project aquila - auburn · pdf filesection 1.2: launch vehicle summary ... project aquila...
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PROJECT AQUILA
211 ENGINEERING DRIVE
AUBURN, AL 36849
FLIGHT READINESS REVIEW REPORT
MARCH 14, 2016
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Table of Contents
Table of Contents ............................................................................................................................ 2
List of Figures ................................................................................................................................. 7
List of Tables ................................................................................................................................ 10
Section 1: Summary of FRR Report ........................................................................................ 12
Section 1.1: Team Summary................................................................................................ 12
Section 1.2: Launch Vehicle Summary ............................................................................... 12
Section 1.3: Payload Summary ............................................................................................ 13
Section 2: Changes Since CDR ................................................................................................ 15
Section 2.1: Vehicle Changes .............................................................................................. 15
Section 2.2: Payload Changes.............................................................................................. 15
Section 2.3: Project Plan Changes ....................................................................................... 15
Section 3: Vehicle Design ........................................................................................................ 16
Section 3.1: Mission Statement ........................................................................................... 16
Section 3.2: Structural Elements.......................................................................................... 16
Section 3.2.1: Structure ...................................................................................................... 17
Section 3.2.2: Propulsion ................................................................................................... 20
Section 3.2.3: Aerodynamics ............................................................................................. 24
Section 3.3: Design Integrity ............................................................................................... 26
Section 3.3.1: Fin Shape and Style ..................................................................................... 26
Section 3.3.2: Materials in Fins, Bulkheads and Structural Elements ............................... 27
Section 3.4: Drawings and Schematics ................................................................................ 29
Section 3.5: Testing ............................................................................................................. 37
Section 3.5.1: Materials Testing ......................................................................................... 37
Section 3.5.2: Wind Tunnel Testing................................................................................... 40
Section 3.6: Workmanship................................................................................................... 41
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Section 3.7: Mass Report ..................................................................................................... 42
Section 3.8: Requirement Verification ................................................................................ 42
Section 4: Recovery ................................................................................................................. 60
Section 4.1: Recovery System Overview ............................................................................ 60
Section 4.2: Structural Elements.......................................................................................... 61
Section 4.3: Electrical Elements .......................................................................................... 63
Section 4.4: Parachutes ........................................................................................................ 66
Section 4.4.1: Parachute Sizing .......................................................................................... 66
Section 4.4.2: Manufacturing ............................................................................................. 70
Section 4.4.3: Deployment Process .................................................................................... 71
Section 4.4.4: Drift ............................................................................................................. 73
Section 4.5: Testing ............................................................................................................. 74
Section 4.5.1: Subscale Testing.......................................................................................... 74
Section 4.6: Requirement Verification ................................................................................ 76
Section 5: Full Scale Results .................................................................................................... 80
Section 5.1: Project Aquila Test Launch 1 .......................................................................... 80
Section 5.2: Project Aquila Test Launch 2: ......................................................................... 80
Section 5.3: Project Aquila Test Launch 3: ......................................................................... 80
Section 5.4: Project Aquila Test Launch 4: ......................................................................... 81
Section 6: Payload Fairing ....................................................................................................... 82
Section 6.1: Design Overview ............................................................................................. 82
Section 6.2: Payload Fairing Materials ................................................................................ 85
Section 6.3: Payload Fairing Testing ................................................................................... 85
Section 6.3.1: Aerodynamic Design Testing ...................................................................... 85
Section 6.3.2: Charge Chamber Strength Testing .............................................................. 86
Section 6.3.3: Ground Testing............................................................................................ 88
Section 6.3.4: Water Seal Test ........................................................................................... 89
Section 6.3.5: Full Scale Testing ........................................................................................ 89
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Section 6.4: Payload Fairing Requirements......................................................................... 90
Section 7: Aerodynamic Analysis Payload - WAFLE ............................................................. 92
Section 7.1: Experiment Concept ........................................................................................ 92
Section 7.2: Science Value .................................................................................................. 92
Section 7.2.1: Payload Objectives ...................................................................................... 92
Section 7.2.2: Mission Success Criteria ............................................................................. 92
Scientific .......................................................................................................... 94
Section 7.3: Experiment....................................................................................................... 94
Section 7.4: Flight Performance Predictions ..................................................................... 121
Section 7.5: Payload Design .............................................................................................. 123
Section 7.6: Requirement Verification .............................................................................. 129
Section 7.7: Payload Integration ........................................................................................ 132
Section 8: Safety .................................................................................................................... 137
Section 8.1: Safety Officer ................................................................................................ 137
Section 8.2: Airframe Hazard Analysis ............................................................................. 137
Section 8.2.1: Airframe Failure Modes ............................................................................ 138
Section 8.2.2: Airframe Risk Mitigation – Testing Systems............................................ 149
Section 8.3: Scientific Payloads Hazard Analysis ............................................................. 150
Section 8.3.1: Scientific Payload Risk Mitigation – Payload Fairing .............................. 151
Section 8.3.2: Scientific Payload Risk Mitigation – WAFLE ......................................... 160
Section 8.4: Recovery Hazard Analysis ............................................................................ 179
Section 8.4.1: Recovery Risk Mitigation – Materials ...................................................... 184
Section 8.4.2: Recovery Risk Mitigation - Construction ................................................. 190
Section 8.5: Outreach Hazard Analysis ............................................................................. 195
Section 8.6: Environmental Effects ................................................................................... 200
Section 8.6.1: Vehicle Effects on Environment ............................................................... 200
Section 8.6.2: Environmental Effects on the Vehicle ...................................................... 202
Section 9: Launch Operations Procedures ............................................................................. 205
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Section 9.1: Parts Checklists.............................................................................................. 205
Section 9.2: Final Assembly Checklists ............................................................................ 207
Section 9.3: Motor Preparation .......................................................................................... 211
Setup for Launch/ .......................................................................................... 212
Section 9.4: Igniter Installation.......................................................................................... 212
Section 9.5: Launch Procedures ........................................................................................ 213
Section 9.6: Troubleshooting ............................................................................................. 215
Section 9.7: Post-Flight Inspection .................................................................................... 217
Section 10: Project Plan ........................................................................................................... 218
Section 10.1: Budget ............................................................................................................ 218
Section 10.2: Funding Plan .................................................................................................. 219
Section 10.3: Timeline ......................................................................................................... 219
Section 11: Educational Engagement ...................................................................................... 223
Section 11.1: Drake Middle School 7th Grade Rocket Week .............................................. 223
Section 11.1.1: Rocket Week Plan of Action ................................................................... 224
Section 11.1.2: Rocket Week Launch Day ...................................................................... 225
Section 11.1.3: Rocket Week Learning Objectives.......................................................... 226
Section 11.1.4: Gauging Success ..................................................................................... 227
Section 11.2: Samuel Ginn College of Engineering E-Day ................................................ 227
Section 11.3: Boy Scouts of America: Space Exploration Badge ....................................... 227
Section 11.3.1: Space Exploration Merit Badge Requirements ....................................... 228
Section 11.3.2: Boy Scouts of America - AUSL Requirements ...................................... 229
Section 11.3.3: Boy Scouts of America - Plan of Action ................................................. 230
Section 11.3.4: Boy Scouts of America: Goals ................................................................ 231
Section 11.4: Girl Scouts of the USA - Space Badge .......................................................... 231
Section 11.5: Auburn Junior High School Engineering Day ............................................... 232
Section 12: Conclusion ............................................................................................................ 233
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List of Figures
Figure 1.1: Aerodynamic Grid Fin................................................................................................ 13
Figure 3.1: Full Rocket Rendering................................................................................................ 16
Figure 3.2: 5 Inch Braided Isogrid ................................................................................................ 18
Figure 3.3: Motor Tube Rendering ............................................................................................... 20
Figure 3.4: Aerotech L1520T Thrust Curve ................................................................................. 20
Figure 3.5: Aeropack Motor Retention ......................................................................................... 23
Figure 3.6: Fin Rendering ............................................................................................................. 24
Figure 3.7: Fin Shapes. ................................................................................................................ 27
