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AUBURN UNIVERSITY STUDENT LAUNCH
Project Nova
211 Davis Hall
AUBURN, AL 36849
FLIGHT READINESS REVIEW
MARCH 5, 2018
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Table of Contents
Table of Contents ...........................................................................................................................2
List of Figures .................................................................................................................................8
List of Tables ................................................................................................................................12
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 FRR 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 CDR ....................................................................................21
Section 3.1: Vehicle Changes ...............................................................................................21
Section 3.2: Payload Changes...............................................................................................21
Section 3.3: Project Plan Changes ........................................................................................22
Section 4: Launch Vehicle .......................................................................................................22
Section 4.1: Vehicle Changes ...............................................................................................22
Section 4.1.1: Since CDR ......................................................................................................22
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Section 4.1.2: During Construction .......................................................................................23
Section 4.2: Systems Component Analysis ..........................................................................23
Section 4.2.1: Structure ..........................................................................................................23
Section 4.2.2: Propulsion .......................................................................................................26
Section 4.2.3: Aerodynamics .................................................................................................29
Section 4.2.4: Schematics ......................................................................................................32
Section 4.3: Flight Reliability Confidence ...........................................................................40
Section 4.3.1: Fin Shape and Style ........................................................................................40
Section 4.3.2: Materials .........................................................................................................42
Section 4.3.3: Assembly Procedures ......................................................................................42
Section 4.4: Mass Statement .................................................................................................44
Section 4.5: Manufacturing Process .....................................................................................44
Section 5: Recovery System Design ........................................................................................49
Section 5.1: Structural Elements...........................................................................................50
Section 5.2: Materials ...........................................................................................................51
Section 5.3: Ejection .............................................................................................................55
Section 5.4: Parachutes .........................................................................................................59
Section 5.5: Altimeters .........................................................................................................63
Section 5.6: Recovery System Technical Drawings .............................................................66
Section 6: Mission Performance Predictions .........................................................................68
Section 6.1: Simulations .......................................................................................................68
Section 6.1.1: Motor Thrust Curve ........................................................................................71
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Section 6.1.2: Component Weights .......................................................................................71
Section 6.1.3: Stability ...........................................................................................................72
Section 6.1.4: Computational Fluid Dynamics ......................................................................73
Section 6.2: Kinetic Energy ..................................................................................................76
Section 6.3: Drift ..................................................................................................................77
Section 6.4: Simulation Verification ....................................................................................80
Section 7: Full-Scale Flight Results ........................................................................................81
Section 7.1: Flight Data ........................................................................................................81
Section 7.2: Launch Day Simulation and Analysis ..............................................................81
Section 7.3: Full-scale Analysis ...........................................................................................82
Section 7.3.1: Post-Launch Simulation ..................................................................................83
Section 7.4: Full-Scale and Subscale Comparison ...............................................................83
Section 8: Rover .......................................................................................................................84
Section 8.1: Design Changes ................................................................................................84
Section 8.1.1: Since CDR ......................................................................................................84
Section 8.1.2: During the Construction Process ....................................................................88
Section 8.2: Structural Elements...........................................................................................89
Section 8.3: Electrical Elements ...........................................................................................97
Section 8.4: Drawings and Schematics ...............................................................................102
Section 8.5: Flight Reliability Confidence .........................................................................104
Section 8.5.1: Deployment ...................................................................................................104
Section 8.5.2: Navigation .....................................................................................................104
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Section 8.5.3: Communication .............................................................................................104
Section 9: Altitude Control Module .....................................................................................105
Section 9.1: Design Changes ..............................................................................................105
Section 9.1.1: Since CDR ....................................................................................................105
Section 9.1.2: During the Construction Process ..................................................................105
Section 9.2: Design .............................................................................................................106
Section 9.2.1: Structural Elements .......................................................................................106
Section 9.2.2: Electrical Elements .......................................................................................107
Section 9.2.3: Drawings and Schematics .............................................................................108
Section 9.3: Flight Reliability Confidence .........................................................................110
Section 10: Safety .....................................................................................................................113
Section 10.1: Personnel Hazard Analysis .............................................................................113
Section 10.2: Failure Modes and Effects Analysis ...............................................................116
Section 10.3: Environmental Hazard Analysis .....................................................................126
Section 10.3.1: Rocket Effects on Environment ..................................................................126
Section 10.3.2: Environmental Effects on Rocket ...............................................................127
Section 11: Launch Operations Procedures ..........................................................................130
Section 11.1: Preassembly Checklists ..................................................................................130
Section 11.1.1: Recovery .....................................................................................................130
Section 11.1.2: Altitude Control ..........................................................................................131
Section 11.1.3: Body ............................................................................................................132
Section 11.1.4: Rover ...........................................................................................................132
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Section 11.1.5: Engine .........................................................................................................133
Section 11.2: Launch Vehicle Assembly and Check ............................................................133
Section 11.3: Launcher Setup and Launch Procedure ..........................................................135
Section 11.3.1: Launcher Setup ...........................................................................................135
Section 11.3.2: Launch Procedure .......................................................................................136
Section 11.4: Post-flight Inspection ......................................................................................136
Section 12: Testing ...................................................................................................................138
Section 12.1: Rover Battery and Motor Test (AU 4.4, 4.5) ..................................................138
Section 12.1.1: Results .........................................................................................................139
Section 12.2: Recovery and Altitude Control Battery Tests (AU 3.1, 6.9) ..........................141
Section 12.2.1: Results .........................................................................................................142
Section 12.3: Full-Scale and Subscale Separation Test (AU 3.2) ........................................142
Section 12.3.1: Results .........................................................................................................144
Section 12.4: Tension Testing of Composite and 3D Printed Material (AU 2.1, 2.2) .........146
Section 12.4.1: Results .........................................................................................................147
Section 12.5: 3-Point Bend Testing of Composite and 3D Printed Material (AU 2.1, 2.2, 6.7)
149
Section 12.5.1: Results .........................................................................................................151
Section 12.6: Compression Testing of Composite and 3D Printed Material (AU 2.1, 2.2, 6.7)
153
Section 12.6.1: Results .........................................................................................................154
Section 12.7: Rover Maneuverability (AU 4.2, 4.3, 4.6, 4.7)...............................................156
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Section 12.7.1: Results .........................................................................................................157
Section 12.8: Altitude Control System (AU 6.4 – 6.8) .........................................................158
Section 12.8.1: Results .........................................................................................................160
Section 13: Project Plan ..........................................................................................................160
Section 13.1: Requirements Verification ..............................................................................160
Section 13.1.1: General Requirements .................................................................................160
Section 1.1.1: Vehicle Requirements ................................................................................166
Section 1.1.2: Recovery Requirements .............................................................................180
Section 13.1.2: Deployable Rover Requirements ................................................................186
Section 13.1.3: Safety Requirements ...................................................................................188
Section 13.2: Team Requirements ........................................................................................189
Section 13.2.1: General Requirements .................................................................................189
Section 13.2.2: Vehicle Requirements .................................................................................190
Section 13.2.3: Recovery Requirements ..............................................................................191
Section 13.2.4: Deployable Rover Requirements ................................................................192
Section 13.2.5: Safety Requirements ...................................................................................195
Section 13.2.6: Altitude Control Requirements ...................................................................196
Section 13.3: Budget .............................................................................................................199
Section 13.4: Funding Plan ...................................................................................................206
Section 13.5: Timeline ..........................................................................................................207
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List of Figures
Figure 1.1: Team Organization Chart ........................................................................................... 17
Figure 4.1: Motor Thrust Curve .................................................................................................... 27
Figure 4.2: Motor Retention ......................................................................................................... 29
Figure 4.3: Fin Rendering ............................................................................................................. 30
Figure 4.4: Lower Booster Section Dimensional Drawing ........................................................... 35
Figure 4.5: Upper Body Tube Dimensional Drawing ................................................................... 36
Figure 4.6: Upper Section Assembly Dimensional Drawing........................................................ 37
Figure 4.7: Centering Ring Dimensional Drawing ....................................................................... 38
Figure 4.8: Bulkhead Dimensional Drawing ................................................................................ 39
Figure 4.9: Fin Dimensional Drawing .......................................................................................... 40
Figure 4.10: Fin Shapes ................................................................................................................ 41
Figure 4.11: Booster Section Diagram.......................................................................................... 45
Figure 4.12: Fin-Jig ....................................................................................................................... 47
Figure 4.13: Isogrid Manufacturing Process ................................................................................. 48
Figure 4.14: Isogrid Manufacturing Process – Filament Winding ............................................... 49
Figure 5.1: Redundant Jolly Logic System ................................................................................... 57
Figure 5.2: Jolly Logic Chute Release .......................................................................................... 58
Figure 5.3: Gore Template ............................................................................................................ 61
Figure 5.4: Parachute Template .................................................................................................... 61
Figure 5.5: Altus Metrum TeleMega Altimeter ............................................................................ 64
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Figure 5.6: Altus Metrum TeleMetrum Altimeter ........................................................................ 64
Figure 5.7: Altimeter Block Diagram ........................................................................................... 65
Figure 6.1: OpenRocket Model..................................................................................................... 68
Figure 6.2: Altitude Vs. Time ....................................................................................................... 69
Figure 6.3: Velocity Vs. Time ...................................................................................................... 70
Figure 6.4: Acceleration Vs. Time ................................................................................................ 70
Figure 6.5: Motor Thrust Curve .................................................................................................... 71
Figure 6.6: Stability Vs. Time....................................................................................................... 73
Figure 6.7: Nosecone Meshing ..................................................................................................... 75
Figure 6.8: Tail-fin Meshing ......................................................................................................... 75
Figure 7.1: Full-scale Flight Data ................................................................................................. 81
Figure 8.1: Rover Coupler Bay ..................................................................................................... 85
Figure 8.2: Servo System .............................................................................................................. 86
Figure 8.3: Bulbous Rover Drive Wheels ..................................................................................... 87
Figure 8.4: Communication Schematic......................................................................................... 88
Figure 8.5: I-beam Rover Body – Top .......................................................................................... 89
Figure 8.6: I-beam Rover Body – Bottom .................................................................................... 90
Figure 8.7: Dual Tread System ..................................................................................................... 91
Figure 8.8: Exploded Dual Tread System ..................................................................................... 92
Figure 8.9: SPDS – Folded ........................................................................................................... 93
Figure 8.10: SPDS - Unfolded ...................................................................................................... 93
Figure 8.11: Solar Panel Deployment Tray .................................................................................. 94
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Figure 8.12: Servo Arm/Jaw ......................................................................................................... 95
Figure 8.13: Open Servo Arm/Jaw ............................................................................................... 96
Figure 8.14: Closed Servo Arm/Jaw ............................................................................................. 96
Figure 8.15: Sample Orientation Check........................................................................................ 97
Figure 8.16: Electrical Schematic ................................................................................................. 98
Figure 8.17: 1000:1 12V Pololu Micro Metal Gearmotors........................................................... 98
Figure 8.18: XCTU Interface ........................................................................................................ 99
Figure 8.19: Arduino Wiring ...................................................................................................... 100
Figure 8.20: Motor Shield Wiring .............................................................................................. 101
Figure 8.21: Rover CONOPS ..................................................................................................... 102
Figure 8.22: Assembled Rover ................................................................................................... 103
Figure 8.23: Rover Dimensions .................................................................................................. 103
Figure 9.1: Bottom-Up view of plates and arms ......................................................................... 109
Figure 9.2:Side View of assembled module ............................................................................... 109
Figure 9.3: IPDS Dimensions ..................................................................................................... 110
Figure 12.1: Subscale Separation Test ........................................................................................ 144
Figure 12.2: Full-Scale Separation Test of the Upper Section ................................................... 145
Figure 12.4: Epoxy - Carbon Fiber Tension Test Results .......................................................... 147
Figure 12.5: Epoxy - Fiberglass Tension Test Results ............................................................... 148
Figure 12.6: Onyx Tension Test Results..................................................................................... 149
Figure 12.7: Epoxy - Carbon Fiber Bend Test Results ............................................................... 151
Figure 12.8: Epoxy - Fiberglass Bend Test Results .................................................................... 152
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Figure 12.9: Onyx Bend Test Results ......................................................................................... 152
Figure 13.1: Spending Comparison ............................................................................................ 206
Figure 13.2: Fall Timeline .......................................................................................................... 207
Figure 13.3: Spring Timeline ...................................................................................................... 208
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List of Tables
Table 1.1: Team Members ............................................................................................................ 17
Table 2.1: Team Information ........................................................................................................ 19
Table 2.2: Mentor Information ..................................................................................................... 20
Table 2.3: Launch Vehicle Information ........................................................................................ 20
Table 4.1: Section Lengths ........................................................................................................... 24
Table 4.2: Motor Specifications .................................................................................................... 28
Table 4.3: Fin Dimensions ............................................................................................................ 31
Table 4.4: Manufacturing Schedule .............................................................................................. 43
Table 4.5: Mass Estimations ......................................................................................................... 44
Table 5.1: Pugh Chart of Carbon Fiber vs. Fiberglass Material for the BAE............................... 50
Table 5.2: Parachute Materials Pugh Chart .................................................................................. 52
Table 5.3: Comparison of Paracord, Tubular Nylon, Kevlar ........................................................ 53
Table 5.4: Comparison of U-bolts and Eye-bolts ......................................................................... 54
Table 5.5: Comparison of Black Powder and CO2 ....................................................................... 55
Table 5.6: Parachute Shape Pugh Chart ........................................................................................ 60
Table 5.7: Parachute Dimensions ................................................................................................. 62
Table 6.1: Flight Simulation Data (Wind = 0 mph) ...................................................................... 68
Table 6.2: Component Weights .................................................................................................... 71
Table 6.3: CFD Drag Coefficient Results ..................................................................................... 76
Table 6.4: Drift Calculations for Upper Section ........................................................................... 79
Table 6.5: Drift Calculations for Lower Section .......................................................................... 79
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Table 9.1: Physical Components................................................................................................. 106
Table 9.2: Electrical Components ............................................................................................... 107
Table 9.3: Mission Success Criteria............................................................................................ 110
Table 13.1: Deployable Rover Requirements Verification ......................................................... 186
Table 13.2: AU General Requirements ....................................................................................... 189
Table 13.3: AU Vehicle Requirements ....................................................................................... 190
Table 13.4: AU Recovery Requirements .................................................................................... 191
Table 13.5: AU Rover Requirements.......................................................................................... 192
Table 13.6: AU Altitude Control Requirements ......................................................................... 196
Table 13.7: Vehicle Costs ........................................................................................................... 199
Table 13.8: Recovery Costs ........................................................................................................ 200
Table 13.9: Budget Allocation .................................................................................................... 205
Table 13.10: Funding Sources .................................................................................................... 206
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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 FRR March 5th, 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 FRR 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 113 in.
