NASA FRR Technical Report

[Type text] I) Summary of FFR Report [Type text]
Flight Readiness
Review
2012-2013 NASA USLI
“We can lick gravity, but sometimes the paperwork is overwhelming.”
- Werner Von Braun

Tarleton State University
Flight Readiness Review
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Note to reader:
To facilitate the reading of the Flight Readiness Review, we have mirrored the Student
Launch Project Statement of Work. In the body of the FRR, 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 250 man hours of work. The PCBs provide major enhancement of the
signal integrity of the sensor data in addition to their space and power efficient qualities.
We have enjoyed finalizing our design for flight readiness and submit this document for
your review.

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Table of Contents
Flight Readiness Review...................................................................................................
2012-2013 NASA USLI ..................................................................................................
I) Summary of FFR Report.............................................................................................. 1
Team Summary ........................................................................................................... 1
Launch Vehicle Summary ............................................................................................ 1
Payload Summary........................................................................................................ 1
II) Changes Made since CDR.......................................................................................... 2
III) Vehicle Criteria........................................................................................................... 4
Design and Construction of the Vehicle ....................................................................... 4
Structural Elements .................................................................................................. 4
Drawings and schematics....................................................................................... 13
Flight Reliability Confidence and Mission Success Criteria .................................... 16
Testing.................................................................................................................... 17
Workmanship.......................................................................................................... 18
Vehicle Safety and Failure Analysis........................................................................ 19
Full-scale Launch Results....................................................................................... 19
Mass Report ........................................................................................................... 27
Recovery Subsystem ................................................................................................. 28
Design Defense ...................................................................................................... 28
Parachute Justification............................................................................................ 43
Safety and Failure Analysis .................................................................................... 43
Mission Performance Predictions............................................................................... 50
Mission Performance Criterion................................................................................ 50
Kinetic Energy Management................................................................................... 59
Altitude and Drift Predictions .................................................................................. 60
Verification of System Level Functional Requirements........................................... 60
Safety and Environment............................................................................................. 65
Safety and Mission Assurance Analysis ................................................................. 65
Payload Integration................................................................................................. 78
Integration of the Payload with the Launch Vehicle ................................................ 78
Compatibility of Elements ....................................................................................... 81
Payload Interface Dimensions ................................................................................ 82

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Payload-Housing Integrity....................................................................................... 82
Integration Demonstrated ....................................................................................... 83
IV) Payload Criteria ....................................................................................................... 86
Experiment Concept .................................................................................................. 86
Science Value............................................................................................................ 86
Payload Objectives................................................................................................. 86
Payload Success Criteria........................................................................................ 86
Test and Measurement, Variables, and Controls.................................................... 89
Payload Design.......................................................................................................... 90
ADGS ..................................................................................................................... 93
ARTCOS................................................................................................................. 99
Precision of Instrumentation and Repeatability of Measurements ........................ 104
Flight Performance Predictions............................................................................. 107
Test and Verification Program .............................................................................. 116
Payload Requirement Verification............................................................................ 117
Verification Statements......................................................................................... 118
Safety and Environment (Payload) .......................................................................... 119
Personnel Hazards ............................................................................................... 122
Environmental Concerns ...................................................................................... 123
V) Launch Operations Procedures .............................................................................. 124
Checklists................................................................................................................. 124
Recovery Preparation Checklist (to be completed before arriving to launch site): 124
Motor Preparation Checklist: ................................................................................ 125
Igniter Installation Checklist: ................................................................................. 126
Launchpad Setup Checklist: ................................................................................. 126
Launch Procedures Checklist: .............................................................................. 127
Troubleshooting Checklist: ................................................................................... 128
Post-flight Inspection Checklist:............................................................................ 129
Payload Preparation (to be completed before arriving at launch field):................. 129
Pre-launch Payload Preparation and Ground Station Setup:................................ 130
Safety Materials Checklist .................................................................................... 130
Safety Checklist.................................................................................................... 130
The following is an itemized components checklist for each subsystem.................. 131
Structure Components:......................................................................................... 131

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Recovery Components: ........................................................................................ 131
Payload and Ground Station Components: .......................................................... 132
Safety and Quality Assurance.................................................................................. 132
Risk Assurance..................................................................................................... 132
Risk Assessment for Launch Operations.............................................................. 132
Environmental Concerns ...................................................................................... 134
Identification of person responsible ...................................................................... 134
VI) Project Plan ........................................................................................................... 135
Budget Summary ..................................................................................................... 135
Funding Plan......................................................................................................... 142
Timeline ................................................................................................................... 142
Outreach Timeline ................................................................................................ 146
Education plan ......................................................................................................... 146
Outreach Plan....................................................................................................... 146
Accomplished Educational Outreach.................................................................... 149
VII) Conclusion............................................................................................................ 161

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Table of Figures
Figure 1: Launch Vehicle Dimensions............................................................................. 4
Figure 2: Nose Cone ....................................................................................................... 5
Figure 3: Ballast and GPS mount.................................................................................... 6
Figure 4: Coupler Frame ................................................................................................. 6
Figure 5: Coupler ............................................................................................................ 7
Figure 6: Avionics Bay..................................................................................................... 7
Figure 7: Avionics Bay Lid............................................................................................... 8
Figure 8: Payload Bulkhead and Avionics Rail System................................................... 8
Figure 9: Avionics Bay Outside of Coupler...................................................................... 9
Figure 10: Avionics Sled Layout...................................................................................... 9
Figure 11: Acrylic Payload Housing Structure............................................................... 10
Figure 12: Coupler Screw in Attachment....................................................................... 10
Figure 13: Weight Simulator.......................................................................................... 11
Figure 14: Booster Section............................................................................................ 11
Figure 15: Fin Dimensions ............................................................................................ 12
Figure 16: Shock Cord Epoxied to Motor Mount Tube .................................................. 12
Figure 17: Dual Attachment Scheme............................................................................. 13
Figure 18: Booster Section and Motor Mount Tube....................................................... 14
Figure 19: Upper Body Airframe with Nose Cone ......................................................... 14
Figure 20: Payload Housing Structure, Aluminum Framework, and Couplers............... 15
Figure 21: Vehicle Assembly......................................................................................... 15
Figure 22: Rendereing of Final Vehicle ......................................................................... 16
Figure 23: Simulation Data............................................................................................ 20
Figure 24: OpenRocket Rendering of Vehicle ............................................................... 20
Figure 25: Simulated Flight Data................................................................................... 20
Figure 26: Drogue Stratologger Flight Data................................................................... 22
Figure 27: Main Stratologger Flight Data....................................................................... 22
Figure 28: Simulation Data............................................................................................ 23
Figure 29: OpenRocket Rendering of Vehicle ............................................................... 23
Figure 30: Simulated Flight Data................................................................................... 24
Figure 31: Drogue Stratologger Data ............................................................................ 25
Figure 32: Drogue Raven3 Flight Data.......................................................................... 25
Figure 33: Main Stratologger Flight Data....................................................................... 26
Figure 34: Main Raven3 Flight Data.............................................................................. 26
Figure 35: Packed Deployment Bag.............................................................................. 33
Figure 36: Deployment Bag Operation.......................................................................... 33
Figure 37: Altimeter Wiring Configuration...................................................................... 37
Figure 38: Stratologger Software Flow Diagram ........................................................... 38
Figure 39: GPS Software Flow Diagram ....................................................................... 39
Figure 40: Avionics Bay Structural Components ........................................................... 40
Figure 41: Attachment Bulkhead Inside Avionics Coupler............................................. 40
Figure 42: Couplers in Profile........................................................................................ 41
Figure 43: Structural Assembly of the Avionics Bay...................................................... 41
Figure 44: GPS Device and Wireless Remote .............................................................. 42
Figure 45: Final Vehicle Simulation............................................................................... 52

