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AUBURN UNIVERSITY STUDENT LAUNCH
Project Nova
211 Davis Hall
AUBURN, AL 36849
PRELIMINARY DESIGN REVIEW
NOVEMBER 3, 2017
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Table of Contents
Table of Contents ...........................................................................................................................2
List of Figures .................................................................................................................................7
List of Tables ................................................................................................................................10
Section 1: General Information ..............................................................................................14
Section 1.1: Team Information .............................................................................................14
Section 1.2: Adult Educators ................................................................................................14
Section 1.3: Safety Officer ...................................................................................................15
Section 1.4: Team Leader .....................................................................................................16
Section 1.5: Project Organization .........................................................................................16
Section 1.6: NAR/TRA Sections ..........................................................................................19
Section 2: Summary of CDR Report ......................................................................................19
Section 2.1: Team Summary.................................................................................................19
Section 2.2: Launch Vehicle Summary ................................................................................20
Section 2.3: Payload Summary .............................................................................................21
Section 3: Changes Made Since PDR .....................................................................................21
Section 3.1: Vehicle Changes ...............................................................................................21
Section 3.2: Payload Changes...............................................................................................21
Section 3.3: Project Plan Changes ........................................................................................23
Section 4: Launch Vehicle .......................................................................................................23
Section 4.1: Mission Statement ............................................................................................24
Section 4.2: System Level Design Review ...........................................................................24
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Section 4.2.1: Structure ..........................................................................................................25
Section 4.2.2: Propulsion .......................................................................................................28
Section 4.2.3: Aerodynamics .................................................................................................31
Section 4.3: Dimensional Drawings .....................................................................................34
Section 4.4: Design Integrity ................................................................................................40
Section 4.4.1: Fin Shape and Style ........................................................................................40
Section 4.4.2: Materials .........................................................................................................41
Section 4.4.3: Assembly Procedures ......................................................................................43
Section 4.5: Mass Statement .................................................................................................44
Section 5: Subscale Flight Results ..........................................................................................45
Section 5.1: Flight Data ........................................................................................................45
Section 5.2: Scaling Factors .................................................................................................46
Section 5.3: Subscale Analysis .............................................................................................46
Section 5.4: Impact on Full-Scale Design ............................................................................48
Section 6: Recovery System Design ........................................................................................49
Section 6.1: Structural Elements...........................................................................................49
Section 6.2: Materials ...........................................................................................................51
Section 6.3: Ejection .............................................................................................................54
Section 6.4: Parachutes .........................................................................................................58
Section 6.5: Altimeters .........................................................................................................62
Section 7: Mission Performance Predictions .........................................................................65
Section 7.1: Simulations .......................................................................................................65
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Section 7.1.1: Motor Thrust Curve ........................................................................................69
Section 7.1.2: Component Weights .......................................................................................70
Section 7.1.3: Stability ...........................................................................................................71
Section 7.1.4: Computational Fluid Dynamics ......................................................................72
Section 7.2: Kinetic Energy ..................................................................................................75
Section 7.3: Drift ..................................................................................................................76
Section 7.4: Simulation Verification ....................................................................................79
Section 8: Safety .......................................................................................................................79
Section 8.1: Pre-Launch Day Items Checklist ......................................................................79
Section 8.2: Preassembly Checklists ....................................................................................81
Section 8.2.1: Recovery .........................................................................................................81
Section 8.2.2: Altitude Control ..............................................................................................82
Section 8.2.3: Body ................................................................................................................82
Section 8.2.4: Rover ...............................................................................................................83
Section 8.2.5: Engine .............................................................................................................84
Section 8.3: Launch Vehicle Assembly and Check ..............................................................84
Section 8.4: Launcher Setup and Launch Procedure ............................................................86
Section 8.4.1: Launcher Setup ...............................................................................................86
Section 8.4.2: Launch Procedure ...........................................................................................87
Section 8.5: Post-flight Inspection ........................................................................................88
Section 8.6: Personnel Safety Hazards .................................................................................89
Section 8.7: Environmental Effects ......................................................................................93
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Section 8.7.1: Effects on Environment ..................................................................................93
Section 8.7.2: Effects of Environment ...................................................................................94
Section 8.8: Failure Modes ...................................................................................................95
Section 9: Rover .....................................................................................................................114
Section 9.1: Mechanical Design .........................................................................................114
Section 9.1.1: Body ..............................................................................................................115
Section 9.1.2: Movement .....................................................................................................118
Section 9.1.3: Solar Panel Deployment System (SPDS) .....................................................120
Section 9.2: Electrical Design .............................................................................................123
Section 9.2.1: Microcontroller .............................................................................................124
Section 9.2.2: Communication Method ...............................................................................124
Section 9.2.3: Orientation ....................................................................................................126
Section 9.2.4: Power ............................................................................................................127
Section 10: Altitude Control Module .....................................................................................128
Section 10.1: Drag Plates ......................................................................................................129
Section 10.1.1: Internal Plate Drag System (IPDS) .............................................................129
Section 10.1.2: Wall Armed Fin-Lattice Elevator (WAFLE) ..............................................131
Section 10.1.3: Spherical Joint Actuator (SJA) ...................................................................133
Section 10.1.4: System Comparison ....................................................................................134
Section 10.2: Drag Plate Deployment ..................................................................................135
Section 10.3: Components ....................................................................................................136
Section 10.3.1: Controller ....................................................................................................136
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Section 10.3.2: Motor ..........................................................................................................141
Section 10.3.3: Electronics ..................................................................................................142
Section 10.3.4: Interfaces .....................................................................................................143
Section 10.3.5: Precision of Instrumentation .......................................................................143
Section 10.4: Dimensional Drawings ...................................................................................144
Section 11: Project Plan ..........................................................................................................144
Section 11.1: Testing ............................................................................................................144
Section 11.1.1: Rover Battery and Motor Test (AU 4.4, 4.5) ..............................................145
Section 11.1.2: Recovery and Altitude Control Battery Tests (AU 3.1, 6.9) ......................147
Section 11.1.3: Full-Scale and Subscale Separation Test (AU 3.2) .....................................148
Section 11.1.4: Tension Testing of Composite and 3D Printed Material (AU 2.1, 2.2) ......151
Section 11.1.5: 3-Point Bend Testing of Composite and 3D Printed Material (AU 2.1, 2.2,
6.7) .......................................................................................................................................154
Section 11.1.6: Compression Testing of Composite and 3D Printed Material (AU 2.1, 2.2,
6.7) .......................................................................................................................................158
Section 11.1.7: Rover Maneuverability (AU 4.2, 4.3, 4.6, 4.7) ...........................................160
Section 11.1.8: Altitude Control System (AU 6.4 – 6.8) .....................................................161
Section 11.2: Requirements Verification ..............................................................................163
Section 11.2.1: General Requirements .................................................................................163
Section 11.2.2: Vehicle Requirements .................................................................................169
Section 11.2.3: Recovery Requirements ..............................................................................179
Section 11.2.4: Deployable Rover Requirements ................................................................185
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Section 11.2.5: Safety Requirements ...................................................................................186
Section 11.3: Team Requirements ........................................................................................188
Section 11.3.1: General Requirements .................................................................................188
Section 11.3.2: Vehicle Requirements .................................................................................189
Section 11.3.3: Recovery Requirements ..............................................................................190
Section 11.3.4: Deployable Rover Requirements ................................................................190
Section 11.3.5: Safety Requirements ...................................................................................192
Section 11.3.6: Altitude Control Requirements ...................................................................193
Section 11.4: Budget .............................................................................................................195
Section 11.5: Funding Plan ...................................................................................................199
Section 11.6: Timeline ..........................................................................................................200
List of Figures
Figure 1.1: Team Organization Chart ........................................................................................... 17
Figure 3.1: Rover Body Comparison ............................................................................................ 22
Figure 3.2: Rover Domes vs Without ........................................................................................... 22
Figure 3.3: Rover Bay Comparison .............................................................................................. 23
Figure 4.1: Vehicle Rendering ...................................................................................................... 23
Figure 4.2: Motor Tube Rendering ............................................................................................... 28
Figure 4.3: Motor Thrust Curve .................................................................................................... 29
Figure 4.4: Motor Retention ......................................................................................................... 31
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Figure 4.5: Fin Rendering ............................................................................................................. 32
Figure 4.6: Booster Tube Dimensional Drawing .......................................................................... 34
Figure 4.7: Engine Block Dimensional Drawing .......................................................................... 35
Figure 4.8: Lower Booster Section Dimensional Drawing ........................................................... 36
Figure 4.9: Upper Body Tube Dimensional Drawing ................................................................... 37
Figure 4.10: Upper Section Assembly Dimensional Drawing...................................................... 38
Figure 4.11: Centering Ring Dimensional Drawing ..................................................................... 39
Figure 4.12: Bulkhead Dimensional Drawing .............................................................................. 39
Figure 4.13: Fin Dimensional Drawing ........................................................................................ 40
Figure 4.14: Fin Shapes ................................................................................................................ 41
Figure 5.1: Subscale Flight Data ................................................................................................... 45
Figure 5.2: Initial Subscale Design ............................................................................................... 46
Figure 6.1: Redundant Jolly Logic System ................................................................................... 56
Figure 6.2: Jolly Logic Chute Release .......................................................................................... 57
Figure 6.3: Gore Template ............................................................................................................ 60
Figure 6.4: Parachute Template .................................................................................................... 60
Figure 6.5: Altus Metrum TeleMega Altimeter ............................................................................ 63
Figure 6.6: Altus Metrum TeleMetrum Altimeter ........................................................................ 63
Figure 6.7: Altimeter Block Diagram ........................................................................................... 64
Figure 7.1: OpenRocket Model..................................................................................................... 65
Figure 7.2: Altitude Vs. Time ....................................................................................................... 67
Figure 7.3: Velocity Vs. Time ...................................................................................................... 68
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Figure 7.4: Acceleration Vs. Time ................................................................................................ 69
Figure 7.5: Motor Thrust Curve .................................................................................................... 70
Figure 7.6: Stability Vs. Time....................................................................................................... 72
Figure 7.7: Nosecone Meshing ..................................................................................................... 74
Figure 7.8: Tail-fin Meshing ......................................................................................................... 74
Figure 9.1: Rover Overview “top side” ...................................................................................... 115
Figure 9.2: Rover Overview “bottom side” ................................................................................ 115
Figure 9.3: First Iteration of the Rover Body “top side” ............................................................ 116
Figure 9.4: First Iteration of the Rover Body “bottom side” ...................................................... 116
Figure 9.5: Rover Body Dimensions .......................................................................................... 118
Figure 9.6: Rover Wheels and Tread .......................................................................................... 118
Figure 9.7: MATLAB Script....................................................................................................... 119
Figure 9.8: Rover Motor Inset .................................................................................................... 120
Figure 9.9: SPDS Stowed vs Deployed ...................................................................................... 121
Figure 9.10: SPDS Deployed Position Overview ....................................................................... 121
Figure 9.11: Solar Panel Tray, Motor, and Gear......................................................................... 123
Figure 9.12: Rover Electrical Layout.......................................................................................... 124
Figure 9.13: X-Bee Module ........................................................................................................ 125
Figure 9.14: X-Bee Shield Module on Arduino .......................................................................... 125
Figure 9.15: X-Bee Setup ........................................................................................................... 126
Figure 9.16: Rover Battery Setup ............................................................................................... 128
Figure 10.1: IPDS System Concept ............................................................................................ 130
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Figure 10.2: IPDS Bottom-Up View .......................................................................................... 130
Figure 10.3: WAFLE .................................................................................................................. 132
Figure 10.4: Spherical Joint Actuator ......................................................................................... 133
Figure 10.5: SJA External Fairing .............................................................................................. 134
Figure 10.6: Arduino Uno Microcontroller ................................................................................ 137
Figure 10.7: IMU Breakout......................................................................................................... 137
Figure 10.8: AndyMark NeveRest 40 DC Motor ....................................................................... 141
Figure 11.1: Subscale Separation Test ........................................................................................ 150
Figure 11.2: Epoxy - Carbon Fiber Tension Test Results .......................................................... 152
Figure 11.3: Epoxy - Fiberglass Tension Test Results ............................................................... 153
Figure 11.4: Onyx Tension Test Results..................................................................................... 154
Figure 11.5: Epoxy - Carbon Fiber Bend Test Results ............................................................... 156
Figure 11.6: Epoxy - Fiberglass Bend Test Results .................................................................... 157
Figure 11.7: Onyx Bend Test Results ......................................................................................... 157
Figure 11.8: Spending Comparison ............................................................................................ 198
Figure 11.9: Fall Timeline .......................................................................................................... 200
Figure 11.10: Spring Timeline .................................................................................................... 201
List of Tables
Table 1.1: Team Members ............................................................................................................ 17
Table 2.1: Team Information ........................................................................................................ 19
Table 2.2: Mentor Information ..................................................................................................... 20
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Table 2.3: Launch Vehicle Information ........................................................................................ 20
Table 4.1: Section Lengths ........................................................................................................... 25
Table 4.2: Motor Specifications .................................................................................................... 29
Table 4.3: Fin Dimensions ............................................................................................................ 32
Table 4.4: Manufacturing Schedule .............................................................................................. 43
Table 4.5: Mass Estimations ......................................................................................................... 44
Table 5.1: Subscale Launch Data vs Simulation .......................................................................... 47
Table 6.1: Pugh Chart of Carbon Fiber vs. Fiberglass Material for the BAE............................... 50
Table 6.2: Parachute Materials Pugh Chart .................................................................................. 51
Table 6.3: Comparison of Paracord, Tubular Nylon, Kevlar ........................................................ 53
Table 6.4: Comparison of U-bolts and Eye-bolts ......................................................................... 54
Table 6.5: Comparison of Black Powder and CO2 ....................................................................... 54
Table 6.6: Parachute Shape Pugh Chart ........................................................................................ 59
Table 6.7: Parachute Dimensions ................................................................................................. 61
Table 7.1: Flight Simulation Data (Wind = 0 mph) ...................................................................... 66
Table 7.2: Component Weights .................................................................................................... 70
Table 7.3: CFD Drag Coefficient Results ..................................................................................... 75
Table 7.4: Drift Calculations for Upper Section ........................................................................... 78
Table 7.5: Drift Calculations for Lower Section .......................................................................... 78
Table 10.1: Drag System Comparison ........................................................................................ 134
Table 10.2: Arduino Options ...................................................................................................... 136
Table 10.3: Microcontroller Data ............................................................................................... 138
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Table 10.4: Normalization of Data ............................................................................................. 138
Table 10.5: Trade Study Weighting Factors ............................................................................... 139
Table 10.6: Sensor Trade Study (Normalized Values) ............................................................... 140
Table 10.7: Sensor Trade Study (Weighted Values) .................................................................. 140
Table 10.8: AndyMark NeveRest 40 .......................................................................................... 141
Table 10.9: Uxcell 40 RPM 24V Motor (Alternative) ................................................................ 142
Table 10.10: New Guanlian RE 40 (Alternative) ....................................................................... 142
Table 11.1: General Requirements Verification ......................................................................... 163
Table 11.2: Vehicle Requirements Verification ......................................................................... 169
Table 11.3: Recovery Requirements Verification ...................................................................... 179
Table 11.4: Deployable Rover Requirements Verification ......................................................... 185
Table 11.5: AU General Requirements ....................................................................................... 188
Table 11.6: AU Vehicle Requirements ....................................................................................... 189
Table 11.7: AU Recovery Requirements .................................................................................... 190
Table 11.8: AU Rover Requirements.......................................................................................... 190
Table 11.9: AU Altitude Control Requirements ......................................................................... 193
Table 11.10: Vehicle Costs ......................................................................................................... 195
Table 11.11: Recovery Costs ...................................................................................................... 196
Table 11.12: Rover Costs ............................................................................................................ 197
Table 11.13: Budget Allocation .................................................................................................. 198
Table 11.14: Funding Sources .................................................................................................... 199
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Section 1: General Information
Section 1.1: Team Information
General Team Information
Team Affiliation Auburn University
Mailing Address 211 Engineering Drive
Auburn, AL 36849
Title of Project Project Nova
Date of CDR January 12th, 2018
Experiment Option 2: Deployable Rover
Section 1.2: Adult Educators
Contact Information
Name Dr. Brian Thurow
Title Aerospace Engineering Department Chair, Faculty
Advisor
Email [email protected]
Phone 334-844-4874
Address 211 Davis Hall
Auburn, AL 36849
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Contact Information
Name Dr. Eldon Triggs
Title Lecturer, Aerospace Engineering, Mentor
Email [email protected]
Phone 334-844-6809
Address 211 Davis Hall
Auburn, AL 36849
Section 1.3: Safety Officer
Safety Officer – Contact Information
Name Corey Ratchick
Title Senior in Aerospace Engineering
Auburn University
Email [email protected]
Corey Ratchick will be the Safety Officer for the Auburn Student Launch team this year. It is his
third year on the team. His goal for the year is to provide more exhaustive checklists than the team
has had access to in the past in an attempt to minimize human error.
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Section 1.4: Team Leader
Student Team Lead – Contact Information
Name Tanner Straker
Title Senior in Aerospace Engineering
Auburn University
Email [email protected]
Phone 847-507-1193
Address 211 Engineering Dr.
Auburn, AL 36849
Tanner Straker will be the student team leader for this year’s competition team. This is Tanner’s
third year on the team. In the previous year, Tanner served as the Recovery team leader and
oversaw the successful recovery of the team’s rocket throughout the year and during the
competition flight. He enjoys long walks across launch sites and sewing parachutes. Tanner is level
one high power rocket certified through Tripoli Rocketry Association.
Section 1.5: Project Organization
The Auburn Student Launch team is broken into five major sub-teams: vehicle body design,
payload, electronic systems, testing, and recovery. Safety and educational engagement also exist
as sub-teams composed of students from the five primary groups. Each sub-team has at least one
member dedicated to identifying safety concerns and acting as the point of contact (POC) for the
safety officer. In addition, all members of the Auburn Student Launch Team are required to
participate in at least one educational engagement event and each event has its own coordinator,
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all of whom are working members of other sub-teams. Figure 1.1: Team Organization Chart shows
the hierarchy of project management with all teams reporting to their team leads, the student
project manager, and the safety officer, who in turn report to the adult educators.
Figure 1.1: Team Organization Chart
Table 1.1: Team Members
Name Role Team
Dr. Eldon Triggs Adult Educator Overall Management
Dr. Brian Thurow Adult Educator Overall Management
Tanner Straker Project Manager Overall Management
Corey Ratchick Safety Officer Overall Management/Safety
Reilly B. Team Lead Vehicle Body
Tanner O. Team Lead Electronic Systems
David T. Team Lead Payload
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Ben C. Team Lead Recovery
Bryce G. Team Lead Testing
Kate M. Team Lead Education
Nick R. Team Member Testing
Zac B. Team Member Education
Icis M. Team Member Education
Jaylene A. Team Member Education
Jake R. Team Member Safety
Rhett R. Team Member Safety
Ruth A. Team Member Safety
Sydney F. Team Member Safety
Zach W. Team Member Recovery
Paul L. Team Member Recovery
Omkar M. Team Member Recovery
Jaysal S. Team Member Recovery
Bill M. Team Member Vehicle Body
Adam B. Team Member Vehicle Body
CJ L. Team Member Vehicle Body
Matthew D. Team Member Vehicle Body
Logan J. Team Member Vehicle Body
Anthony G. Team Member Vehicle Body
Victor D. Team Member Vehicle Body
Rhett R. Team Member Payload
Stephen S. Team Member Payload
Zach S. Team Member Payload
Kevin H. Team Member Payload
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Matthew W. Team Member Payload
Landon B. Team Member Payload
Salaar K. Team Member Electronic Systems
Andrew R. Team Member Electronic Systems
Ruth A. Team Member Electronic Systems
Michael C. Team Member Electronic Systems
Austen L. Team Member Electronic Systems
Matthew H. Team Member Electronic Systems
Section 1.6: NAR/TRA Sections
The Auburn Student Launch team is planning on attending launches hosted by Southern Area
Rocketry (SoAR) at Phoenix Missile Works (PMW) in Sylacauga Alabama (NAR Section #571).
The team also occasionally attends launches with the Music City Missile Club (MC2) in
Manchester, Tennessee (NAR Section #589) and the South Eastern Alabama Rocket Society
(SEARS) in Samson, Alabama (NAR Section #572/TRA Prefect 38). We will also be partnering
with SEARS through Christopher Short. Chris provides technical experience and serves as a
reliable rocketry vendor for the team.
Section 2: Summary of CDR Report
Section 2.1: Team Summary
Table 2.1: Team Information
Team Information
Team Name Auburn University Student Launch (AUSL)
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Mailing Address 211 Engineering Drive
Auburn, AL 36849
Project Name Project Nova
Table 2.2: Mentor Information
Mentor Information
Mentor Name Eldon Triggs
TRA Number 12159
Certification Level 2
Contact Information Email: [email protected]
Phone: 334-844-6809
Section 2.2: Launch Vehicle Summary
Table 2.3: Launch Vehicle Information gives the basic details of the launch vehicle. More
information regarding the launch vehicle can be found in Section 4: Launch Vehicle of this report.
Table 2.3: Launch Vehicle Information
Launch Vehicle Information
Total Length 108 in.
Estimated Mass 40.5 lbm
Motor Selection Aerotech L1420R
Recovery System Double Separation, Dual Deployment
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Rail Size 12 ft. 1515
Section 2.3: Payload Summary
Auburn University's Student Launch team will be completing the deployable rover experiment.
After landing under parachute, the rover that will be completely housed inside the rocket will be
remotely deployed from the rocket. It will autonomously travel at least five feet from the rover and
deploy foldable solar panels. A dual-tread design deployed from an orientation-independent
containment bay has been chosen to minimize the possibility of error and risk within the system
Section 3: Changes Made Since PDR
Section 3.1: Vehicle Changes
The vehicle design has not changed substantially since PDR. The team has stopped considering a
grid fin design for the tail fins and will continue with a clipped delta fin planform. This decision
was made due to the performance of clipped delta fins and in favor of focusing efforts on the
isogrid weave body tube structure. Mass estimates have also been adjusted due to experience with
gained from constructing a subscale vehicle.
Section 3.2: Payload Changes
Due to extensive trade studies done during PDR, minimal changes have been made to the overall
design of the rover. However, based on further research, some components have been changed or
removed entirely. The following figures show changes made to the mechanical design of the rover.
The body design was changed to provide more space for electronics, the treads had an external
dome removed that was deemed unnecessary, and the rover bay was modified to account for
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different orientations upon landing. Some electronics that had scored comparably in the PDR trade
studies have also been substituted for the team’s initial choices due to experience with them.