Figure 3.8: Patran Tube Model ..................................................................................................... 28
Figure 3.9: Upper Section Dimensions ......................................................................................... 29
Figure 3.10: Lower Section Dimensions ...................................................................................... 30
Figure 3.11: Booster Tube ............................................................................................................ 31
Figure 3.12: Upper Body Tube ..................................................................................................... 32
Figure 3.13: Fins ........................................................................................................................... 33
Figure 3.14: Filament Wound Body Tube .................................................................................... 34
Figure 3.15: Bulkhead ................................................................................................................... 35
Figure 3.16: Motor Tube Bulkhead .............................................................................................. 36
Figure 3.17: Carbon Fiber Test Data ............................................................................................ 38
Figure 3.18: Carbon Fiber Test Results ........................................................................................ 38
Figure 3.19: HIPs Data ................................................................................................................. 39
Figure 3.20: HIPS Test Results ..................................................................................................... 39
Figure 3.21: Three Point Bending Test ......................................................................................... 39
Figure 3.22: FPS vs Lb Force ....................................................................................................... 40
Figure 3.23: Wind Tunnel Test ..................................................................................................... 40
Figure 4.1: Parachute Configuration ............................................................................................. 60
Figure 4.2: BAE Bottom View ..................................................................................................... 62
Figure 4.3: BAE Cutaway View ................................................................................................... 62
Figure 4.4: BAE Side View .......................................................................................................... 63
Figure 4.5: Altus Metrum Telemega ............................................................................................. 64
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Figure 4.6: AltusMetrum Telemetrum .......................................................................................... 64
Figure 4.7: Taoglas FXP240 433 MHz ISM Antenna .................................................................. 65
Figure 4.8: Parachute Shape Parameters ....................................................................................... 67
Figure 4.9: Main Parachute Visualization .................................................................................... 69
Figure 4.10: Pictures of Tender Descender in Undeployed and Deployed Configurations .......... 71
Figure 4.11: Subscale Recovery Configuration ............................................................................ 75
Figure 7.1: Connectors and Points on a Fin in Pointwise ............................................................. 97
Figure 7.2: Domain on a Fin in Pointwise .................................................................................... 97
Figure 7.3: Domain on a Grid Fin in Pointwise ............................................................................ 98
Figure 7.4: Rocket with Mesh and All Domains in Pointwise ..................................................... 99
Figure 7.5: Final Mesh and Boundaries on Rocket in Pointwise .................................................. 99
Figure 7.6: Starting points created over the top face of the grid fin to demonstrate the starting point
of the flow. .................................................................................................................................. 101
Figure 7.7: Flow is directed over the grid fins and represented by arrows and lines. The lines and
arrows are representatives of pressure due to flow. .................................................................... 102
Figure 7.8: A flow representation of 0.2 Mach flow over a grid fin. The arrows and lines represent
the pressure over the grid fin. ..................................................................................................... 103
Figure 7.9: Total Fin Axial Force Coefficient versus Angle of Attack Mach 8 Low Pressure, Low
Temperature ................................................................................................................................ 106
Figure 7.10: Fin Normal Force Coefficient versus Angle of Attack Mach 8 High Pressure, High
Temperature ................................................................................................................................ 107
Figure 7.11: Fin Normal Force Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low
Temperature ................................................................................................................................ 107
Figure 7.12: Fin Normal Force Coefficient versus Angle of Attack Mach 0.1 High Pressure, High
Temperature ................................................................................................................................ 108
Figure 7.13: Fin Moment Coefficient versus Angle of Attack Mach 8 Low Pressure, Low
Temperature ................................................................................................................................ 108
Figure 7.14: Fin Moment Coefficient versus Angle of Attack Mach 8 High Pressure, High
Temperature ................................................................................................................................ 109
Figure 7.15: Fin Moment Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low
Temperature ................................................................................................................................ 109
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Figure 7.16: Fin Moment Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low
Temperature ................................................................................................................................ 110
Figure 7.17: Vortex Shedding Testing visualization .................................................................. 119
Figure 7.18: WAFLE system ...................................................................................................... 123
Figure 7.19: Arduino Uno ........................................................................................................... 124
Figure 7.20: Savox SV-1270TG ................................................................................................. 125
Figure 7.21: 10-DOF IMU .......................................................................................................... 126
Figure 7.22: Grid Fin Fairing ...................................................................................................... 127
Figure 7.23: Aerodynamic Grid fin ............................................................................................ 128
Figure 7.24: WAFLE electronics schematic ............................................................................... 129
Figure 7.25: LANTERN Configuration ...................................................................................... 134
Figure 7.26: Servo Bracket ......................................................................................................... 135
Figure 11.1: Picture from Rocket Week 2014 ............................................................................ 224
Figure 11.2: A photo taken from DMS 7th Grade Rocket Week in April 2014 ......................... 226
Figure 11.3: Space Exploration Merit Badge ............................................................................. 230
Figure 11.4: A Photo taken from Auburn Junior High School Engineering Day ....................... 232
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List of Tables
Table 1.1: General Team Information .......................................................................................... 12
Table 1.2: Team Leadership ......................................................................................................... 12
Table 1.3: Launch Vehicle Summary ........................................................................................... 13
Table 3.1: Vehicle Length............................................................................................................. 17
Table 3.2: Aerotech L1520T Motor Specifications ...................................................................... 21
Table 3.3: Fin Dimensions ............................................................................................................ 25
Table 3.4: Final Mass .................................................................................................................... 42
Table 3.5: Verification Plan .......................................................................................................... 43
Table 4.1: Parachute Shape Pugh Chart ........................................................................................ 68
Table 4.2: Main Parachute Dimensions ........................................................................................ 69
Table 4.3: Kinetic Energy Calculations ........................................................................................ 70
Table 4.4: Drift Calculations......................................................................................................... 74
Table 4.5: Recovery Requirement Validation .............................................................................. 76
Table 7.1: Aerodynamic Payload Success Criteria ....................................................................... 92
Table 7.2: Aerodynamic Payload Simulation List ........................................................................ 94
Table 7.3: SolidWorks simulation run cases............................................................................... 103
Table 7.4: Aerodynamic Payload Fortran- Flight and Dynamic model...................................... 104
Table 7.5: Sample Data Mach=0.8 Low Pressure, Low Temperature ........................................ 110
Table 7.6: Sample Data at Mach=0.8 High Pressure, High Temperature................................... 112
Table 7.7: Sample Data at Mach 0.1 Low Pressure, Low Temperature ..................................... 113
Table 7.8: Sample Data at Mach 0.1 High Pressure. High Temperature .................................... 115
Table 7.9: Calculated Drag and Acceleration Values ................................................................ 117
Table 7.10: WAFLE Constants ................................................................................................... 121
Table 7.11: Vehicle Constants ................................................................................................... 121
Table 7.12: Estimations from Miller's Document ...................................................................... 122
Table 7.13: Constants from Transition Open Rocket ................................................................. 122
Table 7.14: Calculated Drag and Acceleration Values ............................................................... 122
Table 7.15: Aerodynamic Payload System Validation Table ..................................................... 129
Table 8.1: Risk Mitigation Table – Airframe ............................................................................. 140
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Table 8.2: Risk Mitigation Table – Autoclave............................................................................ 142
Table 8.3: Risk Mitigation Table - Filament Winder ................................................................. 144
Table 8.4: Risk Mitigation Table - Carbon Fiber ....................................................................... 146
Table 8.5: Risk Mitigation Tables – Epoxy ................................................................................ 146
Table 8.6: Risk Mitigation Table - Flight Recovery Operations ................................................ 179
Table 8.7: Risk Mitigation Tables - Shear Pin Test Rig ............................................................. 183
Table 8.8: Risk Mitigation Table - Outreach Operations ........................................................... 196
Table 8.9: Risk Mitigation Table - Outreach Construction ........................................................ 198
Table 10.1: Final Budget............................................................................................................. 218
Table 10.2: Funding Sources ...................................................................................................... 219
Table 10.3: Launches and Vehicle Timeline .............................................................................. 221
Table 10.4: Subsystem Timeline................................................................................................. 221
Table 10.5: Competition Timeline .............................................................................................. 222
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Section 1: Summary of FRR Report
Section 1.1: Team Summary Table 1.1: General Team Information
Team Affiliation Auburn University
Mailing Address 211 Engineering Drive Auburn, AL 36849
Title of Project Project Aquila
Table 1.2: Team Leadership
Student Team Lead Cassandra Seelbach
Safety Officer Austin Phillips
Academic Advisor Dr. Joseph Majdalani
NAR/Tripoli Advisor Dr. Eldon Triggs
Section 1.2: Launch Vehicle Summary Table 1.3 gives the basic details of the launch vehicle. The vehicle was designed to
accommodate the chosen payloads and electronics while simultaneously providing stability and
proper weight for reaching the competition altitude. More information regarding the launch
vehicle can be found in Section 3 of this report.
