................................... 66
Figure 49: Test Flight Eleven Vehicle............................................................................ 67
Figure 50: Simulated Flight Eleven Data....................................................................... 68
Figure 51: Test Flight Eleven GPS Data ....................................................................... 68
Figure 52: Test Flight Twelve Vehicle ........................................................................... 69
Figure 53: Simulated Test Flight Twelve ....................................................................... 70
Figure 54: Raven3 Test Flight Twelve Data .................................................................. 71
Figure 55: Test Flight Thirteen Vehicle.......................................................................... 71
Figure 56: Test Flight Thirteen Simulation..................................................................... 72
Figure 57: Test Flight Thirteen Stratologger Data ......................................................... 73
Figure 58: Test Flight Fourteen Vehicle ........................................................................ 74
Figure 59: Test Flight Fourteen Simulation ................................................................... 75
Figure 60: Test Flight Fourteen Stratologger Data ........................................................ 76
Figure 61: Test Flight Fifteen Vehicle............................................................................ 77
Figure 62: Test Flight Fifteen Simulation....................................................................... 78
Figure 63: Test Flight Fifteen Stratologger Data ........................................................... 78
Figure 64: Ejection Canister.......................................................................................... 80
Figure 65: 3F Black Powder .......................................................................................... 80
Figure 66: Astro 320 GPS System ................................................................................ 80
Figure 67: SkyAngle XXLarge Deployment Freebag..................................................... 82
Figure 68: Main Parachute Attachment Scheme........................................................... 83
Figure 69: Attachment Scheme to Couplers.................................................................. 84
Figure 70: Drogue Parachute Attachment Scheme....................................................... 85
Figure 71: Altimeter Electronics Schematics................................................................. 87
Figure 72: Raven3 Software Flow Diagram................................................................... 89
Figure 73: Stratologger Software Flow Diagram ........................................................... 91
Figure 74: Example Drogue/Main Avionics Bay ............................................................ 93
Figure 75: Drawing of Avionics Sleds............................................................................ 94
Figure 76: GPS Software Flow Diagram ....................................................................... 95
Figure 77: Launch Vehicle Prototype ............................................................................ 96
Figure 78: Final Vehicle Simulation............................................................................... 96
Figure 79: Final Vehicle Simulation............................................................................. 117
Figure 80: Input Parameters for Final Simulation ........................................................ 118
Figure 81: L1720-WT Actual Thrust Curve.................................................................. 118
Figure 82: Rear Payload Bulkhead to Frame Connection ........................................... 122
Figure 83: Telemetry Verification GUI ......................................................................... 123
Figure 84: SMD Payload ............................................................................................. 124
Figure 85: SMD Payload attached with Avionic Bays.................................................. 125
Figure 86: Aluminum Angle ......................................................................................... 127
Figure 87: Altimeter Wiring Diagrams.......................................................................... 131
Figure 88: Materials and Components (Image obtained from the Cesaroni Pro 75 mm
Motor Assembly Kit Instructions)................................................................................. 133
Figure 89: Payload ...................................................................................................... 149
Figure 90: Upper Payload Circuit Boards.................................................................... 150