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Figure 46: Input Parameters for Final Simulation .......................................................... 52
Figure 47: L1720-WT Actual Thrust Curve.................................................................... 53
Figure 48 - PDF of Apogee for Competition Flight ........................................................ 55
Figure 49 - PDF of Lateral Distance for Competition Flight........................................... 55
Figure 50: Center of Gravity and Center of Pressure .................................................... 57
Figure 51: Rear Payload Bulkhead ............................................................................... 79
Figure 52: Bracket on Forward Payload Bulkhead ........................................................ 79
Figure 53: Securing Payload Framework to Rear Bulkhead.......................................... 80
Figure 54: Framework Fitting into the Bracket............................................................... 80
Figure 55: Fully Assembled SMD Payload Section ....................................................... 81
Figure 56: Payload Framework Secured to Rear Payload Bulkhead ............................ 83
Figure 57: Payload Framework Secured....................................................................... 84
Figure 58: Payload Integration ...................................................................................... 84
Figure 59: Payload Integration ...................................................................................... 84
Figure 60: Integrated Payload ....................................................................................... 85
Figure 61: Payload Design ............................................................................................ 91
Figure 62: Constructed Payload.................................................................................... 92
Figure 63: ADGS PCB Schematic................................................................................. 94
Figure 64: ADGS PCB................................................................................................... 95
Figure 65: Power PCB Schematic................................................................................. 96
Figure 66: Power Board ................................................................................................ 97
Figure 67: Solar Sensor Array....................................................................................... 98
Figure 68: Real Time Sensor Readings ........................................................................ 98
Figure 69: ARTCOS Control PCB Schematic................................................................ 99
Figure 70: ARTCOS Powerboard PCB Schematic...................................................... 100
Figure 71: ARTCOS Explosion Assembly ................................................................... 101
Figure 72: ARTCOS Assembly.................................................................................... 102
Figure 73: Component Connections............................................................................ 103
Figure 74: SU-100 Spectral Response........................................................................ 105
Figure 75: Acrylic Sheet UV Transmission Curves...................................................... 105
Figure 76: Spectral Response for Apogee Pyranometer............................................. 106
Figure 77: Pressure Data ............................................................................................ 107
Figure 78: CDR Pressure Data ................................................................................... 108
Figure 79: Launch Pad Temperature........................................................................... 108
Figure 80: Temperature Data...................................................................................... 109
Figure 81: CDR Launch Pad Humidity ........................................................................ 110
Figure 82: Lux Sensor Data ........................................................................................ 110
Figure 83: Solar Irradiance.......................................................................................... 111
Figure 84: CDR Solar Irradiance Data......................................................................... 112
Figure 85: GPS Flight Data ......................................................................................... 113
Figure 86: CDR GPS Flight Data ................................................................................ 114
Figure 87: ARTCOS Flight and Landing Photos.......................................................... 115
Figure 88: President's Circle Awarding the Aeronautical Team................................... 142
Figure 89: Gantt Testing Timeline ............................................................................... 145
Figure 90: Outreach Timeline...................................................................................... 146
Figure 91: Students Launching a Water Rocket.......................................................... 147

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Figure 92: Team Members Educate and Entertain Acton Students ............................ 149
Figure 93: Educational Outreach Survey Results........................................................ 150
Figure 94: Presentation Learning Outcomes............................................................... 151
Figure 95: Students' Favorite Portion of Presentations ............................................... 152
Figure 96: Students Received Stickers for Correct Responses................................... 155
Figure 97: Students Engaged in Learning................................................................... 156
Figure 98: Team Members Addressed Large Audiences of Students ......................... 157
Figure 99: Students Learning at the Recovery Station at Dublin Middle School ......... 159
Figure 100: Students Enjoying the Art Station, Decorating Parachutes ...................... 159

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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 Recovery.............................................................................. 3
Table 5: Launch Conditions........................................................................................... 19
Table 6: Simulated Flight Data ...................................................................................... 21
Table 7: Recovery Electronics Data .............................................................................. 21
Table 8: Launch Conditions........................................................................................... 23
Table 9: Simulated Flight Data ...................................................................................... 24
Table 10: Recovery Electronics Data ............................................................................ 24
Table 11: Mass Summary ............................................................................................. 27
Table 12: Mass by Subsection ...................................................................................... 28
Table 13: Test Flight 1/26/13......................................................................................... 44
Table 16: Test Flight 3/7/13........................................................................................... 47
Table 20 - Confidence Intervals for Apogee and Drift.................................................... 54
Table 21: Calculated versus Simulated CG and CP Measurements ............................. 59
Table 22: Kinetic Energy by Section During Flight ........................................................ 60
Table 23: Altitude Predictions........................................................................................ 60
Table 24: Drift Predictions............................................................................................. 60
Table 25: Vehicle Verification Table.............................................................................. 64
Table 26: Potential Failure Modes for the Design of the Vehicle................................... 67
Table 27: Potential Structure Failure Modes and RPNs................................................ 68
Table 28: Potential Propulsion Failure Modes aand RPNs............................................ 69
Table 29: Potential Recovery Failure Modes and RPNs ............................................... 73
Table 30: Top 10 Failure Modes by RPN ...................................................................... 74
Table 31: Potential Hazards to Personnel..................................................................... 75
Table 32: Legal Risks.................................................................................................... 77
Table 33: Effects of Materials used in Construction and Launch................................... 78
Table 34: Environmental Factors Hendering Recovery................................................. 78
Table 35: Procedures for Installing Payload.................................................................. 81
Table 36: Payload Interface Dimensions....................................................................... 82
Table 37: Payload Objectives........................................................................................ 87
Table 38: FR-4 Characteristics...................................................................................... 93
Table 39: Payload Electrical Components................................................................... 104
Table 40: Payload Sensor Precision ........................................................................... 106
Table 41: Payload Functional Requirements............................................................... 118
Table 42: Potential Failure Modes for Payload Section During Integration ................. 119
Table 43: Potential Failure Modes for Payload During Launch Operations................. 121
Table 44: Top Five Potential Failure Modes by RPN .................................................. 121

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Table 45: Potential Hazards to Personnel................................................................... 123
Table 46: Risk Assessment for Launch Operations .................................................... 134
Table 47: Known Project Costs ................................................................................... 135
Table 48: Team Budget............................................................................................... 135
Table 49: Structure/Propulsion System Budget........................................................... 136
Table 50: Recovery System Budget............................................................................ 137
Table 51: Payload Budget (Through-Hole PCB) ......................................................... 138
Table 52: Payload Budget (PCB/Non PCB) ................................................................ 141
Table 53: Educational Outreach Survey Results......................................................... 150
Table 54: Presentation Learning Outcomes................................................................ 151
Table 55: Students' Favorite Portion of Presentations ................................................ 152
Table 56: Educational Outreach Stations.................................................................... 155
Table 57: Educational Outreach Stations.................................................................... 158

I) Summary of CDR Report
I) Summary of FFR 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 97.5 inches
Outer Diameter 5.525 inches
Mass 35.9 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, 700 foot AGL Deployment
Avionics
Primary PerfectFlite Stratologger Altimeter,
Backup PerfectFlite Stratologger Altimeter, and Garmin GPS Tracking
Rail Size
Rail 1010 and 1515
Milestone Review Flysheet – see link on website
Table 1: Vehicle Size and Mass
Payload Summary
Title Experiment
Science Mission
Directorate (SMD)
Payload
Sponsored by NASA; Gather Atmospheric and GPS Data, Autonomously Orient
Photographic Camera, Capture Video for Public Outreach
Table 2: Experiment Summary

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II) Changes Made since PDR
II) Changes Made since CDR
Changes Made to Vehicle Criteria
Structure Rationale
Avionics Port Holes at 5/32 inch
Stratologger Manual
Specification
Booster Length is 32 inches (approved*) Weight Reduction
GPS location to nosecone
Poor attachment scheme on
shock chord
Payload Housing is 27 inches (approved*) Weight reduction
Avionics Sled Guides Changed Aluminum material strength
Payload Framework is 20.5 inches Payload integration
Payload Framework to Bulkhead Mount (Bracket) Payload integration
Payload Framework is 0.5 by .75 inch aluminum angle Added strength
*These changes were submitted via email directly
after CDR on 3/4/2013
Recovery Rationale
4 Strattologgers are used for all Altimeters
Power problems with Raven
3s
1 battery is used for each avionics bay Power indicator
1 LED per altimeter Power Indicator
1 second delay set on backup apogee altimeter Redundancy
700 foot main parachute deployment setting; 650 foot
backup setting
Redundancy
J-Tek 10 E-matches with plastic test tubes Reliability
Ejection charges use clip connectors Safety
Kevlar TAC-9B Deployment Bag Ease of Packing
24 inch Pilot drogue for deployment bag Required for Deployment bag
Deployment Bag Attachment Scheme Required for Deployment bag
Recovery Rationale
1/4 inch to 3/8 inch tubular Kevlar shock
chord (approved**)
Greater FOS; Packing constraints
Main Parachute packing scheme Prevent entanglement
**RSO approved on 3/13/2013
Changes Made to Payload Criteria
Payload Rationale
LCD Screen removed Problems in testing, power consumption
4th
battery added Power Activation Circuit
Table 3: Changes Made to Vehicle Criteria

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II) Changes Made since PDR
4 UV sensors 90 degree intervals; better data acquisition
4 Solar Irradiance 90 degree intervals; better data acquisition
Minor PCB Changes Layout Optimization
The team made no significant changes to the project plan.
Table 4: Changes Made to Recovery