Estimated Mass 38.4 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.
The rover has been named Monica. After landing under parachute, Monica will be completely
housed inside the rocket will be remotely deployed from the rocket. She 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 CDR
Section 3.1: Vehicle Changes
The team has decided not to continue with braided carbon fiber as a body tube material for this
year. The team successfully built and flew a braided carbon fiber rocket, but the weight reduction
of the structure was minimal and the rugged shape of the skin overlaying the braid induced
additional drag. With additional research, this construction style could improve rocket
performance, but given time constraints the team constructed and flew a solid carbon fiber airframe
rocket for competition purposes.
Section 3.2: Payload Changes
The rover design has not changed significantly. Domed wheels are being used in place of flat
wheels, and the body of the rover has been modularized for ease of interaction. These changes
22
came as a result of experience gained during the construction process. More information on the
rover changes can be seen in Section 8.1: Design Changes.
The larger change to the payloads is the decision not to fly at competition with the airbrake system.
The system was constructed and functioned, but due to an unavailability of larger motors could
not be tested in flight as the full-scale flights never flew above 5280 feet. Since the airbrakes
modify the external flight profile of the vehicle, the team understands this means the system cannot
be used at the competition. The data tracking components of the system will still be flown to assess
how the algorithm would function at the competition, as they have been flown with the full-scale
vehicle and do not affect the external aerodynamics of the rocket.
Section 3.3: Project Plan Changes
The project plan has been updated to account for the completion of the full-scale rocket and
completion of the Flight Readiness Review. The final hurdle for the team is the competition itself.
The team has also come in under budget, and has reflected that in the budget portion of the report.
Spare funds will be allocated to supporting materials for launch week and the Rocket Fair.
Section 4: Launch Vehicle
Section 4.1: Vehicle Changes
Section 4.1.1: Since CDR
The vehicle design has not changed substantially since CDR. Minor length additions have been
added to accommodate unexpected length in the motor tube section. Additionally, the team has
decided to not pursue the isogrid structure for the competition vehicle. This is mainly due to the
initial structure having a much higher weight than expected. This problem was compounded by
23
approaching due dates which made analysis and implementing a solution infeasible. Mass
estimates have also been adjusted due to final construction weights
Section 4.1.2: During Construction
Upon completion of construction of an isogrid tube, the team realized that the design did not afford
the performance improvements expected. For this reason, the team decided to default to standard,
rolled carbon fiber tubes as these have a lower coefficient of drag.
Section 4.2: Systems Component Analysis
The vehicle was 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
ensured adequate space for avionics, payload equipment, and electronics. These systems are
crucial to the success of the mission. The vehicle design was also heavily driven by manipulating
weight and length to control altitude and stability. These factors determined the success of the
flight itself. The vehicle design was 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 was designed to be able to withstand the forces the rocket will
experience during operation. The launch vehicle body was strong enough to maintain stable flights,
while accommodating all other subsystems, and ensuring they had adequate space and protection.
The design of the structure requires heavy tradeoffs between strength, space, and weight.
The total length of the rocket is 113.0 inches. Component lengths are shown in Table 4.1: Section
Lengths.
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Table 4.1: Section Lengths
Body Tubes:
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 was 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.
Initially, it was planned that the tubes would be constructed using carbon fiber braiding, a process
that involved taking individual strands of carbon fiber and stitching them into a tightly-wound
braid. The carbon fiber braids produced, were formed into an isogrid structure around a 6-inch
mandrel, which was 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 basalt fiber was filament wound 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
were projected to 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. The isogrid tubes were generously manufactured and
provided by Highland Composites.
Section Length (in.)
Nose Cone 20
Avionics Section 36
Rover Section 18
Airbrake Section 12
Booster Section 27
Total 113
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Both the isogrid and solid-tube airframes were manufactured and flown over the course of the
project, and despite promising initial estimates the isogrid structure came in at a much higher
weight than was expected. The team believes that the added weight came from issues in the method
used to wrap the airframe. Mostly due to time constraints and the end of available launch
opportunities, the team was unable to manufacture a newer iteration of the isogrid structure, but
the team feels confident that this manufacturing method will provide large weight savings for
future airframes. After much discussion, it was decided to select the solid carbon tube design due
to the simplicity of manufacturing and quickly approaching deadlines.
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 was attached to the booster section with 3 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 was ideal for making the coupler which was holding our electronics, the avionics bay. 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.
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 were made from pre-impregnated carbon
26
fiber. This was chosen due to the simplicity of manufacturing with pre-impregnated carbon fiber.
The interior diameter for the circular cross-sectional rocket was 6 inches and the bulkheads were
designed to fit perfectly into this size. All bulkheads for this rocket was 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 uses three centering rings. These 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 4.3: Motor Tube Rendering
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Motor:
The motor selected for the competition is the Aerotech L1420R. This is the same motor that was
used in CDR, 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 4.1: Motor Thrust Curve.
Figure 4.1: 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.
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Table 4.2: Motor Specifications
Motor Specifications
Manufacturer Aerotech
Motor Designation L1420R
Diameter 2.95 in
Length 26.2 in
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 was
made by braiding carbon fiber strands and then filament winding around a mandrel that was 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 is 0.1-inch-thick and was designed to fit around an Aerotech L1420R motor.
To mount the motor tube, three centering rings were epoxied to the outer diameter of the motor
tube and the inner diameter of the lower section tube. The epoxy was a 24-hour epoxy, which will
create a permanent bond between the components. A bulk plate was 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.
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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 chose a commercially bought
Aeropack bolt-on motor retention system, as shown in Figure 4.2: Motor Retention. This is a
simple system with two components. One component was bolted 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 4.2: 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.
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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
machine. A clipped delta planform was selected for the fins. Four fins was machined from 0.2-
inch-thick carbon fiber flat plates. A rendering of the fin design is shown in Figure 4.3: Fin
Rendering.
When attached, the trailing edge of each fin is located slightly forward of the end of the body tube.
This design feature provides 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 the chance of fin flutter which
increases the vehicles chance of success during flight. Each fin has 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 2.7 calibers. This
Figure 4.3: Fin Rendering
31
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
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.
32
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:
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 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 was also changed to be fiber
glass, as this is the most common material sold by these vendors.
Section 4.2.4: Schematics
Figure 4.7: Final Rocket Design
33
Figure 4.7: Booster Tube Dimensional Drawing
34
Figure 4.8: Engine Block Dimensional Drawing
35
Figure 4.4: Lower Booster Section Dimensional Drawing
36
Figure 4.5: Upper Body Tube Dimensional Drawing
37
Figure 4.6: Upper Section Assembly Dimensional Drawing
38
Figure 4.7: Centering Ring Dimensional Drawing
39
Figure 4.8: Bulkhead Dimensional Drawing
40
Figure 4.9: Fin Dimensional Drawing
Section 4.3: Flight Reliability Confidence
Section 4.3.1: Fin Shape and Style
The three most common planforms are clipped delta, trapezoidal and elliptical, as shown in Error! R
eference source not found.. 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.
41
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
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 4.10: Fin Shapes
The fins were manufactured from the same carbon fiber plates as the bulkheads and centering
rings. The same data used to verify the bulkheads and centering rings was 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 were also been two subscale flights which further verified the simulation data. Multiple full-
scale test flights were performed to visually verify no anomalies are present on the fins during
flight.
42
Section 4.3.2: Materials
Body Tubes:
The structural tubes of the launch vehicle were constructed using a rolled braided carbon fiber
structure. As this is something the team has not done in past years, structural data was collected
for this structure. To do this, using the same material and manufacturing method, a test sample
was made consisting of an equal diameter of the tubes that was used on the launch vehicle. This
sample was then placed into a load cell to determine the maximum load of the structure. This
allowed us to determine that the structure is capable of safely completing the mission. The structure
experiences 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 were 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 were made. These samples were manufactured using the same material that the
bulkheads and centering rings was 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 was performed. Once
proven on the ground, a subscale flight test using this coupler component was used.
Section 4.3.3: Assembly Procedures
Manufacturing of the vehicle generally takes two weeks to produce and assemble the components.
To account for this the team started manufacturing three weeks prior to any scheduled launches.
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This allowed for one extra week if any problems arise during the manufacturing process. The
typical manufacturing schedule can be seen in Error! Reference source not found..
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. 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. Luckily, Highland Composites generously agreed to
produce the braided carbon fiber tubes. Unfortunately, these tubes arrived far too late into the
project for in-depth analysis and revisions. To mitigate the effects of a total loss crash, multiple
tube sections were produced, which allowed for the construction of three full scale rockets, one
with braided body tubes and two with non-braided body tubes.
Several flat plates of carbon fiber were produced at various thicknesses. The plates were placed in
a CNC router to be shaped into flat components. These components include fins, bulk plates and
centering rings.
44
Section 4.4: 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 was
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 final weights of the rocket and its components are displayed in Table
4.5.
Table 4.5: Mass Estimations
Section Mass (lb) Percentage
Structure 15.4 40.10%
Motor 10.1 26.30%
Rover 3.9 10.15%
Recovery 7 18.23%
Airbrake 2 5.21%
Total 38.4 100%
Section 4.5: Manufacturing Process
With the success of subscale rocket, the team’s initial design was validated and construction on
the full sale rocket began. The primary iteration of the rocket was constructed using solid carbon
45
fiber tubes in order to quickly create a rocket to test the various electrical subsystems of the rocket.
This would allow the team to validate all the sub-systems without waiting for the complex and
time-intensive isogrid airframe. These solid carbon tubes were manufactured in the student
laboratory using pre-preg carbon fiber and a steel mandrel. The team has a large wealth of
experience with this manufacturing method, so confidence was high in the quality of these tubes.
Figure 4.11: Booster Section Diagram
The most time intensive portion of the airframe to manufacture is the booster section, regardless
of whether the tube is isogrid or solid carbon. This is primarily due to the precision necessary to
align the fins and the time it takes for the 24-epoxy to set. The section must be assembled in stages
to ensure proper alignment. Small deviations in the orientation of the fins or motor tube can cause
the rocket to be unstable upon launch. The process begins with the installation of the bulkhead that
will rest at the front of the motor tube. This takes place but first using 5-minute epoxy to secure
the plate, and then adding filets of 24-hour epoxy around the edges of the bulk-plate to reinforce
46
the plate and ensure stability. These filets are placed on the front and back portions of the bulk-
plate. This fileting procedure will be repeated for many of the other components of the rocket.
Next, the motor tube will be installed into the rocket. Prior to insertion into the booster section, a
centering ring was secured to the motor tube to ensure alignment. This centering ring was
positioned such that, once inserted, it will align with the leading edge of the fins. It was secured
using 5-minute epoxy and then 24-hour filets. It was checked with a level to ensure correct
orientation. Once the motor tube / centering ring assembly was inserted into the rocket, it was
secured using epoxy filets.
After the motor tube is secured in the assembly, the fins can be inserted and epoxied. This is the
most tedious and time-consuming process as each fin must be inserted and secured individually
and checked against a fin-jig that was cut out using the laser cutter and a sheet of ply-wood. The
fin-jig is used to ensure proper alignment of each fin which in turn helps guarantee the stability
of the rocket.
Each fin is secured using epoxy filets on each contact point in the rocket. For example, each fin
will have 6 epoxy filets, 2 on the motor tube, two on the interior of the airframe, and two on the
exterior of the airframe. This is to ensure the fins are locked into place and to guarantee
structural stability. This extra epoxy does cause a decent increase in weight, but the team is
confident that the added weight is worth the stability.