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Figure 91: UV Sensor Mounting .................................................................................. 150
Figure 92: ARTCOS .................................................................................................... 151
Figure 93: Test Flight Data.......................................................................................... 153
Figure 94: Test Flight Humidity Data........................................................................... 154
Figure 95: Launch Pad Humidity Data......................................................................... 155
Figure 96: Test Flight Temperature Data .................................................................... 156
Figure 97: Launch Pad Temperature Data.................................................................. 156
Figure 98: Correlation between Temperature and Humidity........................................ 157
Figure 99: Test Flight Pressure Data........................................................................... 158
Figure 100: Test Flight Altitude Data........................................................................... 158
Figure 101: Test Flight GPS Data ............................................................................... 159
Figure 102: Test Flight Solar Irradiance Data.............................................................. 160
Figure 103: ARTCOS Image ....................................................................................... 161
Figure 104: BMP 180 Pressure Sensor Wiring............................................................ 164
Figure 105: BMP 180 Software Flowchart................................................................... 165
Figure 106: TSL2561 Pyranometer Wiring.................................................................. 166
Figure 107: TSL2561 Pseudo Code............................................................................ 167
Figure 108: TSL2561 Lux Conversion Factors............................................................ 167
Figure 109: BMP 180 and TSL2561 Wiring................................................................. 168
Figure 110: HIH4030 Humidity Sensor Wiring............................................................. 168
Figure 111: HIH4030 Software.................................................................................... 169
Figure 112: HIH4030, BMP180, and TSL2561 Wiring................................................. 169
Figure 113: HH10D Humidity Sensor Wiring............................................................... 170
Figure 114: HH10D Humidity Calculation Algorithm.................................................... 170
Figure 115: HH10D, HIH4030, BMP180, and TSL2561 Wiring ................................... 171
Figure 116: SU100 Testing ......................................................................................... 172
Figure 117: SU100 UV Sensor Wiring......................................................................... 172
Figure 118: SU100 Software ....................................................................................... 173
Figure 119: GPS Wiring .............................................................................................. 174
Figure 120: MicroSD Wiring ........................................................................................ 175
Figure 121: XBee Wireless Transmitter Wiring ........................................................... 176
Figure 122: Digi Technical Support Forum Post.......................................................... 177
Figure 123: De-Soldering LED from XBee Adapter..................................................... 177
Figure 124: XBee Range Test..................................................................................... 178
Figure 125: Ground Station GUI.................................................................................. 179
Figure 126: ADGS Wiring Schematic .......................................................................... 180
Figure 127: VC0706 Camera Wiring ........................................................................... 181
Figure 128: VC0706 Configuration GUI....................................................................... 182
Figure 129: ARTCOS Mounting .................................................................................. 183
Figure 130: ARTCOS Mounting .................................................................................. 184
Figure 131: ARTCOS Orientation Algorithm................................................................ 184
Figure 132: ARTCOS Wiring Schematic ..................................................................... 186
Figure 133: Payload Block Diagram............................................................................ 188
Figure 134: Breakout Board Compatible PCB............................................................. 190
Figure 135: Surface Mount PCB ................................................................................. 191
Figure 136: Bulkhead Aluminum Frame Interface ....................................................... 192

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Figure 137: Bulkhead Recessed Slot .......................................................................... 193
Figure 138: Telemetry Verification GUI ....................................................................... 193
Figure 139: SU-100 Spectral Response...................................................................... 195
Figure 140: SP-110 Spectral Response...................................................................... 196
Figure 141: Clean Room ............................................................................................. 197
Figure 142: ARTCOS Epoxy Mounting Failure............................................................ 198
Figure 143: Post-Flight Payload .................................................................................. 198
Figure 144: GPS Mounting Failure.............................................................................. 199
Figure 145: PCB Board ............................................................................................... 200
Figure 146: Self-Leveling Camera System.................................................................. 201
Figure 147: Allocated Funds ....................................................................................... 213
Figure 148: Budget Status........................................................................................... 214
Figure 149: Vehicle Budget Status.............................................................................. 215
Figure 150: Payload Budget Status............................................................................. 215
Figure 151: Propulsion Budget Status......................................................................... 216
Figure 152: Outreach Budget Status........................................................................... 216
Figure 153: Early Funding........................................................................................... 225
Figure 154: Project Timeline ....................................................................................... 226
Figure 155: Testing Gantt Chart.................................................................................. 227
Figure 156: Outreach Timeline.................................................................................... 228
Figure 157: Acton Middle School ................................................................................ 229
Figure 158: Team Members Educate and Entertain Acton Students .......................... 231
Figure 159: Subject Interest ........................................................................................ 232
Figure 160: Presentation Learning Outcomes............................................................. 233
Figure 161: Favorite Part............................................................................................. 234
Figure 162: Students won NASA stickers for answering questions............................. 237
Figure 163: Interactive Physiics at Morgan Mill ........................................................... 238
Figure 164: Preparing to Launch at BluffDale ............................................................. 239
Figure 165: Students Learning at the Recovery Station at Dublin Middle School ....... 242
Figure 166: Students Enjoying the Art Station, Decorating Parachutes ...................... 242