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III) Vehicle Criteria
Design and Construction of the Vehicle
The entire launch vehicle, as shown in Figure 1, is 97.5 inches long. The fin span
measures 13.525 inches. This includes the 5.525 inch width of the airframe. Each
section of the launch vehicle will be described in more detail in the following
subsections.
Structural Elements
Upper Body Airframe
This section contains the nose cone and upper body airframe.
Airframe
The upper body airframe is 28.0 inches long. It has a 5.525 inch outside diameter and
5.375 inch inside diameter. These dimensions were chosen in order to fully house the
payload and main recovery system. Fiberglass was chosen as the material due to its
Figure 1: Launch Vehicle Dimensions

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strength and durability. This allows it to withstand the forces experienced during
acceleration and landing.
Nose Cone
The elliptical nose cone (Figure 2) is attached to the upper body airframe with
four screws and drywall anchors. It is 8.5 inches long and has a 4.6 inch
shoulder. Fiberglass was chosen for the nose cone for its strength and durability.
Ballast system/Attachment/GPS
The ballast system is built into the upper body airframe and contained in the shared
space of the upper body airframe and the nose cone. A bulkhead is epoxied into the
airframe where the shoulder of the nose cone sits. A hole is drilled through the center of
the bulkhead and a welded 5.5 inch eyebolt is secured through it. The eyebolt serves as
the mounting point for the main parachute on the opposite side of the bulkhead. It is
important to note that all eyebolts used for attachment hardware are welded to increase
the maximum load rating. The threaded portion of the eyebolt serves as a location to
secure both the ballast weights and the Garmin Astro DC40 GPS.
An assortment of washers serves as the ballast weight as seen below in Figure 3. To
secure different sized washers, they are staggered in different directions and use a
spacer to distribute the force.
Figure 2: Nose Cone

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Couplers
The couplers are handmade to fit the SMD payload housing to both the upper body
airframe and the booster section. The coupler frame can be seen in Figure 4. Each
section will remain attached to the payload housing throughout the entire flight.
Commercially available couplers are not used because of the varying inside diameters
of the fiberglass and the acrylic sections. Thus, two different couplers are overlapped to
make one single coupler measuring 11.25 inches in length. One side of the coupler is
5.372 inches in diameter and the other side is 5.248 inches in diameter, which can be
seen in Figure 4.
Each coupler has a collar of fiberglass epoxied into place where the two sections meet.
In Figure 5, the fiberglass collar goes where the purple section meets the red section.
Four 0.15625 inch portholes are located on the collar to allow for easier vehicle
assembly by preventing any misalignment. An LED is also mounted to the coupler on
the collar. The LED provides visual verification when the avionics switch is activated.
Figure 3: Ballast and GPS mount
Figure 4: Coupler Frame

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Shear pins are used in the coupler between the upper body airframe and SMD payload
housing to prevent premature separation at apogee when the drogue parachute‟s
charges fire.
Avionics Bays
The avionics bays are housed within each coupler. They are separated from the SMD
payload by bulkheads and the other sections by bay lids. Each avionics sled is
supported by a rail system consisting of two 10 inch rods of 0.25 inch steel all thread.
The inside of each avionics bay is lined with aluminum flashing to shield the avionics
from outside interference, as seen in Figure 6.
Each avionics bay lid, shown in Figure 7, consists of two fiberglass bulkheads. One
bulkhead fits the inside diameter of the airframe, and the other fits the coupler to keep
the bulkhead centered. This prevents the lid from shifting and becoming lodged within
the bay. A welded eyebolt is mounted in the center and used as an attachment point for
the recovery system. Two 0.25 inch holes are drilled through the lid on either side of the
eyebolt to slide onto the rail system. A 0.125 inch hole is drilled through the lid to allow
the ejection charge leads from the altimeters to pass.
Figure 5: Coupler
Figure 6: Avionics Bay

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The bulkheads separating each avionics bay from the SMD payload housing were
machined out of PVC using a lathe. They are each one inch thick with material lathed
out to leave a cup shape on the side facing the avionics. Six 0.1875 inch holes are
drilled through the side of the bulkhead at 60 degree intervals for screws. Two 0.125
inch holes are offset from each screw hole to secure self locking nut plates inside the
bulkhead using rivets. Two 0.25 inch holes are drilled through the bulkhead for the rail
system, which is secured using washers, lockwashers, and nuts. This design is
displayed in Figure 8.
Figure 7: Avionics Bay Lid
Figure 8: Payload Bulkhead and Avionics Rail System

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Each avionics sled is made out of 0.25 inch thick plywood. Aluminum tubes are
attached to the sled using coax cable mounts as shown in Figure 9. These are used to
easily slide the sled onto the rail system of the avionics bay. The layout of the avionics
sleds is displayed in Figure 10. Each component is mounted to the sled using screws
and nuts. The battery is further secured to the battery mount using cable ties to prevent
them from becoming dislodged during flight.
SMD Payload Housing
The payload housing structure is 27 inches long. It is made of clear acrylic for visible
verification that the payload is powered on and functioning as illustrated in Figure 11.
This also allows UV sensors and cameras to be mounted internally thus not requiring
payload ejection. As seen in Figure 12, the couplers are mounted with six screws
containing nylon grommets to prevent the metal screws from contacting the acrylic. The
Figure 9: Avionics Bay Outside of Coupler
Figure 10: Avionics Sled Layout

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payload mounting rail is attached to one coupler with screws and held in place by a
bracket on the other. Portholes for the payload are located on the acrylic eight inches
from upper body airframe. Five 0.25 inch holes are drilled to provide proper venting.
Mass Simulator
For the full-scale flight, the payload was omitted and a mass simulator was used in
place. A similar payload mounting rail was lined with aluminum framework and washers
were attached to simulate the weight of the components of the payload. Weights were
placed as shown to accurately represent mass locations of the actual payload. The
simulator weighed 2.59 pounds. The completed simulator can be seen in Figure 13.
Figure 11: Acrylic Payload Housing Structure
Figure 12: Coupler Screw in Attachment

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Booster section
The booster section of the vehicle is 28 inches long. Mounted inside the booster section
is a 20 inch motor mount tube as pictured in Figure 14. The motor mount tube is three
inches in diameter to accommodate a 75 millimeter motor. Slots measuring 0.125 inch
wide 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.
Each fin tab is made of 0.125 inch thick fiberglass and is secured with epoxy. The
fiberglass can has high durability and increases the chance that they can withstand
impact and be reusable. The fin dimensions can be seen in Figure 15.
Figure 13: Weight Simulator
Figure 14: Booster Section

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For safety of the recovery system, the booster section contains dual recovery
attachment. The drogue parachute attaches via shock cord to a larger 0.375 inch
welded eyebolt attached to the motor casing as well as a shock cord epoxied along the
entire length of the motor mount tube as seen in Figure 16. The dual attachment
scheme is shown in Figure 17.
Figure 15: Fin Dimensions
Figure 16: Shock Cord Epoxied to Motor Mount Tube

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Drawings and schematics
The vehicle design is illustrated in the following figures. These show the various internal
characteristics to each section in light gray lines.
Figure 18 shows the booster section and motor tube layout.
Figure 17: Dual Attachment Scheme

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Figure 19 shows the layout of the upper body airframe with nose cone attached.
Figure 20 shows the payload housing structure with aluminum framework installed and
couplers that house the avionics bays.
Figure 18: Booster Section and Motor Mount Tube
Figure 19: Upper Body Airframe with Nose Cone

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Figure 21 shows the layout of the entire vehicle assembly.
Figure 22 is a CAD rendering of the final vehicle.
Figure 20: Payload Housing Structure, Aluminum Framework, and Couplers
Figure 21: Vehicle Assembly

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Flight Reliability Confidence and Mission Success Criteria
The confidence in the design of the vehicle is based off of two major factors, testing and
assembly. Careful assembly and precision in manufacturing provide confidence that
vehicle will withstand all loading during launch and throughout the flight. With the high
number of test launches conducted, the structural integrity and aerodynamic profile of
the vehicle has proven very reliable. Considering motor choice, fin and nose cone
shape, the resulting stability margin, total weight, and the final design of the launch
vehicle has demonstrated that a safe and reliable flight can occur.
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. The vehicle design is fully verified in Table 25
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
conditions are met and an apogee of exactly 5,280 feet is achieved, plus or minus 0.1%
plus one foot due to precision of the scoring altimeter.
Figure 22: Rendereing of Final Vehicle