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Table 1.3: Launch Vehicle Summary
Total Length 75.125 inches
Final Mass 31.3 lbs
Motor Selection Loki L1482
Section 1.3: Payload Summary One of the payloads chosen is to perform aerodynamic analysis on structural protuberances. The
structural protuberance was chosen to be a grid fin control surface. Grid fins are control surfaces
with a lattice as the main structure, as show in Figure 1. Grid fins are a new type of control surface
in the realm of aerospace, therefore there is minimal public data on how the control surface reacts
in flight. This will provide a challenge for the team but will also provide valuable data to the public
domain.
Figure 1.1: Aerodynamic Grid Fin
The grid fin will be mounted to the side of the air frame exposed to the flow. After the boost phase
and as the rocket travels to apogee, the grid fins will be deployed at indicated times to correct the
rockets trajectory. An Arduino will determine the deployment of the grid fins. The Arduino will
interpret data from sensors stored in the rocket body and deploy the fins to slow the velocity of the
rocket and decrease the height, insuring that the rocket reaches the mile height requirement. As the
rocket travels through apogee and initiates the decent, the fins will store flat to the rocket. The grid
fins will remain stored throughout the decent phase and landing phase.
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The second payload chosen was a payload fairing. The fairing serves as the nosecone for the rocket
and separates at apogee, deploying the drogue parachute and upper main in a bag using the tender
descender system.
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Section 2: Changes Since CDR
Section 2.1: Vehicle Changes The motor selection has been changed from an Aerotech L1520T to a Loki L1482-LB as a result
of multiple motor CATOs in full scale flight testing.
Section 2.2: Payload Changes The following changes have been made to the payload fairing:
1. The outer lip has been extended to protect seam from being forced open by forces of
launch.
2. The entire fairing will be sealed in wax prior to launch to ensure no leaks.
The following changes have been made to the subsystems of the Wall Armed Fin-Lattice
Elevator (WAFLE) system.
1. Savox servos have replaced the HiTec servos. The Savox servos produce much more
torque while maintaining the same dimensions and minimal weight addition.
2. A RF tracker will replace the GPS integrated into the WAFLE. The RF tracker will work
to locate the WAFLE section.
3. A 10 DOF IMU sensor will replace the ADXL335 Accelerometer and the GPS. The 10
DOF will read the acceleration and altitude of the WAFLE.
Section 2.3: Project Plan Changes The team will be rebuilding and launching a fifth full scale rocket on April 1st as a result of the
multiple motor CATOs that occurred during full scale flight testing.
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Section 3: Vehicle Design
Section 3.1: Mission Statement The Auburn University Student Launch team (AUSL) is determined to design and manufacture
an effective and unique launch vehicle. Learning from past experiences and Auburn’s history
with the competition, AUSL has re-examined every component of the launch vehicle. In order to
reach the goals set by NASA in this year’s competition, the team must achieve the highest
possible quality for all components.
Section 3.2: Structural Elements The vehicle has been designed to satisfy mission requirements set forth by NASA in the 2015-
2016 NASA Student Launch Handbook, as well as requirements set by the team. These
requirements are detailed in Section 6. The vehicle design must ensure adequate space for
avionics and payload equipment and electronics. These systems are vital to the success of the
scientific mission. The vehicle design is also heavily driven by manipulating weight and length
to control altitude and stability. These factors determine the success of the flight itself. The
vehicle design is separated into three major divisions: structure, propulsion and aerodynamics.
Figure 3.1: Full Rocket Rendering
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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 3.2.1: Structure
The structure of the launch vehicle must be able to withstand the forces the rocket will
experience during operation. The launch vehicle body must be strong enough to maintain stable
flights. Additionally, the vehicle structure must accommodate all other subsystems, ensuring they
have adequate space and protection. The design of the structure requires heavy tradeoffs between
strength, space, and weight.
The total length of the rocket is 73.125 inches. Component lengths are shown in Table 3.2. Table 3.1: Vehicle Length
Component Length (Inches)
Nose Cone (Fairing) 13.125
Upper Tube 25
Booster Section 37
Total 75.125
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, with
the exception of the motor. To ensure mission success, it is critical to select and design body
tubes that can survive the stresses of high-powered flight while still remaining light enough to
achieve the mission success.
The body tubes will be constructed using carbon fiber braiding, a process that involves taking
individual strands of carbon fiber and stitching them into a tightly-wound braid. The carbon fiber
braids that are produced will be formed into an isogrid structure around a 5 inch mandrel. Isogrid
structures are a lighter alternative to using a solid tube structure. For aerodynamic purposes, a
Kevlar “sock” will be placed over the braiding providing an exterior skin. By giving the structure
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this skin, the result is a high tensile strength-to-weight ration and a lightweight aerodynamic
body. Using this wrapped isogrid method, the mass of the body tubes will be decreased by
approximately 20 to 30 percent compared to tubes constructed using only filament wound carbon
fiber, while also maintaining the same compressive strength properties as a carbon fiber tube.
This mass reduction was confirmed using tube samples constructed by team members using final
production methods. An image of a sample of the braided isogrid structure without the
aerodynamic skin can be seen in Figure 3.2.
Figure 3.2: 5 Inch Braided Isogrid
Coupler:
The coupler serves as a joint between the two body tube sections. The coupler is designed to
separate during the recovery phase of the flight. To accomplish this, the lower body tube is
attached to the coupler using two nylon machine screws which will function as shear pins during
separation. The upper end of the coupler will remain fixed to the upper body tube using four
aluminum bolts.
The coupler will be constructed by rolling a carbon fiber tube with the required length and
thickness. The team has vast experience with this method of construction working on both
previous USLI competition rockets and personal projects, and can create couplers with the
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required degree of accuracy. By creating the coupler entirely out of carbon fiber, we can ensure
the structure is capable of withstanding the expected forces during flight.
Ballast Tank:
The ballast tank is used to hold additional mass if balance corrections must be made. The design
allows for easy mass addition and reduction as needed to account for variations in mass
predictions and launch day conditions. One ballast tank is placed forward of the avionics bay to
add to the stability of the vehicle. Another Ballast tank will be placed forward of the motor tube.
As the tank will not be subjected to a large force, the team is confident that the pins will hold the
tank securely without potential for a shear failure. The tank was constructed using high impact
polystyrene (HIPS) on a TAZ4 3D printer.