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Index of Tables
Table 1: Vehicle Size and Mass ...................................................................................... 1
Table 2: Experiment Summary........................................................................................ 1
Table 3: Changes Made to Vehicle Criteria..................................................................... 2
Table 4: Changes Made to Payload Criteria.................................................................... 3
Table 5: Project Milestones Continued............................................................................ 8
Table 6: Fin Force Resistance ...................................................................................... 15
Table 7: Motor Specifications........................................................................................ 17
Table 8: Vehicle Verification Table................................................................................ 23
Table 9: Mass Summary ............................................................................................... 38
Table 10: Mass by Subsection ...................................................................................... 40
Table 11: Flight Data..................................................................................................... 42
Table 12: Test Flight One Conditions............................................................................ 43
Table 13: Test Flight Two Conditions............................................................................ 46
Table 14: Test Flight Three Conditions ......................................................................... 49
Table 15: Test Flight Four Conditions ........................................................................... 51
Table 16: Test Flight Five Conditions............................................................................ 54
Table 17: Test Flight Six Conditions.............................................................................. 56
Table 18: Test Flight Seven Conditions......................................................................... 58
Table 19: Test Flight Eight Conditions........................................................................... 60
Table 20: Test Flight Nine Conditions ........................................................................... 62
Table 21: Test Flight Ten Conditions............................................................................. 65
Table 22: Test Flight Eleven Conditions........................................................................ 67
Table 23: Test Flight Twelve Conditions ....................................................................... 69
Table 24: Test Flight Thirteen Conditions...................................................................... 72
Table 25: Test Flight Fourteen Conditions .................................................................... 75
Table 26: Test Flight Fifteen Conditions........................................................................ 77
Table 27: Kinetic Energy Summary............................................................................... 97
Table 28: Static Tests.................................................................................................. 102
Table 29: Safety and Failure Analysis 11-30-12.......................................................... 103
Table 30: Safety and Failure Analysis 12-5-12............................................................ 104
Table 40: Safety and Failure Analysis 1-5-13.............................................................. 114
Table 43: Calculated versus Simulated CG and CP Measurements ........................... 121
Table 44: Payload Preparation Steps.......................................................................... 124
Table 45: Payload Integration Steps ........................................................................... 126

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Table 46: Payload Framework Dimensions................................................................. 127
Table 47: Potential Failure Modes for Design of the Vehicle....................................... 138
Table 48: Potential Failure Modes during Payload Integration.................................... 139
Table 50: Potential Hazards to Personnel................................................................... 144
Table 51: Summary of Legal Risks ............................................................................. 146
Table 52: Effects of Materials used in Construction and Launch................................. 147
Table 53: Environmental Factors ................................................................................ 148
Table 54: Payload Functional Requirements............................................................... 163
Table 55: XBee XSC S3B Specifications .................................................................... 176
Table 56: Payload Components and Qualities ............................................................ 189
Table 57: Payload Preparation Steps.......................................................................... 194
Table 58: Payload Sensor Precision ........................................................................... 196
Table 60: Potential Failure Modes during Payload Integration .................................... 206
Table 61: Potential Failure Modes during Launch ....................................................... 209
Table 62: Potential Hazards to Personnel................................................................... 211
Table 63: Preliminary Budget Summary...................................................................... 213
Table 64: Structure/Propulsion System Budget........................................................... 218
Table 65: Recovery System Budget............................................................................ 219
Table 66: Payload Budget (Through-Hole PCB) ......................................................... 221
Table 67: Payload Budget (Surface Mount PCB) ........................................................ 224
Table 68: Accomplished Educational Outreach........................................................... 232
Table 70: Favorite Part................................................................................................ 234
Table 71: Educational Outreach Stations.................................................................... 237
Table 72: Educational Outreach Stations.................................................................... 241