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Testing
Static testing on individual components, such as airframe strength, was completed prior
to CDR. The results of the static tests were conclusive that each component was strong
enough to withstand all expected loads, both during liftoff and at landing. This has been
further verified through the large number of test flights conducted to date. Since CDR, 7
test flights have demonstrated a very durable vehicle design.
Functional testing has been achieved through extensive test flights. Sub-scale flight
testing was used to minimize losses while testing the parachute attachment scheme
and avionics. Once each subsystem was perfected, the first prototype was built to
perfect the integration of all components and test the design of the airframe and
aerodynamics. Before the first prototype was completed, simulation using OpenRocket
provided a method of testing the size, shape, and weight of the rocket.
Each component in the final build must work together perfectly for a successful flight
and recovery. This was demonstrated in the first full-scale launch of the final build. This
launch contained a mass simulator for the payload. The flight was within 30 feet of the
simulation, and was recovered with no damages. This flight provided verification of a
successful design, assembly, and launch procedure.
As documented in the CDR, the results of the static tests on the vehicle components are
included below.
All Structure Tests were done using a hydraulic press. Force is applied in ~10 pound
increments. 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.
These tests were done at the Polen Facility in Granbury, Texas. The tests were done in
a way to minimize destruction to both the components and the facility. They were
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.
Airframe
To test the strength of the 0.125 inch thick bulkhead, a 2 inch by 3 inch block was
placed on a bulkhead and a hydraulic press was used to apply force to it. The result
was that it held the maximum amount of weight, 1500 pounds, that the scale could
measure. Using this number and dividing by the contact area, the flat sheet of fiberglass
can hold over 250 pounds per square inch.
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. The coupler,

18
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.
The fins for the final vehicle are twice as thick as the fiberglass bulkhead. Thus, the
shear strength of a fin is greater than 250 pounds per square inch, the minimum tested
strength of the bulkhead.
Epoxy
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, pressure was applied to the bulk head. 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.
To test the epoxy for mounting, 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. 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. Using τ = F x d,
a torque of 550 inch-pounds is the maximum force applicable before the epoxy is
compromised. After the CDR it became necessary to trim the fins by one inch to obtain
better stability witch changes the maximum amount of perpendicular force on the fin
before the epoxy is compromised. Using 550 in-pounds as the torque, and 4 inches as
the distance, the maximum amount of force perpendicular to the fin is 137.5 pounds.
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
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.

19
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 measured 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 or Polen employees.
Vehicle Safety and Failure Analysis
Please refer to the Safety and Environment (Vehicle) section for the vehicle safety and
failure analysis.
Full-scale Launch Results
March 14, 2013
Changes made to design:
 None
Changes made to procedures:
 None
Date March 14, 2013
Location Hunewell Ranch, 32.2N, -98.2E, 1309MSL
Launch Rod 10 foot, button
Wind Speed 15 mph
Wind Direction SSW
Temperature 75° F
Pressure 30.21 inHg
Table 5: Launch Conditions

20
Figure 23: Simulation Data
Figure 24: OpenRocket Rendering of Vehicle
Figure 25: Simulated Flight Data

21
Stage Time Velocity Height AGL Lateral Displacement
Off-Rail 0.290 s 76.766 fps 10 ft 0 ft
Drogue
Deployment
17.708 s 3.493 fps 4925 ft 627 ft
Main
Deployment
64.471 s 92.478 fps 676 ft 135 ft
Impact 119.550 s 11.323 fps 0 ft 1377 ft
Component Apogee Time to Apogee Time of Flight
Drogue
Raven 3
Corrupted Data Corrupted Data Corrupted Data
Drogue
Stratologger
4947ft AGL 18.40 s 97.30 s
Main Raven
3
Corrupted Data Corrupted Data Corrupted Data
Main
Stratologger
4935ft AGL 18.95 s Corrupted Data
GPS* 4977ft AGL 45.00 s 120.00 s
*The Garmin Astro DC40, which is secured inside the nose cone, reported a landing
distance of 1,056 feet from the launch rail. The simulation predicted an apogee of 4,925
feet AGL, the average recorded apogee was 4,953 feet AGL, which is within 28 feet. A
mass simulator was used to represent the weight of the final SMD payload. The vehicle
was fully ballasted to 10% of the total vehicle weight, resulting in a total weight of 39.5
pounds. This was the first flight with the final vehicle design and is used for the full scale
demonstration flight. The flight was a success and the vehicle was safely recovered and
reusable at landing.
The drogue Stratologger (secondary) flight data is shown in Figure 26.
Table 6: Simulated Flight Data
Table 7: Recovery Electronics Data

22
The main Stratologger (primary) flight data is shown in Figure 27.
Figure 26: Drogue Stratologger Flight Data
Figure 27: Main Stratologger Flight Data

23
March 15, 2013
Changes made to design:
 None
Changes made to procedures:
 None
Date 15 March 2013
Location Hunewell Ranch, 32.2N, -98.2E, 1309MSL
Launch Rod 10 foot, button
Wind Speed 19 mph
Wind Direction SSW
Temperature 80 F
Pressure 33.2 inHg
Table 8: Launch Conditions
Figure 28: Simulation Data
Figure 29: OpenRocket Rendering of Vehicle

24
Stage Time Velocity Height AGL Lateral Displacement
Off-Rail 0.280 s 78.994 fps 10 ft 0 ft
Drogue
Deployment
18.202 s 1.501 fps 5306 ft 517 ft
Main
Deployment
70.344 s 89.326 fps 690 ft 242 ft
Impact 128.980 s 11.120 fps 0 ft 1299 ft
Component Apogee Time to Apogee Time of Flight
Drogue
Raven 3
5255ft AGL 18.96s 72.68s
Drogue
Stratologger
5282ft AGL 19.60s 73.10s
Main Raven
3
5207ft AGL 18.72s 72.68s
Main
Stratologger
5269ft AGL 19.65s 74.90s
GPS* Not Reliable Not Reliable Not Reliable
*The Garmin Astro DC40 reported a landing distance of 681ft from the launch rail.
While the simulation predicted an apogee of 5306ft AGL, the average recorded apogee
was 5253ft AGL, within 27ft of the desired apogee of one mile. The ballast was
optimized for the purpose of reaching an apogee of one mile, at 1.3 pounds. The SMD
payload was on board, and while the rocket did become treed, it was recovered. The
SMD payload maintained telemetry throughout the flight and after landing. Due to the
amount of time that was taken to recover the rocket from the tree, it was confirmed that
recovery and payload electronics can maintain power for much longer than one hour.
Figure 30: Simulated Flight Data
Table 9: Simulated Flight Data
Table 10: Recovery Electronics Data

25
For a more detailed analysis of the SMD flight, please refer to the Payload Criteria
Section.
The drogue Stratologger data is shown in Figure 31.
Drogue Raven3 is shown in Figure 32.
Figure 31: Drogue Stratologger Data
Figure 32: Drogue Raven3 Flight Data

26
Main Stratologger is shown in Figure 33.
Main Raven3 flight data is shown in Figure 34.
Figure 33: Main Stratologger Flight Data
Figure 34: Main Raven3 Flight Data

27
Mass Report
The unballasted mass summary of the launch vehicle and its three main sections are
located in Table 11. Each vehicle section is also broken down into its individual
components in Table 12. All listed mass calculations for the vehicle, sections, and
individual components were obtained through the weighing of the components on two
separate scales. All items weighing less than 2.2045 pounds were weighed on a JS-
1000V digital scale with an accuracy of 0.0002 pounds. For heavier components, a
Tanita 1583 baby scale was used. The baby scale has a weight capacity of 40 pounds
with a graduation of 0.5 ounces for up to 20 pounds and one ounce for 20 to 40 pounds.
The final launch vehicle has an unballasted mass of 35.9 pounds on-the-rail and a fully
ballasted mass of up to 39.5 pounds. The variability of up to 10 percent of the vehicle
mass for the ballast system allows for greater control of altitude for a wider range of
atmospheric conditions. It was also found during testing that the mass of each motor
varies slightly from the listed amount and is accounted for in each simulation with a
variable mass component.
Mass Summary
Section Mass (oz) Mass (lb)
Upper Body 148.99 9.31
Payload Housing 186.69 11.67
Booster 238.22 14.89
Total Mass (Launch) 573.90 35.87
Total Mass (Apogee) 570.03 32.00
Mass per Section
Upper Body
Component Mass (oz) Mass (lb)
Ballast System Bulkhead 4.85 0.303
Ballast System/Upper Eye Bolt1
2.54 0.159
Dog Tracker GPS 4.80 0.300
Main Deployment System 64.0 4.00
Main Ejection Charge 1 0.192 0.012
Main Ejection Charge 2 0.208 0.013
Main Shock Cord 9.30 0.581
Nose Cone 16.5 1.03
Upper Body Airframe 45.1 2.82
Upper Dog Barf 1.50 0.0940
Subtotal 148.99 9.31
Payload Housing
Acrylic Airframe 40.96 2.56
Drogue Avionics 5.952 0.372
Table 11: Mass Summary