Bulkheads:
Bulkheads are typically flat plates used to increase the structural strength of a rocket. They are
also used to create airtight spaces and to divide the body into separate compartments. In rockets,
they are commonly used to separate payload bays and to mount equipment for avionics and
payloads. For rockets similar in size to the Project Aquila rocket, the material used varies from
fiberglass to plywood to carbon fiber. The bulkheads for this rocket will be made from pre-
impregnated carbon fiber and manufactured using the Computer Numerical Control (CNC)
machine at Auburn University Aerospace Design Lab due to the availability and the teams
experience with pre-impregnated carbon fiber. The interior diameter for the circular cross-
sectional rocket will be 5 inches and the bulkheads are designed to fit perfectly into this size. All
bulkheads for this rocket will be 0.125 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 Aquila rocket will be using three centering rings.
These centering rings are located in the engine tube and serve to attach to the fin set and to the
motor retention. Like the bulkheads, the centering rings are made of carbon fiber and
manufactured using the Auburn University Aerospace Design Lab’s Computer Numerical
Control (CNC) machine. The centering rings have an outer diameter of 5 inches with an inner
diameter of 3 inches. The thickness of each ring is approximately 0.125 inches.
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Section 3.2.2: Propulsion
The propulsion system includes the motor, motor tube and motor retention. These parts must
function flawlessly to ensure a safe and stable launch. An initial rendering of the propulsion
system can be viewed in Figure 3.3.
Figure 3.3: Motor Tube Rendering
Motor:
The motor selected for the competition is the Aerotech L1520T. The specifications are listed
below in Table 3.4. Additionally, the thrust curve for this motor is shown in Figure 3.4.
Figure 3.4: Aerotech L1520T Thrust Curve
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This motor was chosen based on OpenRocket simulations, as it provides the roughly 13-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 second, thus reaching the required 13-to-1 thrust ratio in about
one second. Based on OpenRocket simulations, the motor provided an apogee in excess of 5479
feet with a max acceleration of 427 ft/s2 which delivers a max velocity of 857 ft/s or close to
Mach = 0.77. Table 3.2: Aerotech L1520T Motor Specifications
Motor Specifications
Manufacturer Aerotech
Motor Designation L1520T
Diameter 2.95 in
Length 20.9 in
Impulse 3769 N-s
Total Motor Weight 128 oz
Propellant Weight 62.8 oz
Average Thrust 340 lbs
Maximum Thrust 382 lbs
Burn Time 2.49 s
Due to the failure rate the team has experienced with the Aerotech L-1520T, the team has opted to switch to a Loki L-1482. The Loki motor will provide a very similar performance to the Aerotech motor. With the Loki L-1482 the vehicle is simulated to reach a velocity of 780 feet per second or Mach 0.7. The simulations indicate that apogee will be 5400 feet.
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Motor Specifications
Manufacturer Loki
Motor Designation L1482
Diameter 2.95 in
Length 19.6 in
Impulse 3882 N-s
Total Motor Weight 7.78 lbs
Propellant Weight 4.05 lbs
Average Thrust 339 lbs
Maximum Thrust 407 lbs
Burn Time 2.6 s
Motor Tube:
To contain the motor on the rocket, a carbon fiber motor tube is being used. The motor tube will
be made by braiding carbon fiber strands and then filament wound around a mandrel that is the
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same diameter of the motor. The 3D braided carbon fiber material was chosen for its strength-to-
weight ratio when compared to a solid carbon tube. Basalt fiber was considered to be used for the
motor tube for its high heat resistance properties, but the team decided the weight of the basalt,
which was approximately 50% heavier when compared with the carbon fiber was not worth the
tradeoff. The tube will be 0.1 inch thick and is designed to fit around an Aerotech L1520T-P
motor. With these specifications, the motor tube will be ideal for the rocket and mission success.
To mount the motor tube, three centering rings will be epoxied to the outer diameter of the motor
tube and the inner diameter of the lower section tube. The epoxy will be a 24-hour epoxy, which
will create a permanent bond between the components. A bulk plate will be epoxied forward of
the motor tube. This is to provide extra strength to hold the motor in place as well as separate the
motor from the internal components of the rocket.
Motor Retention:
The purpose of the motor retention system is to secure the rocket motor during launch and flight
and to be easily removable for subsequent flights. The team has chosen a commercial bought
Aeropack motor retention system, Figure 3.5. This is a simple system with two components. One
component will bolt directly into a centering ring, using aluminum bolts. The other component
threads onto the part that is bolted onto the structure. This allows for a fast replacement of a used
motor. The team chose a commercial motor retention system due to past reliability and to avoid
the time requirements of designing and manufacturing a custom system.
Figure 3.5: Aeropack Motor Retention
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Section 3.2.3: Aerodynamics
The aerodynamics system requires the rocket remain stable during flight. The placement and
design of the aerodynamic surfaces determines the center of pressure along the length of the
rocket.
Fins:
The stability of the rocket is controlled by the fins. The primary purpose of the fins is to keep the
center of pressure aft of the center of gravity. The greater drag on the fins will keep them behind
the upper segments of the vehicle, thus allowing the rocket to fly straight along the intended
flight path. They are also helpful in minimizing the chances of weather-cocking. Fins serve as an
ideal addition to the vehicle body as they are lightweight and easy to manufacture with the tools
the team has available. A clipped delta planform was selected for the fins. A rendering of the fin
design is shown in Figure 3.6.
Figure 3.6: Fin Rendering
The trailing edge of each fin are located one inch forward of the end of the body tube. This
design feature will theoretically provide some impact protection for the fins when the rocket hits
ground. Each fin will have a surface area of 54 in2 (summing both sides), making the fin surface
area total equal to 216 in2. The total component mass is 13.5 ounces. These dimensions provide
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the vehicle with a projected stability of 2.25 calibers. This level of stability is close to ideal, as it
is well above stable, yet still below over-stable. Detailed fin dimensions are provided in Table
3.4.
Total CP location calculated from separate locations Xi and normal force coefficient derivatives
𝑋𝑋 =∑ 𝑋𝑋𝑖𝑖(𝐶𝐶𝑁𝑁𝛼𝛼)𝑖𝑖𝑛𝑛𝑖𝑖=1
∑ (𝐶𝐶𝑁𝑁𝛼𝛼𝑛𝑛𝑖𝑖=1 )𝑖𝑖
Single fin 𝐶𝐶𝑁𝑁𝛼𝛼 at subsonic speeds:
(𝐶𝐶𝑁𝑁𝛼𝛼)1 =2𝜋𝜋 𝑠𝑠2
𝐴𝐴𝑟𝑟𝑟𝑟𝑟𝑟
1 + �1 + ( 𝛽𝛽𝑠𝑠2𝐴𝐴𝑟𝑟𝑖𝑖𝑛𝑛 cos Γ𝑐𝑐
)^2
Table 3.3: Fin Dimensions
Trapezoidal Fin Dimensions
Root Chord 6.25 in
Tip Chord 2.5 in
Height 6 in
Sweep 3.68 in
Sweep Angle 31.5 °
Thickness 0.33 in
Aero-elastic flutter has been considered as a potential failure mode for the rocket structure. At a
particular 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 equation below represents the NACA flutter boundary equation with thin plate
theory included.
𝑉𝑉𝑟𝑟 = 𝑎𝑎�𝐺𝐺
1.337𝐴𝐴𝑅𝑅3𝑃𝑃(𝜆𝜆 + 1)
2(𝐴𝐴𝑅𝑅 + 2)(𝑡𝑡𝑐𝑐)3
26
The flutter velocity is directly reflective of the aero-elastic conditions of the structure/fin system.
The catastrophic flutter phenomenon results from coupling of aerodynamic forces creating a
positive feedback loop. The increase in either torsion or bending drives an infinitely looped
increase in the other motion. Since it is assumed that the fins are rigidly fixed and cantilevered to
an infinitely stiff rocket body, the fin twist (torsion) and fin plunge (bending) are the only two
degrees of freedom.
Once this flutter velocity is exceeded, the air, inversely, amplifies the oscillations and
significantly increases the energy within the respective fin. As velocity increases, the fin twist
and plunge are no longer damped. At this velocity, known as the divergent speed, one degree of
freedom usually diverges while the other remains neutral. Structural failure usually occurs at or
just above this velocity. Due to certain failure of the structure associated with potential aero-
elastic flutter, the flutter velocity is applied to the design as a “never-to-exceed” parameter.