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Figure 4.12: Fin-Jig
Once the fins are secured, the final two centering rings were inserted and secured using epoxy
filets. One centering ring butts up against the trailing edge of the tail fins, the other is flush with
the back of the rocket.
All other sections of the rocket are fairly easy to produce as they are just tubes with minor
modifications and length differences. Any bulkhead attachments are preformed using the same
method of epoxy fileting described previously.
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Figure 4.13: Isogrid Manufacturing Process
This method is the same for the isogrid airframe, but with an initial filament winding portion.
The primary objective of the filament winding was to ensure air tightness for pressurization, as
well as for aerodynamics. For the isogrid airframe, the team decided to use basalt fiber for its
weight and strength properties. A filament winder was used to apply the fiber to the exterior of
the isogrid airframe. The individual sections of the airframe were then placed in the oven at 60°C
to accelerate the curing process. Issues were encountered with the fiber width, but due to time
constraints, no analysis or solution was able to be implemented.
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Figure 4.14: Isogrid Manufacturing Process – Filament Winding
Section 5: Recovery System Design
The overall design of the Auburn Student Launch team’s recovery system has not changed at all
since its original as it has worked as designed on the subscale flight and both full-scale test flights.
This design used an augmented dual-stage recovery system with a drogue parachute deployed at
an apogee height of 4433 ft, an upper main parachute deployed at 1000 ft. and a lower main
parachute deployed at 750 ft. At apogee, the nose cone, drogue parachute and upper main parachute
were all ejected using redundant black powder charges. When the upper main parachute was
ejected, it was held closed by the Jolly Logic Chute Release System. At 1000 ft. the Jolly Logic
System released and deployed the upper main parachute with the rocket still in one piece. After
the rocket had coasted down to 750 ft. a second set of redundant black powder charges detonated
and separated the rocket, pulling the lower main parachute from its housing. Using this
configuration, the entire rocket fell under a single drogue parachute from apogee until the second
50
event occurred where the rocket separated into two separate pieces and fell under two separate
main parachutes. The nose cone and drogue parachute remained attached via shock cord to the
upper section after separation.
Section 5.1: 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 Error! Reference source not found.. As for the design o
f 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 3D printed 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 hole has two charge wire
pairs 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. This
isolation of pressure also helped to reduce the amount of black powder needed for each charge.
Table 5.1: 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
51
Ease of
Manufacturing 1 3 2
Total 16 9
The BAE served as the coupler between the upper and the lower parachute housings. Each section
was secured to the BAE with three machine bolts per section. Neither of these sections separated
once the rocket was 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 was done so the tube connections
between the upper parachute housing, the BAE, and the lower parachute housing were continuous
and smooth, minimizing the impact on the aerodynamic performance of the rocket. 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 allowed the team to
externally arm the altimeters while the rocket was assembled and sitting on the launch rail.
A single U-bolt was 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 was connected to a second U bolt
mounted to the bottom of the Rover section. This 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 of the
rover getting stuck when it was being deployed after landing.
Section 5.2: Materials
The materials that were chosen to create the team's recovery system had a direct effect on the
success of the system. Failure of the recovery system materials to survive the ejection charges
would have caused an inadequate landing of the vehicle, and severe damage to the rover payload.
52
The chosen materials for the parachute, shock cord, and bulkhead attachment were rip-stop nylon,
paracord, and U-bolts, respectively.
Ideally, the chosen material for the parachute allow it to be sturdy to be used in multiple tests and
launches. The factors that were 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
weighs about 2.75 oz. per square yard while cotton fabric weighs about 4.3 oz. per square yard.
Rip-stop nylon 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 5.2: Parachute Materials
Pugh Chart as well as our two successful full-scale test flights confirmed the choice of rip-stop
nylon for the parachute material.
Table 5.2: Parachute Materials Pugh Chart
Rip-stop Nylon Cotton fabric
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 was that it can only withstand 550 pounds of force before failure.
However, the other benefits of paracord made it an ideal material for the drogue shroud lines, since
the drogue did not create much drag force and did not put as much tension on the shroud lines.
53
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 force. For the main parachutes, which create more drag force, the shroud lines were
made from this material. Tubular nylon has a much higher strength and also allowed 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 ensured failure is less likely to happen
in this component. The pliability and strength of tubular nylon led to the team choosing to use it
as our material of choice for shock cords. The Pugh chart that was used to evaluate which material
was chosen for the parachute is shown in Table 5.3: Comparison of Paracord, Tubular Nylon,
Kevlar.
Table 5.3: 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
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Total 14 13 11
Nylon shear pins were used to attach the nosecone to the BAE and the lower parachute housing to
the rover section to prevent drag separation. In the team’s configuration, #4-40 nylon screws were
used to secure these sections. These machine screws have a double shear strength of 50 lbs. Ground
testing was performed on these shear pins prior to launching to ensure safety and eliminate the
possibility of manufacturing discrepancies.
The materials used to attach the shock cord to the bulk plates was chosen appropriately. The team
used 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 can 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 were used to assess the suitability of different bulk plate attachment materials to 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 5.4:
Comparison of U-bolts and Eye-bolts.
Table 5.4: Comparison of U-bolts and Eye-bolts
U-bolt Eye-bolt
Attachment Strength 3 2
Limits cord tangles 3 1
Totals 6 3
All components that were mentioned were flown on both of the teams’ full-scale test flights and
performed as designed. The three of the parachutes and their shock cord provided the drag required
55
to slow down the separate sections of the rocket to put it well below the required kinetic energy
limit all while remaining usable for following launches.
Section 5.3: Ejection
4F black powder was 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 and worked to a
superb level again this year. 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 5.5: Comparison of Black Powder and CO2.
Table 5.5: 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
Reliability 2 1
Total 12 8
For the both events, two charges were placed within 3D printed charge cups and armed with
electronic matches: the first charge is the primary means of ejection, and the other is a backup
charge for redundancy and to decrease the chances for failure of the recovery system. The
redundant apogee charge was set to fire 1 second after the first charge. The charges are not ignited
56
at the same time as that has the potential to cause damage to the airframe, shock cord and
parachutes.
Black powder, when ignited, can be approximated as an ideal gas. Testing has proven to be
successful and helped guarantee the validity of the team’s calculations. The Ideal Gas Law was
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 = 700R°
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 14 inches as:
𝑉 = 𝐴 × 𝐿 =𝜋
4× 𝑑2 ×= 395.8 𝑖𝑛3
Using this equation, the team calculated the amount of black powder needed to produce pressures
sufficient to shear the nylon screws attaching the nose-cone to the upper section. Assuming 3
nylons screwed each rated for 50 psi of shear, pressure required to shear the nylon screws is:
𝑃 =3 × 𝐹
𝜋4 × 𝑑2
= 5.3 𝑝𝑠𝑖
Which, in conjunction with the previous equations, yields a charge size for 4F black powder of:
𝑛 =𝑃 × 𝑉
𝑅 × 𝑇×
453.592𝑔
𝑙𝑏𝑚= 4.74 𝑔
Applying a factor of safety of 1.5 to the above calculations, 7 grams of black powder was estimated
as the required amount of black powder needed per charge. During testing, a 6 gram charge of
black powder separated both the upper and lower sections without any trouble so the team was
confident a 7 gram charge would be sufficient on launch day. This was confirmed through both
sections successfully separating on both full-scale test flights.
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Both charge sections were filled with fireproof cellulose insulation (colloquially known as “barf”)
to protect the parachutes from the heat of the ejection charges. The first event deployed the drogue
parachute and the rifled upper main parachute, held closed by our Jolly Logic chute release system.
The second event consisted of the second set of charges igniting to separate the rocket and deploy
the lower main parachute as well as the mechanical release of a pin in the Jolly Logic system, thus
allowing the upper main parachute to deploy.
The AUSL team utilized the Jolly Logic system in series for redundancy in the recovery systems.
It is an efficient system that doesn’t require any black powder charges and has its own self-
contained batteries which decreased amount of wires connected to the altimeter bay, as well as
reduced risk of entangled shock cord in the main body.
Figure 5.1: Redundant Jolly Logic System
The Jolly Logic system allowed for the deployment of both a drogue and main parachute
simultaneously in a single separation. This system deployed more parachutes with fewer
separations, thus reducing the chance of failure of the recovery portion of flight. The only
parachute that utilized the Jolly Logic Chute Release System was the upper main parachute.
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Figure 5.2: 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 was connected to the Jolly Logic
device via a rubber band which provided tension on the pin so that when it was released, it was
pulled from its socket. Since the Jolly Logic system consisted of rubber bands which could easily
break if put under stress, the recovery system was set up so that tension from the shock cord was
never transferred to the Jolly Logic devices. To achieve this, the main parachute was not folded
with the shock cord inside as is normal. Instead the parachute was located on the shock cord so
that the shroud lines were at full extension and so that the drogue line, which ran through the main
parachute spill hole was able to deploy to its full length during the initial separation. The main
parachutes were then carefully gathered and positioned in the body tubes so that the drogue line
pulled it out without putting any forces on the actual parachutes and Jolly Logic system.
To preserve the redundancy of the recovery system, two Jolly Logic chute releases were used in
series to gather the upper main parachute. If either of the Jolly Logic devices happened to
malfunction, the other still would have released, thus allowing the main parachutes to deploy.
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Section 5.4: Parachutes
Auburn’s augmented dual deploy recovery approach made 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 is 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 deployed from the top of the rocket. This stabilized
descent until main deployment at the second event. The drogue parachute was designed to bring
the rocket down at a velocity of approximately 100 ft/s. This velocity minimized the drift of the
rocket while still having a stable descent. The drag coefficient of 0.8 for a fully inflated circular
parachute that was determined from research was confirmed in both full-scale flights.
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 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 yielded a circular drogue with a 24.9-inch diameter. The actual descent velocity
of the rocket under drogue for the full-scale test flight was 94.3 ft/s.
The recovery system has two main parachutes constructed of rip-stop nylon with 0.5-inch tubular
nylon shroud lines. The main parachutes are a hemispherical shape. The hemispherical shape can
be more difficult to manufacture, but produced the most drag, allowing the rocket to have
maximum drag with minimum weight. A Pugh chart comparing different parachute shapes can be
seen below in Table 5.6: Parachute Shape Pugh Chart.
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Table 5.6: 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 5.3: Gore Template and
Figure 5.4: Parachute Template. When the rocket reaches 750 feet in altitude, a second charge
separates the top section of the rocket to release the lower main chute, and the jolly logic chute
release system releases the bundled upper main chute. A spill hole was added to the main
parachutes. This spill hole is necessary with our configuration of dual-deployment from the same
compartment at the top of the rocket body. The diameter of the spill holes were made to be 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 while marginally altering the descent rate.
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Figure 5.3: Gore Template
Figure 5.4: Parachute Template
Parachute areas for hemispherical shaped chutes were 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 used this equation to calculate an appropriate area for the main parachutes so that the kinetic
energy of either section of the rocket did not exceed 75 ft-lbs during recovery and remained within
safe limits.
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𝐴𝑈𝑝𝑝𝑒𝑟 =2 × 14𝑙𝑏𝑚 × 32.2
𝑓𝑡𝑠2
0.076474𝑙𝑏𝑚
𝑓𝑡3 × 1.31 × (15𝑓𝑡𝑠 )
2= 40.00 𝑓𝑡2
𝐴𝐿𝑜𝑤𝑒𝑟 =2 × 24𝑙𝑏𝑚 × 32.2
𝑓𝑡𝑠2
0.076474𝑙𝑏𝑚
𝑓𝑡3 × 1.31 × (11.5𝑓𝑡𝑠 )
2= 116.66 𝑓𝑡2
Table 5.7: Parachute Dimensions
Upper Chute Lower Chute
Area 40.00 ft2 116.66 ft2
Diameter 60.5 in 103.4 in
Diameter of Spill hole 12.1 in 20.7 in
Number of Gore 6 8
Width of Each Gore at Base 31.7 in 40.7 in
Height of Each Gore 47.5 in 81.15 in
Circumference at Base 190.1 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 of the parachute to a true
hemisphere.
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Section 5.5: Altimeters
The avionics bay houses two altimeters to satisfy redundant system requirements. Both altimeters
fired the apogee charge at the apogee height of 4433 ft to eject the nosecone and thus the drogue
and bundled upper main parachute. Then both altimeters fired the main deployment charges at an
altitude of 750 ft. Neither set of charges are fired at the exact same time. A one and three second
delay has been placed between the firing of the apogee and main charges respectively to ensure
the structural integrity of the rocket and parachutes are maintained.
The team used one Altus Metrum TeleMega as the primary altimeter and one Altus Metrum
TeleMetrum as the secondary altimeter. 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. All the data gathered
from the launch was taken from this altimeter and confirmed by the secondary one. Using two
Altus Metrum altimeters made programming quick and easy, as they share an interface program.
This made 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 5.5: Altus
Metrum TeleMega Altimeter and Figure 5.6: Altus Metrum TeleMetrum Altimeter shown below
are pictures of the altimeters the team used.