Figure 3.1: Rover Body Comparison
Figure 3.2: Rover Domes vs Without
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Figure 3.3: Rover Bay Comparison
Section 3.3: Project Plan Changes
The team has secured additional support from Lockheed Martin, so the budget and funding plan
have been updated accordingly. Several tests, including a subscale flight, have also been completed
and this is reflected in the team’s updated timeline.
Section 4: Launch Vehicle
Figure 4.1: Vehicle Rendering
The Auburn University Student Launch team (AUSL) is determined to design and manufacture an
effective and unique launch vehicle. Learning from past experiences and Auburn’s history with
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the competition, AUSL has re-examined every component of the launch vehicle. AUSL requires
the highest quality of all components to reach the goals set by NASA in this year’s competition.
Section 4.1: Mission Statement
The mission of the AUSL launch vehicle is to design and construct a lightweight, safe, and reliable
vehicle. These motivators will ensure a vehicle that will allow for successful launches and the
flexibility to adapt the design to various experimental payloads. Three driving requirements from
the NASA Student Launch Handbook have been chosen to guide the success of this mission:
1. The vehicle must have an apogee of 5280 feet AGL
2. The vehicle must be recoverable and reusable.
3. The vehicle will not exceed Mach 1 during flight.
These requirements serve as minimum standards set to achieve mission success. AUSL has also
determined three team-specific requirements to drive the design of the launch vehicle:
1. The vehicle must maintain stability of 2 or more calibers.
2. The vehicle must have a factor of safety of at least 2.
3. Structural components must remain attached to launch vehicle.
These team requirements serve as additional constraints to assure the vehicle design is compatible with
the mission.
Section 4.2: System Level Design Review
The vehicle has been designed to satisfy mission requirements set forth by NASA in the 2017-
2018 NASA Student Launch Handbook, as well as requirements set by the team. The vehicle
design must ensure adequate space for avionics, payload equipment, and electronics. These
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systems are crucial to the success of the mission. The vehicle design is also heavily driven by
manipulating weight and length to control altitude and stability. These factors determine the
success of the flight itself. The vehicle design is separated into three major divisions: structure,
propulsion and aerodynamics. These three divisions are all vital to the success of the flight and
recovery of the launch vehicle, as well as the success of the onboard experiments.
Section 4.2.1: Structure
The structure of the launch vehicle has been designed to be able to withstand the forces the rocket
will experience during operation. The launch vehicle body must be strong enough to maintain
stable flights, while accommodating all other subsystems, and ensuring they have adequate space
and protection. The design of the structure requires heavy tradeoffs between strength, space, and
weight.
The total length of the rocket is 108.0 inches. Component lengths are shown in Table 4.1: Section
Lengths.
Table 4.1: Section Lengths
Body Tubes:
Section Length (in.)
Nose Cone 16
Avionics Section 47
Rover Section 16
Airbrake Section 10
Booster Section 19
Total 108
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The body tubes house all subsystems of the launch vehicle. These tubes comprise a majority of the
vehicle body surface exposed to the airflow. Therefore, the aerodynamic properties of the body
tubes are directly related to the altitude gained by the vehicle. Additionally, as the largest structure
in the rocket, the body tubes represent the largest collection of mass in the rocket, except for the
motor. To ensure mission success, it is critical to select and design body tubes that can survive the
stresses of high-powered flight while remaining light enough to achieve the mission altitude.
The body tubes will be constructed using carbon fiber braiding, a process that involves taking
individual strands of carbon fiber and stitching them into a tightly-wound braid. The carbon fiber
braids that are produced will be formed into an isogrid structure around a 6-inch mandrel, which
will be pre-wrapped with a layer of carbon fiber already to make bonding the internal systems to
the body tubes easier. Isogrid structures are a lighter alternative to using a solid tube structure. For
aerodynamic purposes, a thin layer of fiber glass will be cured and wrapped around the structure
to allow for a smooth aerodynamic skin. By giving the structure this skin, the result is a lightweight,
aerodynamic body. Using this wrapped isogrid method, the mass of the body tubes will be
decreased by approximately 20 to 30 percent less than if the tubes were constructed using only
filament wound carbon fiber, while also maintaining the same compressive strength properties as
a carbon fiber tube. This mass reduction was confirmed using tube samples constructed by team
members using final production methods.
Couplers:
The couplers serve as a joint between two body tube sections. The couplers must be able to
withstand forces experienced during rocket ascent to keep the structure of the body attached. The
upper body tube will be attached to the booster section with 4 aluminum bolts. The team has
decided to use fiber glass to create the couplers. This choice of material reduces risks which can
lead to separation of the upper body from the booster section in mid-flight. With the trade-off of
an increase in mass and more difficult construction for strength, fiber glass was considered to be a
very safe and reliable option. Fiber glass also has the additional benefit of being non-conductive,
thus it will be ideal for making the coupler which will be holding our electronics, the avionics bay.
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This will reduce the issues the team has had in the past when it comes to being able to communicate
to the electronics inside the vehicle, such as GPS systems.
Ballast Tank:
The ballast tank is used to hold additional mass if balance corrections must be made. The design
allows for easy mass addition and reduction as needed to account for variations in mass predictions
and launch day conditions. The tank will be placed forward of the grid fin section, near the CG
location, and is secured to the launch vehicle body by two aluminum pins. As the tank will not be
subjected to a large force, the team is confident that the pins will hold the tank securely without
fear of a shear failure. The tank will be constructed using high impact polystyrene (HIPS) on a
TAZ4 3D printer.
Bulkheads:
Bulkheads are typically flat plates used to increase the structural strength of a rocket. They are also
used to create airtight spaces and to divide the body into separate compartments. In rockets, they
are commonly used to separate payload bays and to mount equipment for avionics and payloads.
For rockets similar in size to the Project Nova rocket, the material used varies from fiberglass to
plywood to carbon fiber. The bulkheads for this rocket will be made from pre-impregnated carbon
fiber. This was chosen due to the simplicity of manufacturing with pre-impregnated carbon fiber.
The interior diameter for the circular cross-sectional rocket will be 6 inches and the bulkheads are
designed to fit perfectly into this size. All bulkheads for this rocket will be 0.25 inches thick.
Centering Rings:
The purpose of the centering rings is to center a smaller cylindrical body or tube inside a tube of a
larger diameter. In the case of high powered model rocketry, centering rings can be used as an
engine block in motor mounts. The Project Nova rocket will be using three centering rings. These
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centering rings are located in the engine tube and serve to attach to the fin set and to attach to the
motor retention. The centering rings are made of carbon fiber and manufactured using the
Computer Numerical Control (CNC) machine at Auburn University Aerospace Design Lab due to
the availability and the teams experience with using carbon fiber. The centering rings have an outer
diameter of 6 inches with an inner diameter of 3 inches. The thickness of each ring is approximately
0.25 inch. The centering rings have a mass of 3.65 oz., determined from sample pieces.
Section 4.2.2: Propulsion
Figure 4.2: Motor Tube Rendering
Motor:
The motor selected for the competition is the Aerotech L1420R. This is the same motor that was
used in PDR, and after minor modifications were made to the rocket, it still gave us the needed
altitude for the rocket. The specifications are listed below in Table 4.2: Motor Specifications.
Additionally, the thrust curve for this motor is shown in Figure 4.3: Motor Thrust Curve.
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Figure 4.3: Motor Thrust Curve
This motor was chosen based on OpenRocket simulations, as it provides the roughly 8-to-1 thrust-
to-weight ratio desired for stable and predictable flight.
In addition, as shown in the motor thrust curve above, the motor achieves a higher than average
thrust after approximately one-quarter second, thus reaching the required 8-to-1 thrust ratio in
about one-quarter second. Based on OpenRocket simulations, the motor provided an apogee of
5866 feet with a max acceleration of 275 ft/s^2 which delivers a max velocity of 723 ft/s or close
to Mach = 0.65.
Table 4.2: Motor Specifications
Motor Specifications
Manufacturer Aerotech
Motor Designation L1420R
Diameter 2.95 in
Length 17.4 in
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Total Impulse 1038 lb·sec
Total Motor Weight 10.1 lbm
Propellant Weight 5.69 lbm
Propellant Type Solid
Average Thrust 326 lbf
Maximum Thrust 374 lbf
Burn Time 3.18 sec
Motor Tube:
To contain the motor on the rocket, a carbon fiber motor tube is being used. The motor tube will
be made by braiding carbon fiber strands and then filament wound around a mandrel that is the
same diameter of the motor. The 3D braided carbon fiber material was chosen for its strength
relative to its weight when compared to a solid tube. Basalt fiber was considered to be used for the
motor tube for its high heat resistance properties, but the team decided the weight of the basalt,
which was approximately 50% heavier when compared with the carbon fiber was not worth the
tradeoff. The tube will be 0.1-inch-thick and is designed to fit around an Aerotech L1420R motor.
With these specifications, the motor tube will be ideal for the rocket.
To mount the motor tube, three centering rings will be epoxied to the outer diameter of the motor
tube and the inner diameter of the lower section tube. The epoxy will be a 24-hour epoxy, which
will create a permanent bond between the components. A bulk plate will be epoxied forward of
the motor tube. This is to provide extra strength to hold the motor in place as well as separate the
motor from the internal components of the rocket.
Motor Retention:
The purpose of the motor retention system is to secure the rocket motor during launch and flight
and to be easily removable for subsequent flights. The team has chosen a commercial bought
Aeropack motor retention system, Figure 4.4: Motor Retention. This is a simple system with two
31
components. One component will bolt directly into a centering ring, using aluminum bolts. The
other component threads onto the part that is bolted onto the structure. This allows for a fast
replacement of a used motor. The team chose a commercial motor retention system due to past
reliability and to avoid the time requirements of designing and manufacturing a custom system.
Figure 4.4: Motor Retention
Section 4.2.3: Aerodynamics
The aerodynamics system requires the rocket remain stable during flight. The placement and
design of the aerodynamic surfaces determines the center of pressure along the length of the rocket.
Fins:
The stability of the rocket is controlled by the fins. The primary purpose of the fins is to keep the
center of pressure aft of the center of gravity. The greater drag on the fins will keep them behind
the upper segments of the vehicle, thus allowing the rocket to fly straight along the intended flight
path. They are also helpful in minimizing the chances of weather-cocking. Fins serve as an ideal
addition to the vehicle body as they are lightweight and easy to manufacture using the CNC
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machine. A clipped delta planform has been selected for the fins. Four fins will be machined from
0.2-inch-thick carbon fiber flat plates. A rendering of the fin design is shown in Figure 4.5: Fin
Rendering.
When attached, the trailing edge of each fin will be located slightly forward of the end of the body
tube. This design feature will theoretically provide some impact protection for the fins when the
rocket hits ground. Carbon fiber of 1.03 oz/in3 density has been selected as the material due to its
stiffness, strength, and light weight. The stiffness and strength of carbon fiber reduces change of
fin flutter which increases the vehicles chance of success during flight. Each fin will have a surface
area of 54 in2 (summing both sides), making the fin surface area total equal to 216 in2. The total
component mass is 13.5 ounces. These dimensions provide the vehicle with a projected stability
of 2.1 calibers. This level of stability is close to ideal, as it is well above stable, yet still below
over-stable. Detailed fin dimensions are provided in Table 4.3: Fin Dimensions.
Table 4.3: Fin Dimensions
Fin Dimensions
Root Chord 6.25 in
Tip Chord 2.5 in
Figure 4.5: Fin Rendering
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Height 6 in
Sweep 3.68 in
Sweep Angle 31.5°
Thickness .2 in
Aero-elastic flutter has been considered as a potential failure mode for the rocket structure. At a
particularly high velocity, the air is no longer able to sufficiently dampen the vibrational energy
within the fin. At this flutter velocity, the first neutrally stable oscillations are experienced within
the wings.
The flutter velocity is directly reflective of the aero-elastic conditions of the structure/fin system.
The catastrophic flutter phenomenon results from coupling of aerodynamic forces creating a
positive feedback loop. The increase in either torsion or bending drives an infinitely looped
increase in the other motion. Since it is assumed that the fins are rigidly fixed and cantilevered to
an infinitely stiff rocket body, the fin twist (torsion) and fin plunge (bending) are the only two
degrees of freedom.
Once this flutter velocity is exceeded, the air, inversely, amplifies the oscillations and significantly
increases the energy within the respective fin. As velocity increases, the fin twist and plunge are
no longer damped. At this velocity, known as the divergent speed, one degree of freedom usually
diverges while the other remains neutral. Structural failure usually occurs at or just above this
velocity. Due to certain failure of the structure associated with potential aero-elastic flutter, the
flutter velocity is applied to the design as a “never-to-exceed” parameter.
There are various ways to minimize the chances of experiencing fin flutter. Increasing fin retention
by strengthening the joints between the fins and rocket body is one way to supplement system
stability. Furthermore, additional layers of carbon fiber and epoxy applied to portions of the fins
as well as the joints should provide extra defense against aero-elastic flutter.
Nosecone:
34
The coefficient of drag affects the overall performance of the rocket in flight. The goal for the team
was to select a nose cone shape with a low drag coefficient to maximize performance. Utilizing
the software OpenRocket, the four cone types were compared using the already chosen dimensions
for the rocket. The team has decided to increase the fineness ratio of the nosecone, increasing to a
near 5-to-1 ratio versus the initially designed near 3-to-1 ratio. This change has been made due to
the increased stability that a 5-to-1 provides, and the team can easily acquire these nose cones by
purchasing them from several vendors. The material for the nose cone has also changed to being
fiber glass, as this is the most common material sold by these vendors.
Section 4.3: Dimensional Drawings
Figure 4.6: Booster Tube Dimensional Drawing
35
Figure 4.7: Engine Block Dimensional Drawing
36
Figure 4.8: Lower Booster Section Dimensional Drawing
37
Figure 4.9: Upper Body Tube Dimensional Drawing
38
Figure 4.10: Upper Section Assembly Dimensional Drawing
39
Figure 4.11: Centering Ring Dimensional Drawing
Figure 4.12: Bulkhead Dimensional Drawing
40
Figure 4.13: Fin Dimensional Drawing
Section 4.4: Design Integrity
Section 4.4.1: Fin Shape and Style
The three most common planforms are clipped delta, trapezoidal and elliptical, as shown in Figure
4.14: Fin Shapes. In most situations, the elliptical fin design is the optimal design choice for model
rockets. This is due to their lower drag and superior lift forces compared to various other fin
designs, allowing for the highest altitudes. The downside, however is a far lower stability, and as
previously mentioned, the drag is needed for better stability and a solid flight path. Most
importantly, there is little room for error when it comes to the construction of an elliptical fin. So,
if all four fins are not identical in shape and weight then this will lead to undesirable flight results.
Another variable the team considered while designing the fins is fin flutter. The excessive amounts
of fin flutter in elliptical shaped fins only causes more instability. So, even with the advantage of
an optimal altitude, it does not justify the several other complications that will arise during the
41
construction process. The clipped-delta fin, being a variation of the trapezoidal fin, gives excellent
stability, offering a slight stability advantage over the plain trapezoidal due to having more surface
area aft of the chord of the fins midpoint. The simplicity of the clipped-delta fin design also allows
for an easy and accurate construction process. Though elliptical fins do give a slight advantage
when it comes to altitude, the team has decided to implement the clipped-delta fin design due to
its increased stability.
Figure 4.14: Fin Shapes
The fins will be manufactured from the same carbon fiber plates as the bulkheads and centering
rings. The same data used to verify the bulkheads and centering rings will be used to ensure the
fins can withstand any inflight or landing forces.
To verify that the size and shape of the fins allows for stable flight, simulations were conducted.
There have also been two subscale flights which further verified the simulation data. Multiple full
scale test flights will be performed to visually verify no anomalies are present on the fins during
flight.
Section 4.4.2: Materials
Body Tubes:
42
The structural tubes of the launch vehicle are going to be constructed using a 3D braided carbon
fiber isogrid structure. As this is something the team has not done in past years, structural data will
need to be collected for this structure. To do this, using the same material and manufacturing
method, a test sample will be made consisting of an equal diameter of the tubes that will be used
on the launch vehicle. This sample will then be placed into a load cell to determine the maximum
load of the structure. This will allow us to determine that the structure is capable of safely
completing the mission. The structure will experience a maximum of 300 lbs during flight, to meet
the factor of safety requirements the tube structure must fail at or above 600 lbs of force during
testing.
Bulkheads and Centering Rings:
The bulkheads and centering rings are manufactured by cutting a flat carbon fiber plate with a
CNC machine. To verify these components can handle the expected loads, sample pieces of the
carbon fiber have been made. These samples were manufactured using the same material that the
bulkheads and centering rings will be made of. The samples were placed in a three-point bending
test as well as a tensile stress test.
Coupler:
To verify the coupler functions correctly, ground tests of the separation will be performed. Once
proven on the ground, a subscale flight test using this coupler component will be used.
Ballast Tank:
By running simulations, the team is able to determine where the center of gravity is located. Once
the launch vehicle is manufactured a final simulation will be run using real component weights. If
the center of gravity is not where initially predicted, the ballast tank will be used to correct the
location. Throughout the project, this will be re-examined to ensure stable flight.
43
Section 4.4.3: Assembly Procedures
Manufacturing of the vehicle generally takes two weeks to produce and assemble the components.
To account for this the team plans to start manufacturing three weeks prior to any scheduled
launches. This allows for one extra week if any problems arise during the manufacturing process.
The typical manufacturing schedule can be seen in Table 4.4: Manufacturing Schedule.
Table 4.4: Manufacturing Schedule
Week Events Percentage of
Completion
1
Manufacturing of Major
Components Such as body
tubes, fins, centering rings,
etc.
50%
2
Begin assembly of
subsystems such as the
booster section and the fin
assembly.
90%
3 Assemble completed
rocket 100%
Manufacturing body tubes using braided structures is a very time-consuming process. The team is
in the process of creating the material required to make these body tubes, and the first four weeks
of January will be used to manufacture the tube structures, both braided and non-braided. These
tubes are the most time-consuming component to manufacture and the event of a crash would have
negative effects on the team’s timeline. To mitigate the effects of a total loss crash, six tube
sections and three motor tube sections will be produced during this time, which will allow for the
construction of three full scale rockets, two with braided body tubes and one with non-braided
body tubes.
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Several flat plates of carbon fiber have been produced at various thicknesses. The plates will be
placed in a CNC router to be shaped into flat components. These components include fins, bulk
plates and centering rings.
Section 4.5: Mass Statement
The mass of the rocket and all its subsystems was calculated using optimal mass calculations from
OpenRocket. In addition to using final masses from last year as a basis, a brick sample of carbon
fiber was created to have an accurate density measurement since most of the parts will be
manufactured using carbon fiber. This density test is exceedingly important given the method of
mass estimation. Since construction methods vary drastically from each manufacturer, as well as
different resin and cloth systems varying, it is highly important to get an accurate model of the
density.
Having determined an accurate density for the carbon fiber of the rocket, and the structure of the
rocket being the most significant portion of the weight of the structures of the rocket, the team
used estimates from last year’s rocket to determine the initial size estimate of the rest of the
subsystem components. The team believes that this model presents an estimate that is sufficient.
As the program develops, the model will attain a higher and higher accuracy in its simulation.
Table 4.5: Mass Estimations
Section Mass (lb) Percentage
Structure 15.4 38.02%
Motor 10.1 24.93%
Rover 6 14.81%
Recovery 7 17.28%
Airbrake 2 4.94%
Total 40.5 100%
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Section 5: Subscale Flight Results
Section 5.1: Flight Data
Flight data for the subscale launch can be seen below in Figure 5.1: Subscale Flight Data. This
figure represents the launch data of the full rocket and the recovery data of the upper section of the
rocket. The rocket reached an apogee of 2887 feet, had a maximum velocity of 525 feet per second,
and had a landing velocity of 23 feet per second. The data verifies that the vehicle and recovery
designs are safe and within the limits of the competition.
Figure 5.1: Subscale Flight Data
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Section 5.2: Scaling Factors
A 2/3 scale was applied to the vehicle diameter and approximately applied to the length of the
vehicle. The diameter of the rocket was the primary scaling factor as the diameter controls the area
of the vehicle perpendicular to the flow and as such has a significant impact on aerodynamic
performance. The length was approximated to a 2/3 scale to ensure integrity of the subscale, but
was not exact due to safety concerns. The team was unwilling to compromise the length of the
recovery sections or the booster section due to the potential hazards that could induce.
The scaling factor of 2/3 was chosen to accurately model the geometry of the vehicle and the forces
the full-scale will experience. The smaller size made the subscale easier to manufacture, but the
team decided that any smaller of a scaling factor could result in risk of the recovery system failing.
Section 5.3: Subscale Analysis
Figure 5.2: Initial Subscale Design
Multiple simulations were run before and after the subscale launch. Initial simulations provided a
predicted apogee of 4090 feet, and a static stability margin of 2.1 calibers. The simulated subscale
weighed 11.94 pounds. Although the motor selected for the flight was simulating a lower apogee
than desired, it was flown because it was readily available to the team and, of the alternatives, it
most closely simulated the maximum velocity the full-scale vehicle is expected to experience.
47
The team’s subscale launch took place on November 4th 2017 in Samson, AL under SEARS
(Southeast Alabama Rocketry). Upon arrival, the rocket was assembled and the altimeters and
recovery system were tested to verify integrity. Some issues were encountered due to lack of space
inside the recovery section of the rocket. This was due to not accounting for the size of space the
couplers would take up internally, as well as issues with bolts potentially catching the parachutes
and preventing deployment. Additionally, the team found that the rocket was four pounds over-
weight, bringing the total weight to approximately 16 pounds. The recovery issue was resolved by
carefully packing the parachutes. The rocket was flown slightly overweight and was successfully
launched and recovered. The launch data is shown above in Figure 5.1: Subscale Flight Data.
The flight data did not match the simulated data, but the team believes that this is primarily due to
the change in weight of the rocket. When inputting accurate weight values into OpenRocket, values
similar to those achieved in the flight test were outputted. The results for the simulation with 4
pounds added at approximately the CG, as well as changing the surface finish to increase the drag
coefficient values, are shown in Table 5.1: Subscale Launch Data vs Simulation. Changing the
surface type in OpenRocket resulted in the coefficient of dragging changing from .49 to .54. These
simulated values were then slightly below the flight data. The team performed a CFD analysis that
will be discussed later that yielded a coefficient of drag for the full-scale model of .53, and coupled
with the subscale flight data, that is the value used in the team’s most recent full-scale simulations
discussed in Mission Performance Predictions.
Table 5.1: Subscale Launch Data vs Simulation
OpenRocket
Simulation
Recorded Flight
Data Updated Simulation
Apogee (ft) 4090 2887 2867
Max Velocity (ft/s) 620 525 473
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Section 5.4: Impact on Full-Scale Design
The subscale flight tests determined that the design is very stable and will perform effectively.