28
Drogue Avionics Bay Lid 11.696 0.731
Lower Coupler 22.4 1.4
Lower Payload Bulkhead 10.352 0.647
Main Avionics 5.952 0.372
Main Avionics Bay Lid 11.696 0.731
Upper Coupler 24.496 1.531
Upper Payload Bulkhead 11.744 0.734
USLI Electronics Payload 41.44 2.59
Subtotal 186.688 11.668
Booster
Centering Rings – 4 12.864 0.804
Drogue Ejection Charge 1 0.08 0.005
Drogue Parachute 3.04 0.19
Drogue Shock Cord 4.624 0.289
Fins – 4 21.6 1.35
Lower Body Airframe 44 2.75
Lower Dog Barf 1.056 0.066
Lower Eye Bolt 2.864 0.179
Motor Retaining Ring 4.96 0.31
Motor Tube 25.12 1.57
Motor2 3
117.92 7.37
Subtotal 238.224 14.889
Centering Rings – 4 12.864 0.804
1
Mass of Ballast varies with configuration.
2
Mass listed is for launch. The empty mass is:
3
Mass variation found in each motor.
Recovery Subsystem
Design Defense
The final recovery system design has been tested with sub-scale prototype launches as
well as two full-scale launches.
Table 12: Mass by Subsection

29
The parachute systems are as robust as the shock harnesses. In the upper body
airframe, a layer of 1.5 ounces of dog barf wadding, 33 feet of z-folded 0.375 inch
tubular Kevlar line, and a Kevlar deployment bag protect the main parachute from the
heat of the ejection charge. No heat damage is visible on the main parachute, pilot
parachute, or shroud lines due to test launches. Slight singeing is visible on the front
flap of the Kevlar bag, but its integrity has not been compromised. Similarly, in the
booster section, a layer of 1.05 ounces of dog barf wadding, 16 feet of z-folded 0.375
inch tubular Kevlar line, and a Nomex blanket protect the drogue parachute from the
heat of the ejection charge. No heat damage is visible on any of the drogue parachute
components.
Prior to each launch, the recovery team inspects the shock cord for any fraying or kinks.
All kinks are removed before the line is z-folded and secured with either painter's tape
or a thin rubber band. Fraying has not yet occurred to warrant a new line, but it is
planned to fly with fresh harnesses on launch day.
The manufacturer rating for the 0.375 inch tubular Kevlar from Rocketry Warehouse
used in all test flights is 3,600+ pound-force. Since the design is shedding weight, the
highest load experienced on launch day is expected to be less than 1,750 pound-force,
the highest test load experienced, as analyzed from Featherweight Raven 3 data
captured from the flight on February 19, 2013. This gives a worst-case-scenario factor
of safety rating (FOS) of 3600/1750 = 2.057, which exceeds the 2.000 required.
The highest drogue deployment velocity experienced on flights with the 0.375 inch
shock line used in all test flights occurred on January 26, 2013. After burnout, the
vehicle weighed slightly less than 30 pounds, and experienced 11.68 G on drogue
parachute deployment. This resulted in a maximum shock of 350.4 pound-force, for a
worst-case-scenario FOS of 3600/350.4 = 10.274.
The only knots in the 40 foot shock harness are at the connection point to the nose
cone, and at the connection point to the main avionics bay coupler. The knots in the
drogue (20 foot) shock harness are located at the connection point to the motor casing,
the connection point to the drogue avionics coupler, and four feet from the connection
point to the drogue avionics coupler. This latter knot serves as an attachment point for
the drogue parachute.
All knots are figure-eight knots. No adhesives are used to secure any portion of either
harness. Although the line is heat resistant Kevlar, each harness is shielded from the
blast of the initial ejection charge firing by dog barf. It is recognized that the secondary
charge firing could cause heat damage to the shock cord.
All welded eyebolts have remained intact after each flight. No attachment hardware
including quick links, washers, lock washers, and all-thread, has been damaged as the
result of any test flight.

30
Power loss to the altimeters during flight can result due to the battery coming
temporarily disconnected from the shock force of main parachute deployment. While the
former primary altimeters, the Featherweight Raven3s have shut down and reentered
flight mode due to this power loss, the former secondary altimeters, the Perfectflite
Stratologger SL100s, remain in flight mode and gathering data through the surge. All
four recovery altimeters are now Perfectflite Stratologger SL100s.
The Featherweight magnetic arming switches have not malfunctioned to this point.
Externally visible LED lights wired through each magnetic switch power on only when
the switch is armed by the magnet. These lights confirm power to the recovery
altimeters.
The Garmin Astro DC-40 transmitter in the nose cone is powered on prior to packing the
main parachute components in the upper airframe. Four screws with plastic anchors
hold the nose cone, where the transmitter and the ballast system are housed to the
upper airframe.
Structural Elements
In order to enable deceleration of the three tethered sections on descent, 0.375 inch
tubular Kevlar shock cords have been implemented based on strength and flame-proof
construction. The length of the shock cord for the drogue parachute is 20 feet. Based
on the advice of the team mentor, the main parachute shock cord should be twice as
long as the drogue parachute shock cord. The main parachute shock cord is, in effect,
40 feet long.
The shock cord for each parachute is z-folded and bound with rubber bands. This
reduces the risk of entanglement and saves space in the parachute compartments. The
harnesses are attached to the vehicle via stainless steel quick links through a figure-
eight knot in the end of the line. This connects to welded eyebolts on airframe couplers.
The motor casing within the booster section has a welded eyebolt installed. This eyebolt
is used to attach the booster section to the 20 foot shock cord. A section of shock cord
epoxied to the interior wall of the booster section in case of failure of the eyebolt on the
motor housing is attached to the drogue parachute shock cord. The other end of the
shock cord is attached to a welded eyebolt installed on the exterior bulkhead of the
drogue avionics bay. This effectively tethers the booster section to the payload housing
structure. The drogue parachute shroud lines are attached to a figure-eight knot in the
shroud line four feet below the drogue avionics bay.
The exterior bulkhead of the main avionics bay also has an eyebolt installed. This
eyebolt is used to attach the 40 foot shock cord to the payload housing section. The
other end of this line is attached to a welded eyebolt on the coupler between the nose
cone and the upper body airframe. The main parachute shroud lines are also anchored
here via a swivel between the four lines and the quick link.

31
A „d‟ shaped stainless steel ring at the apex of the main parachute is attached via quick
link to a square stainless steel ring on the end of a nylon strap on the inside of the
deployment bag. A quick link is also used to attach a pilot drogue parachute to a square
stainless-steel ring on the end of a nylon strap on the outside of the deployment bag.
The shroud lines are woven through elastic straps on the front of this bag.
Electrical Elements
In order to stage dual deployment, two sets of altimeters are implemented. Each system
consists of two Perfectflite Stratologger SL100 altimeters mounted to a 0.19 inch thick, 5
inch wide, 8.125 inch long plywood sled and the power supply. The magnetic arming
switch is mounted on the opposing side from the altimeters and sits at the edge of the
inner avionics bay wall.
A nine-volt battery supplies power to each set of altimeters through a Featherweight
Magnetic Arming Switch. Each switch is able to be armed externally using a magnet
provided by the manufacturer. Based on trial and error, the range of the magnet for
arming the switch is best at one inch from the exterior of the airframe. It is confirmed
that the switch has been armed and is supplying power to the altimeters by an
externally visible LED light. This light remains on until the switch is disarmed and power
is no longer being supplied to the altimeters.
Each altimeter is wired to a black powder ejection charge. The charges are each made
from one J-Tek 10 electric match cut to three feet for the booster section and four feet
for the nose cone section. The pyrogen-coated end of the electric match rests at the
bottom of a 13millimeter by 100millimeter Polystyrene test tube, with the leads pushed
through a 0.125-inch hole drilled into the bottom. The electric match is secured to the
tube by hot glue on the exterior of the hole. The exposed leads are secured to a male
crimp connector.
The charge leads from each altimeter are secured to a female crimp connector on the
opposite side of the avionics coupler. Each electric match of a certain black powder
charge weight is connected to the appropriate altimeter terminals. The charge canisters
are placed horizontally at the bottom of the appropriate airframe section. The leads are
secured with a piece of Gorilla tape to ensure that the crimp connection is not
compromised.
Redundancy Features
The recovery system employs a dually redundant ejection charge system to ensure
airframe separation and parachute ejection at the correct flight stage. Two altimeters of
the same make and model are programmed to fire separately. The primary will fire a