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Figure 5.5: Altus Metrum TeleMega Altimeter
Figure 5.6: Altus Metrum TeleMetrum Altimeter
Another reason the Altus Metrum altimeters were used 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
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altitude, battery voltage, igniter status, and GPS status. However, in flight mode, settings can’t be
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 5.7: Altimeter Block Diagram
Isolating one altimeter system (altimeter, battery, and wires) from the other helped prevent any
form of coupling or cross-talk of signals. Isolation was realized via distancing the two systems,
66
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 ~70cm). Avoiding resonant lengths of wire was 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 were mounted
on a carbon fiber board that slides into a set of rails. A small foam plate was placed between the
batteries and the altimeter board to ensure that the batteries did not short out. The altimeters and
batteries were 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 acted as
shielding between the two altimeters and minimize cross-talk as well as near-field coupling. This
board is also easily removable for connecting the altimeters to computers for configuration and for
charging the altimeters’ batteries.
Section 5.6: Recovery System Technical Drawings
Below are technical drawings of different aspects of the recovery system.
Figure 6.8: Technical Drawing of Charge Cup
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Figure 6.9: Technical Drawing of Altimeter Board Rails
Figure 6.10: CAD Model of BAE
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Section 6: Mission Performance Predictions
Section 6.1: Simulations
The launch vehicle was 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 6.1:
OpenRocket Model.
Figure 6.1: 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.
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 6.1: Flight Simulation Data (Wind = 0 mph)
Old Flight Simulation Data (Wind = 0 mph)
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Maximum Velocity 739 ft./s
Maximum Acceleration 283 ft./s2
Launch Weight 38.4 lbm
Burnout Weight 32.71 lbm
Length 113 in
Maximum Diameter 6.25 in
Launch Stability 2.6 calibers
Velocity off Rod 80.9 ft/s
Figure 6.2: Altitude Vs. Time
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Figure 6.3: Velocity Vs. Time
Figure 6.4: Acceleration Vs. Time
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Section 6.1.1: 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 5280 ft.
Figure 6.5: Motor Thrust Curve
Section 6.1.2: Component Weights
Table 6.2: Component Weights
Component Weight (lbm)
Upper Body Tube 4.78
Lower Body Tube 4.64
Nose Cone 2.17
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Fins (4) .91
Centering Rings (3) .6
Motor 10.1
Motor Retention .3
Motor Tube 1
Rover and Bay 4.4
Airbrake System 2
Bulkheads .9
Electronics 1.5
Recovery Section 5
Total 38.4
Section 6.1.3: 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 2.6 calibers at the point of rail exit which is
comfortably above the minimum requirement of 2 calibers. Since the drag plate system will not be
flown, they do not factor into the stability. The team feels confident that having a stability margin
of over 2 calibers will allow the rocket to be stable without becoming overstable.
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Figure 6.6: Stability Vs. Time
Before flight, the center of pressure and the center of gravity are measured to be 89.13 inches and
72.852 inches from the top of the rocket, respectively. In Figure 6.6: 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 6.1.4: 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
74
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.
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Figure 6.7: 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 6.8: 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
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“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 6.3: CFD Drag Coefficient Results
Mach Number (M) Coefficient of Drag (C_d)
0.6 .532
0.725 .5375
0.750 .5425
Section 6.2: 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 14 lbm and a recorded velocity of 15.8
ft./s for the upper section, this equation yields:
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𝐾𝐸 =1
2×
14𝑙𝑏𝑚
32.2𝑓𝑡𝑠2
× (15.8𝑓𝑡
𝑠)
2
= 54.3 𝑓𝑡 ∙ 𝑙𝑏
With a mass of 24 lbm and a recorded velocity of 12.2 ft/s for the lower section, this equation
yields:
𝐾𝑒 =1
2×
24𝑙𝑏𝑚
32.2𝑓𝑡𝑠2
× (12.2𝑓𝑡
𝑠)
2
= 55.5 𝑓𝑡 ∙ 𝑙𝑏
Section 6.3: 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 confirmed this as a
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reasonable value for the coefficient of drag. 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 40.00 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= 22.7
𝑓𝑡
𝑠
Assuming a total upper rocket weight after burnout of 14 lbm and an upper main parachute area
of 40.00 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= 15.0
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 6.4: Drift
Calculations for Upper Section and Table 6.5: 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 6.4: 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 80.70 366.50 760.92
7.5 11.00 470.80 121.11 550.00 1,141.91
10 14.67 627.88 161.52 733.50 1,522.9
12.5 18.33 784.52 201.81 916.50 1,902.83
15 22.00 941.60 242.22 1,100.00 2,283.82
17.5 25.67 1,098.68 282.63 1,283.50 2,664.81
20 29.33 1,255.32 322.92 1,466.50 3,044.74
Table 6.5: 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 80.70 478.04 872.46
7.5 11.00 470.80 121.11 717.42 1,309.33
10 14.67 627.88 161.52 956.78 1,746.18
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12.5 18.33 784.52 201.81 1,195.48 2,181.81
15 22.00 941.60 242.22 1,434.84 2,618.66
17.5 25.67 1,098.68 282.63 1,674.20 3,055.49
20 29.33 1,255.32 322.92 1,912.90 3,491.14
The team understands that these are only predictions and not representative of how the rocket will
actually drift. The team also understands that recovery of our rocket can only be guaranteed up to
3500 ft. at the competition. However, the approximate reported wind conditions for the team’s
final launch at Sylacauga were 50 mph at a height of 1000 ft. Under these conditions, neither piece
of the rocket drifted farther than a mile from the launch pad. Given this data, and the fact that these
calculations were made on an assumption of an apogee of 5280 ft when the true apogee was 4433
ft the team must assume that the rough estimates we have tabulated are rougher than initially
supposed. Unless there is a significant amount of wind, the team believes that the rocket will stay
within the drift limits despite the rough calculations to the contrary.
Section 6.4: 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, full-scale flight, and CFD results to ensure an accurate drag coefficient for the
model. The team is confident that the simulation is accurate, yet understands that the model slightly
overestimates flight characteristics.
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Section 7: Full-Scale Flight Results
Section 7.1: Flight Data
Flight data for the subscale launch can be seen below in Figure 7.1: Full-scale Flight Data. The
flight occurred on the 17th of February 2018. The airframe was made of a solid carbon fiber tube
and flew on an Aerotech L1520T motor. We flew on this motor due to the lack of availability of
our desired motor, the L1420R. 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 4433 feet, had a
maximum acceleration of 350 ft/s2. The data verifies that the vehicle and recovery designs are safe
and within the limits of the competition.
Figure 7.1: Full-scale Flight Data
Section 7.2: Launch Day Simulation and Analysis
The following data is the OpenRocket simulation using the motor that was flown on the day of,
the L1520T. The data is quite similar to the flight results achieved at the launch. The team believes
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that the differences had to due to inaccurate weights obtained during the pre-launch checks and a
faulty scale. We believe that the current simulations will be valid for the competition providing we
ensure that the weights of each component are accurate prior to launch.
OpenRocket
Simulation
Recorded Flight
Data
Apogee (ft) 4552 4433
Max Velocity (ft/s) 620 567
Section 7.3: Full-scale Analysis
The full-scale flight characteristics are very similar to what the simulation model predicted.
Although the values are slightly off, the pre-launch weight values did not match the simulation
weights exactly. The team believes this accounts for the variation in flight data from predicted
data.
The physical subsystems of the vehicle all performed to expectations. The airframe handled all the
experienced loads, and the launch was stable. The winds at the launch were exceeding 50 mph
above the ground, yet the vehicle was neither under- nor over-stable. The vehicle was angled
slightly into the wind to mitigate drift, but flew straight otherwise.
The recovery system performed as intended, safely landing the vehicle. Despite the high winds,
the recovery system did not drift excessively. This leads the team to believe that the calculations
for drift may not be as accurate for a vehicle of this size as expected, and that they are an
overestimation of actual drift. The team decided this is acceptable, as it introduces a factor of
safety.
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The rover retainment system functioned, but the rover was unable to be deployed after landing due
to terrain that was more treacherous than could be expected at the competition. However, ground
tests confirm that Monica is able to traverse the obstacles that she will experience.
Section 7.3.1: Post-Launch Simulation
Using the data from the full-scale launch, the team estimated a coefficient of drag for the rocket of
.53. This value is very similar to the values obtained from CFD and from the subscale flight. A
simulation was performed after the launch with updated coefficient of drag and weight values. The
comparison of these results with the flight data can be seen below.
Updated
OpenRocket
Simulation
Recorded Flight
Data
Apogee (ft) 4489 4433
Max Velocity (ft/s) 586 567
These values more closely approximate the flight data. The team is confident this final simulation
model can accurately simulate the flight of the rocket. Additionally, the team would prefer these
values to overestimate apogee due to the importance of not exceeding the altitude limit.
Section 7.4: Full-Scale and Subscale Comparison
The full-scale rocket performed very similarly to the subscale rocket. Both of these test flights
have assisted the team in improving the simulation model for the competition flight. The most
important lesson learned from these flights is that the team must be careful when estimating the
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mass of different sections of the vehicle. Components tend to weight more than expected, most
likely due to additional weight from the epoxy bonding the carbon fiber composite material.
Both flights, however, did demonstrate the stability and safety of the design. The recovery and
ascent of both flights went perfectly. This is a boon, as it allows the team to iterate and improve
upon designs as often as the team has launch opportunities.
Section 8: Rover
Section 8.1: Design Changes
Section 8.1.1: Since CDR
Design changes to the rover since CDR are reflective of shortcomings noticed during testing; most
are minor tweaks to pre-existing subsystems/components.
The rover bay has been redesigned to be completely modular for ease of access and transportation.
The rover will be housed inside of a coupler with an extension (Figure 8.1: Rover Coupler Bay).
Two slats of wood will act as the platforms for the rover to rest on inside of the coupler. The bay
was changed from 3D printed nylon to wood for cost and time savings. The back of the rover
coupler bay has an eye bolt installed in it that will be used to secure the rover in the bay. Until tests
were performed, the rover was expected to stay in the bay as no forces would be acting on it to
eject it from the bay; the eye bolt acts as an anchor for the rover to latch onto.
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Figure 8.1: Rover Coupler Bay
A number of changes were made to the rover body. Holes were 3D printed into the body of the
rover for ease of component attachment. A panel was added in between the main motor housings
in order to prevent shorting of the motor leads. The solar panel deployment system was moved
closer to the end of the rover to allow for maximum solar exposure of the panels once deployed.
The length of the body was changed from 6.71 in. to 7.51 in. to incorporate a servo that is needed
to the hold the rover into the bay. Attached to the servo is an L-shaped bracket that will clamp
down on the eye-bolt in the back of the rover coupler bay. An arm was added to the back of the
rover for the L-shaped bracket to make a complete closure around the eye-bolt. A detail of the
servo system on the rover can be seen in Figure 8.2: Servo System.
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Figure 8.2: Servo System
After maneuverability tests, the team was satisfied with the tread design. However, the orientation
of the rover and it being able to right itself into a drivable position did present a problem. Going
back to a very early team design idea, bulbous wheels are being used to aid in the “righting” of the
rover after deployment. A top-down view of the wheels can be seen in Figure 8.3: Bulbous Rover
Drive Wheels. These wheels act as a see-saw for the rover so that after deployment, the rover does
not tip onto its side.
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Figure 8.3: Bulbous Rover Drive Wheels
Electrically, there were a number of changes. A new set of communication devices was chosen
after further range testing deemed the XBee Pro 60mW antennas to not perform as well as
expected. To replace the 60mW version (effective communication range of 1 mile), the XBee Pro
900 devices (effective communication range of 6 miles) were selected and have been deemed to
consistently communicate at the maximum expected range of half a mile. Since the new more
powerful XBees did not have antennas built onto them, an antenna system was developed.
Connected to the transmitting XBee is a Yagi-Uda directional antenna. On the rover, an internal
laptop antenna is connected to the XBee and has been proven to successfully receive the
transmitted signal. A communication schematic is shown in Figure 8.4: Communication
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Schematic. For determining the rover orientation in the rocket, the Adafruit 9DOF Breakout board
was deemed too big and was replaced with a Pololu AltIMU-10 v5 inertial measurement unit.
Figure 8.4: Communication Schematic
Section 8.1.2: During the Construction Process
There were unforeseen challenges that we faced during rover manufacturing since CDR. Most of
these challenges were design flaws with tolerances. During assembly, short term fixes included
dremeling/wrapping tape around the component to size it properly. At the first chance, whatever
change needed to be made to the print was updated in SolidWorks.
89
Section 8.2: Structural Elements
Auburn University’s deployable rover has been designed to be housed inside of the rocket during
flight and deploy from the rocket after landing, whereupon it will travel at least five feet from the
rocket and deploy solar panels. All parts of this section (minus the solar panels) have been designed
using SolidWorks and then 3D printed with either nylon or onyx, a chopped-carbon-fiber-infused
nylon composite. Due to its impressive strength properties, onyx was used in parts such as the
rover chassis and the wheels. The only thing not printed with onyx are the treads, which need to
be flexible enough to deform over the wheels of the rover.