There are various ways to minimize the chances of experiencing fin flutter. One of the wasy to
minimize include increasing fin retention by strengthening the joints between the fins and rocket
body is one way to supplement system stability. Furthermore, giving the fins an internal
aluminum honeycomb core increases the fins’ stiffness, preventing it from fluttering and
retaining rigidity.
Section 3.3: Design Integrity Section 3.3.1: Fin Shape and Style
The three most common planforms are clipped delta, trapezoidal and elliptical, as shown in
Figure 3.22. During subsonic flight, the differences in drag characteristics of the planforms are
negligible at this scale. The clipped delta offers a slight stability advantage over the trapezoidal
fins due to having more surface area aft of the chord of the fins midpoint. This extra surface area
provides increased induced drag, allowing for more rapid and effective course correction. The
elliptical planform can create manufacturing difficulties due to its complex shape that are not
present in the manufacturing methods of a clipped delta planform. Elliptical fins also provide
diminished surface area to counteract course change. Therefore, a clipped delta fin shape was
chosen.
27
Figure 3.7: Fin Shapes.
To verify that the size and shape of the fins allows for stable flight, simulations were conducted.
There has also been two subscale flights which further verified the simulation data. Multiple full
scale test flights will be performed to visually verify no anomalies are present on the fins during
flight.
Section 3.3.2: Materials in Fins, Bulkheads and Structural Elements
Body Tubes:
The structural tubes of the launch vehicle are going to be constructed using a 3D braided carbon
fiber isogrid structure. As this is something the team has not done in past years, structural data
will need to be collected for this structure. To do this, using the same material and manufacturing
method, a test sample will be made consisting of an equal diameter of the tubes that will be used
on the launch vehicle. This sample will then be placed into a load cell to determine the maximum
load of the structure. This will allow us to determine that the structure is capable of safely
completing the mission. The structure will experience a maximum of 175 lbs during flight, to
meet the factor of safety requirements the tube structure must fail at or above 350 lbs of force
during testing. The team is also creating Patran models to perform finite element analysis of the
forces along the body tube, as shown in Figure 3.23.
28
Figure 3.8: Patran Tube Model
Bulkheads and Centering Rings:
The bulkheads and centering rings are manufactured by cutting a flat carbon fiber plate with a
CNC machine. To verify these components are able to handle the expected loads, sample pieces
of the carbon fiber have been made. These samples were manufactured using the same material
that the bulkheads and centering rings were 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 coupler separation were performed.
Once proven on the ground, a subscale flight test using this coupler component was also
performed successfully.
Ballast Tank:
By running simulations, the team is able to determine where the center of gravity is located. Now
that the launch vehicle has been manufactured a final simulation was run using real component
weights. It was found that the new, reinforced avionics bay held had enough mass already, and
that the use of the ballast tank was unnecessary. Throughout the project, this will be re-
examined to ensure stable flight.
29
Section 3.4: Drawings and Schematics
Figure 3.9: Upper Section Dimensions
30
Figure 3.10: Lower Section Dimensions
31
Figure 3.11: Booster Tube
32
Figure 3.12: Upper Body Tube
33
Figure 3.13: Fins
34
Figure 3.14: Filament Wound Body Tube
35
Figure 3.15: Bulkhead
36
Figure 3.16: Motor Tube Bulkhead
37
Section 3.5: Testing Section 3.5.1: Materials Testing
In order to ensure that the composite material used in the rocket body is capable of handling the
stresses involved in the launch, the material 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 composite the team will be using must be
determined via testing.
A universal testing machine in the Auburn University Aerospace Department was used to
determine the material properties of the composite material. A standard in materials testing, the
universal testing machine can test both the tensile and compressive properties of a material
through a variety of methods. Several specimens were produced for use with the universal tester.
The specimens were placed under great tensile loading in the universal tester, with the load
increasing slowly until the specimen fractured. By comparing the force loaded onto the specimen
to the elongation of the specimen prior to fracture, a stress-strain relationship was plotted and the
tensile properties of the material determined. The compressive properties were determined using
a similar method, utilizing an increasing compressive load upon the specimen.
The first test completed was a three-point bending test, which was completed on October 22,
2015. The test was done to address the infill of the 3D print and to determine how many layers of
carbon fiber would be required to handle the load with an appropriate safety factor during flight.
The results of the test have shown that during the plastic stage of stress, the infill had little effect
on the results for the 3D print. However, the infill did have a noticeable effect on the maximum
load recorded, as the solid infill recorded an average maximum load of 32.575 lb, while the 50%
infill had an average maximum load of only 28.300 lb. The solid infill test pieces had an average
weight of 0.0144 lb, while the 50% infill had an average weight of 0.0117 lb. This meant that the
23.1% increase in weight caused by increasing the infill from 50% to a full 100% was
responsible for only a 15.1% increase in performance.
The carbon fiber samples showed a much more drastic improvement in strength with additional
layers, as shown in the following figure. The carbon fiber samples were all 3 in long by .5 in
wide with a variable thickness depending on how many layers were used to create the sample.
38
On average, the 6 layer samples of carbon fiber weighed 0.03128 lb, while the 10 layer samples
weighed 0.03467 lb. The 6 layer samples recorded an average yield force of 144.2 lb, while the
10 layer samples recorded an average yield force of 295.3 lb. By increasing the number of layers
of carbon fiber from 6 to 10, a 104.8% increase in performance was recorded, at the expense of
only a 10.8% increase in weight.
Figure 3.17: Carbon Fiber Test Data
Figure 3.18: Carbon Fiber Test Results
39
Figure 3.19: HIPs Data
Figure 3.20: HIPS Test Results
Figure 3.21: Three Point Bending Test
40
Section 3.5.2: Wind Tunnel Testing
Wind tunnel tests have been conducted to better understand the aerodynamics of the launch
vehicles unique shape. The team is unable to simulate the grid fins effects on the rockets flight
through our available software. To account for this, the team constructed a one fifth model that
was placed in a sub sonic wind tunnel at Auburn University. From this, the team gathered
significant data on the aerodynamic effects of the grid fins. The data in Figure 3.20 will be
compared with CFD analysis to verify the accuracy of simulations. The wind tunnel test model
can be seen in Figure 3.21.
Figure 3.22: FPS vs Lb Force
020406080
100120140160180
0 0.2 0.4 0.6 0.8 1 1.2
Feet
Per
Sec
ond
Lb Force Drag
Figure 3.23: Wind Tunnel Test
41
Section 3.6: Workmanship The Auburn University Student Launch Team is confident in the design of the launch vehicle.
Through several iterations and months of planning, the team has developed a rocket capable of
achieving a successful mission.
Every component of the rocket has been examined to ensure the best possible performance.
Every structural material has been tested for strength to make sure all components are capable of
handling the expected loads.
The rockets flight has been simulated on a simulation software OpenRocket. To verify the results
of the simulation, wind tunnel testing.
The Auburn University Student Launch team (AUSL) strives for success by minimizing risk
through proactive means. AUSL is determined to design and manufacture a uniquely effective
launch vehicle to achieve our goals. AUSL has used former launch vehicle data, design faults
and failures as examples to anticipate and mitigate any potential failures with construction of this
year’s launch vehicle.
Therefore, the fabrication and workmanship of the launch vehicle, and payload bay are overseen
by the engineering faculty advisors Professor Eldon Triggs and Professor Joe Majdalani, as well
as the graduate technical advisors, Benjamin Bauldree and Mariel Shumate. To ensure that our
workmanship is of top tier for each category, all assembly tasks are initially identified, inspected
and analyzed before any process of fabrication begins. This aspect of the team can be noticed in
the design of the launch vehicle grid fins. The testing of the launch vehicle, the payload bay, and
all components also helps to reduce the possibility of unforeseen failures or problems that may
arise on competition day.