32
charge of the calculated size for the compartment at the desired time, while the
secondary will fire a slightly larger charge at a delayed time.
In the drogue parachute avionics bay, the primary SL100 is programmed to fire at
barometric detection of apogee. A 2.6 gram FFFg black powder charge is to be ignited
and cause separation of the booster section from the forward airframe. The secondary
SL100 is programmed to fire at one second past barometric detection of apogee. A 2.8
gram FFFg black powder charge is to be ignited in case of failure of the primary charge
to ensure separation and ejection of the drogue parachute.
In the main parachute avionics bay, the primary SL100 is programmed to fire at 700 feet
above ground level (AGL). The secondary SL100 is programmed to fire at 650 feet
AGL. A time delay is not used for main parachute ejection, since the time delay feature
on these altimeters is only available for apogee. However, the 50 foot delay should be
sufficient to ensure that both altimeters in the main avionics bay do not fire
simultaneously.
The manufacturer-reported calibration accuracy of these altimeters is within 0.05%.
This corresponds to an error of 0.35 feet at 700 feet AGL between two typical SL100
altimeters in the main parachute avionics bay, supposing they entered flight mode at the
same time. Furthermore, the manufacturer-reported measurement precision of these
altimeters is 0.1% of the reading, plus a foot. This corresponds to an error of 1.7 feet at
700 feet AGL between two typical SL100 altimeters in the main parachute avionics bay,
supposing they entered flight mode at the same time. A difference of 2.05 feet between
the main avionics bay altimeters is typical, which is well within the 50 foot delay
employed.
As mentioned in the previous section on structural elements, the recovery system also
features redundant tethering of the booster section. A 24 inch section of 0.375 inch
tubular Kevlar with a loop at the forward end is epoxied to the interior wall. The drogue
parachute shock cord is attached to this loop via quick link, then to the eyebolt on the
motor housing. This serves as a backup connection of the drogue parachute harness to
the booster section, ensuring tethering of the booster section to the forward airframe in
the event of failure of the eyebolt on the motor housing.
Parachutes
The final recovery design consists of the ejection of a drogue parachute at apogee,
followed by the ejection of a main parachute at a height of between 700 and 650 feet
AGL on descent. The drogue parachute is protected from ejection charge firing by a
Nomex cloth and dog barf wadding. The main parachute is also protected from ejection
charge firing by dog barf. Additionally, the main parachute is packed into a Kevlar TAC-
9B Deployment Bag to reduce the risk of entanglement of the shroud lines on ejection.
To further ensure deployment of the main parachute, a pilot drogue parachute is
attached to the bottom of the deployment bag to help pull the bag out of the airframe on

33
separation. The deployment bag is shown in Figure 35. The operation of the
deployment bag is specified in Figure 36.
Figure 35: Packed Deployment Bag
Figure 36: Deployment Bag Operation

34
As in the CDR, the choice of drogue parachutes remains the Sky Angle CERT-3 24",
with a drag coefficient of 1.16. Likewise, the choice of main parachute remains the Sky
Angle CERT-3 XX Large 120", with a drag coefficient of 2.92. Each parachute is cross-
form and constructed of 1.9 ounce silicon-coated rip-stop Nylon. All seams are
reinforced with Nylon webbing to reduce the probability of shroud line disconnection on
deployment. With four shroud lines, each made of 5/8 inch woven tubular nylon for
durability, these parachutes have a low probability of entanglement. The length of the
shroud lines on the main parachute is ten feet, while that on each drogue parachute is
two feet. Further reducing the risk of entanglement upon deployment is the 7,000
pound tested swivel between the main parachute shroud lines and shock cord.
The total launch vehicle weight is 35.9 pounds, but the team has tested the maximum
(10%) ballast for a launch vehicle weight of up to 39.49 pounds. At apogee, burnout will
have occurred already, to give a vehicle weight of up to 35.62 pounds. The following
calculations give the maximum descent rate upon landing for the vehicle to have a
kinetic energy of less than 75 foot-pound force.
√
From the above calculations, it can be concluded that the launch vehicle will experience
a combined impact force of less than 75 foot-pounds should it land at a speed of no
more than 11.645 feet per second. This result is used below to calculate the minimum
diameter of the main parachute. Assumptions in the calculation are that the vehicle is
descending at a constant speed, and that the downward motion is simple, along the z-
axis.

35
( )
√ √
The minimum diameter of the main parachute is thus 2a = 9.918ft. Based on
commercial availability of parachutes, a 10 foot diameter parachute theoretically suffices
for the main. The drag coefficient of 2.92 is provided by the manufacturer for the chosen
and tested main parachute. The calculation for the descent rate of the vehicle under the
10 foot main parachute follows.
√ √
The calculated descent rate of the vehicle after main parachute deployment at 700 feet
AGL is thus 11.645 feet per second. With this figure, it is possible to calculate the total
drift of the vehicle after main parachute deployment until landing. The drift of the vehicle
from main parachute deployment until landing follows, assuming a 15 mile per hour
horizontal wind.
With a fifteen mile per hour wind, the launch vehicle will have drifted approximately
1,322 feet from the launch pad under the main parachute before landing. Since the
maximum allowable drift is 2,500 feet in a fifteen mile per hour wind, the drift between
drogue and main parachute deployment can be at most 1,178 feet. Using this maximum
drift between drogue and main parachute deployment, the minimum descent rate for the

36
vehicle under the drogue parachute is calculated, assuming an apogee of one mile, or
5280 feet AGL.
Thus the minimum descent rate from apogee to main parachute deployment to ensure a
total drift of no more than 2,500 feet by landing is 85.535 feet per second. With this
figure, it is possible to determine the appropriate size of the drogue parachute. The
maximum diameter of the drogue parachute that would allow for this descent rate from
apogee is determined below. Assumptions made in this calculation are that the vehicle
is descending at a constant rate, and that this downward motion is simple, lying solely
along the z-axis.
( )
√ √
The minimum diameter of the drogue parachute is thus 2a = 2.142ft. Based on
commercial availability of parachutes, a two foot diameter parachute theoretically
suffices for the drogue. The drag coefficient of 1.16 is provided by the manufacturer for
the chosen and tested drogue parachute.
Drawings and Schematics
Electrical

37
The wiring configuration of the altimeters is given in Figure 37.
Figure 38 shows the Stratologger Software flow diagram.
Figure 37: Altimeter Wiring Configuration

38
Figure 39 shows the GPS software flow diagram.
Figure 38: Stratologger Software Flow Diagram

39
Figure 39: GPS Software Flow Diagram

40
Structural
The following drawings are of the structural components of the avionics bay. The
avionics bay are shown below in Figure 40.
The features of the attachment bulkhead within the avionics coupler is shown below in
Figure 41.
Figure 40: Avionics Bay Structural Components
Figure 41: Attachment Bulkhead Inside Avionics Coupler

41
The dimensions of the couplers that house the avionics bays are shown in Figure 42.
Figure 43 contains an exploded view for the structural assembly of the avionics bay.
Recovery Tracking
The recovery system includes a GPS transmitter and a handheld receiver, as seen in
Figure 44. The transmitter, as stated in the CDR, is the Garmin Astro DC-40, which runs
up to 48 hours on a rechargeable lithium-ion AA battery. This unit is attached to the all-
thread support structure of the ballast system in the nose cone.
Figure 42: Couplers in Profile
Figure 43: Structural Assembly of the Avionics Bay

42
The handheld receiver is the Garmin Astro 320, which runs on two AA batteries for up to
20 hours. This unit is located at the ground station with a member of the recovery team.
The range of transmission accuracy between the DC-40 and the 320 is reported by the
manufacturer to be up to nine miles in clear, unobstructed terrain.
The output power of the DC-40 is two Watts. Use of any high-powered radio near the
320 could cause irreparable damage to the unit. Note that a high-powered radio is
defined as having a power output of at least five Watts.
Transmitting frequencies picked up by the receiver are in the Multi-Use Radio Service
(MURS) band. The MURS band consists of 151.8200MHz, 151.8800MHz,
151.9400MHz, 154.5700MHz, and 154.6000MHz. The default setting on the 320 is
151.8200MHz, but any of the five MURS frequencies may be chosen.
The recovery team conducted a series of range tests. These included leaving the DC-40
in the rocketry lab at the university math building and driving off with the 320 until the
signal was lost, and leaving the DC-40 at the launch site on Hunewell Ranch and driving
or walking off until the signal was lost. It was found that in obstructed terrain, including
neighborhoods, 0.792 miles (4181.760 feet) was the shortest range of accuracy. Thus
the system should be sufficient to locate the vehicle within the 2500 foot allowable
landing radius on launch day.
System Sensitivity
Aluminum flashing (0.01 inches in thickness) lines all surfaces of each avionics bay.
Since aluminum is a conductor of electricity, this forms a Faraday cage around the
recovery altimeters. Thus, not only are the avionics bays shielded from any onboard or
nearby radio frequencies reflected by the flashing, but also from magnetic waves
generated by onboard electronics.
Figure 44: GPS Device and Wireless Remote