During the design process, size was the biggest limiting factor the team faced. An I-beam structure
(Figure 8.5: I-beam Rover Body – Top) was chosen as the frame to support components on
Figure 8.5: I-beam Rover Body – Top
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both sides of the web while keeping height to a minimum. Over the course of the design process,
multiple 3D printed compartments were added to the body to allow for placement of individual
components. In Figure 8.5: I-beam Rover Body – Top, two motor sections and a servo
compartment can be seen. The wall towards the front of the rover allows placement of a 9V battery,
which is friction fit between the wall and the motor housing. In Figure 8.6: I-beam Rover Body –
Bottom, (a bottom-up view of the rover, “Monica”) the solar
Figure 8.6: I-beam Rover Body – Bottom
panel deployment system (SPDS) motor compartment and another 9V battery station are visible.
The use of friction fitting for the batteries allows for easy removal and replacement.
In both figures, holes can be seen on the web of the body. The holes mark the mounting location
of the Arduino UNO in Figure 8.5: I-beam Rover Body – Top and the Adafruit motor shield in
Figure 8.6: I-beam Rover Body – Bottom.
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A dual tread system was selected as the best way to maneuver across the unknown terrain of the
launch field. The system, Figure 8.7: Dual Tread System, not only allows for impeccable traction
on most surfaces
Figure 8.7: Dual Tread System
that we have tested on, but also acts as a guide for placement of components inside the structure;
anything outside of the treads will be dragging on the ground. The treads are 3D printed nylon.
The bulbous wheels prevent the rover from tipping onto its side and resting there. If tipped upon
deployment or during the travel to reach five feet from the rocket, the wheels act as a pivot point
for the rover, which will right itself; “righting” tests have been successfully performed on the rover
to confirm this passive orientation correction method. The drive wheels are secured to the body
directly through the drive motors’ shafts while the idle wheels on the rear of the rover spin freely
around separate bolts. The idle wheels are secured with nuts. An exploded view of this setup is in
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Figure 8.8: Exploded Dual Tread System. Once the wheels are attached, the treads are stretched
trightly across them.
Figure 8.8: Exploded Dual Tread System
To complete the solar panel deployment aspect of the competition, an array of solar panels are
installed on the back of the rover in an accordion fashion (Figure 8.9: SPDS – Folded and Figure
8.10: SPDS - Unfolded). After the rover reaches
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Figure 8.9: SPDS – Folded
Figure 8.10: SPDS - Unfolded
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five feet from the rocket, a rack and pinion system is utilized to deploy the solar panels. Attached
to the motor is a gear that spins to deploy a tray (Figure 8.11: Solar Panel Deployment Tray) out
of the
Figure 8.11: Solar Panel Deployment Tray
back of the rover. The solar panels, which are anchored to the rover and attached to the front of
the solar panel depolyment tray, are unfolded and held in place.
To secure the rover into the rover coupler bay, a jaw was designed to fit onto a servo arm. When
the servo arm/jaw (Figure 8.12: Servo Arm/Jaw) closes, it clamps onto an eyebolt in the back of
the rover
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Figure 8.12: Servo Arm/Jaw
coupler bay and is supported undereneath by an arm that protrudes from the back of the rover. To
allow for proper pressurization of the recovery section directly above the rover, a carbon fiber door
covers the rover bay. However, it does nothing to prevent the rover from prematurely deploying
from the rocket; it is simply pulled from the bay when the second main parachute deploys. That is
solely resting on the jaw-like servo. The open and closed servo jaw can be seen in Figure 8.13:
Open Servo Arm/Jaw and Figure 8.14: Closed Servo Arm/Jaw.
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Figure 8.13: Open Servo Arm/Jaw
Figure 8.14: Closed Servo Arm/Jaw
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Section 8.3: Electrical Elements
The electrical elements of the rover include everything that does not provide a structural purpose.
This includes the motors/servos, circuit boards, communication and power that ensure the rover
operates as designed. Once the rocket lands under parachute, the team will send a signal to the
rover via XBee modules; this signal will initialize the deployment sequence. First, the Pololu
Altimu-10 v5 will determine the rover orientation in the rocket based on the z-axis of the
Altimu/rover. This is a critical step as an improper orientation detection will send the rover
backwards into the rocket. A sample of orientation detection code is seen in Figure P. Once the
Figure 8.15: Sample Orientation Check
orientation is determined, the servo unclamps to allow the rover to deploy. The drive motors then
spin the direction determined by the uploaded script for a predetermined amount of time. Once the
rover exits the rocket (after that predetermined amount of time) one motor will stop spinning while
the other continues, turning the rover perpendicular to the rocket. This is done to avoid running
into the parachute that might be directly in front of the open rover bay. At this point, both motors
will spin another predetermined amount of time, driving the rover forward. Upon reaching that
time, the drive motors will stop and the solar panel motor will engage the solar panel tray to deploy
the foldable panels. At various times throughout this process, the Arduino pulls the orientation
from the AltIMU to ensure that the rover has not flipped over, thus ensuring the rover travels the
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same direction throughout the course of the deployment. An electrical schematic is shown in
Figure 8.16: Electrical Schematic.
Figure 8.16: Electrical Schematic
The motors used for driving the rover and for the SPDS are 1000:1 12V Pololu Micro Metal
Gearmotors. These motors (Figure 8.17: 1000:1 12V Pololu Micro Metal Gearmotors) provide 125
oz.-in. of torque and have a cross-sectional
Figure 8.17: 1000:1 12V Pololu Micro Metal Gearmotors
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area of 0.4 in. x 0.45 in., small enough to fit on the rover, but strong enough to provide enough
torque to overcome difficult obstacles once deployed. The wheels were designed so that the keyed
shaft of these motors would fit snuggly into the wheel holes.
The servo used for securing the rover into the coupler bay is an Adafruit TowerPro SG-5010. Since
the jaw of the servo will not be experiencing much lateral force during the flight (only longitudinal
due to the nature of the servo orientation) the strength of the servo was not critical. After several
successful shake tests and a successful flight using the servo-secured method, the team has proven
the feasibility of this servo.
To communicate with the rover, XBEE Pro 900s were selected. Previous models of the XBees did
not provide the communication range required, so the 900s, a more powerful model that has an
effective range up to six miles, was chosen instead. On the transmitting side, one XBee is
connected to a laptop via USB dongle. Through a software called XCTU, the XBees can be
programmed and a team member can communicate with the other paired XBee through a serial
monitor (Figure 8.18: XCTU Interface). The transmitting antenna is connected to the XBee
through a series of connections from the U.Fl antenna on
Figure 8.18: XCTU Interface
wires from the U.Fl antenna on the XBee to a male N-type connection on a Yagi-Uda directional
antenna. The Yagi’s transmitting frequency is 900 MHz as is the XBee Pro 900. The rover antenna
is an internal Bluetooth laptop antenna that sticks through a hole in the body of the rocket and
wraps around the outside of the rocket. The team decided to have the antenna on the outside of the
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rocket in order to avoid attenuation problems from the electrically conductive carbon fiber body.
Once the rover receives the signal and begins the exit the coupler bay, the antenna will get pulled
off of the rover XBee, allowing the rover to proceed unhindered by the 27 in. Bluetooth antenna.
The black piece on the XBee in Figure 8.19: Arduino Wiring ensures the antenna is disconnected
when the rover starts to deploy.
The core of the rover electronics is the Arduino UNO. This was chosen due to previous experience
with Arduino Unos and to the accessibility of sample codes and libraries online. Using the Uno
also meant that we did not need a large source of power; one 9V is able to successfully power the
Arduino during the launch and deployment. The wiring for the Arduino Uno can be seen in Figure
8.19: Arduino Wiring.
Figure 8.19: Arduino Wiring
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In order to ensure proper powering of the three motors and the servo, an Adafruit Motor Shield v2
was used; this was chosen based on previous experience with the unit. Similarly to the Arduino,
there are numerous resources to aid in the utilization of this unit. It also is powered simply with a
9V battery. The motor shield wiring can be seen in Figure 8.20: Motor Shield Wiring.
Figure 8.20: Motor Shield Wiring
To determine the rover orientation in the rocket after landing under parachute, the team decided
on using a Pololu AltIMU-10 v5. This unit includes a 3-axis gyroscope, accelerometer,
magnetometer and digital barometer on a board no larger than a quarter.
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Section 8.4: Drawings and Schematics
Figure 8.21: Rover CONOPS
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Figure 8.22: Assembled Rover
Figure 8.23: Rover Dimensions
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Section 8.5: Flight Reliability Confidence
Based on ground tests and successful flight tests, Auburn University Student Launch team has high
hopes for the successful launch, deployment, and recovery of “Monica”. Ground tests have
individually validated aspects of the rover such as being able to navigate different terrains,
communicate effective and successfully up to the range specified by the competition, and right
itself in the event that the rover tips to one side during its deployment. Flight tests have given us
the opportunity to successfully demonstrate the ability to overcome all of these individual
challenges.
Section 8.5.1: Deployment
The eyebolt added to the bay introduces a static attachment point for the rover to ensure it will not
deploy during ascent. Flight tests have verified this capability, and ground tests have ensured the
clasp will separate properly after landing when the deployment signal is sent.
Section 8.5.2: Navigation
The dual-tread design has been confirmed to be able to traverse the terrain that the rover will
encounter. The treads also enable travel regardless of the landing orientation, ensuring that the
rover will be able to travel at least five feet from the rocket to deploy solar panels. The domed
wheels prevent the rover from tipping onto a side which is the only orientation that the treads
cannot function.
Section 8.5.3: Communication
The rover is required to communicate remotely with the team for a successful and safe flight. All
communication systems have been thoroughly tested and proven to function from a distance. All
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communication systems are also isolated from other transmitting devices by carbon fiber plates to
ensure there is no cross-talk between devices, leading to unsafe and unexpected events. This
shielding has been proven in flight.
Section 9: Altitude Control Module
Section 9.1: Design Changes
Section 9.1.1: Since CDR
From CDR, the Internal Plate Drag System (IPDS) has remained relatively unchanged with regard
to the module itself. A few tweaks were made to the dimensions of the support plates and arms
within the system to retrofit the module to the inside of the rocket body. Also, assumptions with
securing the module within the rocket body have changed. Due to design changes from other
systems, the module took up roughly half the space it was allotted within the body. A new direction
was then taken to secure the module within its section. Three screws are now being used to hold
the module stable during flight. In addition to this, the upper half of the module bay is being used
to hold key switches, which are able to power on the system once the rocket is mounted to the
launch rail before a flight. These changes are required both for system integrity and ease of
integration.
Section 9.1.2: During the Construction Process
As detailed in the previous section, several changes were made to the system overall since CDR.
With regard to the dimensional changes of the system, several different components had to be
resized and printed again to get a better fit for system integration. The arms that act as the interface
between the motor shaft and the drag plates had to be made longer to achieve a higher drag
coefficient during flight. Making this change results in a longer actuation from within the body.
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Also with earlier iterations, our 6-inch diameter support plates were too large to be printed with
the departmental 3D printer. In order to address this issue, we split the support plates in two within
the design in order to make the module pieces snap into place upon assembly. Also mentioned in
Section 4.1.1, the original iteration of the system was intended to be secured by the bottom bulk
plate separator of the section above it. Upon implementation the dimensions were off. This led to
the addition of securing screws with washers to hold the module down. The upper portion of the
body was then used for drilling in key switches to help power on the IPDS. This significantly cut
down on the time spent integrating and prepping the system prior to launch.
Section 9.2: Design
Section 9.2.1: Structural Elements
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
components of the design in order to prove the feasibility of using additive techniques to produce
a prototype and ultimately a final product.
Table 9.1: Physical Components
Component Responsibility
(3) Drag Plates Actuate out of IPDS to produce drag on the
vehicle perpendicular to the air flow.
(3) Support Plates Form the backbone of the module. These hold
the system together.
(3) Threaded Support Rods Form the backbone of the module. These hold
the system together.
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(1) Shaft-to-Plate Interface This connects to the shaft of the motor and the
support arms that actuate the drag plates.
(3) Plate Interface Arms These connect the Shaft-to-Plate Interface to
the three drag plates to help carry out actuation.
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.
Section 9.2.2: Electrical Elements
The electronics for this system combine to provide all of the inputs necessary to provide a continual
calculation for projected altitude. The controller will act as the decision maker for the system,
determining the rate at which the vehicle will self-correct its trajectory.
Table 9.2: Electrical Components
Component Responsibility
(1) Arduino Uno Controller for the IPDS. Responsible for
coordinating all calculations and guidance.
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(1) AndyMark NeveRest 40 DC Motor Handles uniform actuation of the drag plates.
(1) Adafruit Motor Shield v.2.3 Allows communication and power distribution
to the DC motor.
(1) Adafruit accelerometer Calculates triple-axis acceleration needed for
controls.
(1) Adafruit Micro SD card breakout Destination for importation of all flight data.
(1) Adafruit barometric pressure sensor Measures altitude needed for calculations in
controls.
(2) 9V Battery Distributes power to system.