The team’s belief is that extreme care and precision to detail be taken at each step of the design,
fabrication and testing processes in order to achieve a superior mission success. If any team
member has any question or doubt about any part of the process, an answer is sought from either
the team’s faculty advisors, safety advisors or any other reliable source before any proceeding of
activities.
42
Section 3.7: Mass Report The mass of the launch vehicle was determined by weighing each individual component before
final assembly. Table 3.4: Final Mass
Section Mass (lb) Percentage
Structure 10.8 34.5%
Recovery 4.51 14.4%
Grid Fins 3.00 9.58%
Electronics 1.52 4.85%
Motor 7.90 25.24%
Ballast 5.00 15.97
Total 31.3 100%
Section 3.8: Requirement Verification Vehicle:
1. The vehicle must maintain stability of 2 or more calibers.
2. The vehicle must have a factor of safety of at least 2.
3. Structural components must remain attached to launch vehicle.
Grid Fins:
1. Grid Fin payload is self-contained within a separate segment of the rocket.
2. Aerodynamic fairing is firmly adhered to the gird fin segment.
3. Bulkheads sealing the ends of the segment are stationary throughout flight
4. Grid fins must stay deployed during the decent phase of the trajectory.
5. Grid fins must stow away at touch down.
Fairing:
43
1. The deployment charge shall induce separation without harming the structural integrity of
the PLF.
2. The deployment charge shall not harm the recovery payload contained within the PLF.
3. The retaining clips shall break into no more the 2 individual pieces.
The team has developed a set of requirements that covers all points addressed in the 2015-2016
NASA Student Launch Handbook as well as requirements set forth by the team leadership to
ensure a unique and successful product. Table 3.6 outlines all requirements and how the team
plans to address them. Table 3.5: Verification Plan
Team
Requirement
NASA
Requirement
Section/Number
Requirement
Statement
Verification
Method
Execution of
Method
AU – 1 Vehicle 1.1
The vehicle shall
deliver the payload to
an apogee altitude of
5,280 feet above
ground level (AGL).
Analysis
Demonstration
Testing
Launch vehicle
and check
altimeters
AU – 2 Vehicle 1.2
The vehicle shall
carry one
commercially
available, barometric
altimeter for
recording the official
altitude used in the
competition scoring.
Inspection
Demonstration
Purchase and
calibrate one
commercially
available
altimeter
44
AU – 3 Vehicle 1.2.1
The official scoring
altimeter shall report
the official
competition altitude
via a series of beeps
to be checked after
the competition
flight.
Inspection
Testing
Test the
altimeter to
verify it creates
audible beeps
AU – 4 Vehicle 1.2.2
Teams may have
additional altimeters
to control vehicle
electronics and
payload
experiment(s).
Demonstration
The team may
use additional
altimeters.
AU – 5 Vehicle 1.2.2.1
At the Launch
Readiness Review, a
NASA official will
mark the altimeter
that will be used for
the official scoring
Inspection
Demonstration
Complete
safety check at
LRR
AU – 6 Vehicle 1.2.2.2
At the launch field, a
NASA official will
obtain the altitude by
listening to the
audible beeps
reported by the
official competition,
marked altimeter.
Inspection
Demonstration
Ensure beeps
are audible,
launch
successfully
45
AU – 7 Vehicle 1.2.2.3
At the launch field, to
aid in determination
of the vehicle’s
apogee, all audible
electronics, except for
the official altitude-
determining altimeter
shall be capable of
being turned off.
Inspection
Demonstration
Testing
Ensure all
electronics can
be turned off
and back on
AU – 8 Vehicle 1.2.3.1
The official, marked
altimeter will not be
damaged
Inspection
Analysis
Testing
Design the
electronics
housing to
prevent
damage to
altimeter
AU – 9 Vehicle 1.2.3.2
The team will report
to the NASA official
designated to record
the altitude with their
official, marked
altimeter on the day
of the launch.
Demonstration
The team is
timely and
organized in
gathering data
and reporting
to NASA
official
AU – 10 Vehicle 1.2.3.3
The altimeter will not
report an apogee
altitude over 5,600
feet AGL.
Demonstration
Testing
Design and test
launch vehicle
to meet altitude
requirement
46
AU – 11 Vehicle 1.2.3.4
The rocket will be
flown at the
competition launch
site.
Demonstration
Team will
launch the
rocket at the
appropriate site
on launch day
AU – 12 Vehicle 1.3
The launch vehicle
shall be designed to
be recoverable and
reusable. Reusable is
defined as being
able to launch again
on the same day
without repairs or
modifications
Testing
Analysis
Demonstration
Inspection
Trajectory
simulations
and testing will
ensure the
launch vehicle
is recoverable
and reusable
AU – 13 Vehicle 1.4
The launch vehicle
shall have a
maximum of four (4)
independent sections.
An independent
section is defined as a
section that is either
tethered to the main
vehicle or is
recovered separately
from the main vehicle
using its own
parachute
Demonstration
Team will
design and
build launch
vehicle that
can have, but
does not
require, four
independent
sections
47
AU – 14 Vehicle 1.5
The launch vehicle
shall be limited to a
single stage
Demonstration
Team will
design and
build a single-
stage launch
vehicle
AU – 15 Vehicle 1.6
The launch vehicle
shall be capable of
being prepared for
flight at the launch
site within 2 hours,
from the time the
Federal Aviation
Administration flight
waiver opens.
Demonstration
Team will be
timely and
organized to
ensure vehicle
is prepared on
time
AU – 16 Vehicle 1.7
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.
Testing
Batteries shall
be tested with
full electronics
to verify their
life
48
AU – 17 Vehicle 1.8
The launch vehicle
shall be capable of
being launched by a
standard 12 volt
direct current firing
system. The firing
system will be
provided by the
NASA-designated
Range Services
Provider
Demonstration
Vehicle will be
designed and
tested to be
launched by
the standard 12
volt DC system
AU – 18 Vehicle 1.9
The launch vehicle
shall use a
commercially
available solid motor
propulsion system
using ammonium
perchlorate composite
propellant (APCP)
which is approved
and certified by the
National Association
of Rocketry (NAR),
Tripoli Rocketry
Association (TRA),
and/or the Canadian
Association of
Rocketry (CAR).
Demonstration
Vehicle will be
designed
around
commercially
available,
certified
motors
49
AU – 19 Vehicle 1.9.1
Final motor choices
must be made by the
Critical Design
Review (CDR).
Demonstration
CDR will
determine
which motor
the team will
use for
competition
AU – 20 Vehicle 1.9.2
Any motor changes
after CDR must be
approved by the
NASA Range Safety
Officer (RSO), and
will only be approved
if the change is for
the sole purpose of
increasing the safety
margin.
Demonstration
If the change is
made to
increase safety
margin, NASA
RSO will allow
the change
AU – 21 Vehicle 1.10
The total impulse
provided by a launch
vehicle shall not
exceed 5,120
Newton-seconds (L-
class).
Demonstration
Launch vehicle
impulse will be
designed to not
exceed 5,120
Newton-
seconds.
AU – 22 Vehicle 1.11
Pressure vessels on
the vehicle shall be
approved by the RSO
Analysis
Testing
Inspection of
pressure vessel
by RSO
standards by
testing.
50
AU – 23 Vehicle 1.11.1
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
Testing of the
low-cycle
fatigue.
AU – 24 Vehicle 1.11.2
Each pressure vessel
shall include a
pressure relief valve
that sees the full
pressure of the tank.
Inspection
Analysis
Testing
Inspection of
each pressure
vessel and
testing of the
pressure relief
valve to see
does it work as
inspected.
AU – 25 Vehicle 1.11.3
Full pedigree of the
tank shall be
described, including
the application for
which the tank was
designed, and the
history of the tank,
including the number
of pressure cycles put
on the tank, by
whom, and when.