I) Summary of CDR Report
I) Summary of CDR Report
Team Summary
Tarleton Aeronautical Team
Box T-0470
Stephenville, Texas 76402
Team Mentor: Pat Gordzelik.
Past and Present Credentials:
Tripoli Amarillo #92 Board of Directors Member, Technical Advisor Panel
Panhandle of Texas Rocketry Society Inc. – Founder, President, Prefect
TRA 5746 L3 NAR 70807 L3CC Committee Chair
Married to Lauretta Gordzelik, TRA 7217, L2.
Launch Vehicle Summary
Size and Mass
Length 109.25 inches
Outer Diameter 5.525 inches
Mass 37.125 pounds
Motor
Selection Cesaroni L1720-WT-P
Recovery
Drogue 24” Silicone Coated Rip stop Nylon Parachute, Apogee Deployment
Main 120” Silicone Coated Rip stop Nylon Parachute, 500 foot AGL Deployment
Avionics
Primary Featherweight Raven3 Altimeter,
Backup PerfectFlite Stratologger Altimeter, and Garmin GPS Tracking
Rail Size
Rail 1010 Aluminum
Milestone Review Flysheet – see Appendix B
Table 1: Vehicle Size and Mass
Payload Summary
Title Experiment
Science Mission
Directorate (SMD)
Payload
Sponsored by NASA; Gather Atmospheric and GPS Data, Autonomously
Orientate Photographic Camera, Capture Video for Public Outreach
Table 2: Experiment Summary

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II) Changes Made since PDR
Changes Made to Vehicle Criteria
Structure Rationale
Drogue avionics relocated to rear coupler
from booster section
Ease of construction and accessibility
Ballast system relocated from nose cone to
upper body airframe
Ease of construction
Coupler port hole rings added
Eliminates need of lining up port holes
through body and coupler
Centering ring in fin tab relocated to front of
fin tab
Ease of construction
Bulkhead at upper end of motor tube
replaced with centering ring
Allows access to anchor point on motor
housing for shock cord
Centering ring added to lower end of motor
tube
Used to secure motor retaining ring
Nose cone length changed from 7.5 inch to
8.5 inch
Manufactured at this length
Payload bulkheads reduced to 1 inch
thickness from 2 inch
Reduces weight without compromising
integrity
Payload bulkheads epoxied to couplers
Creates seals between avionics bays and
payload compartment
Coupler bulkheads added to avionics bay
lids
Reinforce bay lids in event of failed main
chute deployment
Recovery Rationale
U-bolts changed to welded eyebolts
Reduce weight without compromising
integrity
Removed deployment bag for drogue chute Unnecessary
Changed to XL ejection canisters
Reduce friction by allowing chute more
room lengthways
Changed to 3F black powder from 4F Availability
Switched to 3 portholes in each bay(sizing
in SFR)
Following recommendations of
manufacturers
LEDs added to allow visible confirmation of
altimeter activations
Eliminates need of audio confirmation of
altimeters
GPS relocated from nose cone to drogue
shock cord
Easy to secure
Changed from Big Red Bee GPS to Garmin
Astro DC40/320 system
Easy to implement
Increased deployment bag size of main
chute to XXL from XL
Increase ease of deployment
Avionics Bays changed to standard sled
containing design
Modular and easy to access
Table 3: Changes Made to Vehicle Criteria

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Recovery Rationale
Tubular Kevlar shock cords reduced to .25
inch from .5 inch
Weight reduction and increase space in
upper body section
Swivel removed from drogue chute Unnecessary
Backup shock cord (.25 inch, 4.5 feet)
epoxied along motor tube
Safety/Redundancy
Shock cord of main chute length changed
to 40 feet from 20 feet
limit multiple section collision
Shock cord of drogue chute length
changed to 20 feet from 25 feet
limit multiple section collision
Increased size of all ejection charges Necessary for proper separation
Secondary charges have .4 grams more
black powder than primary
Simple Logic
Switched to cross-form rip stop nylon
parachutes
Availability and durability
Separation now occurs between the upper
body airframe and payload section instead
of at the nose cone
Allows for easier transportation and
preparation
Changes Made to Payload Criteria
Payload Rationale
Rail changed to .5 inch x .5 inch x .0625
inch aluminum angle from .5 inch x .125
inch flat aluminum
Add rigidity and reduce weight
Payload centered in payload section
Allows avionics bays to have uniform
dimensions
Rear coupler removable to access payload Easier to access
Port holes changed to 5 evenly spaced .25
inch holes
Provide adequate ventilation
Reduced 9V battery count from 8 to 4 Unnecessary
MS5611 pressure sensor removed Availability
HH10D humidity sensor removed Simplify circuit
Video camera changed to Keyfob from
VCC-003-MUVI-BLK
Availability and cost
Arduino Mini added to ARTCOS Dedicated for video processing
Moved HIH4030 to ARTCOS
The reference voltage required by the
SU100 and SP110
Changed to buck converters from linear
regulators
Power efficiency
Mounted ARTCOS to fiberglass brackets More secure installation
BMP180 placed in between circuit boards Shields the sensor from light
Magnetic switch connected to relay to
activate entire payload
Simplify and speed up launch preparation
Table 4: Changes Made to Payload Criteria