The IPDS is completely powered via two 9V batteries. One battery pack powers the controller,
while the other powers the motor shield. Both power sources are linked to the two key switches
towards the top of the system. This circuit is completed via the switches directly prior to launch in
order to conserve battery life. This ensures the system is not in danger of losing power. All sensors
and the Micro SD card breakout are located on top of the module, soldered directly to a breadboard.
The placement and location of the sensors and breadboard allow the correct axis data to be
collected during launch.
Section 9.2.3: Drawings and Schematics
This section includes pictures, renderings, and dimensions of the Internal Plate Drag System
(IPDS).
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Figure 9.1: Bottom-Up view of plates and arms
Figure 9.2:Side View of assembled module
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Figure 9.3: IPDS Dimensions
Section 9.3: Flight Reliability Confidence
This system’s mission success criteria are listed in the table below.
Table 9.3: Mission Success Criteria
Criteria Number Criteria Method of Validation
AU1
All aerodynamic data must be
validated through analytical
and experimental testing.
An aerodynamic analysis of
the drag plates and the internal
system was conducted through
pid controller implementation
and simulations.
AU2 Drag plates must stay static
throughout launch.
Code implemented phases will
be implemented through the
controller that will prevent
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any system action prior to the
end of the boost phase.
AU3 Electronics must stay secured
throughout flight.
Proper screws and fasteners
are used to secure the system
in place.
AU4
Subsystem components must
engage accordingly after
boost phase and stay online
for the remainder of the flight.
Testing will verify that the
code implemented phases will
keep the system static until
motor burn.
AU5
Controller and IMU must be
able to correctly predict the
projected altitude of the
launch vehicle.
Test data will be logged to a
Micro SD card to review post-
launch.
AU6
Drag plates must deploy after
boost phase in order to self-
correct the trajectory of the
launch vehicle.
PID controller simulations and
test flights will validate
deployment precision.
AU7
Drag plates must be able to
withstand the perpendicular
force of the airflow.
Wind tunnel testing and
structural testing will be
conducted to ensure the
integrity of the plates and the
materials used to construct
them.
Mission success criteria 2, 4, and 6 (AU 2, 4, 6) are satisfied via the implementation of flight
phases. In the controller script, phases keep unwanted output from ever occurring in any case.
Phase 1 waits for the launch event of the launch vehicle. Phase 2 awaits the end of the motor burn
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event. Phase 3 runs all calculations and will feature action from the drag plates and self-correction
of trajectory. The last phase shuts the system down once the system has done its job. Mission
success criteria 1 and 7 (AU1 and AU7) is satisfied through simulations, computational fluid
dynamics, and wind tunnel testing. Finally, securing the module in the body with screws fulfills
criteria 3 (AU3). The system that has been implemented is extremely fault-tolerant given many
common potential circumstances and can adjust accordingly.
Section 10: Safety
Section 10.1: Personnel Hazard Analysis
Hazard Cause Result Severity Probability Combined
Risk Mitigation Verification
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
Trainings and safety
measures taken as
described in proposal
Soldering fumes
from heated
metals or plastics
Use soldering
tool on improper
surfaces or for
excessive time
Toxic fumes created
are inhaled and cause
the team member to
become sick from
prolonged exposure
3 3 9
Work with any soldering
tools will be done in a
ventilated area and all
unsupervised work will
be performed by trained
individuals
Trainings and safety
measures taken as
described in proposal
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
Trainings and safety
measures taken as
described in proposal
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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
Trainings and safety
measures taken as
described in proposal
Burning surface
of soldering tool
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
Trainings and safety
measures taken as
described in proposal
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Noxious fumes
from body
manufacturing
Improper
ventilation of the
workstation
where 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, work
will not continue to avoid
a buildup of fumes and
allow dispersion of the
Trainings and safety
measures taken as
described in proposal
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.
Trainings and safety
measures taken as
described in proposal
Checklists shown in
Launch Operations
Procedures are
followed
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.
Trainings and safety
measures taken as
described in proposal
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Improper use of
man powered
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
Trainings and safety
measures taken as
described in proposal
Section 10.2: Failure Modes and Effects Analysis
Hazard Cause Result Severity Probability Combined
Risk Mitigation Verification
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.
Wiring noted in
Electrical Elements
Parachute failure
after deployment
(tangled)
The parachute is
not packed
properly
The parachute
becomes tangled on
descent resulting in an
erratic and fast
moving 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.
Noted in Preassembly
Checklists
Parachute design
noted in Parachutes
117
Parachute
deploys early
A faulty
altimeter fails to
detect the
altitude at which
the parachute
should deploy or
improper wiring
switches event
order
The rocket's ascent
will be compromised
and its descent will
result in the rocket
drifting for a very
long distance
3 3 9
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.
Noted in
Preassembly
Checklists
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.
Proper charge size
noted in Preassembly
Checklists and
Ejection
118
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.
Noted in Preassembly
Checklists and
Launch Vehicle
Assembly and Check
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.
3 2 6
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.
Noted in Flight
Reliability
Confidence
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.
Guidelines state all
motor work is to be
done by team mentor
119
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.
Parachute materials
noted in Materials
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.
Testing – Recovery
battery tests verifies
Mitigation A
Testing – Full scale
and Subscale
separation testing
verifies Mitigation B
120
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.
Parachute sized noted
in Parachutes
Deployment height
noted in Recovery
System Design
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.
All parts tested pre-
flight in Launch
Vehicle Assembly
and Check
121
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.
Verified in Flight
Reliability
Confidence
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.
Tested pre-flight in
Launch Vehicle
Assembly and Check
Rocket descends
too 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.
Parachute sizes noted
in Parachutes
Kinetic energy
calculations located in
Kinetic Energy
122
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
and clamp 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.
Design changes to
maintain steadiness
noted in Rover
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.
Testing sections on
materials testing in
Testing
123
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.
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.
Prelaunch checks can
be found in Launch
Vehicle Assembly
and Check
124
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.
Testing sections on
materials testing in
Testing
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.
Testing sections on
materials testing in
Testing
Noted in Preassembly
Checklists
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.
Noted in Preassembly
Checklists
125
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.
First step in each
preflight checklist
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.
Model simulation
noted in Simulations
Motor selection noted
in Propulsion
126
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
Checked in Engine
Section 10.3: Environmental Hazard Analysis
Section 10.3.1: Rocket Effects on Environment
Hazard Cause Result Severity Probability Combined
Risk Mitigation Verification
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.
Noted in checklist in
safety section
Launcher Setup and
Launch Procedure
127
Debris from
ballistic return
or explosion
A recovery
system failure
cause the rocket
to impact the
ground at high
velocity
scattering debris
or an engine
CATO forms
debris either in
the air or on the
launch pad
Heated materials are
spread on flammable
surfaces causing fires
or debris is spread as
a contaminant
4 2 8
This situations are
mitigated by other
precaution shown in
failure modes and proper
recovery techniques are
followed according to
NAR guidelines
Noted in checklist in
safety section Post-
flight Inspection
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
Trainings and safety
measures taken as
described in proposal
Epoxy gas and
chemical release
Epoxy releases
volatile
chemicals and
gasses as it cures
This small release can
be vented into the
environment causing
pollution of air or
water
1 4 5
With our rather small
scale of our construction,
the impact of our
ventilation is minimal
and all solid or liquid
contaminates are
disposed of in a proper
fashion
Trainings and safety
measures taken as
described in proposal
Section 10.3.2: Environmental Effects on Rocket
128
Hazard Cause Result Severity Probability Combined
Risk Mitigation Verification
Cross winds in
flight
Rocket is
launched when
cross winds are
faster than
desired 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. The
rocket is also designed to
be stable under cross
wind conditions
Stability margin noted
in Stability
Fin design noted in
Aerodynamics
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
Natural
environment
effects recovery
of rocket
The rocket drifts
into difficult
terrain or foliage
when landing
Finding or accessing
the rocket during
recovery is made
difficult or even
dangerous for team
members
3 3 9
The size of the main
parachutes are minimized
to prevent drift while still
ensuring the safe decent
of rocket sections. The
event heights for
separation and
deployment of parachutes
is lowered for similar
reasons
Parachute sized noted
in Parachutes
Deployment height
noted in Ejection
129
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 any corrosion
Air temperature
and sun exposure
Prolonged
exposure to the
sun heats
components of
the launch
vehicle beyond
standard
temperatures for
manufactured
materials
Epoxies and other
resins may loosen or
melt causing the
separation of the
launch vehicle body
3 2 6
On-site work and
assembly of the launch
vehicle are done under a
tent or in another form of
shaded area
Unstable launch
surface
The surface,
below launch
rails, is loosely
packed or wet
The weight of the
rocket causes a tilt of
the launch rail system
resulting in a skewed
trajectory
2 2 4
All launch locations are
checked by the safety
officer, project lead, and
the RSO prior to launch
setup
Noted in safety
checklist Launcher
Setup and Launch
Procedure
130
Section 11: Launch Operations Procedures
Section 11.1: Preassembly Checklists
Section 11.1.1: Recovery
Task Completed (Initialed)
Ensure all necessary items are transported
Check primary and back-up batteries
Attach RF trackers to both the upper and
lower sections with electrical tape
Open up the BAE, plug batteries into
altimeters and ensure all lead wires are
plugged into correct altimeter points
Check both altimeters and jolly logic
systems for functionality
Folded parachutes
Attach RF trackers to both the upper and
lower sections with electrical tape
Turn on both Jolly Logic Chute Releases,
ensuring that the batteries are at full
charge and have a deploy height of 1000ft
set
131
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
Section 11.1.2: Altitude Control
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
132
Place system in housing and secure
attachment screws
Section 11.1.3: Body
Section 11.1.4: Rover
Task Completed (Initialed)
Ensure all necessary items are transported
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
Check engine housing for structural
integrity
Gather all shear pins and attachment
screws
133
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
Section 11.1.5: Engine
Task Completed (Initialed)
Use engine transported by team mentor or
purchase engine from authorized range
store
Engine prepared by licensed team mentor
Section 11.2: 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
134
Attach upper main parachute assembly to
avionics bulkhead
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
135
Section 11.3: Launcher Setup and Launch Procedure
Section 11.3.1: 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
Have all viewing personnel move to safe
distance
Place ignitor against engine fuel grain
Attach launch controller to ignitor
Check for proper connection
136
Section 11.3.2: 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
Section 11.4: Post-flight Inspection
Task Completed (Initialed)
Track both halves of rocket with RF
tracker
Approach the upper section of the rocket
and turn off the altimeters with the key
switches
137
Check to see if all black powder charges
have detonated
Remove undeployed black powder charges
by cutting the wires to the e-match and
placing the charge in a bucket of water
Check altimeter beeps for launch altitude
Separate the upper parachute assembly
from the upper half of the rocket
Recover any vehicle components
Remove any environmental hazards
Disconnect the upper and lower parachute
housings from the BAE
Open up the BAE and remove the
altimeter board
Plug both altimeters into a laptop, read the
flight data and determine the kinetic
energy of the upper section impact
Team Lead Safety Officer
x___________________ x___________________
Section 12: 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.
Section 12.1: 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
139
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.
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.
Section 12.1.1: 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
140
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
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.
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
141
Section 12.2: 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
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.
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4. Take voltage measurements every thirty minutes, ending the test if any components
stopped functioning.
Section 12.2.1: Results
Due to the failure of the multimeter during the recovery test, voltage readings are not available.
Both systems outlasted the required number of hours, however.
Design Changes due to Test
No design changes were necessary, as both systems exceeded the time requirements.
Section 12.3: 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.
5. Place the assembled system on the ground or a test stand away from all personnel.
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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 12.1: Subscale Separation Test
Error! Reference source not found. shows the subscale lower section prior to conducting s
eparation 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.
Section 12.3.1: 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.
Full scale separation testing was performed at an off-site testing location on February 13, 2018,
and utilized seven grams of black powder per charge cup. The amount of black powder used was
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determined via the aforementioned equations, and confirmed with estimations based on previous
projects. These determined amounts proved to be acceptable, as the separation test was completely
successful. Both the nose cone and the recovery section separated cleanly. Additionally, both
sections separated successfully during flight testing.
Figure 12.2: Full-Scale Separation Test of the Upper Section
Design Changes due to Test
For subscale, 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.
Additionally, full scale testing with seven grams per charge cup confirmed the effectiveness of the
system. The separation system will remain unchanged in the final rocket.
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Section 12.4: 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
-Onyx 3D-printed Carbon Fiber
-Epoxy Carbon Fiber
-Epoxy Fiberglass
Procedure
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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.
Section 12.4.1: Results
Figure 12.3: Epoxy - Carbon Fiber Tension Test Results
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Figure 12.4: Epoxy - Fiberglass Tension Test Results
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Figure 12.5: 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 12.5: 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.
Section 12.5.1: Results
Figure 12.6: Epoxy - Carbon Fiber Bend Test Results
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Figure 12.7: Epoxy - Fiberglass Bend Test Results
Figure 12.8: 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 12.6: 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
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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 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.