Because of the complex aerodynamic shape of the rocket, simulations were not exactly accurate
at first, but with the subscale data the team is confident the current simulation iteration is far more
accurate than before. More tests will be run to further improve and verify the accuracy of the
simulation model. The design of the rocket will not change because of the latest subscale flight.
The major insight from the subscale launch came from the vast underestimates of the weight of
the different subsystem components. To ensure that this does not happen in the full-scale, each
subsystem will be re-examined to ensure that the proper weight is simulated for each section.
As for recovery, all components of the subsystem worked exactly as designed in the subscale test
of the Auburn Student Launch team’s rocket. The subscale was 2/3 the scale of the full-scale
rocket, giving it a diameter of 4 inches. The subscale used a nosecone of a similar design to the
full-scale rocket, so ejection of the nosecone provided data proving the efficacy of the ejection
system. The subscale, like the full scale, had three parachutes, a drogue deployed at apogee, an
upper main parachute deployed at 1000ft and a lower main parachute deployed at 750 ft. The
drogue parachute was a standard 31-inch diameter circular parachute and both main parachutes
were 17.5-inch diameter hemispherical parachutes. Black powder was used as the ejection system
for both the lower and upper parachutes in the subscale rocket. The avionics bay board was
modified slightly to fit into the approximately 3.75” interior diameter of the avionics bay, and
employed a more compressed assembly of altimeters and batteries than will be used in the full
scale. The board had one altimeter and its corresponding battery mounted on each side of the
carbon fiber avionics bay board. This helped save space and reduce interference between the
altimeters. Charge cups filled with black powder were mounted directly adjacent to the outsides
of both the upper and lower bulk plates of the avionics bay. Electronic matches attached to the
black powder charges were run through holes the bulk plates and inserted to the appropriate ports
in the altimeters. In summary, the recovery design will not change as a result of subscale
experience.
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Section 6: Recovery System Design
The Auburn Student Launch team is using an augmented dual-stage recovery system with a drogue
parachute deployed at apogee (target height of 5280 ft.), an upper main parachute being deployed
at 1000 ft. and a lower main parachute deployed at 750 ft. At apogee, the nose cone, drogue
parachute and upper main parachute are all ejected using redundant black powder charges. When
the upper main parachute is ejected though, it is held closed by the Jolly Logic Chute Release
System. At the beginning of the second event (at 1000 ft.), The Jolly Logic System will release
and deploy the upper main parachute with the rocket still in one piece. At the end of the second
event (at 750 ft.), a second set of redundant black powder charges will detonate and separate the
rocket, pushing the lower main parachute from its housing. Using this configuration, the entire
rocket will fall under a single drogue parachute from apogee until the second event occurs where
the rocket separates into two separate pieces and falls under two separate main parachutes. The
nose cone and drogue parachute will remain attached via shock cord to the upper section after
separation. This augmented dual-stage recovery system was chosen to reduce the likelihood of the
upper section drifting into and collapsing the lower main parachute after deployment.
Section 6.1: Structural Elements
The centerpiece of Auburn's recovery system is the Barometric Avionics Enclosure (BAE). Every
recovery subsystem is either attached to or contained inside the BAE. The BAE is formed by a 12-
inch-long cylinder of fiber glass. A comparison showing why fiber glass was chosen over carbon
fiber can be seen in a Pugh chart below in Table 6.1: Pugh Chart of Carbon Fiber vs. Fiberglass
Material for the BAE. As for the design of the BAE, there is one inner bulk plate attached inside
of the BAE that serves as the top cap of the avionics bay. Inside the avionics bay there are two sets
of rails to secure the avionics board. Both altimeters and their batteries are mounted to this board.
The bottom of the BAE is closed off by another bulk plate. The two bulk plates are linked by two
rods and secured by locking nuts on the outside of the BAE. Both bulk plates have a single open
hole to allow the ejection charge wires to run from the altimeters to their proper e-matches. Each
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hole has two charge wires that run through it to minimize the chance of the sensitive recovery
electronics being damaged by the pressurization that occurs when the black powder charges are
deployed.
Table 6.1: Pugh Chart of Carbon Fiber vs. Fiberglass Material for the BAE
Weight Fiberglass Carbon Fiber
Altimeter Signal
Disruption
3 3 1
Strength 2 2 3
Ease of
Manufacturing 1 3 2
Total 16 9
The BAE serves as the coupler between the upper and the lower parachute housings. Each section
is secured to the BAE with three machine bolts per section. Neither of these sections separate once
the rocket is assembled. On the outside of the BAE is a ring of the vehicle body tube taken from
the same tube the upper section is constructed from. This is done so the tube connections between
the upper parachute housing, the BAE, and the lower parachute housing are continuous and
smooth, minimizing the impact on the aerodynamic performance of the rocket due. This ring is the
only surface of the BAE that is on the outside of the rocket, so two key switches and two pressure
holes are located along this ring. The key switches located on the ring allow the team to externally
arm the altimeters while the rocket is assembled.
A single U-bolt is mounted to the top bulk plate of the BAE for the upper parachute assembly to
be mounted to. This location was chosen as it is the only point in the upper section where the
parachutes can be tethered to keep the rocket in an upright position and minimize the chance of
the parachute tearing when deployed. The lower main parachute is connected to a second U bolt
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mounted to the bottom of the Rover section. This is mounting location was chosen because it is
the only point where the lower parachutes can be mounted while also keeping the charge cups next
to the BAE, thus minimizing the potential for error with the e-matches and decreasing risk.
Section 6.2: Materials
The materials chosen to create the team's recovery system have a direct effect on the success of
the system. Specifically, the parachutes, shock cords, and bulkhead attachments must be strong
enough to withstand the recovery procedures and also be light enough to avoid hindering the
launch. Failure of the recovery system materials to survive the ejection charges would cause an
inadequate landing of the vehicle, and could severely damage the rover payload. The success of
recovery systems materials is essential for proper payload delivery, so they are chosen
appropriately. The chosen materials for the parachute, shock cord, and bulkhead attachment are
rip-stop nylon, paracord, and U-bolts, respectively.
The parachute must withstand ejection charges during the recovery process. Ideally, the chosen
materials for the parachute allow it to be sturdy to be used in multiple tests and launches. The
factors that should be used to assess the suitability of different parachute materials to allow for
recovery system requirement satisfaction are strength, durability, and weight. Rip-stop nylon was
chosen over cotton fabric as our parachute material for several reasons. Rip-stop nylon weights
about 2.75 oz. per square yard while cotton fabric weighs about 4.3 oz. per square yard. It also
has a tensile strength of 1500 psi while cotton fabric has a tensile strength of 400 psi. Additionally,
rip-stop nylon will stretch up to 40 percent of its length before breaking while cotton will stretch
up to 10 percent. The Pugh chart shown below in Table 6.2: Parachute Materials Pugh Chart
confirms the choice of rip-stop nylon for the parachute material.
Table 6.2: Parachute Materials Pugh Chart
Rip-stop Nylon Cotton fabric
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Strength 3 1
Durability 3 1
Weight 3 2
Totals 9 4
The shock cord must be able to survive the same events as the parachute. The materials that were
considered for shock cord and shroud lines are Paracord, 1-inch Tubular Nylon, and Kevlar.
Paracord is extremely light, strong for its weight, and does not take up much space. The only
possible problem with paracord is that it can only withstand 550 pounds. However, the other
benefits of paracord make it an ideal material for the drogue shroud lines, since the drogue will
not create much drag force and so will not put as much tension on the shroud lines.
Tubular nylon consists of a nylon tube which is made from exceptionally high strength material
which is both light and strong. Tubular nylon is easy to handle and cost efficient. The wrap around
webbing increases the overall strength per inch. Tubular nylon is highly flexible and pliable. Due
to its pliability, it tends to glide better over rough or jagged surfaces preventing the wear and tear
that occurs more with Kevlar. One inch width of tubular nylon webbing can withstand about 4000
pounds of pressure. For the main parachutes, which create more drag force, the shroud lines will
be made from this material. Nylon has a much high strength and will also allow the team to use a
double seam across either side of the shock cord, ensuring the stitch is stronger as well.
While tubular Kevlar is stronger, the strength of tubular nylon is more than sufficient for the needs
of the mission. In addition, the pliability of nylon will allow the shock cord to better absorb the
shock of ejection and ensure smooth movement of the potentially rough surfaces inside the upper
section. By using a material known to be strong, the team ensures failure is less likely to happen
in this component. The pliability and strength of tubular nylon have led to the team choosing to
use it as our material of choice for shock cords.
53
The factors that should be used to assess the suitability of different shock cord materials to allow
for recovery system requirement satisfaction are volume, strength, durability, pliability, flexibility,
and weight. The Pugh chart that was used to evaluate which material was chosen for the parachute
is shown in Table 6.3: Comparison of Paracord, Tubular Nylon, Kevlar
Table 6.3: Comparison of Paracord, Tubular Nylon, Kevlar
Paracord 1-inch Tubular Nylon Kevlar
Volume 3 1 2
Weight 3 2 2
Strength 1 2 3
Durability 1 2 1
Pliability 3 3 1
Flexibility 3 3 2
Total 14 13 11
Nylon shear pins will be used to attach the nosecone to the BAE and to prevent drag separation.
In the team’s configuration, #4-40 nylon screws are used to secure the sections. These machine
screws have a double shear strength of 50 lbs. Ground testing was performed on these shear pins
to ensure safety and eliminate the possibility of manufacturing discrepancies.
The materials used to attach the shock cord to the bulk plates should be chosen appropriately. The
team will be using U-bolts to attach the parachutes to the bulk plates, because U-bolts have proven
to be more reliable in the team’s past. Shock cord cannot become easily tangled, and an eye bolt
is more susceptible to failure for this reason, as the shape allows cord to wrap around it. U-bolts
also provide two points of attachment to the bulk plate while eye bolts only provide one. The
factors that should be used to assess the suitability of different bulk plate attachment materials to
54
allow for recovery system requirement satisfaction are strength and an ability to limit cord tangles.
The Pugh chart that was used to evaluate which material was chosen for the parachute is shown in
Table 6.4: Comparison of U-bolts and Eye-bolts.
Table 6.4: Comparison of U-bolts and Eye-bolts
U-bolt Eye-bolt
Attachment Strength 3 2
Limits cord tangles 3 1
Totals 6 3
Section 6.3: Ejection
4F black powder has been chosen for the ejection of the team's parachutes. Black powder is an
effective, reliable means of pressurization that the team has had success with in the past. Compared
to CO2 ejection, black powder can produce greater pressures per cubic inch required to house the
system. A comparison between CO2 and black powder can be seen in Table 6.5: Comparison of
Black Powder and CO2.
Table 6.5: Comparison of Black Powder and CO2
Criteria Black Powder 𝐂𝐎𝟐
Pressure Produced per
Volume 3 1
Damaging Heat Produced 1 3
Cost 3 2
Ease of Integration 3 1
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Reliability 2 1
Total 12 8
For the both events, two charges are placed within 3D printed charge cups and armed with
electronic matches: the first charge is the primary means of ejection, and the other will be a backup
charge for redundancy and to decrease the chances for failure of the recovery system. The
redundant apogee charge is set to fire 3 seconds after the first charge. The charges are not ignited
at the same time as that has the potential to cause damage to the airframe and failure of the system.
Black powder, when ignited, can be approximated as an ideal gas. While further testing will be
performed to guarantee the validity of the team’s calculations, the Ideal Gas Law can be used to
calculate an estimate for the amount of black powder needed for ejection. The Ideal Gas Law is as
follows:
𝑃×𝑉 = 𝑛×𝑅×𝑇
𝑅 = 265.9𝑖𝑛 − 𝑙𝑏𝑓
𝑙𝑏𝑚×𝑅 𝑇 = 3300R°
Ignoring the volume of the parachutes and other recovery systems contained within the upper
section, a conservative volume for the upper section can be calculated with an inner diameter of 6
inches and a length of 24 inches as:
𝑉 = 𝐴×𝐿 =𝜋
4×𝑑2×= 678.6 𝑖𝑛3
Using this equation, the team can calculate the amount of black powder needed to produce
pressures sufficient to shear the nylon screws attaching the nosecone to the upper section.
Assuming 3 nylons screwed each rated for 50 psi of shear, pressure required to shear the screws
is:
𝑃 =3×𝐹
𝜋4 ×𝑑2
= 5.3 𝑝𝑠𝑖
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Which, in conjunction with the previous equations, yields a charge size for 4F black powder of:
𝑛 =𝑃×𝑉
𝑅×𝑇×
453.592𝑔
𝑙𝑏𝑚= 1.86 𝑔
1.86 grams of black powder will serve as a starting estimate and will be verified via ground testing.
Both charge sections are filled with fireproof cellulose insulation (colloquially known as “barf”)
to protect the parachutes from the heat of ejection. The first event deploys the drogue parachute
and the rifled upper main parachute, held closed by our Jolly Logic chute release system. The
second event consists of the second set of charges being deployed to separate the rocket and deploy
the lower main parachute as well as a mechanical release of a pin in the Jolly Logic system,
allowing the upper main parachute to deploy.
The AUSL team is utilizing the Jolly Logic system in series for redundancy in the recovery
systems. It is an efficient system that doesn’t need any black powder charges, decreased amount
of wires connected to the altimeter bay, as well as reduced risk of entangled shock cord in the main
body. The only cons of this system would be a Jolly Logic system malfunction.
Figure 6.1: Redundant Jolly Logic System
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The Jolly Logic system allows for the deployment of both a drogue and main parachute
simultaneously in a single separation. This system deploys more parachutes with fewer
separations, reducing the chance of failure of the recovery portion of flight. The only parachute
that will utilize the Jolly Logic Chute Release System will be the upper main parachute.
Figure 6.2: Jolly Logic Chute Release
The Jolly Logic system consists of an independent altimeter device which releases a mechanical
pin an altitude that is preset at (target height 1000 ft.). The pin is connected to the Jolly Logic
device via a rubber band which provides tension on the pin so that when it is released, it is pulled
from its socket. Since the Jolly Logic system consists of rubber bands which can easily break if
put under stress, the recovery system is set up so that tension from the shock cord is never
transferred to the Jolly Logic devices. To do this, the main parachute is not folded with the shock
cord inside as is normal. Instead the parachute is located on the shock cord so that the shroud lines
are at full extension and so that the drogue line, which runs through the main parachute spill hole
is able to deploy to its full length during the initial separation. The main parachutes are then
carefully gathered and positioned in the body tubes so that the drogue line will pull it out without
putting any forces on the actual parachutes and Jolly Logic system.
58
To preserve the redundancy of our recovery system, we are using two Jolly Logic chute releases
in series to gather the main parachutes. If either of the Jolly Logic devices happens to malfunction,
the other will still release, allowing the main parachutes to deploy.
Section 6.4: Parachutes
Auburn’s augmented dual deploy recovery approach will make use of three separate parachutes
designed and constructed in house by the AUSL team. The team has been constructing its own
parachutes in house for five years and has refined its manufacturing process to produce quality,
custom chutes that produce the desired drag and drift for all sections of the rocket.
The drogue parachute will be a small, circular parachute constructed of rip-stop nylon with
paracord shroud lines. Following the first event at apogee, the drogue along with the upper main
parachute bundled with a jolly logic chute release system will be deployed from the top of the
rocket. This will stabilize descent until main deployment at the second event. The drogue parachute
is designed to stabilize the descent of the rocket from apogee to main parachute deployment at a
descent velocity of approximately 100 m/s. This velocity was chosen to minimize the drift of the
rocket while still having a stable descent. A drag coefficient of 0.8 for a fully inflated circular
parachute was determined from research. Further testing will be done later to ensure the validity
of this coefficient.
𝐴 = 2×𝐹
𝜌×𝐶𝐷×𝑉2
Where F is force, ρ is density of the air, CD is the drag coefficient and V is descent velocity. The
team has used this equation to calculate an appropriate area for the drogue parachute.
𝐴 = 2×32𝑙𝑏𝑚×32.2
𝑓𝑡𝑠2
0.076474𝑙𝑏𝑚
𝑓𝑡3 ×0.8× (100𝑓𝑡𝑠 )
2 = 3.37 ft2
These calculations yield a circular drogue with a 24.9-inch diameter.
59
The recovery system will have two main parachutes constructed of rip-stop nylon with 0.5-inch
tubular nylon shroud lines. The main parachutes will be hemispherical. The hemispherical shape
can be more difficult to manufacture, but will produce the most drag, allowing the rocket to have
maximum drag with minimum weight. A Pugh chart can be seen below comparing different
parachute shapes.
Table 6.6: Parachute Shape Pugh Chart
Baseline Square Circular Hemispherical
Drag Produced 3 1 1 2
Ease of
Manufacturing
2 1 2 1
Stability 1 2 1 1
Total 7 8 9
The shape of the main parachutes and their gores can be seen in Figure 6.3: Gore Template and
Figure 6.4: Parachute Template. When the rocket reaches 750 feet in altitude, a second charge will
separate the top section of the rocket to release the lower main chute, and the jolly logic chute
release system will release the bundled upper main chute. A spill hole will be added to the main
parachutes. This spill hole will be necessary with our configuration of dual-deployment from the
same compartment at the top of the rocket body. In accordance with the general rule of thumb, the
spill holes will be close to 20% of the total base diameters of the chutes. The 20% diameters of the
spill holes are chosen because it only reduces the areas of the parachutes by about 4%, allowing
enough air to go through the spill hole to stabilize the rocket without drastically altering the descent
rate.
60
Figure 6.3: Gore Template
Figure 6.4: Parachute Template
Parachute areas for hemispherical shaped chutes are determined using the following equation:
𝐴 = 2×𝐹
𝜌×𝐶𝐷×𝑉2
Where F is force, ρ is density of the air, CD is the drag coefficient and V is descent velocity. The
team has used this equation to calculate an appropriate area for the main parachutes so that the
kinetic energy of either section of the rocket does not exceed 75 ft-lbs during recovery and remains
within safe limits.
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𝐴𝑈𝑝𝑝𝑒𝑟 =2×8𝑙𝑏𝑚×32.2
𝑓𝑡𝑠2
0.076474𝑙𝑏𝑚
𝑓𝑡3 ×1.31× (20𝑓𝑡𝑠 )
2= 12.86 𝑓𝑡2
𝐴𝐿𝑜𝑤𝑒𝑟 =2×24𝑙𝑏𝑚×32.2
𝑓𝑡𝑠2
0.076474𝑙𝑏𝑚
𝑓𝑡3 ×1.31× (11.5𝑓𝑡𝑠 )
2= 116.66 𝑓𝑡2
Table 6.7: Parachute Dimensions
Upper Chute Lower Chute
Area 12.86 ft2 116.66 ft2
Diameter 34.3 in 103.4 in
Diameter of Spill hole 6.9 in 20.7 in
Number of Gore 6 8
Width of Each Gore at Base 18.0 in 40.7 in
Height of Each Gore 26.9 in 81.15 in
Circumference at Base 107.8 in 27.1 ft.
The team decided to move from 6 gores to 8 gores for the lower main parachute. The main reason
for this change is that the template the team would have to print out for a 6-gore hemispherical
parachute would simply be wider than any printer available to the team would be able to print.
Moving to an 8-gore configuration also increases the accuracy the accuracy of the parachute to a
true hemisphere.
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Section 6.5: Altimeters
The avionics bay will house two altimeters to satisfy redundant system requirements. Both
altimeters will fire the apogee charge at the apogee height of 1 mile (5280 feet) to eject the
nosecone and thus the drogue and bundled upper main parachute. Then both altimeters will fire
the main deployment at an altitude of 750 ft. Neither set of charges are fired at the exact same
time. A three second delay has been placed between the firing of each charge to ensure the
structural integrity of the rocket is maintained.
The team will be using one Altus Metrum TeleMega as the primary and one Altus Metrum
TeleMetrum as the secondary altimeters. The other altimeter that was being considered was the
PerfectFlite Stratologger. The TeleMega has 4 additional sets of pyro connectors, allowing for
future expansion if necessary. It can also have a second battery easily installed into dedicated screw
terminals for additional power for pyro ignition purposes. The TeleMega also has a more advanced
accelerometer for more detailed flight data acquisition. The PerfectFlite Stratologger requires a
standard 9V battery, which is larger and heavier than the battery used for the Altus Metrum
altimeters. Additionally, using two Altus Metrum altimeters will make programming quicker and
easier, as they share an interface program. This makes any last minute or on-site adjustments across
both boards simpler. Should one of the Altus Metrum altimeters fail, the PerfectFlite Mawd or
Stratologger can be used as additional backup. All altimeters are capable of tracking in flight data,
apogee and main ignition, GPS tracking, and accurate altitude measurement up to a maximum of
25,000 feet. Figure 6.5: Altus Metrum TeleMega Altimeter and Figure 6.6: Altus Metrum
TeleMetrum Altimeter shown below are pictures of the altimeters the team will be using.
63
Figure 6.5: Altus Metrum TeleMega Altimeter
Figure 6.6: Altus Metrum TeleMetrum Altimeter
Another reason the Altus Metrum altimeters are preferred are their radio frequency (RF)
communication abilities. Both TeleMega and TeleMetrum are capable of communicating with a
Yagi-Uda antenna operated by the team at a safe distance at any point during the launch. It can be
monitored while idle on the ground or while in flight. While on the ground, referred to as “idle
mode”, the team can use the computer interface to ensure that all ejection charges are making
proper connections. Via the RF link, the main and apogee charges can be fired to verify
functionality, which was used to perform ground testing. The voltage level of the battery can also
be monitored, and should it dip below 3.8V, the launch can be aborted in order to charge the battery
to an acceptable level. Additionally, the apogee delay, main deploy height, and other pyro events
can be configured. The altimeter can even be rebooted. While in flight, referred to as “flight mode”,
the team can be constantly updated on the status of the rocket via the RF transceiver. It will report
altitude, battery voltage, igniter status, and GPS status. However, in flight mode, settings can’t be
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configured, and the communication is one way from the altimeter to the RF receiver. Figure 5.11
shown below demonstrates the process the altimeters go through to deploy the charges at each
event.