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The team made no significant changes to the project plan.

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III) Vehicle Criteria
Design and Verification of Launch Vehicle
Launch Vehicle Mission Statement
The mission is to design, build, and launch a reusable vehicle capable of delivering a
payload to 5,280 feet above ground level (AGL). The vehicle will carry a barometric
altimeter for official scoring and the Science Mission Directorate (SMD) payload. The
design of the vehicle ensures a subsonic flight and must be recoverable and reusable
on the day of the official launch. The launch vehicle meets the customer prescribed
requirements set forth in the Statement of Work (SOW) of the NASA 2012-2013 Student
Launch Projects (SLP) handbook.
Launch Vehicle Requirements
The vehicle adheres to the following primary requirements. The complete list of
requirements is in the Vehicle Verification Table (Table 8).
 Vehicle shall carry a scientific or engineering payload. (Requirement 1.1)
 Vehicle shall reach an apogee altitude of one mile above ground level.
(Requirement 1.1)
 Vehicle shall carry one official scoring altimeter. (Requirement 1.2)
 Vehicle must remain subsonic from launch until landing. (Requirement 1.3)
 Vehicle must be recoverable within a 2500 foot radius from the launch pad and
reusable on the day of the official launch. (Requirement 2.3)
 Vehicle must use a commercially available APCP motor with no more than 5,120
Newton-seconds of impulse. (Requirement 1.11, 1.12)
Mission Success Criteria
The project defines the mission as a vehicle flight with a payload onboard where both
the vehicle and SMD payload are recovered and able to be reused on the day of the
official launch. Moreover, the vehicle will not exceed 5,600 feet of altitude, and the
official scoring altimeter will be intact, audible, and report altitude. The recovery system
stages a deployment of the drogue parachute at apogee and deploys the main
parachute at 700 feet. After apogee and descent, the entire vehicle lands within 2,500
feet of the launch pad.
If the above conditions are met, the mission will be considered partially successful in
that requirements have been met by the vehicle design. However, because the actual
altitude of the vehicle at apogee is scored based on comparison to one mile above
ground level, a successful mission would be warranted only if the aforementioned

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conditions are met and an apogee of exactly 5,280 feet is achieved, plus or minus 0.1%
plus 1 foot due to precision of the scoring altimeter.
Major Milestone Schedule
Significant milestones of the project from initiation to final launch day and
announcement of contest winners are detailed in Table 5. Each date has a description
as well as the completion status of each event up to the time that the CDR is submitted
(Jan 14). Additionally, the type of event is specified as either being provided by the
NASA USLI SOW, a test date, a deadline for verification, or a deadline for
manufacturing/assembly of the vehicle.