Section 12.6.1: Results
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Onyx 3D-printed Carbon Fiber
Epoxy Carbon Fiber and Epoxy Fiberglass
After discussing testing methods with a materials testing oriented professor and consulting the
ASTM procedure for compression testing on unidirectional composites, it was determined that the
team’s home university did not have the proper equipment to conduct the test. However, based on
a combination of experience from previous projects and consultation with the assisting professor,
the team is confident in the materials’ abilities to withstand compressive forces during launch.
Potential design changes
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Although unlikely, if this year’s data did not vary significantly from the previous structural data
used to design the rocket, so no design changes will be made to reinforce vulnerable rocket
components.
Section 12.7: 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
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-Rocket rover bay (optional)
-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.
Section 12.7.1: Results
By orienting the rover bay with regards to the fins at the rocket rear, instead of needing to deploy
from any orientation the rover only needed to be able to deploy when lying at a 45-degree angle,
with either side on the top. This was them verified with a completed assembly of the entire rocket
aft of the rover bay and the rover for the four possible orientations.
Prior to the full-scale launch on February 17 in Sylacauga, the rover was placed out in the launch
field and activated remotely. Team members placed the rover in front of various terrain obstacles,
such as the aforementioned slopes, dips, mud patches, and tufts of grass. The rover was able to
contend satisfactorily with all terrain obstacles of a size similar or smaller than itself without issue.
Dried corn stalks and ruts deeper than the rover body with steep sides proved impassable, and no
attempt was made to ford deep standing water.
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Possible design impacts
As the rover was able to exit the rocket from any landed orientation, the design is considered
verified.
Section 12.8: 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.
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.
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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
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.
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Section 12.8.1: Results
The altitude control system was implemented on the March 3, 2018, full scale launch in Samson,
Alabama. The system was active, correctly predicted the altitude for the launch vehicle, and stood
by to deploy the altitude control surfaces to control the vehicle’s energy. However, due to other
circumstances with the launch inside the isogrid rocket, the apogee was only around 3200 feet, and
therefore the altitude control system did not actuate to control the rocket’s altitude. Therefore, AU
6.1-6.5 were verified, as they only relate to the electronics’ ability to function during a launch and
predict the altitude of the rocket, but AU 6.6 to 6.8, which relate to the performance of the drag
plates themselves were not verified.
Design Impact
The altitude control system electronics are verified, but since it did not manage the energy of the
rocket successfully (it did not get a chance to), the altitude control system will not be active at
competition. However, since the design is complete and modular, it can easily be implemented
into future Auburn designs.
Section 13: Project Plan
Section 13.1: Requirements Verification
Section 13.1.1: General Requirements
Table 10.1: General Requirements Verification
Requirement
Number Requirement Statement Verification Method
Execution of
Method
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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
Throughout the
competition, student
members of USLI
have handled all
tasks beyond those
specifically restricted
to the team’s mentor.
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 provided
and maintained a
project plan
throughout the
duration of the
competition.
1.3
Foreign National (FN) team
members must be identified
by the Preliminary Design
Review (PDR) and may or
may not have access to
Demonstration
FN team members
were identified by the
PDR.
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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 has
identified all
members attending
launch week by
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
requirement, all events must
occur between project
Demonstration
The team has
completed the
Educational
Engagement
requirements,
engaging over 400
participants on Junior
E-Day alone.
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acceptance and the FRR due
date
1.6
The team will develop and
host a Web site for project
documentation.
Demonstration
The team developed
and hosts 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 posted 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 are
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 contains
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 includes
page numbers.
1.11
The team will provide any
computer equipment
necessary to perform a video
teleconference with the
review panel. This includes,
but is not limited to, a
Demonstration
The team provided
all necessary
equipment for the
video
teleconferences.
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computer system, video
camera, speaker telephone,
and a broadband Internet
connection. Cellular phones
can be used for speakerphone
capability only as a last
resort.
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 used 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
implemented the EIT
Accessibility
Standards.
1.14
Each team must identify a
“mentor.” A mentor is
defined as an adult who is
included as a team member,
who will be supporting the
Demonstration
The team has
identified a mentor,
Dr. Eldon Triggs.
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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
supports. The stipend will
only be provided if the team
passes FRR and the team and
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mentor attends launch week
in April.
Section 1.1.1: Vehicle Requirements
Table 10.2: 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
The vehicle has been
launched, and upon
recovery the altimeters will
be checked.
2.2
The vehicle shall
carry one
commercially
available,
barometric
altimeter for
recording the
official altitude
used in
determining the
Inspection
Demonstration
A commercially available,
barometric altimeter has
been purchased and
calibrated for the vehicle.
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altitude award
winner.
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 altimeters are designed
and constructed to be armed
from outside the vehicle
2.4
Each altimeter
will have a
dedicated power
supply.
Inspection
Demonstration
Each altimeter in use on the
rocket has a dedicated
power supply.
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
Each arming switch cannot
be disabled by launch
forces, as seen by the
successful full scale flight.
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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 have
indicated that the rocket
will be reusable. Flight
testing has demonstrated the
launch vehicle is
recoverable and reusable
2.7
The launch
vehicle shall have
a maximum of
four (4)
independent
sections.
Demonstration
The team has designed a
launch vehicle that has two
independent sections.
2.8
The launch
vehicle shall be
limited to a single
stage.
Demonstration As designed, the launch
vehicle is a single stage.
2.9
The launch
vehicle shall be
capable of being
prepared for flight
at the launch site
within 3 hours,
from the time the
Demonstration
Subscale and full-scale
launches days have shown
that the rocket can be
readied for launch in the
required 3 days.
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Federal Aviation
Administration
flight waiver
opens.
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
Testing
All components of the
launch vehicle can remain
in a pad ready position for
over an hour without losing
functionality, as verified
through component testing.
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
Demonstration
The rocket can be flown
using the launch controller
provided by the Range
Services Provider, as
proven by the successful
full-scale launches.
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Range Services
Provider.
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 rocket, as designed,
does not require any extra
launch equation.
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
Demonstration
Vehicle is designed around
commercially available,
certified motors
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Rocketry (NAR),
Tripoli Rocketry
Association
(TRA), and/or the
Canadian
Association of
Rocketry (CAR).
2.13.1
Final motor
choices must be
made by the
Critical Design
Review (CDR).
Demonstration
The motor used for the
competition has been
determined, an Aerotech L-
1420R
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 No change will be
necessary after CDR
2.14
Pressure vessels
on the vehicle
shall be approved
by the RSO and
Analysis
Testing
No pressure vessels were
used on the rocket this year.
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shall meet the
following criteria:
2.14.1
The minimum
factor of safety
(Burst or Ultimate
pressure versus
Max Expected
Operating
Pressure) shall be
4:1 with
supporting design
documentation
included in all
milestone reviews.
Inspection
Analysis
Testing
No pressure vessels were
used on the rocket this year.
2.14.2
Each pressure
vessel shall
include pressure
relief valve that
sees the full
pressure of the
tank.
Inspection
Analysis
Testing
No pressure vessels were
used on the rocket this year.
2.14.3
Full pedigree of
the tank shall be
described,
including the
application for
Inspection
Demonstration
No pressure vessels were
used on the rocket this year.
173
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.
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 chosen a
motor with a total impulse
that does not exceed 5,120
Newton-seconds (L-class),
an Aerotech L-1420R with
an impulse of 4616
Newton-seconds.
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 designed and
tested the vehicle to ensure
that it has a stability margin
of greater than 2.0 at the
point of rail exit.
2.17
The launch
vehicle shall
accelerate to a
minimum velocity
Demonstration
Analysis
Testing
The team designed and
flight tested the vehicle to
ensure that its minimum
velocity at rail exit is at
least 52 fps.
174
of 52 fps at rail
exit.
2.18
All teams shall
successfully
launch and
recover a subscale
model of their
rocket prior to
CDR.
Demonstration
Testing
A subscale model of the
rocket was flown and
recovered successfully prior
to CDR.
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 and constructed to
resemble and perform
similarly to the full-scale
model.
2.18.2
The subscale
model shall carry
an altimeter
capable of
reporting the
Demonstration
An altimeter capable of
reporting the model’s
apogee altitude was
implemented on the
subscale model.
175
model’s apogee
altitude.
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 flown
on launch day.
The following
criteria must be
met during the
full-scale
demonstration
flight:
Testing
Demonstration
Analysis
A successful launch and
recovery of a full-scale
rocket was accomplished
prior to FRR.
2.19.1
The vehicle and
recovery system
shall have
functioned as
designed.
Testing
Demonstration
The recovery system
functioned as designed.
176
2.19.2.1
If the payload is
not flown, mass
simulators shall be
used to simulate
the payload mass.
Testing
Demonstration
Analysis
The rover (payload) was
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
By flying the actual
payload, the mass was
located exactly where
planned for the competition
launch.
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
activated during
the full-scale
demonstration of
flight.
Demonstration
Testing
The rover payload does not
affect the total energy of the
vehicle.
177
2.19.4
The full-scale
motor does not
have to be flown
during the full-
scale test flight.
Inspection
Demonstration
The full-scale motor was
not flown during the full-
scale test due to extremely
low availability, as only one
could be secured.
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 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,
Demonstration
The full-scale flight will be
completed by March 6,
2018. If for some reason the
full-scale flight is not
178
2018). If
necessary, an
extension to
March 28th, 2018
will be granted.
Only granted for
re-flights.
completed by March 6,
2018, NASA will grant an
extension to March 28,
2018 for a re-flight attempt.
2.20
Any structural
protuberance on
the rocket shall be
located aft of the
burnout center of
gravity.
Demonstration
Team designed and
constructed rocket such that
all structural protuberances
on the vehicle to be aft of
the burnout center of
gravity.
2.21 Vehicle
Prohibitions
2.21.1
The launch
vehicles shall not
utilize forward
canards.
Demonstration The Launch vehicle does
not use forward canards.
2.21.2
The launch
vehicle shall not
utilize forward
firing motors.
Demonstration
The team designed the
vehicle so that it does not
utilize forward firing
motors.
179
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 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 has flight tested
the rocket to demonstrate to
ensure that the vehicle does
not exceed Mach 1 at any
point during flight
180
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 1.1.2: Recovery Requirements
Table 10.3: Recovery Requirements Verification
Requirement
Number
Requirement
Statement Verification Method
Execution of
Method
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
Design
Testing
The team has staged
the deployment of the
recovery devices with
a drogue parachute
deployed at apogee
(5280 ft.), and two
main parachutes
deployed at 750 ft.
181
that kinetic energy
during drogue-stage
descent is reasonable,
as deemed by the
Range Safety Officer.
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 has performed a
ground ejection test
for both the drogue
and upper main
parachute section as
well as the lower
main parachute
section.
3.3
At landing, each
independent sections
of the launch vehicle
shall have a
maximum kinetic
energy of 75 ft-lbf.
Design
Demonstration
The team has flight
test and calculated
both sections of our
launch vehicle to
ensure that a
maximum energy of
182
75 ft-lbf at landing is
not exceeded.
3.4
The recovery system
electrical circuits
shall be completely
independent of any
payload electrical
circuits.
Design
The team has created
independent circuits
for the recovery
system so that they
are independent of all
payload electrical
circuits.
3.5
All recovery
electronics will be
powered by
commercially
available batteries.
Design
The recovery system
utitlizes
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
sophisticated flight
computers.
Design
The recovery system
includes a
TeleMetrum and
TeleMega altimeter
each with their own
independent set of
charges for each
section.
183
3.7
Motor ejection is not
a permissible form of
primary or secondary
deployment.
Design
The team does not
use motor ejection as
a primary or
secondary
deployment. An
electronic form of
recovery deployment
is used.
3.8
Removable shear
pins will be used for
both the main
parachute
compartment and the
drogue parachute
compartment.
Design
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.
Design
The main parachutes
deploy from a low
enough altitude so
that the rocket will
not drift more than
2500 ft.
3.10
An electronic
tracking device shall
be installed in the
launch vehicle and
shall transmit the
position of the
tethered vehicle or
Design
The team installs a
tracking device on
both sections of the
launch vehicle so that
the location of both
pieces can be
184
any independent
section to a ground
receiver.
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.
Design
All separating
sections of the rocket
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 tests and
makes sure the
electronic tracking
devices will be fully
functional during the
official flight before
each launch.
3.11
The recovery system
electronics shall not
be adversely affected
by any other on-
board electronic
devices during flight
Testing
The team has
demonstrated through
flight testing that the
recovery system will
not be adversely
affected by any other
185
(from launch until
landing).
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.
Design
The recovery system
altimeters are 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.
Design
The recovery
electronics are 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
Design
The recovery
electronics are sealed
in their own separate
compartment separate
186
which may generate
magnetic waves
(such as generators,
solenoid valves, and
Tesla coils) to avoid
inadvertent excitation
of the recovery
system.
from all other
magnetic wave
inducing devices in
the rocket.
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.
Design
The recovery
electronics are 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 13.1.2: Deployable Rover Requirements
Table 13.1: Deployable Rover Requirements Verification
187
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
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
188
Section 13.1.3: 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
the FRR report and used
during the Launch Readiness
Review (LRR) and any
launch day operations.
Demonstration The team will use
checklists.
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
Demonstration The team will follow the
rules.