Inspection
Demonstration
The team will
inspect the
tank along with
documentation
of testing and
history.
51
AU – 26 Vehicle 1.12
All teams shall
successfully launch
and recover a
subscale model of
their full-scale rocket
prior to CDR. The
subscale model
should resemble and
perform as similarly
as possible to the full-
scale model,
however, the full-
scale shall not be
used as the subscale
model.
Demonstration
Testing
A subscale and
full scale
launch will be
completed.
AU – 27 Vehicle 1.13
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 to be flown on
launch day.
Testing
Demonstration
Testing
A test of the
rocket will be
exhibit
demonstration
all hardware
functions
properly.
52
AU – 28 Vehicle 1.13.1
The vehicle and
recovery system shall
have functioned as
designed.
Testing
Testing of
vehicle will
show how
recovery
system
functions.
AU – 29 Vehicle 1.13.2.1
If the payload is not
flown, mass
simulators shall be
used to simulate the
payload mass.
Inspection
Demonstration
Payload will be
flown.
AU – 30 Vehicle 1.13.2.2
The mass simulators
shall be located in the
same approximate
location on the rocket
as the missing
payload mass.
Inspection
Inspection of
the rocket
payload will be
done by the
team to ensure
it is properly
placed.
53
AU – 31 Vehicle 1.13.2.3
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
active during the full-
scale demonstration
flight
Demonstration
Testing
Demonstration
of the
adaptability of
the systems
notice to
payload
changes of the
external
surfaces
through
testing.
54
AU – 32 1.13.3
The full-scale motor
does not have to be
flown during the full-
scale test flight.
However, it is
recommended that the
full-scale motor be
used to demonstrate
full flight readiness
and altitude
verification. If the
full-scale motor is not
flown during the full-
scale flight, it is
desired that the motor
simulate, as closely as
possible, the
predicted maximum
velocity and
maximum
acceleration of the
competition flight.
Inspection
Demonstration
Inspection of
the motor will
be done by the
team to ensure
it is flown
through full-
scale testing.
55
AU – 33 Vehicle 1.13.4
The vehicle shall be
flown in its fully
ballasted
configuration during
the full-scale test
flight. Fully ballasted
refers to the same
amount of ballast that
will be flown during
the competition
flight.
Testing
Demonstration
Testing of the
vehicle will
demonstrate it
being fully
ballasted.
AU – 34 Vehicle 1.13.5
After successfully
completing the full-
scale demonstration
flight, the launch
vehicle or any of its
components shall not
be modified without
the concurrence of
the NASA Range
Safety Officer (RSO).
Demonstration
The team will
demonstrate
that it did not
alter any
components or
vehicle after
demonstration
flight.
AU – 35 Vehicle 1.14
Each team will have a
maximum budget of
$7,500 they may
spend on the rocket
and its payload(s).
Demonstration
The team will
demonstrate its
budget of the
competition
rocket to
validate its
cost.
56
AU – 36 Vehicle 1.15.1
The launch vehicle
shall not utilize
forward canards.
Demonstration
The team will
demonstrate
how the launch
vehicle does
not utilize
canards.
AU – 37 Vehicle 1.15.2
The launch vehicle
shall not utilize
forward firing
motors.
Demonstration
A
demonstration
of the launch
vehicle will
demonstrate it
not utilizing
forward firing
motors.
AU – 38 Vehicle 1.15.3
The launch vehicle
shall not utilize
motors that expel
titanium sponges
(Sparky, Skidmark,
MetalStorm, etc.)
Demonstration
The team will
demonstrate
that the motor
does not expel
titanium
sponges
through test
flight.
AU – 39 Vehicle 1.15.4
The launch vehicle
shall not utilize
hybrid motors.
Demonstration
The team will
exhibit how the
launch vehicle
does not utilize
hybrid motors.
57
AU – 40 Vehicle 1.15.5
The launch vehicle
shall not utilize a
cluster of motors.
Demonstration
The team will
demonstrate
and inspect the
launch vehicle
to validate it
does not use a
cluster of
motors.
• To ensure compliance with requirement AU-1, the vehicle will have a test launch with the
goal of attaining the 5280 ft apogee requirement of the competition. After the launch, the
altimeter will be checked; should the vehicle fail to adhere to the requirement,
modifications to the design will be made to correct any issues and the vehicle will be
retested.
• To ensure compliance with requirement AU-3, the altimeter will be checked after a test
launch of the vehicle to ensure that the altimeter reports the altitude reached via a series of
beeps.
• To ensure compliance with requirement AU-7, the switch that controls the vehicle's
electronics shall be activated and deactivated to ensure that the electronics properly turn
on and off on command.
• To ensure compliance with requirement AU-8, the altimeter shall be checked for damage
after each test launch of the vehicle. Should any damage occur to the altimeter, the housing
for the altimeter will be modified to ensure the altimeter will survive future flights, and the
vehicle will undergo an additional test flight.
• To ensure compliance with requirement AU-10, the vehicle's altitude will be monitored
during test launches. If the vehicle exceeds 5,600 ft. AGL during test flight, steps will be
taken as necessary to bring the vehicle's flight back into the acceptable altitude range. This
may include adding/removing ballast weight, choosing a different engine, or similar
measures.
• To ensure compliance with requirement AU-12, the vehicle will undergo a test launch, and
must be recovered intact and in a reusable condition. If the vehicle is not
58
recoverable/reusable after this test launch, design changes will be made as necessary to
ensure future iterations meet the requirement.
• To ensure compliance with requirement AU-16, the vehicle will be placed on its launch
pad in launch-ready configuration for at least one hour as a test of the electronic system's
battery life.
• To ensure compliance with requirement AU-22, any pressure vessels on the launch vehicle
will have to meet the RSO's standards through standard testing.
• To ensure compliance with requirement AU-23, any pressure vessels on the launch vehicle
will be put through testing to ensure that they meet a minimum factor of safety of four. The
results of these tests will be well documented and presented during milestone reviews.
• To ensure compliance with requirement AU-24, any pressure vessels must have solenoid
pressure relief valves; these valves must be tested to ensure they function as intended.
• To ensure compliance with requirement AU-26, a subscale model of the launch vehicle
shall be built and launched before CDR. This model will be a separate vehicle from the
actual launch vehicle, and will be designed to be as close to the actual launch vehicle in
performance as possible.
• To ensure compliance with requirement AU-27, the final version of the launch vehicle will
be completed before FRR, and will go through at least one full, successful launch to
demonstrate the vehicle's adherence to general competition requirements.
• To ensure compliance with requirement AU-28, the recovery systems shall be fully
demonstrated during the test flight listed under AU-27.
• To ensure compliance with requirement AU-32, if the payload changes the external surface
of final vehicle design or alters the total energy of the vehicle, then those systems will be
active during the test under AU-27.
• To ensure compliance with requirement AU-33, the vehicle must be fully ballasted during
the full-scale test under AU-27.
Due to multiple unforeseen failures, including two motor-related catastrophe at take-off events,
acquiring data to assemble an accurate prediction of the full-scale rocket performance with
working payloads has become a challenge. The challenges faced along with the data that will be
used to predict the performance of the full-scale rocket are featured throughout this section.
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The 3D carbon fiber braided rocket isogrid body tube takes much longer to construct than a solid
carbon fiber body tube. Thus, only one isogrid structure was able to be constructed over the
course of the project's life cycle. Actual data from the test launch with the isogrid structure was
unable to be acquired due to a motor failure that only allowed the fully-assembled rocket to
ascend two feet off the launch rail.
Prior to this aforementioned Cato, multiple full-scale models with functioning payloads and solid
carbon fiber body tubes rather than carbon fiber braided tubes, were launched. The first of these
test launches also experienced a motor failure, disallowing the team to become aware of any
malfunctions within the payload systems. During the second of these launches, a malfunction
within the fairing payload caused the drogue parachute to deploy maturely and wrap into the
motor burn. This failure caused the rocket to reach an altitude of approximately only 1,040 feet
and also prevented the main chutes from ever deploying.