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Date Milestone Description Status Type
8/31/12 Proposal Due Met USLI
9/27/12 Schools Notified Met USLI
10/4/12 Team Teleconference Met USLI
10/11/12 PDR Q&A Met USLI
10/22/12 Web Presence est. Met USLI
10/23/12 SMD Award 1 ($780) Met USLI
10/28/12 Subscale Dual Deployment Test Met Test Launch
10/29/12 PDR due Met USLI
11/7-16/12 PDR Presentations Met USLI
11/17/12 Dual Deployment Test Not Met Test Launch
11/30/12 Subscale Launch Met Test Launch
11/30/12 Lab Prototyping Met Verification
12/1/12 Low Altitude Flight Met Test Launch
12/3/12 Post Launch Failure Analysis Met Verification
12/3/12 Full Scale Prototype Assembly 1 Not Met Manufacturing
12/3/12 CDR Q&A Not Met USLI
12/18/12 Range Radio Testing Met Verification
12/20/12 E-match Testing Met Verification
12/22/12 Subscale Low Altitude Flight Met Test Launch
12/31/12 PCB Testing Not Met Verification
12/31/12 Programming Met Verification
1/5/13 Static Black Powder Testing Met Verification
1/5/13 Static Ejection Test Met Verification
1/5/13 Low Altitude, Full Scale Launch (w/o
SMD)
Not Met Test Launch
1/5/13 Low Altitude, Full Scale Launch (w/
SMD)
Met Test Launch
1/6/13 Alternative Launch Day-Used Met Test Launch
1/6/13 Post Launch Failure Analysis Met Verification
1/7/13 Low Altitude, Full Scale Launch (w/o
SMD)
Met Test Launch
1/7/13 Static Motor Test Met Verification
1/12/13 Alternative Launch Day TBD Test Launch
1/14/13 Post Launch Failure Analysis TBD Verification
1/14/13 CDR due Met USLI
1/14/13 Spring Semester Begins …. ….
1/19/13 Low Altitude, Full Scale Launch (w/ TBD Test Launch

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Date Milestone Description Status Type
SMD)
1/22/13 Full Scale Prototype Assembly 2 TBD Manufacturing
1/26/13 High Altitude, Full Scale Launch (w/o
SMD)
TBD Test Launch
1/27/13 High Altitude, Full Scale Launch
(w/SMD)
TBD Test Launch
2/1/13 CDR Presentations TBD USLI
2/2/13 SMD Award 2 ($1400) TBD USLI
SMD)
TBD Test Launch
2/11/13 FRR Q&A TBD USLI
SMD)
TBD Test Launch
(w/SMD)
TBD Test Launch
(w/SMD)
TBD Test Launch
3/1/13 Final Vehicle Assembly TBD Manufacturing
3/2/13 Final Demonstration Flight TBD Verification
3/9/13 Final Demonstration Flight (alt) TBD Verification
3/16/13 Final Demonstration Flight (alt) TBD Verification
3/18/13 FRR due TBD USLI
3/25-4/3/13 FRR Presentations TBD USLI
4/17/13 LRR Begin TBD USLI
4/18-19/13 Welcome Day TBD USLI
4/20/13 Launch Day TBD USLI
4/21/13 Launch Rain Day TBD USLI
5/6/13 PLAR due TBD USLI
5/17/13 Winners Announced TBD USLI
Table 5: Project Milestones Continued

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Review the design at a system level
Final Drawings and Specifications
The overall launch vehicle, as shown in Figure 1, is 109.25 inches long. The fin span is
15.525 inches. This includes the 5.525 inch width of the airframe. Each section of the
launch vehicle will be further specified below.
Figure 1: Launch Vehicle Specifications

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

11
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

12
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

13
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

14
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

15
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

16
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

17
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

18
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

19
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

20
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

21
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

22
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

23
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

24
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.

25
Root Chord: 12 inches
Tip Chord: 0 inches
Height: 5 inches
Sweep Length: 9.8 inches
Sweep Angle: 63 degrees
Proper use of materials in fins, bulkheads, and structural elements
Airframe/Motor Tube/Nosecone
The upper airframe, booster section, motor tube, and nose cone will be made of
fiberglass. The durability of fiberglass improves the chances of the rocket being
reusable (Requirement 1.4). Fiberglass is also readily available.
This material was chosen for structural components due to strength and ease of
manufacturing. The consistency in component materials allows for use of a single type
of epoxy manufactured especially for fiberglass. The epoxy used for attaching structural
components is Proline 4500 epoxy. This will provide a strong and uniform bond
throughout the vehicle’s structural components.
Bulkheads/centering rings
The bulkheads and centering rings are made of fiberglass due to the material’s
superior strength to mass ratio as well as adhesive qualities between components.
A .2 inch thickness was chosen to provide a larger surface area to epoxy to the
airframe and provide adequate strength of mounting hardware during parachute
Figure 12: Fin Dimensions