Figure 6.7: Altimeter Block Diagram
Isolating one altimeter system (altimeter, battery, and wires) from the other helps prevent any form
of coupling or cross-talk of signals. Isolation is realized via distancing the two systems, avoiding
parallel wires, and twisting wires within the same circuit. Additionally, the most apparent form of
radio-frequency interference, the antenna, will resonate on wires any multiple of ¼ λ (1/4 of
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~70cm). Avoiding resonant lengths of wire will be done wherever possible. Should a wire happen
to be a resonant length and is unable to be shortened or lengthened, a low-pass filter can be
implemented to block the high frequency noise. Both altimeters and their batteries will be mounted
on a carbon fiber board that slides into a set of rails. The altimeters and batteries will be mounted
on opposing sides of the board, with one battery and altimeter per side. Since carbon fiber is an
effective shielding material (50dB attenuation), this board will act as shielding between the two
altimeters and minimize cross-talk as well as near-field coupling. This board will also be easily
removable for connecting the altimeters to computers for configuration and for charging the
altimeters’ batteries.
Section 7: Mission Performance Predictions
Section 7.1: Simulations
The launch vehicle has been simulated using OpenRocket, an open source rocket simulation
software. The team is confident in this software’s ability to simulate a rockets flight due to past
years success with using OpenRocket. The rocket as it is modeled in OpenRocket is shown in
Figure 7.1: OpenRocket Model.
Figure 7.1: OpenRocket Model
The multiple runs of the simulation were conducted using various calculation methods and
assumptions such as wind speed as well as using different approximations for the earths shape.
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Wind speeds were tested up to a maximum of 25 mph, which yielded only nominal changes in the
expected velocity, acceleration, or apogee. The same held for the different earth approximations.
Additionally, simulations were run where the temperature was varied between 50°F to 80° F. These
changes had no major impact on the projected values for velocity, acceleration and apogee.
Between all the different simulations the team ran, the results differed by approximately 1%.
Therefore, the following data is presented from a simulation run with 0 mph winds and standard
sea level conditions.
Table 7.1: Flight Simulation Data (Wind = 0 mph)
Flight Simulation Data (Wind = 0 mph)
Maximum Velocity 708 ft./s
Maximum Acceleration 269 ft./s2
Launch Weight 39.9 lbm
Burnout Weight 34.4 lbm
Length 108 in
Maximum Diameter 6.25 in
Launch Stability 2.4 calibers
Velocity off Rod 78.6 ft/s
67
Figure 7.2: Altitude Vs. Time
68
Figure 7.3: Velocity Vs. Time
69
Figure 7.4: Acceleration Vs. Time
Section 7.1.1: Motor Thrust Curve
The simulated motor thrust curve for the selected motor can be seen below. The team is confident
that this motor selection will provide adequate thrust to propel the launch vehicle to above the
target altitude to enable the altitude control module to adjust the final altitude to 5280 ft.
70
Figure 7.5: Motor Thrust Curve
Section 7.1.2: Component Weights
Table 7.2: Component Weights
Component Weight (lbm)
Upper Body Tube 4.78
Lower Body Tube 4.64
Nose Cone 2.17
Fins (4) .91
Centering Rings (3) .6
Motor 10.1
Motor Retention .3
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Motor Tube 1
Rover and Bay 4.5
Airbrake System 2
Bulkheads .9
Electronics 1.5
Recovery Section 7
Total 40.5
Section 7.1.3: Stability
For the rocket to be stable during flight the center of pressure must be located aft of the center of
gravity. It is recommended for this size of rocket that the stability margin should be 1-2 calibers.
The stability margin for the rocket is predicted to be 2.4 calibers at the point of rail exit which is
comfortably above the minimum requirement of 2 calibers. This number was calculated without
considering the drag plate system, which will move the center of pressure forward. However, the
drag plates will not deploy until after motor burnout. The team feels confident that having a
stability margin of over 2 calibers will allow the rocket to still remain stable with the inclusion of
the drag plates.
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Figure 7.6: Stability Vs. Time
Before flight, the center of pressure and the center of gravity are measured to be 84.13 inches and
69.992 inches from the top of the rocket, respectively. In Figure 7.6: Stability Vs. Time, the change
in the locations of the center of pressure and center of gravity can be seen for the entirety of the
flight. After apogee, the data for the stability drops off due to the deployment of the drogue
parachute. The increase in stability is due to the motor losing propellant mass due to motor burnout
as well as the decrease in velocity which moves the center of pressure away from the center of
gravity.
Section 7.1.4: Computational Fluid Dynamics
Computational Fluid Dynamics (CFD) is a branch of fluid mechanics that utilizes numerical
analysis methods and high-performance parallel computing to best analyze the flow properties and
fluid dynamic interactions of aerospace systems. Through CFD, the Navier-Stokes equations are
73
solved numerically across a finite volume that encompasses the boundary conditions that comprise
the tested system. The finite volume is constructed through a series of successive steps, beginning
with the creation of a surface geometry through computer-aided design (CAD) software. The
surface data is then imported into mesh generation software, where domains are constructed
according to the specific boundary conditions inherent in the desired simulation. The exported grid
is then processed through a CFD software package, where flow parameters and boundary
conditions are specified, and a specific flow solution is chosen to account for various aerodynamic
phenomena (vorticity, viscosity, turbulence, and more), and numerical data approaches
convergence.
A new shell was constructed in SolidWorks to simulate the exterior of the rocket body, comprised
of the revolved nose-cone, uniform cylindrical body, and trapezoidal fin control surfaces. The
SolidWorks geometry was then exported in the form of a .STP file in order to be imported into
Pointwise mesh generation software. A farfield domain, in the form of a cylinder, was constructed
around the rocket body, extending roughly five rocket-body lengths in the aft direction, and two
lengths in the forward direction, while having a diameter many multiples of the rocket-body
diameter in order to maintain a farfield condition. The connection point discontinuities, apparent
after importation into Pointwise, between the nose-cone and the cylindrical body, and between the
trapezoidal fins and the cylindrical body, were resolved and smoothed.
74
Figure 7.7: Nosecone Meshing
Grid spacing on the fins was also applied more uniformly through the leverage of anisotropic
tetrahedral meshing, and is evidenced in the following figure.
Figure 7.8: Tail-fin Meshing
The entire domain was processed through an anisotropic tetrahedral mesh extrusion, with the
cylindrical farfield as a “farfield” boundary condition, and the entire rocket-body defined as a
75
“wall” condition, with a ∆s value of roughly 1.0x10^-7 (a value calculated as a function of the
Reynolds number of the simulated transonic flow).
Upon exportation of the grid, the TENASI CFD software package, developed by UT Chattanooga
and used by Auburn University, was utilized to construct separate parameter and boundary
condition files to test the freshly created grid with. Current research efforts are focusing on the
comparison between the use of an arbitrary mach flow regime with the Menter Scale-Adaptive
Simulation (SAS single equation) and the Spalart-Allmaras (SA single equation) flow solution
types. Preliminary analysis of the rocket-body grid flying at roughly Mach 0.65, with a post-
burnout center of gravity occurring roughly 1.25 meters aft of the origin of the coordinate axis
defined at the center of the nose-cone connecting ring produced coefficient of drag values, and are
shown in the following table. These drag coefficients were then inputted into OpenRocket to
ensure the use of accurate drag coefficients in the team’s simulations.
Table 7.3: CFD Drag Coefficient Results
Mach Number (M) Coefficient of Drag (C_d)
0.6 .532
0.725 .5375
0.750 .5425
Section 7.2: Kinetic Energy
The kinetic energy for the rocket upon impact can be calculated using the following formula:
𝐾𝐸 =1
2𝑚×𝑉2
Where m is mass and V is descent velocity. With a mass of 8 lbm and a calculated velocity of 20
ft./s for the upper section, this equation yields:
76
𝐾𝐸 =1
2×
8𝑙𝑏𝑚
32.2𝑓𝑡𝑠2
× (20𝑓𝑡
𝑠)
2
= 49.7 𝑓𝑡 ∙ 𝑙𝑏
With a mass of 24 lbm and a calculated velocity of 11.5 ft/s for the lower section, this equation
yields:
𝐾𝑒 =1
2×
24𝑙𝑏𝑚
32.2𝑓𝑡𝑠2
× (11.5𝑓𝑡
𝑠)
2
= 49.3 𝑓𝑡 ∙ 𝑙𝑏
Section 7.3: Drift
The distance the rocket will drift during descent can be estimated with the following equation.
𝐷𝑟𝑖𝑓𝑡 = 𝑊𝑖𝑛𝑑 𝑆𝑝𝑒𝑒𝑑×𝐴𝑙𝑡𝑖𝑡𝑢𝑑𝑒 𝐶ℎ𝑎𝑛𝑔𝑒
𝐷𝑒𝑠𝑐𝑒𝑛𝑡 𝑉𝑒𝑙𝑜𝑐𝑖𝑡𝑦
However, this drift estimation assumes wind speed and descent velocity are constant and does not
account for the horizontal distance the rocket travels during ascent. There will be two stages of
descent. First, the rocket will descend under the drogue parachute from an altitude of 5280 ft. to
750 ft. At 750 ft. a second black powder series will occur separating the rocket into two pieces,
and the Jolly Logic parachute release system will separate and another event will occur releasing
a lower main parachute to safely escort the rover to the ground. The rate of descent under drogue
can be calculated with the following equation:
𝐷𝑒𝑠𝑐𝑒𝑛𝑡 𝑉𝑒𝑙𝑜𝑐𝑖𝑡𝑦 = √2×𝐹𝑜𝑟𝑐𝑒
𝐴𝑖𝑟 𝐷𝑒𝑛𝑠𝑖𝑡𝑦×𝐷𝑟𝑎𝑔 𝐶𝑜𝑒𝑓𝑓𝑖𝑐𝑖𝑒𝑛𝑡×𝑃𝑎𝑟𝑎𝑐ℎ𝑢𝑡𝑒 𝐴𝑟𝑒𝑎
Since this is a variation on the formula used to calculate the parachute areas, the resulting velocities
are the team’s desired descent velocities. However, as the competition progresses, this formula can
be used to update our predicted velocities with different drag coefficients, weights, or parachute
areas. This descent velocity will then be used to ensure drift is kept to a reasonable amount. An
assumed drag coefficient of 0.8 was estimated from research. Testing will be performed later to
77
ensure the validity of this coefficient. Assuming a total rocket weight after burnout of 32 lbm and
a drogue diameter of 12.43 inches (3.37 ft2), the descent velocity under drogue is:
𝐷𝑒𝑠𝑐𝑒𝑛𝑡 𝑉𝑒𝑙𝑜𝑐𝑖𝑡𝑦 = √2×32𝑙𝑏𝑚×32.2
𝑓𝑡𝑠2
0.076474𝑙𝑏𝑚
𝑓𝑡2 ×0.8×3.37 𝑓𝑡2= 100
𝑓𝑡
𝑠
Assuming a total rocket weight of 32 lbm and an upper main parachute area of 12.86 ft2, the descent
velocity of the entire rocket before separation is:
𝐷𝑒𝑠𝑐𝑒𝑛𝑡 𝑉𝑒𝑙𝑜𝑐𝑖𝑡𝑦 = √2×32𝑙𝑏𝑚×32.2
𝑓𝑡𝑠2
0.076474𝑙𝑏𝑚
𝑓𝑡2 ×1.31×12.86 𝑓𝑡2= 39.9
𝑓𝑡
𝑠
Assuming a total upper rocket weight after burnout of 8 lbm and an upper main parachute area of
12.86 ft2, the descent velocity of the upper section after separation is:
𝐷𝑒𝑠𝑐𝑒𝑛𝑡 𝑉𝑒𝑙𝑜𝑐𝑖𝑡𝑦 = √2×8𝑙𝑏𝑚×32.2
𝑓𝑡𝑠2
0.076474𝑙𝑏𝑚
𝑓𝑡2 ×1.31×12.86 𝑓𝑡2= 20.0
𝑓𝑡
𝑠
Assuming a total lower rocket weight after burnout of 24 lbm and a main parachute area of 116.66
ft2, the descent velocity of the lower section after separation is:
𝐷𝑒𝑠𝑐𝑒𝑛𝑡 𝑉𝑒𝑙𝑜𝑐𝑖𝑡𝑦 = √2×24𝑙𝑏𝑚×32.2
𝑓𝑡𝑠2
0.076474𝑙𝑏𝑚
𝑓𝑡2 ×1.31×116.66 𝑓𝑡2= 11.5
𝑓𝑡
𝑠
Estimated drift distances for a variety of wind speeds are shown below in Table 7.4: Drift
Calculations for Upper Section and Table 7.5: Drift Calculations for Lower Section. These tables
contain the total and the broken-down drift at each wind speed. The drift is broken down into three
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separate sections: Drift under drogue (5280 – 1000 ft), drift under upper main before separation
(1000 – 750 ft) and drift under each respective main after to separation (750 – 0 ft).
Table 7.4: Drift Calculations for Upper Section
Wind Speed
(mph)
Wind
Speed
(ft./s)
Drift Under
Drogue (ft.)
Drift Before
Separation
(ft.)
Drift Under
Upper (ft.)
Total Drift
of Upper
Rocket (ft.)
5 7.33 313.72 45.93 274.88 634.53
7.5 11.00 470.80 68.92 412.50 952.22
10 14.67 627.88 91.92 550.13 1,269.93
12.5 18.33 784.52 114.84 687.38 1,586.74
15 22.00 941.60 137.84 825.00 1,904.41
17.5 25.67 1,098.68 160.84 962.63 2,222.15
20 29.33 1,255.32 183.77 1,099.88 2,538.97
Table 7.5: Drift Calculations for Lower Section
Wind Speed
(mph)
Wind Speed
(ft./s)
Drift Under
Drogue (ft.)
Drift
Before
Separation
(ft.)
Drift Under
Lower (ft.)
Total Drift of
Lower
Rocket (ft.)
5 7.33 313.72 45.93 478.04 837.69
7.5 11.00 470.80 68.92 717.42 1,257.14
10 14.67 627.88 91.92 956.78 1,676.58
12.5 18.33 784.524 114.84 1,195.48 2,094.84
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15 22.00 941.60 137.84 1,434.84 2,514.25
17.5 25.67 1,098.68 160.84 1,674.20 2,933.72
20 29.33 1,255.32 183.77 1,912.90 3,351.99
Section 7.4: Simulation Verification
Multiple instances of the simulation, although not shown, were ran to verify the accuracy of the
simulation. Variables such as ambient temperature and wind speed changed the results by
approximately 1% and so were kept fixed. In addition, the simulation data was compared to the
subscale flight and CFD results to ensure an accurate drag coefficient for the model. The team is
confident that the simulation is reliable currently, but will continue to test and adapt it with full-
scale test flight data.
Section 8: Safety
Section 8.1: Pre-Launch Day Items Checklist
□ All Rocket Components
□ Trailer
□ Generator
□ Tables
□ Insulation Packing (Barf)
□ Power tools
□ Drills
□ Dremels
□ Plug in batteries the night before
□ Recovery Tower
□ Black powder
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□ Charge cups
□ Shear pins
□ Spare electrical wire
□ Spare batteries
□ Wire strippers
□ Recovery Laptop
□ With attachment to talk to
altimeters
□ Systems box of extra electronics
□ Spare batteries
□ Systems laptop
□ With attachment to talk to
electronics
□ Rover box of extra electronics
□ Spare batteries
□ Rover laptop
□ With attachments to talk to rover
□ Silver Sharpies
□ Extra launch buttons
□ Snap rings
□ Axle grease
□ Sandpaper
□ Duct tape
□ Extension cord(s)
□ Tents
□ Old toothbrushes
□ Extra launch controller
□ Allen wrenches
□ Scale
□ At least 48 water bottles
□ Trash bags
□ Chairs
□ Paper towels
81
Section 8.2: Preassembly Checklists
Section 8.2.1: Recovery
Task Completed (Initialed)
Ensure all necessary items are transported
Check primary and back-up batteries
Check both altimeters and jolly logic
systems for functionality
Folded parachutes and pack into sleeves
Attach shock chord for nose cone, drogue
parachute and upper main parachute
Attach key switches
Secure e-match in charge cup
Fill Charge Cups (done or supervised by
authorized personnel):
• Use clear and clean table surface
• Wear provided safety equipment
• Fill cups using clean funnel
• Fill to exact mass (measure with scale)
Attach key switches and e-matches to
altimeter boards
Place system in housing BAE
82
Section 8.2.2: Altitude Control
Section 8.2.3: Body
Task Completed (Initialed)
Ensure all necessary items are transported
Check primary and back-up batteries
Check altimeters for functionality
Check for any damage on plates or housing
Test system for functionality and smooth
extension
Place system in housing and secure
attachment screws
Task Completed (Initialed)
Ensure all necessary items are transported
Check for damage to nose
Check for damage to upper section
Check for damage to lower section
Check fins for correct alignment and any
damage from transportation
83
Section 8.2.4: Rover
Task Completed (Initialed)
Ensure all necessary items are transported
Check primary and back-up batteries
Test rover functionality along with engine
response and functionality
Test radio receiver
Place rover on tracks in housing bay
Secure housing bay door
Check engine housing for structural
integrity
Gather all shear pins and attachment
screws
84
Section 8.2.5: Engine
Safety Officer Signature President Signature
X X
_______________________ __________________________
Section 8.3: Launch Vehicle Assembly and Check
Task Completed (Initialed)
Attach lower main parachute to lower
section attachment ring
Attach upper section to avionics BAE with
screws
Attach upper main parachute assembly to
avionics bulkhead
Task Completed (Initialed)
Use engine transported by team mentor or
purchase engine from authorized range
store
Engine prepared by licensed team mentor
85
Ensure charge cups are in proper locations
indicated by design
Pack insulation (barf) around lower charge
cups and insert lower main parachute and
shock chord
Attach lower section to avionics BAE with
shear pins
Pack insulation (barf) around upper
charge cups and insert upper main
parachute assembly and shock chord
Attach nose cone to upper section with
shear pins
Check all connections for proper alignment
Place and secure engine in engine housing
Test center of gravity
Check key stitches to ensure functionality
of avionics
Check body for flight readiness
Check engine mount for flight readiness
Check fins for flight readiness
If any steps cannot be completed,
disassemble and correct
86
Safety Officer Signature President Signature
X X
_______________________ __________________________
Section 8.4: Launcher Setup and Launch Procedure
Section 8.4.1: Launcher Setup
Task Completed (Initialed)
Test to ensure all weather conditions are
within preset limits
With supervision of range safety officer,
place rocket on launch rail
Ensure angle launch rail is within limits
Turn on avionics
Clear surrounding area of flammable
material
87
Have all viewing personnel move to safe
distance
Place ignitor against engine fuel grain
Attach launch controller to ignitor
Check for proper connection
Section 8.4.2: Launch Procedure
Task Completed (Initialed)
Move setup team to safe launch distance
Initialize mission process with range
officials
Receive all clear from range officials
Initiate motor ignition
Check for proper ignition
Safety Officer Signature President Signature
X X
_______________________ __________________________
88
Section 8.5: Post-flight Inspection
Task Completed (Initialed)
Locate upper and lower sections
Follow guidelines for proper safe recovery
Check altimeter beeps for launch altitude
Remove any vehicle components
Remove any environmental hazards
89
Section 8.6: Personnel Safety Hazards
Hazard Cause Result Severity Probability Combined
Risk Mitigation
Improper
use of small
power tools
Lack of
training,
improper
use, and/or
improper
protection
such as a
lack of
gloves or
safety
glasses
Mild to
severe cuts,
scrapes, and
other
injuries.
Additionally,
reactions can
result in
harm to
rocket
components
being worked
upon.
3 3 9
Demonstration
of proper use by
experienced
team members,
easily accessible
safety materials
and protective
wear, and
securely
fastening the
object being
worked upon
Improper
use of large
power tools
Failure to
pay
attention,
aggressive
use of the
tools, lack
of proper
protective
equipment,
or a lack or
training
Severe cuts,
burns, rashes,
bruises, or
other harm to
fingers,
hands, or
arms
4 2 8
Experienced
team members
will instruct
inexperienced
team members
before the newer
team member is
allowed to use
the tools, and
only those that
display
comprehensive
and safe work
may proceed.
Protective
equipment will
also be easily
accessible
90
Carbon
fiber
particles
Sanding
carbon
fiber or
other
fibrous
material
without
using a
mask or
filter
Mild
coughing and
difficulty
breathing,
irritation in
the eyes and
skin.
2 4 8
When sanding or
cutting tools are
used on carbon
fiber all
members in the
lab, regardless if
they are working
on the carbon
fiber or not, are
required to
utilize a mask to
prevent the
breathing in of
excessive
particles
Improper
use of
soldering
tools
Failure to
pay proper
attention to
the
soldering
tool or lack
of training
Mild to
severe burns
on the fingers
or hands of
the team
member
using it.
Additionally,
could result
in excessive
heat and
damage to
the
component
being worked
on.
3 2 6
Members that
use the soldering
iron are required
to give it their
full attention for
the duration of
their work. They
must turn off
and stow the tool
somewhere
away from the
object being
worked on if
they must attend
to something
else before work
is finished. Only
those who have
been instructed
in the use of the
soldering iron
may work with
the tool
unassisted
91
Noxious
Fumes
Improper
ventilation
of the
workstation
where
soldering or
curing is
taking
place
Excessive
exposure to
toxic fumes
results in
nausea and
irritation.
Reactions
could
potentially
damage the
component
being worked
upon.
3 2 6
The workstation
will be properly
vented and
members using it
are required to
confirm
ventilation is
functioning
periodically.
Additionally, if
ventilation is
malfunctioning,
team members
will not be
allowed to
continue
soldering for an
extended period
of time and must
take a break to
let any buildup
of fumes
disperses
Insecure
Tools
Tools are
left out in
the lab
workspace
and not
returned to
a storage
space
where they
belong
Cuts, pricks,
or tears when
members sort
through
items or
knock loose
tools around
or off tables.
2 3 6
The storage
spaces for all
tools are clearly
marked and easy
to find.
Members are
instructed to
return tools they
find left out to
their storage
spaces. A
checklist must be
finished before
project members
can leave the lab
or start a
different project.
92
Electrical
discharge
from
equipment
a)
Improperly
maintained
equipment
b)
Improper
use of
equipment
Electric
shocks could
occur to team
members
handling the
equipment
4 1 4
Electrical
equipment will
be maintained
regularly.
Electrical
equipment will
only be plugged
in when ready
for immediate
use and will be
promptly
unplugged
afterward. All
wires will be
kept out of the
path of other
team members
or any other
equipment.
Improper
use of
nonpowered
tools
Lack of
training or
discipline
with tools
Damage to
sections of
the rocket or
to team
members and
delays to the
project due
to the need
for
replacements
2 2 4
All team
members will be
trained on the
use and proper
educate of the
lab. If they are
not followed,
members will be
reprimanded
before an
incident occurs
New Mitigation Steps Completed or To Be Completed:
• New, clearer labeling of all tools, tool locations, and hazards associated with the use of
said tools
• Work stations dedicated to one specific job at a time to promote accountability within the
work area
93
Section 8.7: Environmental Effects
Section 8.7.1: Effects on Environment
Hazard Cause Result Severity Probability Combined
Risk Mitigation
Fire on
ignition
Upon
ignition,
the exhaust
from the
rocket may
set fire to
any
vegetation
beneath or
around the
launch site
Fires will
destroy the
vegetation
near launch
and have the
potential to
spread
further from
the site
4 3 12
In accordance
with the NAR
guidelines, the
launch site will
be placed such
that it does not
present the risk
of grass fires.