43
If there is a sufficiently strong source of magnetic waves nearby, it is possible that the
switch could be armed. This would switch the LED(s) of the armed avionics bay(s) on,
however, enabling the team to disarm the affected system(s). Furthermore, the
altimeters will not send voltage to an ejection charge unless they have entered flight
mode, so inadvertent excitation of the recovery altimeters can be safely mitigated
through deactivation.
As discussed in the previous section on recovery tracking, the recovery tracking system
is susceptible to high-power radio devices. None of the onboard radio transmitters has
an output power of greater than two Watts. This can be verified in the description of the
SMD payload.
Parachute Justification
As calculated previously, the most appropriate sizing of the main and drogue
parachutes is 10 and two feet, respectively. These calculations take into account
deployment process and vehicle mass. The attachment scheme of the main parachute
comes directly from the manufacturer. This scheme was validated on the last three test
flights. The attachment scheme for the drogue parachute, as described in the previous
sections on structural elements and redundancy features, has been optimized through
test flights.
As documented in the CDR, one static ejection test has been completed based upon
calculated black powder weights. The ejection charge sizing is appropriate for the length
of the airframe section for each parachute. This has been thoroughly justified through
test flights, as charge sizing has enabled separation upon being prepared properly.
Safety and Failure Analysis
Following is a log of the seven test flights completed since the CDR. For each flight,
failure modes are identified. Each failure mode is then analyzed in terms of
components affected, components responsible, potential causes, and potential hazards.
Mitigations of these hazards are then identified and enacted on subsequent flights.
The first five test flights were conducted using the full-scale Prototype 1, which was built
to specifications from the CDR and gradually modified to incorporate many of the
specifications in the FRR. The last two, namely the full scale demonstration flight
(3/14/2013) along with a full scale test flight with SMD payload (3/15/2013) were
conducted using the finalized design, which was built to specifications in the FRR.
These last two flights reflect expected performance on competition launch day.

44
Test Date: January 26, 2013
Brief Description of Test: A dual deployment test launch was conducted using the
modified full-scale Prototype 1 rocket at Hunewell ranch around noon. The motor used
was a Cesaroni L585. For main parachute ejection the altimeter used was a Perfectflite
Stratologger SL100, equipped with 5.8g packed 3F black powder. A Stratologger was
also used for backup ejection of the drogue parachute, with a 2.8g packed 3F black
powder charge. The primary altimeter used for drogue parachute ejection was a
Featherweight Raven 3, with a 2.6g packed 3F black powder charge. It was shown in
this flight that the attachment scheme for the main parachute deployment bag is
sufficient for successful deployment of the main parachute with minimal entanglement.
Each parachute was ejected fully and as programmed. The launch that vehicle was
recovered in a mostly clear area. The only damage sustained to any component was
that potentially incurred to the Garmin Astro DC40 when it was ejected from its
attachment point.
Failure
Mode
Component(s)
Responsible
Component(s)
Affected
Potential
Cause(s)
Potential
Hazard(s)
Mitigation
Untethered
transmitter
unit
Attachment to
drogue
parachute
shock cord
Direct location
of landed
launch vehicle
Tape
weakened
by heat of
ejection
charge
and broke
on force of
impact
Lost
transmitter
makes it
harder to
find rocket,
could be
disqualified
if this
happens in
competition
New
attachment
scheme
involving
zip ties and
back plate
to secure
DC40 to
drogue
shock cord
Table 13: Test Flight 1/26/13

45
Test Date: February 14, 2013
modified full-scale Prototype 1 rocket at Hunewell ranch at 3 p.m. The motor used was
a Cesaroni L585. For main parachute ejection the altimeters used were a Perfectflite
Stratologger SL100 programmed to fire at 650ft AGL on descent, equipped with a 5.5g
packed 3F black powder charge, and a Featherweight Raven3 programmed to fire at
704ft AGL on descent, equipped with a 5.0g packed 3F black powder charge. A
Stratologger was used for backup ejection of the drogue parachute at 2s past apogee,
with a 2.6g packed 3F black powder charge. The primary altimeter used for drogue
parachute ejection was a Featherweight Raven 3, with a 2.9g packed 3F black powder
charge. Also on board was a small portion of the payload for data gathering. The
Garmin Astro DC40 GPS unit was secured inside the nose cone rather than to the
drogue parachute shock cord as in the last flight. It was shown in this flight that the
main parachute shock cord should be located between the ejection charges and the
deployment bag, rather than between the deployment bag and the avionics coupler, in
order to prevent entanglement. It was also determined that ribbon wire would be used
for all avionics connections to reduce the risk of disconnection before landing.
Failure
Mode
Component(s)
Responsible
Component(s)
Affected
Potential
Cause(s)
Potential
Hazard(s)
Mitigation
Main
Parachute
not
Envelope
Main parachute
shock cord
Main parachute
shroud lines did
not release,
causing main
parachute to
not be released
from bag until
impact
Placement
between
bag and
coupler
caused
shroud
lines to
push into
and tangle
with shock
cord upon
ejection
Ballistic
impact,
damage to
vehicle or
on-board
electronics
Place main
parachute
shock cord
between
bag and
charges

46
Test Date: February 19, 2013
modified full-scale Prototype 1 rocket at Hunewell ranch at 2 p.m. The motor used was
a Cesaroni L585. For main parachute ejection the altimeters used were a Perfectflite
charge. No payload was on board. It was shown in this flight that the new placement of
the main parachute shock cord was successful, and that rubber grommets should be
used to cushion the attachment points of the acrylic payload section to the avionics
couplers.
Failure
Mode
Component(s)
Responsible
Component(s)
Affected
Potential
Cause(s)
Potential
Hazard(s)
Mitigation
Reusability
of acrylic
payload
section
No shock
absorption
between acrylic
and screws
Integrity of
acrylic at
attachment
points to
drogue avionics
coupler
Force of
impact
Damage to
payload
electronics,
reusability
of system
Rubber
grommets
to cushion
acrylic
from
screws
attaching it
to the
avionics
couplers

47
Test Date: March 7, 2013
modified full-scale Prototype 1 rocket at Hunewell ranch at 2p.m. The motor used was
a Cesaroni K510. For main parachute ejection the altimeters used were a Perfectflite
with a 2.9 packed 3F black powder charge. The primary altimeter used for drogue
charge. No payload was on board. It was shown in this flight that the new deployment
bag works, but that the charge leads need to be cut to fit.
Failure
Mode
Component(s)
Responsible
Component(s)
Affected
Potential
Cause(s)
Potential
Hazard(s)
Mitigation
Ignition of
Primary
Drogue E-
Match
Single-core
ground wire
Ground wire
pulled out
Leads were
too long and
became
tangled with
drogue
parachute
shroud lines
Failed
drogue
parachute
ejection if
secondary
charge
doesn't
ignite,
high-
speed
main
parachute
ejection
Cut e-
match
leads to fit
tube length
loosely, try
to find
braided
wire
connectors
Integrity
of Drogue
Parachute
Length of
drogue e-
match leads
Reusability of
parachute
Entanglement
of e-match
leads with
drogue
parachute
shroud lines
Drogue
parachute
not
reusable,
high-
speed
main
parachute
ejection
Cut e-
match
leads to fit
tube length
loosely

48
modified full-scale Prototype 1 rocket at Hunewell ranch at 10a.m. The motor used was
a Cesaroni K555. For main parachute ejection the altimeters used were a Perfectflite
charge. No payload was on board. It was shown in this flight that all eyebolts must be
welded for secure connection to be maintained. Further, it was decided after collecting
the altimeter data from this flight that all four recovery altimeters will be Perfectflite
Stratologger SL100's since the temporary power loss experienced upon airframe
separation does not affect flight mode.
Failure
Mode
Component(s)
Responsible
Component(s)
Affected
Potential
Cause(s)
Potential
Hazard(s)
Mitigation
Vehicle
remaining
tethered
Drogue
avionics
coupler eyebolt
failure
Altimeter power
loss, no main
parachute
deployment,
GPS
transmitter
damage, acrylic
payload section
shattered, one
fin cracked,
one pulled eye
bolt, irreparable
body damage
Eyebolt not
welded,
caused
untethering
of drogue
harness at
apogee
Vehicle not
reusable,
vehicle not
locatable,
damage to
onboard
electronics
Switch all
four
recovery
altimeters to
Perfectflite
Stratologger
SL100's,
weld all
eyebolts