189
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
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 13.2: Team Requirements
Section 13.2.1: General Requirements
Table 13.2: AU General Requirements
Team
Requirement
Requirement
Statement Derivation
Verification
Method
Method of
Execution
AU 1.1
All Educational
Engagement
forms will be
submitted and
This requirement
was made
because last year
a large number
Demonstration
The team will
submit EE
reports within a
week and check
190
verified to have
been received
within 1 week of
an outreach event.
of education
events where
completed, but
the forms where
not submitted on
time
that they have
been properly
received. This
procedure will
be continued to
be upheld going
forward.
Section 13.2.2: Vehicle Requirements
Table 13.3: AU Vehicle Requirements
Team
Requirement
Requirement
Statement Derivation
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.
It is important to
verify the material
properties of rocket
components, and to
maintain the
knowledge base to
conduct materials
testing if new
materials are
introduced in the
future.
Test
All tests are
completed, using
Auburn
University
constructed
samples on AU
testing
equipment.
AU 2.2 3D printed
components will
It is important to
verify the material Test
All tests are
completed, using
191
have strength
comparable to
alternatives and
appropriate for
their role
properties of rocket
components, and to
maintain the
knowledge base to
conduct materials
testing if new
materials are
introduced in the
future.
Auburn
University
constructed
samples on AU
testing
equipment.
Section 13.2.3: Recovery Requirements
Table 13.4: AU Recovery Requirements
Team
Requirement
Requirement
Statement Derivation
Verification
Method Method of Execution
AU 3.1
Recovery
electronics
must be able
function after
being left on
for more than
two hours.
From team
observations of last
year’s competition,
rockets could be
on the standby for
almost an hour,
and we want our
rocket to have a
large margin of
safety over this
time.
Test
The recovery
electronics (excluding
e-matches and black
powder for safety
reasons) have been
assembled and left in
the on/standby
position until loss of
function to determine
system longevity
192
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.
Too much black
powder is
dangerous, but too
little is also
dangerous because
a failed ejection
can result in a
ballistic rocket.
Test
Extensive ground
separation tests were
performed before
launch for both
subscale and full scale.
Section 13.2.4: Deployable Rover Requirements
Table 13.5: AU Rover Requirements
Team
Requirement
Requirement
Statement Derivation
Verification
Method
Method of
Execution
AU 4.1
With the
exception of
electronic parts,
the rover will be
3-D printed in-
house composites
Allows for cost
saving, rapid
design changes
and complex
geometries to
be created in
singular parts.
Demonstration
Rover parts are
designed and printed
in-house so that the
team can continue to
precisely
manufacture rover
parts
193
AU 4.2
Rover activation
signal will
successfully reach
the rover up to a
certain distance
plus a large
tolerance distance
Activating the
rover remotely
is a
requirement,
and doing so at
an additional
distance
provides for a
margin of error
at competition.
Demonstration
Test
XBee has 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)
Launch fields
can be rough,
but the rover
must be
capable of
crossing them.
Demonstration
Test
Rover treads have
tested on various
terrains before the
launches and have
demonstrated
adequate traction
AU 4.4
Rover electronics
must be able
function after
being left on for
more than two
hours.
From team
observations of
last year’s
competition,
rockets could
be on the
standby for
almost an hour,
and we want
our rocket to
have a large
margin of
Test
The rover electronics
were run
continuously for
several hours and
have been proven to
be functional for
more than two hours.
194
safety over this
time.
AU 4.5
Rover motors
must be able to
function after
more than two
hours of battery
drain.
From team
observations of
last year’s
competition,
rockets could
be on the
standby for
almost an hour,
and we want
our rocket to
have a large
margin of
safety over this
time.
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.
There is no
guarantee that
the rocket will
land at any
particular
orientation, but
the rover must
Test
The rover has placed
in section, rotated to
various angles in
increments of 30
degrees from
horizontal, and be
commanded to drive
out.
195
still be
successful
AU 4.7
Rover will
successfully
deploy solar
panels after
travelling at least
5 feet
In order to
complete the
rover
requirements,
this must
occur.
Demonstration,
test
SPDS has been
shown to deploy after
travelling over
various terrain and
travelling varying
distances
Section 13.2.5: Safety Requirements
Team
Requireme
nt
Requirement Statement Derivation 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.
Safety first Demonstratio
n
Team members
have not and will
not work alone.
196
Section 13.2.6: Altitude Control Requirements
Table 13.6: AU Altitude Control Requirements
Team
Require
ment
Requirement
Statement Derivation
Verification
Method Method of Validation
AU 6.1
All
aerodynamic
data must be
validated
through
analytical and
experimental
testing.
In order to have an
accurate computer
model, the drag
contribution of each
plate must be
computed
Analysis
An aerodynamic analysis
of the drag plates and the
internal system was
conducted through
computational fluid
dynamics (CFD) and
sensor tests.
AU 6.2
Drag plates
must stay static
throughout
launch
Deploying the drag
plates before the
boost phase is over is
against competition
rules and is
dangerous.
Demonstration
A timer has been
implemented through the
controller that prevented
any system action prior
to the end of the boost
phase.
AU 6.3
Electronics
must stay
secured
throughout
flight.
Loose electronics,
are vulnerable to
damage from launch
forces, as learned
with previous
launches.
Demonstration
A housing was made for
all electrical components
to keep everything in
place.
197
AU 6.4
Subsystem
components
must engage
accordingly
after boost
phase and stay
online for the
remainder of
the flight.
If the subsystem
components are not
engaged, they cannot
act to control the
rocket energy.
Test
Flight testing has
verified that the
behavioral integrity of
the system remains
intact.
AU 6.5
Controller and
IMU must be
able to
correctly
predict the
projected
altitude of the
launch vehicle.
The controller and
IMU must be able to
predict the max
altitude in order to
then act to control
the altitude.
Test
Flight testing has
confirmed that the IMU
can 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.
Deploying the drag
plates before the
boost phase is over is
against competition
rules and is
dangerous.
Test
“Hardware In-The-
Loop” testing has
occurred, but test flights
were unable to validate
deployment precision.
AU 6.7
Drag plates
must be able to
withstand the
If the drag plates
cannot withstand the
force required, they
Test
Wind tunnel testing and
structural testing has not
been able to ensure the
198
perpendicular
force of the
airflow.
cannot perform the
altitude control
functions.
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.
If the system is not
accurate, it could
hurt the team’s
ability to hit the
altitude requirement
instead of helping.
Analysis
Demonstration
The subsystem was
designed to conform to
accuracy requirements
and but could not 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.
From team
observations of last
year’s competition,
rockets could be on
the standby for
almost an hour, and
we want our rocket
to have a large
margin of safety over
this time.
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.
199
Section 13.3: Budget
The budgets displayed in Table 13.9: Budget Allocation are the final costs allocated to the full-
scale vehicle, research and development, the subscale (rounded up for a factor of safety), travel,
and test flight fees. As the team will continue to perform outreach events, the educational
engagement budget is a close approximation of the expected final expenditures. 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 team paid $2,200 for these rooms, and thanks to the team’s close proximity to
Huntsville, other travel costs should be negligible. 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 $1,000 so far, and is estimated to cost
$500 more. The team has spent $3,410.67 for the rocket on the pad, $2,000 for the sub-scale
vehicle, and $2,200 for travel, $1,794 for research and development, and $600 on test flights. The
team plans to spend a total of $1,500 on educational outreach, leaving $2,995.33 for promotional
items and additional costs, based on the $16,500 amount for total funding presented in Table 13.10:
Funding Sources.
Overall, the team is quite happy with how the finances were managed this year. Thanks to the
recovery of all components from all launches this year, no additional funds had to be allocated to
replacing broken or missing materials. The additional funds will be spent on accruing materials
for level 2 certification flights, additional research for future endeavors, and updated facilities after
the competition has been completed.
Below are the budgets for the on-the-pad rocket.
Table 13.7: Vehicle Costs
Vehicle (Full Scale)
200
Item Cost Per Unit Unit Quantity Total
Pre-preg Carbon
Fiber $118 Per yard 6 $708
Aerotech L1420R $259.99 Per unit 1 $259.99
RMS 75/5120
Motor Case and
Associated
Hardware
$390 Per unit 1 $390
Fiberglass
Coupler $69 Per unit 3 $207
Rail Buttons $3 Per unit 2 $6
Total $1,571
Table 13.8: 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
201
Jolly Logic $130 Per unit 2 $260
Total $1,039
Rover (Full Scale)
Item Cost Per Unit Unit Quantity Total
1000:1 Micro
Metal Gearmotor
HPCB 12V
$24.95
Per Unit 12 $299.4
2 X 6inch IPX /
U.fl to to RP-
SMA Male Plug
Straight Wifi
Antenna
Extension Cable
Pigtail Cable
1.13mm 15cm
for Wireless 6"
$9.99
Per Unit
1
$9.99
AltIMU-10 v5
Gyro,
Accelerometer,
Compass, and
Altimeter
(LSM6DS33,
LIS3MDL, and
LPS25H Carrier)
$22.95
Per Unit
1
$22.95
202
Arduino
Wireless SD
shield
$17.49
Per Unit 1
$17.49
Continuous
Rotation Servo -
FeeTech
FS5103R
$11.95
Per Unit 1
$11.95
Digi
International
XBP9B-DPUT-
001
$39.99
Per Unit 2 $79.98
Elegoo UNO
project $35.00 Per Unit 1 $35.00
New Pair Laptop
Wireless Mini
PCI PCI-E WIFI
Bluetooth
Internal
$8.99 Per Unit 1 $ 8.99
Qunqui L298N
Motor Driver $7.00 Per Unit 1 $7.00
Roll of Onyx $189.00 Per Roll 1 $189.00
TRENDnet Low
Loss Reverse
SMA Female to
N-Type Male
Weatherproof
$22.99 Per Unit 1 $22.99
203
Connector Cable
(8M, 26.2ft.)
TEW-L208
U.FL Mini PCI
to RP-SMA
Pigtail Antenna
WiFi Cable Pack
of 2
$4.99 Per Unit 1 $4.99
Xbee Explorer
dongle $24.95 Per Unit 2 $49.90
Xbee Pro 60 mW
Wire Antenna $37.95 Per Unit 2
$75.90
$800.67
Below are costs associated with research and development. As the altitude control module will be
unable to be on the pad, those costs are factored in here as well. The majority of the R&D expenses
went towards constructing the isogrid airframe to assess its performance compared to pre-preg
carbon fiber rolls, and the construction of a second test rocket as an asset and a new member
development tool.
Research and Development Costs
Cost of Rocket On the Pad $3,410.67
204
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 3 $354
Aerotech L1520T $199.99 Per unit 1 $199.99
RMS 75/3840
Motor Case and
Associated
Hardware
$385 Per unit 1 $385
Rail Buttons $3 Per unit 2 $6
Stratologger CF $55 Per unit 2 $110
Roll of Onyx $189.00 Per Roll 1 $189.00
Total $1,689
Altitude Control Module (R&D)
Item Cost Per Unit Unit Quantity Total
Arduino Uno $8 Per unit 1 $22
AndyMark
NeveRest 40 DC
Motor
$8 Per unit 1 $28
205
Adafruit Motor
Shield v.2.3 $1 Per unit 1 $20
Adafruit
accelerometer $5 Per unit 1 $15
Adafruit Micro
SD card
breakout
$200 Per unit 1 $8
Adafruit
barometric
pressure sensor
$300 Per unit 1 $10
9V Battery $130 Per unit 2 $2
Total $105
Table 13.9: Budget Allocation
Item Cost
Full-Scale $3,410.67
Sub-Scale $2,000
Travel $2,200
Educational Outreach $1,500
Test flights (3) $600
Research and Development $1,794
206
Promotional Items $1,000
Total $12,504.7
Figure 13.1: Spending Comparison
Section 13.4: Funding Plan
The team has secured funding from the sources presented in Table 13.10: 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. This funding was more than sufficient for the cost of the project, but the team hopes to
accrue additional sponsors to increase the scope of future projects.
Table 13.10: Funding Sources
Budget Breakdown
Full-Scale Sub-Scale Travel
Educational Outreach Test Flights (5) Development
Promotional Items
207
Source Amount
Alabama Space Consortium $12,000
Dynetics $2,500
Lockheed Martin $2,000
Total Funding $16,500
Section 13.5: 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 it helps to remove
clutter. These can be found in below in Figure 13.2: Fall Timeline and Figure 13.3: Spring
Timeline.
Figure 13.2: 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
208
Figure 13.3: Spring Timeline
As seen in the timelines, the team has completed all research and construction of the full-scale
rocket. The team attended two separate launch dates and recovered the rocket at both.
Unfortunately, the dates of Rocket Week had to be pushed back to accommodate Drake Middle
School; however, the team still plans to go ahead with this outreach event even though it occurs
past the educational engagement deadline. For now, all that remains for the team is to get ready
and excited for Huntsville, and complete the Post-Launch Assessment Review afterwards.
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
Two Day Full-Scale Flight Opportunity
Emergency Full-Scale Flight
Competition Preparation
Huntsville Competition
Post-Launch Assessment Review