Cured
epoxy in
landfill
Cups used
to cure
epoxy are
thrown into
normal
trash bins
and taken
to landfills
The epoxy
breaks down
and releases
harmful
chemicals
into the
ground
2 4 8
Epoxy and
epoxy stirring
cups will be
disposed of
separately into a
bin by the work
station. The
contents will be
taken to an
approved
chemical
disposal site
Epoxy gas
and
chemical
release
Epoxy
releases
volatile
chemicals
and gasses
as it cures
This small
release can
be vented
into the
environment
1 5 5
With our rather
small scale of
our construction,
the impact of our
ventilation is
minimal
94
Section 8.7.2: Effects of Environment
Hazard Cause Result Severity Probability Combined
Risk Mitigation
Exposure to
humidity
(corrosion)
Exposure to
humidity
can cause
metals
within
certain
systems to
corrode
Corroded
metals do not
have the
same
integrity as
their original
states leading
to potential
damage
3 2 6
Our system
components will
be kept in a
dehumidified
space at room
temperature to
avoid and
corrosion
Exposure to
humidity
(electronics)
Exposure to
humidity
can cause
wires or
electronic
boards to
corrode
Corroded
wires can
cause
electrical
signals to not
be
transmitted
leading to a
loss of
avionics or
rover control
3 3 9
Our system
components will
be kept in a
dehumidified
space at room
temperature to
avoid and
corrosion
95
Cross winds
in flight
Rocket is
launched
when cross
winds are
faster than
allowed
limits
Winds can
cause the
rockets
trajectory to
change or its
flight to
become
unstable.
This can
further cause
the rocket to
land in an
unanticipated
or unsafe
location
3 4 12
In accordance
with NAR
regulations, the
rocket will not
be launched if
wind speeds
exceed 20 miles
per hour. The
safety officer
and range safety
officer will
monitor the wind
speed prior to
launch
Section 8.8: Failure Modes
Hazard Cause Result Severity Probability Combined
Risk Mitigation
Improper
wiring
Faulty
connections or
mistaken
placement of
connections
The electronic
payload does
not behave as
expected or
does not
function at all
4 4 16
Wires will be
color-coded to
communicate
their function
and a specific
checklist will
be used to
correct wiring.
The checklist
will be
doubled
checked by the
systems lead
and recovery
lead.
96
Parachute
failure after
deployment
(tangled)
The parachute
is not packed
properly
The parachute
becomes
tangled on
descent
resulting in an
erratic and
fastmoving
projectile that
endangers
personnel and
property below
5 3 15
Packing of the
parachute will
be performed
by dedicated
members of the
recovery team
who will have
practiced
previously.
Parachute
deploys early
A faulty
altimeter fails
to detect the
altitude at
which the
parachute
should deploy
The rocket's
ascent will be
compromised
and its descent
will result in
the rocket
drifting for a
very long
distance
4 3 12
The altimeter
will be
thoroughly
tested prior to
its use in a
full-scale
capacity to
confirm that it
will function
as intended.
Safety
procedures will
be followed as
directed by the
USLI
handbook for
the recovery of
the vehicle.
97
Fire in
recovery
system
Excessive
black powder
and
insufficient
flame retardant
wadding
The descent of
the rocket is
accelerated and
the external
and internal
structure of the
rocket is
jeopardized.
Upon landing,
a flaming
parachute or
chords could
ignite brush.
4 3 12
Testing will be
done to
determine the
exact amount
of black
powder and
wadding
needed to
safely deploy
the parachute.
No more than
is needed will
be used. A fire
extinguisher
will be
available to
combat any
fires that may
occur once the
rocket lands. If
the fire spreads
or is
significantly
large,
authorities will
be contacted.
Safety
procedures
will be
followed as
directed by the
USLI
handbook for
the recovery of
the vehicle.
98
Rocket
sections poorly
coupled
The use of
weak bolts or
poorly
designed
manufactured
couplers
between
sections
Sections of the
rocket may
wobble and the
trajectory of
the rocket
could be
affected during
ascent or
during
recovery.
5 2 10
Extensive
testing of the
coupler will be
done prior to
sub-scale and
full-scale
launches. The
coupler will be
visually
inspected
before and
after assembly
by the team
leads, safety
officer, and
Range Safety
Officer.
99
Holes in the
airframe
Insufficient
communication
in addition to
excessive
drilling or
work on
components or
failure to
notice missing
pins or screws
The hole could
result in an
improper
reading of air
pressure by the
altimeter and
result in
premature
activation of
the recovery
system.
5 2 10
All sections of
the rocket will
be visually
inspected
immediately
after
construction,
before
transport to the
launch site,
and on
assembly.
Duplicates of
objects such as
pins, screws,
etc. will be
available to
replace any
missing ones.
Prelaunch
machining will
be kept to an
efficient
minimum with
our preflight
checklists and
assembly plan.
Motor fails on
launch
(explosion)
Manufacturing
defect
The rocket is
destroyed on
the launch pad
or shortly after
launch
5 2 10
Rocket motors
will only be
purchased
from a certified
source and will
be handled
with extreme
care
exclusively by
the team
mentor or by
someone with
permission of
the team
mentor.
100
Parachute
failure after
deployment
(tear)
Defects in the
parachute or
parachute bay
occurred
during
construction
The rocket's
descent will
not be slowed
as effectively
and could
endanger the
rocket or
personnel
5 2 10
The parachute
will be visually
inspected and
tested prior to
its utilization
in a full-scale
capacity and
upon assembly
on launch day.
The container
holding
parachute will
be smoothed to
not contain any
sharp edges.
All parachutes
will be
reinforced at
any potential
tear location.
101
Parachute
fails to deploy
a) A faulty
altimeter fails
to detect the
altitude at
which the
parachute
should deploy
b) Not enough
black powder
is used in the
recovery
system
The rocket
descends
chaotically at a
speed that his
extremely
dangerous to
both the rocket
and personnel.
5 2 10
a) A reliable
altimeter will
be selected
during the
PDR phase
and will be
tested prior to
launch in a
full-scale
capacity.
b) The amount
of black
powder that
will be used
will be
calculated by
team members
beforehand.
Calculations
will include
the amount
necessary and
the amount
allowable with
the final
amount used
lying
somewhere
within the
range. Use and
final
preparation of
black powder
charges will be
monitored by
the safety
officer.
102
Rocket blown
off course on
descent
a) Strong
winds on the
day of the
launch affect
descent more
than expected
b) A premature
parachute
deployment
causes the
rocket to be
subject to
more drift
Rocket could
become lost,
damaged, or
could endanger
observers.
3 3 9
The rocket will
not be
launched if
weather
conditions are
considered
dangerous by
either the team
or the range
safety officer.
All parts of the
rocket will
have a GPS
locater device
securely
attached to
facilitate
tracking during
and after
descent. Safety
procedures will
be followed as
directed by the
USLI
handbook for
the recovery of
the vehicle.
103
Insecure
aerodynamic
attachments
such as fins or
brakes
a) The epoxy
was mixed or
cured
improperly
b) The epoxy
used was not
strong enough
to withstand
forces
encounter in
flight
Fins may
vibrate and
cause
unexpected or
erratic changes
to the course of
the rocket.
This could
cause mission
failure and
potentially
endanger
personnel
4 2 8
Proper
procedures
regarding the
mixing and
curing of
epoxy will be
strictly
followed
during
construction of
the rocket.
During
assembly, team
members will
apply pressure
to the fins to
confirm they
do not move
and will not
move during
flight. If the
epoxy is not
sufficient,
steps will be
taken to fully
secure the fins
and if they
cannot, the
safety officer
will deem the
rocket unsafe
to launch.
104
Improper
coding
Improper
coding of the
microcontroller
controlling the
airbrake
system
Airbrakes do
not actuate as
expected
compromising
our maximum
altitude and
overall mission
4 2 8
The code that
will drive the
microcontroller
will be written
and reviewed
by multiple
team members
and tested on
the ground to
ensure that it
reacts in ways
it is meant to.
Improper
soldering or
board
manufacturing
Too much or
too little solder
is used when
constructing
the electrical
equipment
Electrical
malfunctions
and a loss of
system
integrity
leading to a
loss of
whichever
system the
electronics are
used in
4 2 8
The electrical
equipment will
be visually
inspected by
multiple team
members and
tests run to
ensure that it
carries
electrical
signals as
intended.
105
Rocket
descends to
rapidly
Design
oversight
causes the
rocket to fall
faster than
desired
The body of
the rocket will
be damaged
and potentially
the internal
components
damaged as
well. This
could violate
vehicle
requirement
1.4 and
jeopardize
mission
success.
4 2 8
The exact size
of the
parachute
needed to slow
down the
descent of the
rocket and the
timing of its
release will be
calculated and
sufficient
leeway given
to ensure that
recovery will
not threaten the
rocket or
personnel. All
members and
observers will
remain vigilant
until the rocket
is recovered
after landing.
106
Payload
(rover)
becomes
unstable
during flight
Payload is not
properly
secured within
its
compartment
With a moving
payload inside
the rocket, the
center of
gravity would
be constantly
changing. This
would cause
the rockets
flight to
become
unstable and
potentially
damage the
payload and
other
components
within the
rocket
4 2 8
The payload
will be housed
within a secure
bay inside the
rocket. It will
be placed
inside this bay
with a bulk
plate door to
hold it in place.
The payload is
also at the
bottom of the
rocket with the
opening facing
upward. If the
payload
dislodges, the
inertial forces
of launch
would keep it
in place inside
the bay.
107
Structure is
dropped and
damaged
during
construction,
assembly, or
in transport
Distracted or
clumsy
handlers that
are not aware
of their
surroundings
The body of
the rocket or
components in
the rocket may
be damaged by
the impact and
may require
replacement
3 2 6
Great care will
be taken when
working on
components
under all
conditions.
The
transportation
vehicle will
have a stable
carrying
structure for
the launch
vehicle.
During
transportation,
multiple
personnel will
carry the
rocket slowly
and carefully
while an
additional
team member
removes
obstacles or
opens doors as
necessary.
Replaceable
parts such as
pins, screws,
and the nose
cone will have
duplicate parts
available
during
assembly.
108
Motor fails on
launch (fails to
ignite)
a)
Manufacturing
defect
b) Failure of
the ignition
system
c) Delayed
ignition
a) and b) The
motor will not
fire and the
rocket will not
launch
c) The motor
will fire and
the rocket will
launch at an
unknown
amount of time
after the button
is pressed
3 2 6
In accordance
with the NAR
Safety Code,
the safety
interlock will
be removed or
the battery will
be
disconnected
and no team
member will
approach the
rocket for 60
seconds. After
60 seconds
without
activity the
safety officer
will approach
and check the
ignition
systems. In the
event that the
ignition
systems are not
at fault, the
motor will be
removed and
replaced with a
spare. A
second launch
will be
attempted if
there is time to
do so.
109
Servos used in
aerodynamic
systems do not
actuate
smoothly
a) Internal
electrical
failure
b) Worn gears
or slide rails
The science
payload may
not respond as
accurately as
expected
3 2 6
The servos that
will be used
for flight will
be purchased at
the beginning
of the project
and will be
stored in a
space away
from any
chemicals or
excessive
humidity. The
servos will be
tested before
transport,
before and
after assembly
to confirm that
they actuate
properly. The
airbrake
system will be
tested before
launch to
ensure the
accurate
response from
our system as a
whole.
110
Structural
integrity of
rocket
compromised
in flight
Excessive
aerodynamic
loading on the
airframe of the
rocket
Rocket may be
entirely lost
after flight
becomes
unstable and
recovery
systems may
be
compromised
5 1 5
Extensive
testing of the
materials and
structural
architecture of
the rocket body
will be done
before sub-
scale and full-
scale launches
to confirm that
the design will
withstand
forces that it
will encounter.
Cracked
airframe
Excessive
physical or
thermal
loading to the
rocket body
during storage
or transport
The rocket
body fractures
on launch or on
ascent
releasing
debris in the
immediate area
5 1 5
Sections of the
rocket body
will be kept in
a dry location
at room
temperature.
The rocket
body will be
visually
inspected
before and
after transport,
during
assembly, and
immediately
prior to launch
to confirm that
there are no
cracks. If
cracks are
found, the
launch vehicle
will be deemed
unsafe for
launch by the
safety officer.
111
Center of
gravity or
center of
pressure
misplaced
Rocket
property
calculations
were made
incorrectly or
with improper
data
The rocket's
ascent will be
unstable and
potentially
dangerous to
personnel and
equipment
5 1 5
Calculations
will be
checked
multiple times
prior to launch.
Subscale
launches will
provide an
opportunity to
confirm these
calculations
prior to full
scale launch.
Additionally,
the center of
gravity will be
physically
checked prior
to launch.
Misplaced or
lost
components
A messy or
disorganized
work
environment
leads to poor
tracking and
storage of
pertinent
rocket
components
An incomplete
rocket body is
transported or
ready for
assembly on
launch day.
Segments may
need to be
remanufactured
if they cannot
be located.
4 1 4
Once segments
of the rocket
body are
completed they
will be
immediately
stored in a
location
exclusively for
launch-ready
components.
For subscale
launches, some
components
will be
manufactured
twice so that
one may serve
as a backup.
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Rocket
exceeds mach
1 on ascent
The rocket
motor utilized
in the design is
too powerful
for the mass of
the rocket
Vehicle
requirement
1.19.7 is
violated,
compromising
the validity of
the mission
and an
infraction of
our launch
licensing
4 1 4
Team members
will
analytically
evaluate the
expected speed
of the rocket
prior to testing
and will
confirm these
results in sub-
scale and full-
scale testing.
In the event
that the mass
of the rocket is
too low,
additional
mass will be
added to the
inside of the
rocket to
ensure it does
not exceed
Mach 1. If the
mass cannot be
fixed in a safe
manner, the
rocket will be
deemed unsafe
to launch by
the safety
officer.
113
Motor is
physically
damaged
Motor was
damaged
during
handling or
transport due
to a drop or
insecure
transportation
Motor can
function
improperly or
potentially
explode due to
pressure forces
on damaged
area
3 1 3
The motor will
be primarily
handled by the
team mentor or
another
member
certified to do
so with the
permission of
the mentor or
safety officer.
The motor will
be checked
multiple times
before flight to
ensure no
damage has
been done
114
Section 9: Rover
The following deployable rover system and component designs have been chosen at this point in
the competition based on trade studies of alternative designs and testing that has proven the validity
of the designs. Most designs for the rover have stayed similar to the designs chosen and laid out in
the PDR.
The design of the rover has been broken down into mechanical and electrical sub-teams. This team
format has allowed those members who are interested in those aspects of design dedicate their time
to those fields. Integration of the mechanical and electrical sub-teams' work has been placed on
the deployable rover team lead.
Section 9.1: Mechanical Design
The rover mechanical design entails any part of the rover that is a structural component. This
design group includes the rover body, rover movement system and the SPDS.
An overview of the rover assembly is shown in figures Figure 9.1: Rover Overview “top side” and
Figure 9.2: Rover Overview “bottom side” below.
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Figure 9.1: Rover Overview “top side”
Figure 9.2: Rover Overview “bottom side”
Section 9.1.1: Body
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Figure 9.3: First Iteration of the Rover Body “top side”
Figure 9.4: First Iteration of the Rover Body “bottom side”
An I-beam shape has been chosen for the shape of the rover body. The initial design employed a
2-layered sandwich design but did not have enough space for the necessary components. A
comparison of the designs can be seen in Figure 3.1: Rover Body Comparison. The components
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that take up the most space on the rover are the electrical components. The I-beam design allows
for the placement of components on both sides of the I-beam web, giving the team more room to
work with to place components. The “top side” (orientation is dependent on how the rocket lands)
will house the Arduino used for control, the communication system, two motors for driving the
rover, and two 9V batteries for powering the Arduino and the motor shield (Figure 9.3: First
Iteration of the Rover Body “top side”). The “bottom side” will house the SPDS (panel tray, motor,
and solar panels) and the motor driver. Wiring from one side of the rover will be routed through a
hole located under the motor shield (Figure 9.4: First Iteration of the Rover Body “bottom side”).
Placement of the components was based on fitting the parts together closely and so that they would
not move around no matter the deployment orientation of the rover. On the top side, the motors
will stick through holes in the side wall of the rover and will be secured with a shell that the motors
will slip into. The Arduino will rest on the body with the communications system on top of that
(communication system is on a shield that attaches to the Arduino). The 9V batteries will be
slipped between two “compression walls” at the back end of the rover. On the bottom side, the
motor driver will rest on racks above the wire hole while the SPDS motor will also rest on a raised
platform housed inside another shell (shells not included on first iteration of rover body). The
SPDS tray will slide under rails seen closest to the motor platform in Figure 9.2: Rover Overview
“bottom side”. Since the body is a carbon fiber composite, the electrical conductivity of the Onyx
was tested. The carbon fiber based composite was determined to not be electrically conductive at
all.
Onyx, a printing material comprised of chopped carbon fiber in a nylon matrix, was chosen as the
material for the rover body as it is stronger than nylon and easier to manufacture than standard
carbon fiber. The team decided that, although weaker than carbon fiber, the strength of this material
is sufficient for the rover as it is not expected to experience large loads.
Dimensions of the rover were constrained by the size of the bay allotted for the rover. The team
decided that the rover needed to be able to deploy either right-side-up or upside-down, so the body
had to fit inside the height of the wheels/treads and be shorter than 10 in. (allotted space). The
initial body length of the rover did have to be modified due to “syncing” issues with the tread teeth
and the wheel holes. Dimensions can be seen below in Figure 9.5: Rover Body Dimensions.
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Figure 9.5: Rover Body Dimensions
Section 9.1.2: Movement
Figure 9.6: Rover Wheels and Tread
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Two initial designs where being considered. The first was a tread design and the second was a
four-wheel design. A tread design was chosen over a simple wheel design in order to provide
greater traction on rough terrain. The additional traction and support in between the wheels allows
the rover to easily deploy from the rocket this design can be seen in Figure 9.6: Rover Wheels and
Tread. The treads are 3D printed from a flexible nylon to allow for a customized size and spacing
of the wheels, teeth, and outer tread. The wheels are printed from the Onyx material.
The rover width and the body diameter set the wheel diameter. The teeth size was then determined,
and a MATLAB script used to calculate the wheel and teeth spacing. The script used is in Figure
9.7: MATLAB Script. After a test print of the tread and wheel it was discovered that there was not
enough clearance for the teeth to smoothly fit in the wheel during operation. These clearance issues
were addressed by adjusting the sizing of the holes in the wheels.
Figure 9.7: MATLAB Script
The two front wheels will drive the treads and are powered by two Pololu Micro Metal Gearmotors.
The back two wheels will provide tension to the tread and allowed to freely spin. The separate
motors allow the rover to be differentially steered.
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Figure 9.8: Rover Motor Inset
Section 9.1.3: Solar Panel Deployment System (SPDS)
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Figure 9.9: SPDS Stowed vs Deployed
Figure 9.10: SPDS Deployed Position Overview
The Solar Panel Deployment System (SPDS) consists of "accordion style" folding solar panels
combined with a gear and pinion deployment mechanism, as shown in Figure 9.9: SPDS Stowed
vs Deployed. The components of the SPDS are four solar panels, the solar panel tray, a gear, and
a motor which is identical to those used for rover movement. Solar panels will be deployed from
rear of the rover. The solar panels will be 3D printed panels with solar cells adhered to both sides.
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An "accordion style" folding mechanism was chosen for minimum storage footprint and low
complexity. The four solar panels will be fastened together by pins at each end which allow
rotation. The outermost panel will be pinned to the front of the solar panel tray, while the
innermost panel will be pinned to the rover body. Thus as the solar panel tray is pushed out the
innermost end of the panels will be held stationary while the outermost end is pulled out. This
results in the solar panels extending from the vertical stowed position to the horizontal deployed
position. Each solar panel will span the width of the rover body except for a small margin left to
allow movement, resulting in a panel width of 2.125 inches. A height of 0.75 inches was chosen
for the solar panels to maximize exposed area while maintaining proper clearance between the
body and the ground. Four panels of this height yield a lateral deployment of 3 inches, exceeding
the original Project Nova goal of 2 inches lateral deployment.
A gear and pinion driving mechanism was chosen for the SPDS because this method minimized
the volume footprint of the system. This small footprint allowed the storage of the motor driver
on the same side as the SPDS, reducing the overall length of the rover. The solar panel tray will
be deployed through a geared interface between the SPDS motor and the solar panel tray, as shown
in Figure 9.11: Solar Panel Tray, Motor, and Gear. The solar panel tray was design to pass under
the motor gear to further minimize the space required and to ensure firm contact between the tray
"tail" and motor gear. The motor will be electronically controlled; however, the tray "tail" was
also designed to limit the lateral deployment of the tray. The tail was sized such that the when the
SPDS has reached full deployment (3 inches) the motor will have reached the end of the "tail" and
no longer make geared contact. Thus, there is a mechanical safeguard against control malfunction.
The solar panel tray will slide through channels cut into the wall of the body which will keep the
tray stable and secure during deployment.
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Figure 9.11: Solar Panel Tray, Motor, and Gear
Section 9.2: Electrical Design
The rover electrical design entails any part of the rover that has an electrical aspect to it. This
design group includes the microcontroller, communication system, rover orientation, and how the
rover will be powered.
The rover will remain in the rocket until it is remotely activated from the team’s location. At that
point, the orientation will be read into the Arduino from the Adafruit 9DOF. Based on the
orientation of the rocket, the motors will spin one way or another in order to move the rover out
of the body of the rocket. Upon traveling the distance of the rover (distance needed for the rover
to be outside of the rocket), the rover will turn to the left or right and will then travel a
predetermined amount of time in order to ensure the proper distance has been travelled. This
distance will be measured in the amount of time it takes to travel that far. Once this time has been
reached, the rover will stop and the SPDS will deploy.
The electrical layout is shown in Figure 9.12: Rover Electrical Layout.
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Figure 9.12: Rover Electrical Layout
Section 9.2.1: Microcontroller
An Arduino Uno has been chosen to be the microcontroller, or the “brain”, of the rover based on
past experience with the unit. Powering the Arduino is ideal as it only requires a 9V battery which
does not take up much space and the board in general allows for the attachment of shields which
support our auxiliary components.