26
deployment.
The payload bulkheads are constructed out of PVC. This material was chosen
because of its strength at the desired one inch thickness in addition to its
availability.
Fins
The fin material is fiberglass. This choice is justified for similar reasons to that of the
bulkheads and centering rings. The fins can be properly secured with the adhesive.
The strength of fiberglass also allows for a greater durability upon possible high
impact with the ground.
Couplers
The couplers are handmade to fit the acrylic section to both the upper body airframe
and the booster section. The acrylic section has a different thickness than that of
the fiberglass airframe. This is due to a difference in inside diameter, 5.25 inches
versus 5.375 inches respectively. Thus, two different couplers are overlapped to
make one single coupler. The couplers are made of fiberglass for ease of
integration and strength.
Proper assembly procedures, attachment and alignment of elements, solid
connection points, and load paths
The assembly is carried out in a multistep process which allows for each piece to be
epoxied and cured. This allows sufficient time to focus on proper construction for each
individual component before moving on to the next step. The follow figures are
assembly procedures for the major parts of the launch vehicle.

27
Figure 13: Booster Assembly Steps 1-4

28

29

30
Figure 16: Coupler Assembly Procedure

31
Figure 17: Avionics Assembly Steps 1-3

32
Figure 18: Avionics Assembly Steps 4-6

33
Figure 19: Payload Assembly

34
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

35
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

36
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

37
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.

38
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

39
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
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
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

40
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

41
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.

42
Date Apogee Predicted
Apogee
Drift Video link
November 30,
2012
N/A 401 ft AGL N/A N/A
December 5,
2012
460ft AGL 450 ft AGL 168ft Video
December 7,
2012
December 8,
2012
519ft AGL 600 ft AGL N/A N/A
December 8,
2012
509ft AGL 690 ft AGL N/A N/A
December 14,
2012
December 15,
2012
666ft AGL 615 ft AGL 50ft N/A
December 15,
2012
December 19,
2012
December 19,
2012
1,061ft AGL 1848 ft AGL 342ft Video
December 21,
2012
629ft AGL 780 ft AGL N/A Video
December 21,
2012
4,992ft AGL 5227 ft AGL 2,118ft Video
January 5,
2013
January 6,
2013
January 7,
2013
2,402ft AGL 2908 ft AGL N/A N/A
Table 11: Flight Data

43
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

44
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

45
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

46
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
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
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
70° F 5 mph 31.29 in Hg 1309 ft
Figure 26: Test Flight Two Vehicle
Table 13: Test Flight Two Conditions

47
Simulated Flight
Actual Flight
Figure 27: Simulated Flight Two Data

48
Dual deployment was not achieved due to the main parachute not ejecting.
Post flight analysis revealed the e-match leads were the failure point. The leads were
wired to the ground and main, rather than the being wired to the power and main. As a
result the recovery procedures were modified to include inspection of the ejection
canister connections to the altimeter output channel.
Sub-scale Test Flight Three
Figure 29: Test Flight Three Vehicle

49
were 49 degrees Fahrenheit, 28.43 inches of Mercury, six mile per hour winds, and
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.
(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
49° F 6 mph 28.43 in Hg 1309 ft
Simulated Flight
Table 14: Test Flight Three Conditions

50
Actual Flight
No flight data was recovered or reliable for this test launch.
The dual deployment was only partially successful. This was due to both the drogue
parachute and the main parachute deploying at apogee.
The conclusion of post flight analysis was the main parachute prematurely deployed
because the upper body sections friction fit was not strong enough to withstand the
force of the drogue parachute ejection charge. This resulted in a sheer pins being
implemented to secure the upper body section.
Sub-scale Test Flights 4 and 5
Figure 30: Simulated Test Flight Three

51
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
Flight Four
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
54° F 5 mph 29.96 in Hg 1309 ft
Simulated Flight
Figure 31: Test Flight Four Vehicle
Table 15: Test Flight Four Conditions

52
Actual Flight
Figure 32: Simulated Test Flight Four Data

53
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
Figure 34: Test Flight Five Vehicle