49
fully-ballasted (10%) full-scale Prototype 2 rocket at Hunewell ranch at 5:45p.m. The
motor used was a Cesaroni L1720. For main parachute ejection the altimeters used
were a Perfectflite Stratologger(secondary) SL100 programmed to fire at 650ft AGL on
descent, equipped with a 5.2g packed 3F black powder charge, and a Featherweight
Raven3(primary) programmed to fire at 704ft AGL on descent, equipped with a 4.8g
packed 3F black powder charge. A Stratologger(secondary) was used for backup
ejection of the drogue parachute, with a 2.8g packed 3F black powder charge. The
primary altimeter used for drogue parachute ejection was a Featherweight
Raven3(primary), with a 2.6g packed 3F black powder charge. A mass simulator of the
payload was onboard. This flight was successful.
Failure
Mode
Component(s)
Responsible
Component(s)
Affected
Potential
Cause(s)
Potential
Hazard(s)
Mitigation
None None None None None None

50
Test Date: 15 March 2013
Brief Description of Test: A dual deployment test launch was conducted using the full-
scale Prototype 2 rocket at Hunewell ranch at 6:45p.m. The motor used was a Cesaroni
L1720. For main parachute ejection the altimeters used were a Perfectflite
Stratologger(backup) SL100 programmed to fire at 650ft AGL on descent, equipped with
a 5.2g packed 3F black powder charge, and a Featherweight Raven3(main)
programmed to fire at 704ft AGL on descent, equipped with a 4.8g packed 3F black
powder charge. A Stratologger(backup) was used for backup ejection of the drogue
parachute at 1s after apogee, with a 2.8g packed 3F black powder charge. The primary
altimeter used for drogue parachute ejection was a Featherweight Raven 3(main), with
a 2.6g packed 3F black powder charge. The SMD payload was on board. The ballast
system was optimized for the purpose of hitting an apogee of one mile, at 1.3 pounds.
The results of this flight demonstrated reliability of the simulation.
Failure
Mode
Component(s)
Responsible
Component(s)
Affected
Potential
Cause(s)
Potential
Hazard(s)
Mitigation
Main
Parachute
became
entangled
Pilot drogue for
deployment
bad
Minor acrylic
damage from
impact with
tree
Shock chord
entanglement
Very high
descent
velocity
Verify
attachment
and
packing
scheme
Mission Performance Predictions
Mission Performance Criterion
The project defines a successful mission as a flight with payload, where the vehicle and
SMD payload are recovered and able to be reused on the day of the official launch.
Moreover, the vehicle will not exceed 5,600 feet, and the official scoring altimeter will be
intact and report the official altitude. After apogee and descent, the entire vehicle lands
within 2,500 feet of the launch pad. In addition to meeting the mission success criteria, a
perfect launch will reach apogee at exactly 5,280 feet. With accurate simulation
conditions, an appropriate ballast mast can be determined to achieve a perfect launch.
Flight Profile Simulations, Altitude Predictions, Weights, and Actual Motor
Thrust Curve
Flight simulations are performed on OpenRocket, an open-source software suite for
rocketry. OpenRocket allows the user to completely construct the vehicle profile, define
component materials and masses, select from commercially available motors, and
simulate the entire flight based upon atmospheric conditions.

51
The flight simulation for the final design to be used on the official launch day is shown in
Figure 45. With the vehicle profile, actual weights of all components, and the selected
L1720-WT motor, the predicted altitude of the official flight is 5315 feet AGL with zero
ballast weight in a 15 mile per hour wind. A summary of the input parameters for the
simulation is given in Figure 46. These parameters reflect proposed conditions in
Huntsville, AL.
Careful attention was paid to ensure that simulation weights accurately reflect the actual
weights of all components of the launch vehicle. These weights are completely itemized
in the Mass Report section. The total vehicle mass with motor is 35.9 pounds
unballasted. The overshoot of the target apogee can be reduced by implementing
ballast weight. The most recent simulations suggest that with an added ballast weight of
4 ounces, totaling to 36.1 pounds with motor, the vehicle will reach apogee at
approximately 5277 feet. The optimal vehicle design, for flight in 15 mph winds, will
include the 4 ounces of ballast weight, creating a stability margin of 1.67.
With a gross liftoff weight of 36.1 pounds (this includes total vehicle weight and 4
ounces of ballast), the thrust to weight ratio is 10.9:1 based upon the average thrust of
the L1720 motor. The rail exit velocity is simulated to be 81.5 feet per second. During
the flight, a maximum velocity of 680 feet per second (Mach 0.61) and maximum
acceleration of 365 feet per second squared.
Multiple trials with the same input parameters in OpenRocket show that the predicted
apogee can vary by approximately plus or minus 10 feet, due to turbulence and non-
constant wind speeds. Along with a 10% variability in the commercially available
motors, these represent the most significant factors that are not in control by the team.
Furthermore, these two factors will contribute greatest to error in the predicted versus
actual performance.

52
Figure 45: Final Vehicle Simulation
Figure 46: Input Parameters for Final Simulation

53
As previously presented in the CDR, a static motor test was conducted to obtain an
actual motor thrust curve that is provided in Figure 47. The actual motor thrust curve
closely reflects the neutral thrust curve provided by the motor manufacturer in shape.
Thoroughness and Validity of Analysis, Drag Assessment, and Scale
Modeling Results
As mentioned above, despite thorough detail in modeling the vehicle in OpenRocket,
there still exists a level of uncertainty due to the uncontrollable factors of motor
variability and wind turbulence variability. OpenRocket offers an approximation for the
standard deviation of wind speed based on turbulence intensity; however, the
calculations are based on simulated numbers. These variance values will be compared
against data from the National Oceanic and Atmospheric Administration, as detailed in
the Science Value section, to determine their accuracy. Upon verifying more precise
values for wind speed standard deviation, the margin of error can be reduced and an
apogee of 5,280 can be more easily achieved. Furthermore, the variability of the motor
can only be accurately quantified by numerous thrust tests. Due to the price and
availability of these motors, it is not feasible to burn multiple motors to determine
variance between them. Cesaroni provides an average thrust curve for the L1720, but
without a large amount of sample data these averages lend no insight towards the
standard deviation.
Figure 47: L1720-WT Actual Thrust Curve

54
In lieu of the aforementioned tests and further analysis to determine any correlation
between the variances, a large scale simulation analysis has been devised to observe
overall interference from these factors. Using a custom-built automation macro, 944
simulations were conducted in OpenRocket to simulate the conditions of the official
launch day in Huntsville, AL. The flights, as expected, differed significantly in their
values and ranges across the sample. By conducting the experiment on such a large
scale the unknown variances become easier to analyze via significant trends within the
data pool.
Using data extracted from the simulations confidence intervals were established to
demonstrate the probability of achieving specific goals with the competition vehicle. As
demonstrated in Table 20, different ranges of values offer different probabilities.
According to the simulation analysis the launch vehicle has only a 5.5% chance of
hitting exactly a mile, but as you expand the neighborhood of tolerance the probabilities
rise significantly, resulting in a significantly high chance of near-mile apogee.
Apogee (feet) 5280 ± 5 ± 10 ± 15 ± 20
Probability (%) 5.530 51.374 83.624 96.328 99.465
Lateral Drift (feet) 1900 ± 50 ± 100 ± 150 ± 200
Probability (%) 0.534 49.909 82.174 95.653 99.290
The statistics suite RStudio was used to create the analysis script. The confidence
intervals were built around a basic statistical analysis of the simulation sample. Upon
finding the sample mean and sample variance, those values were used to build
Probability Density Functions (PDF) to give a graphical representation of the likelihood
of specific flight patterns. Figures 48 and 49 illustrate the functions for apogee and
lateral distance. The range value depicts the chances of the variable in question taking
corresponding values. Though the functions demonstrate relatively low probabilities of
landing at any particular value, integrating this function across an interval greatly
increases those probabilities. The table above reiterates this notion, that as the window
of expected values widens, the chance of falling within that range grows exponentially.
Table 20 - Confidence Intervals for Apogee and Drift

55
As the team continues testing with full scale launches, the differences between actual
performance and simulated performance of the vehicle and recovery system will
become clearer. It is very important to understand and analyze these differences in
order to make the simulation more accurate and valuable for altitude predictions. All
parameters must be selected such that they most accurately represent the actual
parameters of the vehicle flight.
The utility of simulations as a performance measure is best illustrated by the successful
launch of March 15, 2013. As the team began testing a full scale rocket with a finalized
Figure 48 - PDF of Apogee for Competition Flight
Figure 49 - PDF of Lateral Distance for Competition Flight

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