One of the biggest benefits of using Arduino is the large amount of codes, or sketches, that already
exist online that can be easily downloaded for different electrical components that we will be using.
Being able to draw on those sketches and simply integrate them to perform the ultimate goal will
save the team time and effort.
Section 9.2.2: Communication Method
The communication system basic requirements consist of only transmitting the initialization
sequence for the rover. This basic requirement indicates that only one-way communication is
required to our rover. An additional requirement is that we can deploy the rover a quarter mile.
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This is in addition to the low power requirements and limited size and weight of the receiver. To
efficiently meet these requirements and provide an easy interface, we decided on the X-Bee
module (Figure 9.13: X-Bee Module) which interfaces easily with our designated control board,
the Arduino Uno, with an easily installable, stackable shield module (Figure 9.14: X-Bee Shield
Module on Arduino).
Figure 9.13: X-Bee Module
Figure 9.14: X-Bee Shield Module on Arduino
The chosen modules have a rated range of one mile, sufficient to overcome our range requirements.
These features and the low power requirements of approximately 3.3V at 215 mA allow our
communication channel to be easily installed. In addition to these basic features, the X-Bee
contains additional optional features and advance networking capabilities. Such feature includes
basic IO for basic switched input and output in addition to serial communication to reprogram the
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two modules or interface with the Arduino. These modules have advance messaging support to
allow for advance networks to be built with routers and repeaters in addition to encryption support
to prevent tampering. These IO channels can be used to simplify communication with our rover
with a simple trigger mechanism for on-field deployment to advance functionalities such as a real-
time communication interface. In addition, these modules contain automatic transmission options
to aid possible transmission failures that may occur from signal degradation. With these modules,
we can program them to interface with each other exclusively despite other similar transceivers in
the area (Figure 9.15: X-Bee Setup). The software we are utilizing to configure these modules,
XCTU, also empowers us to commit range testing for our modules to ensure that these modules
will succeed in the field.
Figure 9.15: X-Bee Setup
Section 9.2.3: Orientation
During landing, the rocket will tip one of four ways due to the fins (the team is making the
assumption that the rocket will not tip over and have a fin stick into the ground perpendicularly).
Based on the orientation of the rocket, the motors will need to spin one way or another to prevent
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the rover from traveling back into the rocket. The team has chosen to use an Adafruit 9DOF
Breakout unit to determine the rover orientation inside the rocket. The team has past experience
with this unit.
The Adafruit 9DOF Breakout board consists of two sensors: an FXOS8700 3-axis accelerometer
and an FXAS21002 3-axis gyroscope. With these sensors, the orientation of the rover in the rocket
will be able to be determined.
Upon landing, the rover will be activated via the XBee communication system. Once the rover
receives the signal from the launch location, the orientation will be read into the Arduino.
The value of the z-vector (positive/negative) will determine the orientation the rocket is in. Based
on this value and the code, the motors will spin one direction or the other to move the rover out of
the rocket body.
Section 9.2.4: Power
To power both the motor shield and the Arduino, two 9V batteries were used. This decision was
based on ease of access and past experience using this model of motor shield and Arduino. An
Arduino Uno’s recommended operating voltage is 7-12V and the recommend voltage for the
Adafruit V2 motor shield is 5V-12V.
Something the team took into account was the possibility that other teams purposefully delay their
own launch in order to drain other teams’ batteries. A set of tests were completed where three
motors were run at full speed until the batteries died. After three hours, the batteries had drained
to a point where the motors were not providing enough torque even though they were still turning;
this was determined to be an acceptable safety factor by the team.
Prior to each launch, the 9V batteries will be replaced to ensure the best performance. The batteries
will be placed between walls that have been printed with the body that will hold them in place no
matter the orientation of the rover. The battery walls can be seen below in Figure 9.16: Rover
Battery Setup.
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Figure 9.16: Rover Battery Setup
Section 10: Altitude Control Module
The primary objective of the team this year is to accomplish precision apogee at a mile high is to
implement a system that provides the highest accuracy of altitude self-correction. Prior teams have
pursued a grid fin design to pursue this. This year’s team will be taking on a new system. In the
vein of variable input drag systems, just like past year’s grid fin designs, the team will be
implementing an entirely internal drag system. This design uses a single motor to actuate out
composite drag plates at variable degrees to produce the desired amount of drag on the vehicle.
All components in this system will be completely encapsulated within the vehicle. Slots in the
body will be cut to provide room for the actuation of the plates.
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Section 10.1: Drag Plates
The Auburn Student Launch Team has decided this year to pursue a different drag-producing
system this year. This year’s drag system will house flat drag plates that will actuate out from the
internals of the rocket. The plates will actuate perpendicular to the airflow and will receive a direct
moment of force from flight.
After boost phase, the internal drag system will read telemetry values and continuously self-correct
the vehicle’s trajectory. This will equalize the rocket’s trip to ascension to apogee. The current
design accounts for 3 internal plates that will carry the load of the system. An Arduino controller
will be the center of system input processing. A reading for the final altitude will continually be
calculated in order to more accurately deliver the vehicle to apogee. Included in the system will be
an IMU capable of reading in acceleration on all 3 axes as well as a barometric altimeter. These
inputs will drive the system’s self-correction. After the final altitude has been reached, the drag
plates will fully retract and the system will remain static.
The new design for this competition season was chosen due to simplicity and practicality. With
the current internal drag system in mind, the only mechanical interface needed is a single motor to
drive plate actuation. While last year’s system was proved to be effective, implementation of this
new design will provide returns in build time.
This year’s drag system was initially analyzed alongside last year’s grid fin design. The following
sections show a preliminary summary of these systems, along with corresponding system
characteristics and benefits.
Section 10.1.1: Internal Plate Drag System (IPDS)
The system that will be pursued for this year’s competition is the Internal Plate Drag System
(IPDS). The concept for this design was centered on the benefits of a modular system. The housing
for this system will be able to be predominantly printed with composite materials. The team printed
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components of the design in order to prove the feasibility of using additive techniques to produce
a prototype and ultimately a final product.
Figure 10.1: IPDS System Concept
Figure 10.2: IPDS Bottom-Up View
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Arms will interface with the drag plates, which revolve around the center point controlled by the
motor. These plates will be able to actuate from a completely revolved position all the way out to
a maximum determined by the controller.
These plates will be actuated out a variable degree in order to accurately bring the rocket to apogee.
The modularity of this design is a prime benefit of using this system. Preparation for launch will
be a small hassle since the housing can be dropped in with relative ease. Another benefit is the
lack of external components. This decreases drag on the system as a whole, since no fins or fairings
have to be attached. Another consideration is the ability to operate with just a single motor. This
allows for all arms to operate from a single stream of inputs.
A potential disadvantage of the system could be the weight added to rocket. Since the IPDS
requires extensive internal housing, the weight could add up quickly depending on the material
used for the structure.
Section 10.1.2: Wall Armed Fin-Lattice Elevator (WAFLE)
The Wall Armed Fin-Lattice Elevator (WAFLE) was the drag producing system of last year’s
AUSL team. This design was used as a primary alternative. Past experience with producing this
system lent a baseline of pros and cons for pursuing this system again. The design was pursued
due to its ability to both pursue the apogee requirement for the launch vehicle as well as the roll
induction experimentation option of last year’s competition. The structure is composed of a body
tube and external fairing.
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Figure 10.3: WAFLE
The inside of the external fairing would be hollow so that multiple servos can be housed and
actuated underneath each of the fairings. The system is comprised of two single-axis drag
producing servo-fin combos as well as two two-axis grid fins, which pursued the roll induction.
Due to the changing requirements of this year’s competition, the need for roll induction has been
taken out of the design considerations for this season.
A benefit of using this system is the weight. Due to the lattice design of the grid fins, the weight
of the system will be less of a factor since a solid fin would be heavier. An external fin system also
removes the need for extensive internals. Most of the effort here could be focused on the external
components of this system.
Since this system is largely external of the rocket, this brings the difficulty of reproducing the
system. This is not a modular design, so implementing it repeatedly every time another rocket is
built would increase overhead with system integration.
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Section 10.1.3: Spherical Joint Actuator (SJA)
A design that was considered as an alternative to the IPDS is the Spherical Joint Actuator (SJA).
The structure is composed of a body tube, spherical rotator, external housing, and an internal
housing.
Figure 10.4: Spherical Joint Actuator
The spherical rotator is placed in between the internal housing and the external structure and is the
main driver of the grid fin. A threaded rod will connect the spherical actuator to the grid fin. Servos
mounted to the internal housing will be directly connected to the spherical actuator by means of
rubber dipped rods and gears. These servos will control the pitch and roll of the grid fins.
Therefore, two servos will be required to control each fin. The external mount is formed into a
fairing shape to allow for uniform flow around the grid fins when they are pitched parallel to the
airframe.
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Figure 10.5: SJA External Fairing
Attaching this system to servos and actuating the system accurately is problematic with this
configuration. If the spherical actuator creates enough friction to prevent the SJA from moving
due to drag force, the servos necessary to move the grid fins on command would have to use more
force than servos used in competing systems. If the spherical actuator does not create enough
friction to prevent the SJA from moving freely, the drag force will prevent the SJA from deploying.
This system, however, allows for the SJA to have two degrees of freedom.
Section 10.1.4: System Comparison
Included below is a table of comparison weighing the value of both systems against each other in
order to determine a result.
Table 10.1: Drag System Comparison
Option 1 2 3
135
System Internal Plate Drag
System (IPDS)
Wall Armed Fin –
Lattice Elevator
(WAFLE)
Spherical Joint
Actuator (SJA)
Overall Volume 2 3 2
Optimization
Quality 3 2 3
Reliability 3 2 2
Cost 3 3 2
Total 11 10 9
The IPDS is the optimal path forward due to its modularity and its use of a single motor to interface
with all mechanical components. The small number of parts will allow the system be very simple
in implementation and reliable in flight. Once the system is tested, it can be used in flight. The
IPDS will be the system used to achieve the altitude goal.
Section 10.2: Drag Plate Deployment
The team is utilizing the single-motor benefit to its advantage. This allows for all 3 plates in the
system to be actuated out at the same degree relative to each other with minimal risk. There is less
risk with this than if each plate had its own mechanical driver (motor or servo). A ring connector
will interface with the rod of the motor, which connects the three arms that are attached to each
plate. When the controller reads input from the flight, this will trigger a responsive from the motor,
which spins a variable degree. This will move the arms that same degree around which will result
in the plates actuating outside of the rocket exterior producing a determined amount of drag on the
vehicle.
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Alternatively, each plate could have had its own motor or servo. As described above, this would
not be ideal. Implementing the system in this manner would add weight, complexity, and risk to
the system; Added weight due to the extra motors/servos, added complexity due to the need to talk
to each different driver, and added risk because of all of the extra components needed.
Section 10.3: Components
Section 10.3.1: Controller
The Arduino Uno was chosen as the right choice to implement into the IPDS. A selection table is
shown below to represent weighing factors of choosing a driver. The chosen controller contains
the right number of I/O pins needed for the project, as well as open source libraries included due
to a large development community. This significantly reduces the overheard when prototyping our
design. The controller will act as the decision maker for the system, determining the rate at which
the vehicle should self-correct its trajectory.
Table 10.2: Arduino Options
Board Operating Voltage
(V) Analog I/O
Digital I/O
Mega 5/7-12 16/0 54/15
Micro 5/7-12 8/0 20/7
Uno 5/7-12 6/0 14/6
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One of the most important considerations when choosing a controller for this system was size. The
Uno has an ideal size for integration with the IPDS. While the Micro takes up less space, it does
not provide the same ease-of-use that the Uno does. The Mega maintains the most I/O ports, but
the board takes up more system space. Due to unnecessary number of pins that comes with the
Mega, the Uno is the ideal choice. With a rocket inner diameter of 6 inches, size is an important
design consideration. A 9V battery will power the board.
Figure 10.6: Arduino Uno Microcontroller
Figure 10.7: IMU Breakout
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Table 10.3: Microcontroller Data
Values Adafruit IMU
Breakout
StratoLogger
Altimeter
SparkFun Triple-
Axis Gyro
Cost $29.95 $54.95 $24.95
Area Covered 1.35 1.68 0.6264
# of Other
Components Needed 0 2 2
Avg. Supply Voltage 7.5 1.5 2.85
Data Logged (DOF) 10 4 3
Table 10.4: Normalization of Data
Values 1 2 3 4 5 6
Cost >300 >200 >100 >50 >25 <25
Area
Covered >2 >1.5 >1.0 >0.5 >0.25 <0.25
# of Other
Components
Needed
2 1 0
Avg. Supply
Voltage >10 >8 >6 >4 >2 <2
139
Data
Logged
(DOF)
3 4 5 6 9 10
Table 10.5: Trade Study Weighting Factors
Item Factor Reason
Cost 3
The team may need other
parts depending on which
unit.
Area Covered 1 Space in the rocket must be
efficiently used.
# of Other Components
Needed 4
It needs to be taken into
consideration because we
might need to purchase
separate units vs just one
IMU.
Avg. Supply Voltage 2 The supply voltage affects the
power drain of the system.
Data Logged (DOF) 5
The more DOF, the more data
in one unit, which is a factor
for our data storage and
process capabilities.
140
Table 10.6: Sensor Trade Study (Normalized Values)
Values Adafruit IMU
Breakout
StratoLoggerCF
Altimeter
SparkFun Triple-
Axis Gyro
Cost 5 4 6
Area Covered 3 2 4
# of Other
Components Needed 6 1 1
Avg. Supply Voltage 3 6 5
Data Logged (DOF) 6 2 1
Table 10.7: Sensor Trade Study (Weighted Values)
Values Adafruit IMU
Breakout
StratoLoggerCF
Altimeter
SparkFun Triple-
Axis Gyro
Cost 15 12 18
Area Covered 3 2 4
# of Other
Components Needed 24 4 4
Avg. Supply Voltage 6 12 10
Data Logged (DOF) 30 10 5
Total 78 40 41
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Section 10.3.2: Motor
For our motor, which is driving the drag plate actuation, we chose to use the NeveRest 40 DC
Motor made from AndyMark. This motor is ideal because of its inexpensiveness coupled with its
effectiveness for our application compared to other DC motors. The NeveRest motor comes pre-
crimped while using eight-teen gauge wire, along with fifth-teen connectors. For the weight and
diameter of the motor it delivers the most power out of any other similar 40 DC motor. The body
and gears are manufactured with top of the line steel, which is vital for the durability of the system.
Figure 10.8: AndyMark NeveRest 40 DC Motor
Table 10.8: AndyMark NeveRest 40
Motor Specifications
Price $28.00
Weight 0.18lb
Pulse/Revolution 28
RPM 160
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Alternatives to our motor choice are displayed below. They were not chosen due to weight and
pricing.
Table 10.9: Uxcell 40 RPM 24V Motor (Alternative)
Motor Specifications
Price $90.00
Weight 2.87lb
Pulse/Revolution 25
RPM 40
Table 10.10: New Guanlian RE 40 (Alternative)
Motor Specifications
Price $90.00
Weight 2.87lb
Pulse/Revolution 25
RPM 40
Section 10.3.3: Electronics
The entire system is powered by 2 9V batteries in parallel. One will provide the power for the
Arduino controller, while the other provides power to the motor shield that will be attached to the
controller. The motor shield serves as a power regulator as well as a medium of communication
for the DC motor. Commands in the Arduino IDE can be used to communicate with the motor and
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control factors such as direction and rotations-per-minute. This will allow for better control of the
launch vehicle in flight.
Section 10.3.4: Interfaces
Three flat plates will be paired with a 9-volt battery, an Arduino Uno, and a single DC motor
embedded within the body of the launch vehicle. The plates as well as the housing for the plates
and electronics will be 3D printed using HIPS (high impact polystyrene). The internal system is
also more durable since it uses a honeycomb shape to maximize strength and minimize weight.
The motor used in this year’s design is a NeveRest Classic 40 Gearmotor. The benefit of using a
single motor is the simpler circuit compared to last year’s design, which used a total of six servos
to control four grid fins. The plates will be actuated outwards by the DC motor outwards in order
to increase drag. The horizontal deployment causes less resistance to the motor since it does not
force the fins against the drag directly.
Section 10.3.5: Precision of Instrumentation
The Arduino Uno will receive the required altitude information from an Adafruit 9-DOF IMU
Breakout unit. This unit was chosen based on its versatility as a comprehensive measurement
system: a 3-axis gyroscope (±250, ±500, or ±2000 degree-per-second scale), a 3-axis compass
(±1.3 to ±8.1 gauss magnetic field scale), a 3-axis accelerometer (±2g/±4g/±8g/±16g selectable
scale), and a barometric (300 - 1100hPa range)/temperature (-40 to 85 °C) gauge. The precision of
the control system, the inertial-measurement-unit, and the production process allows for accurate
repetition of measurement and production in all stages of this project. The data obtained from the
Adafruit 9-DOF IMU Breakout unit will be stored on an 8GB 9p SDHC Class 4 Secure Digital
Card attached to the Arduino Uno.
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Section 10.4: Dimensional Drawings
Figure 6.7: IPDS
Section 11: Project Plan
Section 11.1: Testing
In addition to the creation of requirements, it is essential to verify that they are satisfied. The pre-
competition full scale launches provide a method to check the function of all components, but
testing all systems at once introduces a large degree of risk and reduces the time to make changes
if needed. Therefore, although the goal is for all requirements to be verified through launching a
full-scale rocket with fully functional payloads, when possible components should be tested
previously and separately. Each of the following tests lists first the Auburn University
requirements that the test aims to confirm compliance with, the procedure used, and then reports
whether the test was determined to be a success.
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Section 11.1.1: Rover Battery and Motor Test (AU 4.4, 4.5)
Test Objective
This test aimed to verify the longevity of the electronics of the rover, in accordance with AU
requirements:
AU 4.4 Rover electronics must be able function after being left on for more than two hours.
AU 4.5 Rover motors must be able to function after more than two hours of battery drain.
This test was to be considered successful if the rover electronics were still functional after two or
more hours of intense battery drain.
Justification
Between time spent on the launch pad and time waiting to be deployed after landing, the rover
system will be experiencing power drain for a significant period of time. This power drain will
be less strenuous than two hours of continuous motor operation, however simply leaving the
system in its highest power consuming state provides a useful benchmark for the test. If the Rover
system could not sustain this level of operation, the Rover would need to be redesigned to ensure
mission completion.
Test components
-Arduino Uno
-New 9V batteries
-Rover Motors
-Multimeter
Procedure
1. Assemble the motors, control circuits, and batteries into the planned launch
configuration.
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2. Turn on the Arduino and the motors.
3. Measure the initial voltage from the batteries using the multimeter. Battery 1 was
connected to the Arduino, while battery 2 was providing power to the motors.
4. Take voltage measurements every thirty minutes, ending the test if any components
stopped functioning.
Results
Voltage Reading (Volts)
Time (AM) Battery 1 Battery 2
7:45 9.2 8.8
8:15 8.4 7.6
8:45 8 7
9:15 7.8 6.8
9:45 7.6 6.5
10:15 7.4 6.2
10:45 7 5.9
11:15 7 5.4
11:45 7 5.4
Although very simple, this test has provided some very important data. The motors were able to
run continuously for over three and a half hours off of a single 9V battery. This is a much higher
power drain than expected for simply standing idle in the launch configuration, so even in a case
of extreme battery usage the rover will still exceed AU requirement AU 4.5. The Arduino was
still functional after four hours, double AU 4.4 and four times NASA requirement 2.10 for vehicle
components endurance in the standby launch position.
Design Changes due to Test
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None, the Rover electronics met and exceeded all requirements for the test. This confirms that the
planned electronics setup can be used in the final design.
Section 11.1.2: Recovery and Altitude Control Battery Tests (AU 3.1, 6.9)
Test(s) Objective
Since these two tests are very similar, they have been combined here. They are both intended to
address NASA requirement 2.10, and the more strenuous AU requirements
AU 4.4 Recovery electronics must be able function after being left on for more than two hours.
AU 4.5 Altitude electronics must be able to function after more than two hours of battery drain.
These tests will be considered successful if the respective electronics are still operational after two
or more hours of intense battery drain.
Justification
Both electronics systems, like the rover, may need to standby in the on position for quite some
time on the launchpad before launch. The recovery system must still operate after this time for
mission completion. Although not as critical, the altitude control system must still operate as well
in order to not overshoot the mile altitude goal. If these systems cannot sustain this level of
operation, they will need to be redesigned to incorporate longer lasting batteries.
Test components
-Recovery altimeters
-Fully charged batteries
-Altitude control system Arduino Uno
-Multimeter
Procedure
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1. Assemble the motor (in the case of the altitude control system), control circuits, and
batteries into the planned launch configuration.
2. Turn on the Arduino and the motor or the two altimeters.
3. Measure the initial voltage from the batteries using the multimeter.
4. Take voltage measurements every thirty minutes, ending the test if any components
stopped functioning.
Results
Once the electronics in question are completed for full scale, these tests will be performed
Possible Design Changes due to Test
If any issues arise, they should be solvable by increasing the storage ability of the batteries used
by the systems.
Section 11.1.3: Full-Scale and Subscale Separation Test (AU 3.2)
Test Objective
This test is highly essential for both successful flight execution and safety. Ground Separation
testing was used to verify AU requirement
AU 3.2 Recovery system will be able to separate rocket into desired sections using the minimum
amount of black powder for reliable results, to ensure safety.
and NASA requirement
3.2 Each team must perform a successful ground ejection test for both the drogue and main
parachutes. This must be done prior to the initial subscale and full-scale launches.
Justification
Too little black powder, and the rocket will not separate, preventing parachute deployment, leading
to mission failure and more importantly a dangerous projectile. Too much black powder, however,
provides a fire and explosive hazard to team and launch personnel, and similarly can damage rocket
components. Therefore, it is important to use ground separation testing to determine the minimum
amount of black powder necessary to separate sections and eject the parachutes.
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This test must be performed for both the subscale and full-scale rocket. Although rules exist for
scaling and theoretically determining the required powder, the team prefers to use these methods
as a starting point to be checked with proper testing.
Test components
-Recovery Barometric Avionics Enclosure (BAE) structure
-Launch configuration upper or lower recovery section, secured with shear pins, including
● Main parachute (with drogue in upper section)
● Shock cord
● Recovery wadding
● Assembled black powder charges
● Nose cone or first lower body section coupler (depending on upper or lower
section)
-Electronic matches
-Ignition system
-Fire extinguisher (have not needed to use it, but always will be located beforehand)
Procedure
1. Fill charge cups according to equation established in the Recovery Section, or to
increased amount based on the result of the previous test/launch. (The equation from the
recovery section is an ideal case, and has been found from previous years to undershoot
the required amount).