54
(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
54° F 5 mph 29.96 in Hg 1309 ft
Simulated Flight
Actual Flight
Table 16: Test Flight Five Conditions
Figure 35: Raven3 Test Flight Five Data

55
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
were 58 degrees Fahrenheit, 29.93 inches of Mercury, sixteen mile per hour winds, and
The subscale vehicle, shown above in Figure 36, used a Cesaroni G79 Smoky Sam
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

56
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
58° F 16 mph 29.93 in Hg 1309 ft
Simulated Flight
Table 17: Test Flight Six Conditions
Figure 37: Simulated Flight Six Data

57
Actual Flight
No flight data was recorded or reliable for this test launch.
Dual deployment was unsuccessful. All recovery systems failed, resulting in a ballistic
descent and a lawn dart. The nose cone, upper airframe, and coupler were all
destroyed, though the rest of the rocket was deemed reusable.
Post flight analysis revealed that both altimeters failed to enter flight mode, and no flight
data was recovered. It is suspected the failure point was human error in preparing the
altimeters for launch.
Sub-scale Test Flights Seven and Eight
On December 15, 2012 in Asa, two launches under the supervision of our team mentor,
Pat Gordzelik, at a Hotroc launch event just outside of Waco. Launch conditions for the
first flight were 67 degrees Fahrenheit, 29.93 inches of Mercury, 6 mile per hour winds,
and elevation of 427 feet MSL. Launch conditions for the second flight were 74 degrees
Fahrenheit, 29.74 inches of Mercury, 10 mile per hour winds, and elevation of 427 feet
MSL. A ten foot 1010 launch rail was used for both flights.
Flight 7
motor, with a nine second delay charge for redundant drogue ejection. The recovery
Figure 38: Test Flight Seven Vehicle

58
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.
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
67° F 6 mph 29.93 in Hg 427 ft
Simulated Flight
Table 18: Test Flight Seven Conditions
Figure 39: Test Flight Seven Simulated Data

59
Actual Flight
Flight Analysis and Impact on Design (Flight Seven)
Dual deployment was successful, but the shock cord between the upper airframe to the
booster section failed upon drogue ejection.
Upon apogee, the 5.0 gram black powder charge ignited, with a very loud report, and
broke the shock cord tethering the booster section to the upper airframe. This resulted
in two sections descending; a booster section with a drogue parachute and an upper
airframe descending without decent control. At the pre-programed height of 128 feet
AGL, the main parachute deployed and both sections were recovered.
Post flight analysis revealed damage to the shock cord from heat. This damage and the
5.0 gram black powder charge were concluded to be the point of failure. As a result we
implemented the use of tubular Kevlar shock chord.
Figure 40: Raven3 Test Flight Seven Data

60
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
December 15, 2012 Asa (31.438403, -97.027519) Cesaroni G129 Smoky Sam
Altimeter
Drogue
Charge Size
Main Charge
Size Main Deploy Altitude
74° F 10 mph 29.74 in Hg 427 ft
Simulated Flight
Figure 41: Test Flight Eight Vehicle
Table 19: Test Flight Eight Conditions

61
Actual Flight
Flight Analysis and Impact on Design (Flight Eight)
The flight was a partial success as both parachutes deployed, but both parachutes
deployed at apogee. The rocket was recovered.
Post flight analysis revealed both ejection charges being fired at once and it was
concluded the failure was due to human error.
No flight data was retrieved from the Stratologger SL100 because a DT2U cable was
needed to access the stored flight data, which was not available to the team.
Sub-scale Test Flights Nine and Ten
On December 19, 2012 at Hunewell Ranch, two launches occurred. Launch conditions
were 77 degrees Fahrenheit, 28.87 inches of Mercury, 19 mile per hour winds, and
elevation of 1,309 feet MSL. A ten foot 1010 launch rail was used for this flight.
Figure 42: Simulated Flight Eight Data

62
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
December 19, 2012 Hunewell (32.216114, -98.096019) Cesaroni H125
Altimeter
Drogue
Charge Size
Main Charge
Size
Main Deploy
Altitude
77° F 19 mph 28.87inHg 1309 ft
Simulated Flight
Figure 43: Test Flight Nine Vehicle
Table 20: Test Flight Nine Conditions

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