2. Assemble structural components of the BAE with the electronic matches threaded
through to reach the charge cups. (Using the complete electronics system is possible but
not necessary)
3. Attach recovery section tube to the BAE, packing charges, wadding, parachutes, etc. as
they would be for a launch.
4. Seal the recovery section with the coupler or nose cone that completes the enclosure,
securing with shear pins.
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5. Place the assembled system on the ground or a test stand away from all personnel.
6. Attach the electronic matches to the power supply, and after verifying everyone is at a
safe distance, fire the charges.
7. If the components separate, record the amount of black powder used. If the charges do
not separate the components, disassemble the rocket and increase the charge.
Figure 11.1: Subscale Separation Test
Figure 11.1: Subscale Separation Test shows the subscale lower section prior to conducting
separation testing. The BAE is orange, on the right, and the wires to the electronic matches can
be seen trailing off to a further distance that all members retreated to in order to conduct the test
after this photo was taken.
Results
Separation testing was completed for the subscale rocket at the launch field in Samson, Alabama
on November 4, 2017. Through a series of tests, it was determined that both the upper and lower
sections recovery sections would require 4 grams of black powder to successfully separate the
subscale sections and deploy the parachutes. When flown in this configuration, the subscale was
successfully recovered.
A date and location for full scale separation testing has not yet been determined, as it is reliant on
the completion of several full scale components, such the the BAE.
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Design Changes due to Test
The 4 grams of black powder per charge cup determined still fit within the original charge cup
design, and all recovery components fit into the section, so no changes were necessary.
Section 11.1.4: Tension Testing of Composite and 3D Printed Material (AU 2.1,
2.2)
Test Objective
This test aimed to verify that the characteristics of all materials used to construct any portion of
the rocket are consistent with expected values, in accordance with AU requirements:
AU 2.1 Materials used to construct any portion of the rocket will undergo testing to ensure
…………….that materials characteristics are consistent with expected values.
AU 2.2 Materials used to construct any portion of the rocket will undergo testing to ensure
…………….that materials characteristics are consistent with expected values.
This test was to be considered successful if the materials tested show consistent results with
expected values.
Justification
In order to ensure that the composite materials used in the rocket body are capable of handling the
stresses involved in the launch and recovery, the materials properties must be determined. As the
properties of composite materials vary heavily depending on such factors as matrix orientation,
number of layers, and resin type, the properties of the specific composites the team will be using
must be determined via testing.
Test components
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-Onyx 3D-printed Carbon Fiber
-Epoxy Carbon Fiber
-Epoxy Fiberglass
Procedure
1. For the tensile testing, the team took the plates of Onyx 3D-printed Carbon Fiber, Epoxy
Carbon Fiber, and Epoxy Fiberglass, and measured their length, width, and depth using
dial calipers.
2. Once in the materials testing lab, the team outfitted the Instron Multipurpose testing
machine with the hardware appropriate for the thickness of each material.
3. After proper setup of the machine, measurements were taken again to ensure absolute
accuracy.
4. The material sample dimensions were then inputted into the computer interface.
5. The material was then placed between the grips of the apparatus and the chuck was
torqued to ensure proper grip on the material being tested
6. The load and strain were then zeroed out on the computer interface, after which the test
was commenced.
7. The data was then analyzed and plotted until the max stress and load of each specimen
was reached, and the sample torn, at which point the machine automatically ended the
test.
Results
Figure 11.2: Epoxy - Carbon Fiber Tension Test Results
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Figure 11.3: Epoxy - Fiberglass Tension Test Results
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Figure 11.4: Onyx Tension Test Results
This test provided the data that allows for the calculation of the maximum stress, modulus, tensile
strength and the tensile strain at the maximum load of all of the materials used in the structural
aspects of the rocket. This allows for the determination of the wall dimensions to allow for the
rocket to be stable in flight and during recovery. The stresses that the samples were put under were
much higher than would be seen in a normal flight/recovery pattern, proving the structural integrity
of the materials used.
Section 11.1.5: 3-Point Bend Testing of Composite and 3D Printed Material
(AU 2.1, 2.2, 6.7)
Test Objective
This test aimed to verify that the characteristics of all materials used to construct any portion of
the rocket are consistent with expected values, in accordance with AU requirements:
AU 2.1 Materials used to construct any portion of the rocket will undergo testing to ensure
…………….that materials characteristics are consistent with expected values.
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AU 2.2 Materials used to construct any portion of the rocket will undergo testing to ensure
…………….that materials characteristics are consistent with expected values.
AU 6.7 Drag plates must be able to withstand the perpendicular force of the airflow
This test was to be considered successful if the materials tested show consistent results with
expected values.
Justification
In order to ensure that the composite materials used in the rocket body are capable of handling the
stresses involved in the launch and recovery, the materials properties must be determined. As the
properties of composite materials vary heavily depending on such factors as matrix orientation,
number of layers, and resin type, the properties of the specific composites the team will be using
must be determined via testing.
Test components
-Onyx 3D-printed Carbon Fiber
-Epoxy Carbon Fiber
-Epoxy Fiberglass
Procedure
-The machine used for this test was the Instron Multipurpose testing machine.
1. First, five samples of each type of material were constructed and their length, width, and
thickness were measured using calipers for accuracy.
2. Once in the materials lab, measurements were taken again to ensure absolute accuracy
within the machine.
3. The material sample dimensions were then input into the computer interface.
4. The material sample was then placed within the apparatus and lined up by eye to center it
between two solid contact points. The machine was then positioned with accuracy using
various dials to ensure the contact arm was slightly in contact with the material species.
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5. The load and extension were zeroed then the user began the test within the interface of
the computer.
6. The data was then analyzed and plotted until the max flexure load of each specimen was
reached, at which point the machine automatically ended the test.
Results
Figure 11.5: Epoxy - Carbon Fiber Bend Test Results
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Figure 11.6: Epoxy - Fiberglass Bend Test Results
Figure 11.7: Onyx Bend Test Results
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This test provided the data that allows for the calculation of the maximum stress of all of the
materials used in the structural aspects of the rocket. This allows for the determination of the wall
thicknesses to allow for the rocket to be stable in flight and during recovery. The stresses that the
samples were put under were much higher than would be seen in a normal flight/recovery pattern,
proving the structural integrity of the materials used. This does not result in any design changes,
but instead was useful for design creation.
Section 11.1.6: Compression Testing of Composite and 3D Printed Material
(AU 2.1, 2.2, 6.7)
Test Objective
This test aimed to verify that the characteristics of all materials used to construct any portion of
the rocket are consistent with expected values, in accordance with AU requirements:
AU 2.1 Materials used to construct any portion of the rocket will undergo testing to ensure
…………….that materials characteristics are consistent with expected values.
AU 2.2 Materials used to construct any portion of the rocket will undergo testing to ensure
…………….that materials characteristics are consistent with expected values.
AU 6.7 Drag plates must be able to withstand the perpendicular force of the airflow
This test was to be considered successful if the materials tested show consistent results with
expected values.
Justification
In order to ensure that the composite materials used in the rocket body are capable of handling the
stresses involved in the launch and recovery, the materials properties must be determined. As the
properties of composite materials vary heavily depending on such factors as matrix orientation,
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number of layers, and resin type, the properties of the specific composites the team will be using
must be determined via testing.
Test components
-Onyx 3D-printed Carbon Fiber
-Epoxy Carbon Fiber
-Epoxy Fiberglass
Procedure
-The machine used for this test will be the Instron Multipurpose testing machine. Due to
equipment availability restraints, these tests could not be completed this year, and instead data
from previous years were used during construction.
1. First, five samples of each type of material were constructed and their length, width, and
thickness were measured using calipers for accuracy.
2. Once in the materials lab, measurements were taken again to ensure absolute accuracy
within the machine.
3. The material sample dimensions were then input into the computer interface.
4. The material sample was then placed within the apparatus and lined up by eye to center it
between two solid contact points. The machine was then positioned with accuracy using
various dials to ensure the contact arm was slightly in contact with the material species.
5. The load and extension were zeroed then the user began the test within the interface of
the computer.
6. As the progressively larger load was applied, the data was then analyzed and plotted until
each specimen reached -30% elongation, at which point the machine automatically ended
the test.
Results
Current results are not yet available, as this test is still only planned
Potential design changes
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Although unlikely, if this year’s data varies significantly from the previous structural data used to
design the rocket, design changes to reinforce vulnerable rocket components.
Section 11.1.7: Rover Maneuverability (AU 4.2, 4.3, 4.6, 4.7)
Test Objective
As outlined in requirements
AU 4.2 Rover can be activated remotely from a distance
AU 4.3 Rover must be able to traverse various expected terrains
AU 4.6 Rover must be able to exit the vehicle body from any orientation
AU 4.7 Rover will successfully deploy solar panels after travelling at least 5 feet
This test will aim to determine the cross country performance of the rover, and whether it will be
able to cross the farm field after leaving the rocket from any orientation. While already on testing
all these other aspects of the rover, it makes sense to test the remote activation as well. Success
will be defined as accomplishing all mission objectives.
Justification
The competition launch will take place on a farmer’s field. Cropland, whether left fallow or
planted in, will be very rough terrain for a small rover. Tufts of grass, grooves in the dirt from the
plow, eroded paths left by water runoff, and mud are all major obstacles when compared to the
size of the rover. To reach the desired distance, however, the rover must be able to cross these
obstacles. To complete its objective, the rover must then also be communicated with and deploy
solar panels.
Test Components
-Completed Rover
-Rocket rover bay (optional)
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-Shovel
-Measuring equipment
Procedure
The expected obstacles are characteristics shared by many open fields, including those used as
extra parking on Auburn’s Campus. Therefore, the rover can simply be taken to one of these
locations and activated to see if it can cross the terrain. If the team cannot find particular terrain
features outlined in AU 4.3, the shovel will be used to construct that feature.
After finished testing that capability, the lower section of the rocket, or at least the rover bay, can
be staged on the field as if it had just landed. AU 4.2 will be tested to see if the rover can be
activated remotely while it is inside the rocket. Once remote activation has been verified, the rover
will be remotely activated while the rover bay is rotated off the horizontal at various angles to
determine if it can leave the rocket after landing in any orientation and will be commanded to
deploy solar panels to verify AU 4.6 and 4.7.
Results
The rover is still under construction, so results are not yet available
Possible design impacts
Since the rover is made entirely of 3D printed components, the design can be rapidly iterated in
response to any issues arising from this series of tests. The most likely changes would be
adaptations of the tread design, or changes with regards to the wheels.
Section 11.1.8: Altitude Control System (AU 6.4 – 6.8)
Test Objective
This test aims to address several closely related team derived requirements for the altitude control
system, which are seen in the table below.
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AU
Requirement Description
AU 6.4 Subsystem components must engage accordingly after boost phase and stay online
for the remainder of the flight.
AU 6.5 Controller and IMU must be able to correctly predict the projected altitude of the
launch vehicle.
AU 6.6 Drag plates must deploy after boost phase in order to self-correct the trajectory of
the launch vehicle.
AU 6.7 Drag plates must be able to withstand the perpendicular force of the airflow.
AU 6.8 The altitude control system must be able to correct the vehicles altitude from
overshoot to 5280 ft.
To test all of these requirements at once, a special test 6” rocket will be constructed (time and
funds permitting). This rocket will be 6” diameter, like the actual full scale, but will be shorter,
possessing a weaker motor and greatly simplified recovery. The altitude control system will
control this rocket to a lower altitude of 3200 ft. The test will be considered a success if the altitude
control system can fulfill all the above requirements, albeit to a lower altitude.
If this cannot be tested in a separate rocket, these requirements will be verified by a launch of the
system inside the completed full scale rocket.
Justification
A major component of the competition is the altitude requirement. However, an altitude control
system can also be the most dangerous aspect of a rocket, even in the case of our team’s design
where an engineering control has been applied to prevent asymmetric drag. The system is also
difficult to meaningfully ground test. Therefore, by first testing the system on an expendable
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rocket with a lower apogee, the team can improve safety and avoid risking any of the other full
scale components, such as the isogrid body tube, that are difficult to replace at short notice.
Finally, more construction will provide additional experience for the team’s junior members.
Test Components
-Completed Altitude control system (designed to be easily added and removed from the rocket,
and retargetable for different altitudes)
-6” Diameter, 88” length test rocket with recovery system
-Aerotech K560W-P Motor
Procedure
Due to the similarity of this test to a full scale launch, the procedure for this test will be the same
as the procedure of a full scale launch, as seen in the launch checklist.
Results
This test has not yet been completed
Possible design impacts
If the altitude system over or under corrects, the software controlling the estimated drag plate drag
and therefore position will be adjusted. If the system cannot be safely implemented, it will be
eliminated from the final full-scale rocket.
Section 11.2: Requirements Verification
Section 11.2.1: General Requirements
Table 11.1: General Requirements Verification
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Requirement
Number Requirement Statement Verification Method
Execution of
Method
1.1
Students on the team will do
100% of the project,
including design,
construction, written reports,
presentations, and flight
preparation with the
exception of assembling the
motors and handling black
powder or any variant of
ejection charges, or preparing
and installing electric
matches (to be done by the
team’s mentor).
Demonstration
Students on the team
will do 100% of the
project.
1.2
The team will provide and
maintain a project plan to
include, but not limited to the
following items: project
milestones, budget and
community support,
checklists, personnel
assigned, educational
engagement events, and risks
and mitigations.
Demonstration
The team will
provide and maintain
a project plan.
1.3
Foreign National (FN) team
members must be identified
by the Preliminary Design
Demonstration
FN team members
have been identified
by the PDR.
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Review (PDR) and may or
may not have access to
certain activities during
launch week due to security
restrictions. In addition, FN’s
may be separated from their
team during these activities.
1.4
The team must identify all
team members attending
launch week activities by the
Critical Design Review
(CDR)
Demonstration
The team will
identify all members
attending launch
week by the CDR.
1.5
The team will engage a
minimum of 200 participants
in educational, hands-on
science, technology,
engineering, and mathematics
(STEM) activities, as defined
in the Educational
Engagement Activity Report,
by FRR. An educational
engagement activity report
will be completed and
submitted within two weeks
after completion of an event.
A sample of the educational
engagement activity report
can be found on page 31 of
the handbook. To satisfy this
Demonstration
The team will
complete the
Educational
Engagement
requirements.
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requirement, all events must
occur between project
acceptance and the FRR due
date
1.6
The team will develop and
host a Web site for project
documentation.
Demonstration
The team has
developed and will
host a Web site for
project
documentation.
1.7
Teams will post, and make
available for download, the
required deliverables to the
team Web site by the due
dates specified in the project
timeline.
Demonstration
The team will post
the required
deliverables to the
team Web site by the
due dates specified.
1.8 All deliverables must be in
PDF format Demonstration
All deliverables will
be in PDF format.
1.9
In every report, teams will
provide a table of contents
including major sections and
their respective sub-sections
Demonstration
Every report will
contain a table of
contents.
1.10
In every report, the team will
include the page number at
the bottom of the page.
Demonstration
Every report will
include page
numbers.
1.11
The team will provide any
computer equipment
necessary to perform a video
Demonstration
The team will
provide all necessary
equipment for the
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teleconference with the
review panel. This includes,
but is not limited to, a
computer system, video
camera, speaker telephone,
and a broadband Internet
connection. Cellular phones
can be used for speakerphone
capability only as a last
resort.
video
teleconferences.
1.12
All teams will be required to
use the launch pads provided
by Student Launch’s launch
service provider. No custom
pads will be permitted on the
launch field. Launch services
will have 8 ft. 1010 rails, and
8 and 12 ft. 1515 rails
available for use.
Demonstration
The team will use the
launch pads
provided.
1.13
Teams must implement the
Architectural and
Transportation Barriers
Compliance Board Electronic
and Information Technology
(EIT) Accessibility Standards
(36 CFR Part 1194)
Demonstration
The team will
implement the EIT
Accessibility
Standards.
1.14 Each team must identify a
“mentor.” A mentor is Demonstration
The team has
identified a mentor.
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defined as an adult who is
included as a team member,
who will be supporting the
team (or multiple teams)
throughout the project year,
and may or may not be
affiliated with the school,
institution, or organization.
The mentor must maintain a
current certification, and be in
good standing, through the
National Association of
Rocketry (NAR) or Tripoli
Rocketry Association (TRA)
for the motor impulse of the
launch vehicle and must have
flown and successfully
recovered (using electronic,
staged recovery) a minimum
of 2 flights in this or a higher
impulse class, prior to PDR.
The mentor is designated as
the individual owner of the
rocket for liability purposes
and must travel with the team
to launch week. One travel
stipend will be provided per
mentor regardless of the
number of teams he or she
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supports. The stipend will
only be provided if the team
passes FRR and the team and
mentor attends launch week
in April.
Section 11.2.2: Vehicle Requirements
Table 11.2: Vehicle Requirements Verification
Requirement
Number Requirement Statement
Verification
Method Execution of Method
2.1
The vehicle shall deliver
the science or
engineering payload to
an apogee altitude of
5,280 feet above ground
level (AGL).
Analysis
Demonstration
Testing
The vehicle will be designed
to reach 5,280 ft AGL, test
launches will be performed,
and altimeters will be
assessed post launch.
2.2
The vehicle shall carry
one commercially
available, barometric
altimeter for recording
the official altitude used
in determining the
altitude award winner.
Inspection
Demonstration
The team has purchased and
calibrated several
commercially available
altimeters.
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2.3
Each altimeter will be
armed by a dedicated
arming switch that is
accessible from the
exterior of the rocket
airframe when the rocket
is in the launch
configuration on the
launch pad.
Inspection
Demonstration
The team will make the
altimeters capable of being
armed from the exterior of
the rocket airframe in launch
configuration.
2.4
Each altimeter will have
a dedicated power
supply.
Inspection
Demonstration
The altimeters will have
dedicated power supplies.
2.5
Each arming switch will
be capable of being
locked in the ON position
for launch (i.e. cannot be
disarmed due to flight
forces).
Demonstration
The team will use an arming
switch that cannot be
disabled due to launch
forces.
2.6
The launch vehicle shall
be designed to be
recoverable and reusable.
Reusable defined as
being able to launch
again on the same day
without repairs or
modifications.
Testing
Analysis
Demonstration
Inspection
Trajectory simulations and
testing will ensure the launch
vehicle is recoverable and
reusable
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2.7
The launch vehicle shall
have a maximum of four
(4) independent sections.
Demonstration
Team will design and build a
launch vehicle that can have,
but does not require, four
independent sections.
2.8
The launch vehicle shall
be limited to a single
stage.
Demonstration Team will design and build a
single-stage launch vehicle.
2.9
The launch vehicle shall
be capable of being
prepared for flight at the
launch site within 3
hours, from the time the
Federal Aviation
Administration flight
waiver opens.
Demonstration
Team will be timely and
organized to ensure vehicle
is prepared on time.
2.10
The launch vehicle shall
be capable of remaining
in launch-ready
configuration at the pad
for a minimum of 1 hour
without losing the
functionality of any
critical on-board
component.
Demonstration
Team will design vehicle
with ability to remain
launch-ready for at least one
hour.
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2.11
The launch vehicle shall
be capable of being
launched by a standard
12 volt direct current
firing system. The firing
system will be provided
by the NASA-designated
Range Services Provider.
Demonstration
The team will use systems
compatible with a standard
firing system.
2.12
The launch vehicle shall
require no external
circuitry or special
ground support
equipment to initiate
launch (other than what
is provided by Range
Services).
Demonstration
The team has designed a
vehicle requiring no external
circuitry or special ground
support equipment
2.13
The launch vehicle shall
use a commercially
available solid motor
propulsion system using
ammonium perchlorate
composite propellant
(APCP) which is
approved and certified by
the National Association
of Rocketry (NAR),
Tripoli Rocketry
Association (TRA),
Demonstration
Vehicle will be designed
around commercially
available, certified motors
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and/or the Canadian
Association of Rocketry
(CAR).
2.13.1
Final motor choices must
be made by the Critical
Design Review (CDR).
Demonstration A final motor choice has
been made.
2.13.2
Any motor changes after
CDR must be approved
by the NASA Range
Safety Officer (RSO),
and will only be
approved if the change is
for the sole purpose of
increasing the safety
margin.
Demonstration
If the change is made to
increase safety margin,
NASA RSO will allow the
change
2.14
Pressure vessels on the
vehicle shall be approved
by the RSO and shall
meet the following
criteria:
Demonstration Any pressure vessels will be
approved by the RSO.
2.14.1
The minimum factor of
safety (Burst or Ultimate
pressure versus Max
Expected Operating
Pressure) shall be 4:1
Inspection
Analysis
Testing
Team will design the
pressure vessels to have a
factor of safety of 4:1.
174
with supporting design
documentation included
in all milestone reviews.
2.14.2
Each pressure vessel
shall include pressure
relief valve that sees the
full pressure of the tank.
Demonstration Pressure vessels will include
a pressure relief valve.
2.14.3
Full pedigree of the tank
shall be described,
including the application
for which the tank was
designed, and the history
of the tank, including the
number of pressure
cycles put on the tank, by
whom, and when.
Inspection
Demonstration
The team will inspect the
tank along with
documentation of testing and
history
2.15
The total impulse
provided by a College
and/or University launch
vehicle shall not exceed
5,120 Newton-seconds
(L-class).
Demonstration
Analysis
The team has choosen a
motor with a total impulse
that does not exceed 5,120
Newton-seconds (L-class).
2.16
The launch vehicle shall
have a minimum static
stability margin of 2.0 at
the point of rail exit.
Testing
Demonstration
Analysis
The team will design and test
the vehicle to ensure that it
has a stability margin of 2.0
at the point of rail exit.
175
2.17
The launch vehicle shall
accelerate to a minimum
velocity of 52 fps at rail
exit.
Demonstration
Analysis
Testing
The team will design the and
test the vehicle to ensure that
it’s minimum velocity at rail
exit is at least 52 fps.
2.18
All teams shall
successfully launch and
recover a subscale model
of their rocket prior to
CDR.
Demonstration
Testing
A successful subscale model
has been launched and
recovered.
2.18.1
The subscale model
should resemble and
perform as similarly as
possible to the full-scale
model, however, the full-
scale shall not be used at
the subscale model.
Demonstration
The subscale model was
designed to resemble and
perform similarly to the full
scale model.
2.18.2
The subscale model shall
carry an altimeter
capable of reporting the
model’s apogee altitude.
Demonstration
An altimeter capable of
reporting the model’s apogee
altitude was implemented on
the subscale model.
2.19
All teams shall
successfully launch and
recover their full-scale
rocket prior to FRR in its
final flight configuration.
The rocket flown at FRR
must be the same rocket
Testing
Demonstration
Analysis
A test of the full-scale rocket
will be exhibited,
demonstrating all hardware
functions properly.
176
flown on launch day. The
following criteria must
be met during the full
scale demonstration
flight:
2.19.1
The vehicle and recovery
system shall have
functioned as designed.
Testing
Testing of vehicle will show
how recovery system
functions.
2.19.2.1
If the payload is not
flown, mass simulators
shall be used to simulate
the payload mass.
Testing
Demonstration
Analysis
Payload will be flown.
2.19.2.1.1
The mass simulators
shall be located in the
same approximate
location on the rocket as
the missing payload
mass.
Inspection
Inspection of the rocket
payload will be done by the
team to ensure it is properly
placed.
2.19.3
If the payload changes
the external surfaces of
the rocket (such as with
camera housings or
external probes) or
manages the total energy
of the vehicle, those
systems shall be
Demonstration
The payload does not change
the external surface of the
vehicle.
177
activated during the full-
scale demonstration of
flight.
2.19.4
The full-scale motor does
not have to be flown
during the full-scale test
flight.
Demonstration
The full-scale motor may or
may not be flown during the
full-scale test flight.
2.19.5
The vehicle shall be
flown in its fully
ballasted configuration
during the full-scale test
flight.
Demonstration The vehicle will be fully
ballasted during test flights.
2.19.6
After successfully
completing the full-scale
demonstration flight, the
launch vehicle or any of
its components shall not
be modified without the
concurrence of the
NASA Range Safety
Officer (RSO).
Demonstration
The team will demonstrate
that it did not alter any
components or vehicle after
demonstration flight.
2.19.7
Full scale flights must be
completed by the start of
FRRs (March 6th, 2018).
If necessary, an extension
to March 28th, 2018 will
be granted. Only granted
for re-flights.
Demonstration A full-scale flight will be
completed by FRRs.
178
2.20
Any structural
protuberance on the
rocket shall be located aft
of the burnout center of
gravity.
Demonstration
Any structural protuberances
on the vehicle will be aft of
the burnout center of gravity.
2.21.1
The launch vehicles shall
not utilize forward
canards.
Demonstration The team will not utilize
forward canards.
2.21.2
The launch vehicle shall
not utilize forward firing
motors.
Demonstration The team will not utilize
forward firing motors.
2.21.3
The launch vehicle shall
not utilize motors that
expel titanium sponges
(Sparky, Skidmark,
MetalStorm, etc.)
Demonstration
The team will not utilize a
motor that expels titanium
sponges.
2.21.4 The launch vehicle shall
not utilize hybrid motors. Demonstration
The team will not utilize a
hybrid motor.
2.21.5
The launch vehicles shall
not utilize a cluster of
motors.
Demonstration
A demonstration and
inspection of the launch
vehicle shall be carried out to
179
validate it does not use a
cluster of motors.
2.21.6
The launch vehicle shall
not utilize friction fitting
for motors.
Demonstration
The team will design the
vehicle so that it does not
utilize friction fitting for the
motor.
2.21.7
The launch vehicle shall
not exceed Mach 1 at any
point during flight.
Demonstration
Testing
Analysis
The team will design, test,
and demonstrate vehicle
performance to ensure that
the vehicle does not exceed
Mach 1 at any point during
flight.
2.21.8
Vehicle Ballast shall not
exceed 10% of the total
weight of the rocket.
Demonstration
Testing
Analysis
The team will design ballast
so that it does not exceed
10% of the total weight of
the rocket.
Section 11.2.3: Recovery Requirements
Table 11.3: Recovery Requirements Verification
Requirement
Number Requirement Statement Verification Method
Execution of
Method
180
3.1
The launch vehicle shall
stage the deployment of its
recovery devices, where a
drogue parachute is deployed
at apogee and a main
parachute is deployed at a
much lower altitude. Tumble
recovery or streamer
recovery from apogee to
main parachute deployment
is also permissible, provided
that kinetic energy during
drogue-stage descent is
reasonable, as deemed by the
Range Safety Officer.
Demonstration
Testing
The team will stage
the deployment of
our recovery devices
with a drogue
parachute deployed at
apogee (5280 ft.), and
two main parachute
deployed at 750 ft.
3.2
Each team must perform a
successful ground ejection
test for both the drogue and
main parachutes. This must
be done prior to the initial
subscale and full scale
launches.
Testing
Prior to the initial
subscale and full
scale launches, the
team will perform a
ground ejection test
for both the drogue
and upper main
parachute section as
well as the lower
main parachute
section.
181
3.3
At landing, each independent
sections of the launch vehicle
shall have a maximum
kinetic energy of 75 ft-lbf.
Analysis
Demonstration
The team will
calculate and test
both sections of our
launch vehicle to
ensure that a
maximum energy of
75 ft-lbf at landing is
not exceeded.
3.4
The recovery system
electrical circuits shall be
completely independent of
any payload electrical
circuits.
Demonstration
The team will create
independent circuits
for our recovery
system so that they
are independent of
any payload electrical
circuits.
3.5
All recovery electronics will
be powered by commercially
available batteries.
Demonstration
The recovery system
will feature to
commercially
available 9V batteries
to power the two
altimeters.
3.6
The recovery system shall
contain redundant,
commercially available
altimeters. The term
“altimeters” includes both
simple altimeters and more
Demonstration
The recovery system
will include a
TeleMetrum and
TeleMega altimeter
each with their own
independent set of
182
sophisticated flight
computers.
charges for each
section.
3.7
Motor ejection is not a
permissible form of primary
or secondary deployment.
Demonstration
The team will not use
motor ejection as a
primary or secondary
deployment. An
electronic form of
deployment will be
used.
3.8
Removable shear pins will be
used for both the main
parachute compartment and
the drogue parachute
compartment.
Demonstration
The team will use
removable shear pins
for both the main
parachute
compartment and the
drogue parachute
compartment.
3.9
Recovery area will be limited
to a 2500 ft. radius from the
launch pads.
Demonstration
The main parachutes
will deploy from a
low enough altitude
so that the rocket will
not drift more than
2500 ft.
183
3.10
An electronic tracking device
shall be installed in the
launch vehicle and shall
transmit the position of the
tethered vehicle or any
independent section to a
ground receiver.
Demonstration
The team will install
a tracking device on
both sections of the
launch vehicle so that
the location of both
pieces can be
determined after
landing.
3.10.1
Any rocket section, or
payload component, which
lands untethered to the
launch vehicle, shall also
carry an active electronic
tracking device.
Demonstration
All separating
sections of the rocket
will contain their own
tracking device.
3.10.2
The electronic tracking
device shall be fully
functional during the official
flight on launch day
Verification
The team will test
and make sure the
electronic tracking
devices will be fully
functional during the
official flight.
3.11
The recovery system
electronics shall not be
adversely affected by any
other on-board electronic
devices during flight (from
launch until landing).
Testing
The team will test
and make sure that
the recovery system
will not be adversely
affected by any other
184
on-board electronic
devices during flight.
3.11.1
The recovery system
altimeters shall be physically
located in a separate
compartment within the
vehicle from any other radio
frequency transmitting
device and/or magnetic wave
producing device.
Demonstration
The recovery system
altimeters will be
located in a separate
compartment within
the vehicle from any
radio frequency
transmitting devices,
and magnetic wave
producing devices.
3.11.2
The recovery system
electronics shall be shielded
from all onboard transmitting
devices, to avoid inadvertent
excitation of the recovery
system electronics.
Demonstration
The recovery
electronics will be
sealed in their own
separate compartment
separate from all
other transmitting
devices in the rocket.
3.11.3
The recovery system
electronics shall be shielded
from all onboard devices
which may generate
magnetic waves (such as
generators, solenoid valves,
and Tesla coils) to avoid
Demonstration
The recovery
electronics will be
sealed in their own
separate compartment
separate from all
other magnetic wave
inducing devices in
the rocket.
185
inadvertent excitation of the
recovery system.
3.11.4
The recovery system
electronics shall be shielded
from any other onboard
devices which may adversely
affect the proper operation of
the recovery system
electronics.
Demonstration
The recovery
electronics will be
sealed in their own
separate compartment
separate from all
other devices in the
rocket which could
adversely affect the
proper operation of
the recovery system.
Section 11.2.4: Deployable Rover Requirements
Table 11.4: Deployable Rover Requirements Verification
Requirement
Number
Requirement
Statement
Verification
Method Execution of Method
4.5.1
Teams will design a
custom rover that will
deploy from the internal
structure of the launch
Demonstration,
test, inspection
Demonstration and testing
before the launch will ensure
the rover can fit inside the
rocket and is fit to maneuver
difficult terrain
186
4.5.2
At landing, the team will
remotely activate a
trigger to deploy the
rover from the rocket
Demonstration,
test, inspection
Demonstration and testing will
be done before the flight to
validate that the rover will
receive the deployment signal.
Inspection afterward will also
ensure successful rover
deployment
4.5.3
After deployment, the
rover will autonomously
move at least 5 ft from
the launch vehicle
Demonstration,
test, inspection
Demonstration and testing will
be conducted before the launch
to determine the exact distance
the rover will travel
4.5.4
Once the rover has
reached its final
destination, it will deploy
a set of foldable solar cell
panels
Demonstration,
test, inspection
Demonstration and testing will
be performed on the solar panel
deployment system (SPDS) to
ensure that the system can
successfully deploy the solar
panels
Section 11.2.5: Safety Requirements
Requirement
Number Requirement Statement
Verification
Method Execution of Method
5.1
Each team will use a launch
and safety checklist. The final
checklists will be included in
Demonstration The team will use
checklists.
187
the FRR report and used
during the Launch Readiness
Review (LRR) and any
launch day operations.
5.2
Each team must identify a
student safety officer who
will be responsible for all
items in section 5.3.
Demonstration The team has identified a
safety officer.
5.3
The role and responsibilities
of each safety officer will
include, but not limited to: a
bunch of things.
Demonstration The safety officer is aware
of his responsibilities.
5.4
During test flights, teams will
abide by the rules and
guidance of the local rocketry
club’s RSO. The allowance of
certain vehicle configurations
and/or payloads at the NASA
Student Launch Initiative
does not give explicit or
implicit authority for teams to
fly those certain vehicle
configurations and/or
payloads at other club
launches. Teams should
communicate their intentions
to the local club’s President
or Prefect and RSO before
Demonstration The team will follow the
rules.
188
attending any NAR or TRA
launch.
5.5 Teams will abide by all rules
set forth by the FAA. Demonstration
The team will follow the
rules.
Section 11.3: Team Requirements
The following are tables of team-derived requirements and the associated verification methods.
More information on planned or completed tests necessary to the verification of these requirements
can be seen in Section 11.1: Testing.
Section 11.3.1: General Requirements
Table 11.5: AU General Requirements
Team Requirement Requirement
Statement Verification Method
Method of
Execution
AU 1.1
All Educational
Engagement forms
will be submitted and
verified to have been
received within 1
week of an outreach
event.
Demonstration
The team will
submit EE reports
within a week and
check that they have
been properly
received.
189
Section 11.3.2: Vehicle Requirements
Table 11.6: AU Vehicle Requirements
Team
Requirement Requirement Statement Verification Method
Method of
Execution
AU 2.1
Materials used to
construct any portion of
the rocket will undergo
testing to ensure that
materials characteristics
are consistent with
expected values.
Test
Tension and 3-point
bend tests have
already been
completed.
Compression tests
will be conducted
soon.
AU 2.2
3D printed components
will have strength
comparable to
alternatives and
appropriate for their role
Test
Tension and 3-point
bend tests have
already been
completed.
Compression tests
will be conducted
soon.
AU 2.3
Isogrid structure strength
will be verified to be in
line with or superior to
filament wound carbon
fiber tubes.
Test
Samples of filament
wound material and
Isogrid structures will
be created and tested
under various loading
conditions.
190
Section 11.3.3: Recovery Requirements
Table 11.7: AU Recovery Requirements
Team
Requirement Requirement Statement
Verification
Method Method of Execution
AU 3.1
Recovery electronics must be
able function after being left
on for more than two hours.
Test
The recovery electronics
(excluding e-matches and
black powder for safety
reasons) will be assembled
and left in the on/standby
position until loss of function
to determine system longevity
AU 3.2
Recovery system will be able
to separate rocket into desired
sections using the minimum
amount of black powder for
reliable results, to ensure
safety.
Test
Extensive ground separation
tests will be performed before
launch: subscale tests have
been completed and full-scale
tests will be.
Section 11.3.4: Deployable Rover Requirements
Table 11.8: AU Rover Requirements
Team
Requirement Requirement Statement
Verification
Method Method of Execution
191
AU 4.1
With the exception of
electronic parts, the rover will
be 3-D printed in-house
Demonstration
Rover parts will be
designed to be printed in-
house so that the team can
continue to precisely
manufacture rover parts
AU 4.2
Rover activation signal will
successfully reach the rover
up to a certain distance plus a
large tolerance distance
Demonstration
XBee will be tested at
different distances to ensure
that the signal will reach the
rover
AU 4.3
Rover will be able to traverse
various terrains (examples
include 45 degree inclines
and 3 inch divots)
Demonstration
Test
Rover treads will be tested
on various terrains before
the launches to ensure
adequate traction
AU 4.4
Rover electronics must be
able function after being left
on for more than two hours.
Test
The rover electronics were
run continuously for several
hours and have been proven
to be functional for more
than two hours.
192
AU 4.5
Rover motors must be able to
function after more than two
hours of battery drain.
Test
The rover motors were run
continuously for several
hours and have been proven
to be functional for more
than two hours.
AU 4.6
Rover must be capable of
exiting rocket from any
orientation of rocket body.
Test
Rover will be placed in
section, rotated to various
angles in increments of 30
degrees from horizontal,
and be commanded to drive
out.
AU 4.7
Rover will successfully
deploy solar panels after
travelling at least 5 feet
Demonstration,
test
SPDS will be shown to
deploy after travelling over
various terrain and
travelling varying distances
Section 11.3.5: Safety Requirements
Team
Requirement Requirement Statement
Verification
Method
Method of
Execution
AU 5.1
All team members will work
in groups of at least two,
ensuring immediate assistance
for any team member in need.
Demonstration Team members will
not work alone.
193
Section 11.3.6: Altitude Control Requirements
Table 11.9: AU Altitude Control Requirements
Team
Requirement
Requirement
Statement
Verification
Method Method of Validation
AU 6.1
All aerodynamic data
must be validated
through analytical and
experimental testing.
Analysis
An aerodynamic analysis of the drag plates
and the internal system will be conducted
through computational fluid dynamics
(CFD) and sensor tests.
AU 6.2 Drag plates must stay
static throughout launch Demonstration
Timer will be implemented through the
controller that will prevent any system
action prior to the end of the boost phase.
AU 6.3
Electronics must stay
secured throughout
flight.
Demonstration A housing will be made for all electrical
components to keep everything in place.
AU 6.4
Subsystem components
must engage
accordingly after boost
phase and stay online
for the remainder of the
flight.
Test Testing will verify that the behavioral
integrity of the system remains intact.
194
AU 6.5
Controller and IMU
must be able to
correctly predict the
projected altitude of the
launch vehicle.
Test
Testing will verify that both components
will behave according to the requirement
specification.
AU 6.6
Drag plates must
deploy after boost
phase in order to self-
correct the trajectory of
the launch vehicle.
Test “Hardware In-The-Loop” testing and test
flights will validate deployment precision.
AU 6.7
Drag plates must be
able to withstand the
perpendicular force of
the airflow.
Test
Wind tunnel testing and structural testing
will be conducted to ensure the integrity of
the plates and the materials used to
construct them.
AU 6.8
The subsystem must be
able to correct altitude
by at least 400 feet and
be accurate to 200 feet.
Analysis
Demonstration
The subsystem will be designed to conform
to accuracy requirements and will then be
demonstrated to be accurate in a test flight.
AU 6.9
Altitude control
electronics must be able
to function after being
left on for more than
two hours.
Test
The batteries, motors, and Arduino with all
relevant electronics were assembled and
ran as intended for several hours,
confirming the electronics ability to
function for more than two hours.
195
Section 11.4: Budget
The budgets displayed in Table 11.13: Budget Allocation are an updated approximation of the
costs of the remaining expenditures, and a representation of already allocated funds for the project.
The approximations are conservative, assuming excess quantities of materials and no price breaks.
Hoping to bring a large number of students to the competition this year, the team has already
reserved 6 hotel rooms from the block. The Auburn Aerospace Engineering department has
informed us that we should expect to pay $3,000 for that, but that they will try to lessen that sum
as they can. Thanks to the success of the subscale on its first flight, the team did not have to do
any rebuilding, and all the electronic components are reusable. Factoring this is to the initial
estimate of $2,700 for the subscale, the actual cost of the subscale came out to $2,000. Our
educational outreach has spent $300 so far, and is estimated to cost $1,200 more. Assuming
$3,055.13 for the rocket on the pad, $2,000 for the sub-scale vehicle, and $3,000 for travel, this
leaves $8,444.87 for promotional items, overhead costs and any other testing and development
costs, based on the $16,500 amount for total funding presented in Table 11.14: Funding Sources.
Table 11.10: Vehicle Costs
Vehicle (Full Scale)
Item Cost Per Unit Unit Quantity Total
Carbon fiber and
resin for open
weave structure.
$284 Per tube 2 $568
Fiber glass sleeve $33 Per tube 2 $66
Pre-preg Carbon
Fiber $118 Per yard 5 $590
196
Aerotech L1420R $259.99 Per unit 1 $259.99
RMS 75/3840
Motor Case and
Associated
Hardware
$385 Per unit 1 $385
Rail Buttons $3 Per unit 2 $6
Total $1,575
Table 11.11: Recovery Costs
Recovery (Full Scale)
Item Cost Per Unit Unit Quantity Total
Ripstop Nylon $8 Per yard 25 $200
Nylon Thread $8 Per spool 3 $24
Tubular Nylon $1 Per foot 50 $50
Paracord $5 Per roll 1 $5
Telemetrum $200 Per unit 1 $200
Telemega $300 Per unit 1 $300
Jolly Logic $130 Per unit 2 $260
Total $1,039
197
Table 11.12: Rover Costs
Item Cost Per Unit Unit Quantity Total
1000:1 Micro
Metal Gearmotor
HPCB 12V
$22.95 Per Unit 3 $68.85
Elegoo Uno
Project $35.00 Per Unit 1 $35.00
Xbee Pro 60 mW
Wire Antenna
and dongle
$62.90 Per Pair 2 $125.80
Arduino
Wireless SD
shield
$17.49 Per Unit 1 $17.49
Flexible Solar
Panel 1W 6V $13.99 Per Unit 1 $13.99
Onyx $189.99 Roll .5 $95.00
Nylon $169.99 Roll .5 $85.00
Total $441.13
Cost of Rocket On the Pad $3,055.13
198
Table 11.13: Budget Allocation
Item Cost
Full-Scale $3,055.13
Sub-Scale $2,000
Travel $3,000
Educational Outreach $1,500
Test flights (5) $1,000
Research and Development $3,000
Promotional Items $1,000
Total $14,555.13
Figure 11.8: Spending Comparison
Budget Breakdown
Full-Scale Sub-Scale Travel
Educational Outreach Test Flights (5) Development
Promotional Items
199
The team is very happy with the state of its budget. The success of the subscale vehicle on its first
launch certainly saved the team money and ensured the full-scale vehicle can reuse some expensive
electronics. The team also secured additional sponsorship funds, to be discussed in the following
section. The excess funds have been allocated to research and development projects, which will
help to construct a second test rocket that can fly experimental payloads without risking the
progress of the competition. Even with the extra resources allocated, the team maintains a
comfortable cushion of approximately $2,000 in case of unforeseen expenditures.
Section 11.5: Funding Plan
The team has secured funding from the sources presented in Table 11.14: Funding Sources. This
money will cover the cost of the rocket on the pad, the purchase of capital equipment as needed,
the cost of subscale and full-scale test launch motors, programming and materials for our
educational engagement events, travel and housing for the team at the competition in Huntsville,
Alabama, and any other costs associated with designing, building, and launching our competition
rocket. Since PDR, the team has secured an additional sponsorship from Lockheed Martin totaling
$2,000 and Dynetics has increased their contribution to the team by $500. With these additional
funds, and the subscale test vehicle costing less than budgeted, the team has covered all necessary
expenditures and has allocated additional resources to research. Additionally, all of the anticipated
costs were estimated on the high end to provide a safety factor to the funding plan. The Auburn
University Student Launch team is confident that the current financial situation is quite stable, and
will continue to be for the remainder of the competition.
Table 11.14: Funding Sources
Source Amount
Alabama Space Consortium $12,000
Dynetics $2,500
200
Lockheed Martin $2,000
Total Funding $16,500
Section 11.6: Timeline
Two Gannt charts have been created to illustrate the project timeline. They are broken up by
semester, as this separation most closely imitates how the students operate, and the conclusion of
CDR aligns almost exactly with the start of Auburn University’s Spring semester. These can be
found in below in Figure 11.9: Fall Timeline and Figure 11.10: Spring Timeline.
Figure 11.9: Fall Timeline
8/23/17 9/13/17 10/4/17 10/25/17 11/15/17 12/6/17 12/27/17 1/17/18
Conceptual Design
Proposal
Preliminary Design Review
Materials Testing
CFD Testing
Trade Studies
Detailed Design
Critical Design Review
Sub-Scale Development
Payload Development
First Subscale Launch
Junior E-Day EE Event
Detailed Design Review
201
Figure 11.10: Spring Timeline
As seen in the timelines, the team is slightly behind on materials testing. However, the majority of
the tests are complete and the findings correlate with the research the team has completed, so the
team is not concerned. Those tests will be completed within a week of CDR being submitted.
Conversely, both the construction of the payload and the development of the full-scale vehicle are
ahead of schedule. Overall, the team is quite happy with the progress made so far. A well-
constructed vehicle will most likely not be ready to launch at the first launch opportunity on
February 3rd, so the team will have a full-scale vehicle and a test vehicle ready to launch on
February 17th. Should anything function incorrectly, the team can adjust and relaunch the
following day because this is a two-day launch event. As an emergency measure, there is another
launch opportunity March 3rd. With several opportunities available to launch a full-scale rocket,
and progress on the rocket already ahead of schedule, the team is confident that progress is on
track.
1/10/18 1/31/18 2/21/18 3/14/18 4/4/18 4/25/18
Payload Construction
Full-Scale Development
Flight Readiness Review
Rocket Week EE Event
Auburn E-Day EE Event
Full-Scale Flight
Two Day Full-Scale Flight Opportunity
Emergency Full-Scale Flight
Competition Preparation
Huntsville Competition
Post-Launch Assessment Review