Due to the failures outlined above, data for the launch of a full-scale rendering of the rocket with
isogrid body structure and working payloads has thus far been unattainable. However, data from
a successful launch of the full-scale rocket without working payloads, along with data from
successful subscale launches and component testing are available and will serve as parameters of
performance predictions for this project.
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Section 4: Recovery
Section 4.1: Recovery System Overview The Auburn Student Launch team is using a modified dual-stage recovery system with a drogue
parachute deployed at apogee (target height of 5280 ft.) and two main parachutes deployed at
1000 ft. At apogee the fairings split and the drogue parachute is deployed, along with a main
parachute in a bag using the Tinder Rocketry Tender Descender dual deploy system. At the
second event (at 1000 ft.) two main parachutes are deployed. The upper section, containing the
payload fairing, avionics bay, and ballast falls under the main parachute deployed using the
Tender Descender system. The booster section, including the aerodynamic analysis payload,
separates from the upper section using a proprietary CO2 ejection system and falls under a
second main parachute. The rocket is recovered in two independent sections. The parachute
deployment is shown in Figure 4.1.
Figure 4.1: Parachute Configuration
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Section 4.2: Structural Elements The centerpiece of the Auburn's recovery system is the Barometric Avionics Enclosure (BAE). Every
recovery subsystem (with the exception of parachutes) is either attached to or contained inside
the BAE. The BAE is constructed out of PVC pipe 13 inches long coated in several layers of
pre-impregnated fiberglass composite cloth. The weight of these layers of fiberglass and the
resin binding them together also serves as ballast. This weighs down the top half of our rocket
and moving the center of gravity up the airframe and increasing the rocket's stability. There is
one inner bulk plate attached 4 inches from the top of BAE that serves as the top cap of the
avionics bay. Inside the avionics bay are two sets of rails that secure the avionics board, to
which all recovery electronics are mounted. The bottom of BAE is closed off by another bulk
plate. The two bulk plates are linked together by two ¼ inch rods and secured with ¼ inch
locking nuts. Both of these bulk plates have holes that allow the ejection charge wires to run
from the altimeters to their proper e-matches. The BAE serves as the coupler between the upper
section and the lower section, and each section is secured with machine bolts. Neither of these
sections separate once the rocket has been assembled. On the outside of the BAE is a 2 inch long
ring of Isogrid tube. This ring is taken from the same tube the upper section is constructed. This
is done so the tube connections between the upper section, the BAE, and the lower section is
continuous and smooth, minimizing this connection's 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 on
this ring's surface is where key switches, patch antennae, and pressure holes are
located. Sketches of the BAE can be seen below.
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Figure 4.2: BAE Bottom View
Figure 4.3: BAE Cutaway View
63
Figure 4.4: BAE Side View
Section 4.3: Electrical Elements The BAE houses two altimeters to satisfy redundant system requirements. Both altimeters fire a
charge at apogee (target altitude: 5280 feet) to deploy the fairings and thus the drogue parachute.
Then both altimeters fire the main deployment charges at an altitude of 1000 ft.
The team is using an Altus Metrum TeleMega as the primary and one Altus Metrum TeleMetrum
as the secondary altimeters. Both TeleMetrum altimeters gather flight data via a barometric
pressure sensor and an onboard accelerometer. The TeleMega has an advanced accelerometer for
more detailed flight data acquisition. Additionally, using two Altus Metrum altimeters makes
programming quicker and easier, as they share an interface program. This makes any last minute
or on site adjustments across both boards simpler. Altus Metrum 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. Should one of the Altus Metrum altimeters
encounter unforseen problems, a PerfectFlite Stratologger will be used as additional backup.
64
Figure 4.6: AltusMetrum Telemetrum
Another reason the Altus Metrum altimeters are preferred are their radio frequency (RF)
communication capabilities. Both TeleMega and TeleMetrum are capable of communicating
with a Yagi-Uda antenna operated by the team at a safe distance 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 a more acceptable level. Additionally, the apogee delay, main deploy height, and other
pyro events can be configured to almost any custom configuration. The altimeter can even be
rebooted remotely. 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 reports 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. Both Altus Metrum altimeters
transmit on one of ten channels with frequencies ranging from 434.550 MHz to 435.450 MHz.
Figure 4.5: Altus Metrum Telemega
65
In past years, this radio frequency communication has caused trouble due to signal strength.
Communication could intermittently be established with the rocket while on the ground, and
settings could be configured. Once launched however, connection with the on-board altimeters
was soon lost due to weak signal strength. This is likely due to several causes such as the antenna
not being straight inside the rocket, the conductive carbon fiber body blocking the signal, or low
power output of the altimeter’s whip antenna. To prevent these issues, the team replaced the
altimeters' default antennae with new antennae.
The Altus Metrum altimeters can have their whip antennas replaced with any antenna desired, so
an SMA cable was connected to the board and run to the outside of the rocket. On the outside the
team attached a flexible patch-antenna. The Taoglas FXP240 433 MHz ISM Antenna was the
team's selection and can be seen in Figure 5.13. The advantage of this antenna is it conforms to
the shape of the rocket to have a negligible effect on the aerodynamics of the rocket. Since the
antenna is on the outside of the rocket, the signal is no longer being attenuated by passing
through the carbon fiber body of the rocket and increases connectivity.
Another benefit of removing the antenna from the interior of the avionics bay is the reduced high
power radiated emissions near the altimeters. Due to their delicate sensors, small amounts of
interference can greatly distort measured data from the altimeters. Isolating one altimeter system
(altimeter, battery, and wires) from the other helps prevent any form of coupling or cross-talk of
signals.
Figure 4.7: Taoglas FXP240 433 MHz ISM Antenna
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Isolation is realized via distancing the two systems, avoiding parallel wires, and twisting wires
within the same circuit. Additionally, the most apparent form of radio-frequency interference, the
antenna, will resonate on wires any multiple of ¼ λ (1/4 of ~70cm). Avoiding resonant lengths of
wire was done wherever possible. Within the BAE, the altimeters and batteries are mounted on
opposing sides of the carbon fiber avionics board, with one battery and altimeter per side. Since
carbon fiber is an effective shielding material (50dB attenuation), this board acts as shielding
between the two altimeters and minimizes 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 4.4: Parachutes Auburn’s modified dual deploy recovery approach makes use of three separate parachutes, each
designed and constructed in house by the AUSL team. The team has been making its own
parachutes for three 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.
Section 4.4.1: Parachute Sizing
The drogue parachute is a small, circular parachute constructed of rip-stop nylon with paracord
shroud lines. At apogee, the drogue is deployed from the top of the rocket, out of the payload
fairing. This stabilizes descent until main deployment. A drogue parachute size can be estimated
by the following calculation based on the length and diameter of the rocket body.
4 TUBE TUBEDROGUE
L ddπ
⋅ ⋅=
The team’s rocket has a length of 75.125 in and a diameter of 5.25 in:
𝑑𝑑𝐷𝐷𝑅𝑅𝐷𝐷𝐺𝐺𝐷𝐷𝐷𝐷 = �4 ∙ 75.125 𝑖𝑖𝑖𝑖 ∙ 5.25 𝑖𝑖𝑖𝑖
𝜋𝜋 = 22.4 𝑖𝑖𝑖𝑖
The recovery system involves two main parachutes. Each main parachute is constructed of rip-
stop nylon with 0.5 inch tubular Kevlar shroud lines.
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Both main parachutes are hemispherical. The shape of the main parachutes and their gores can be
seen in Figure 5.3 and Figure 5.4. When the booster section separates, a main is deployed from
the top of that section. The other main parachute deploys through the top of the rocket, following
the drogue. A spill hole was added to both main parachutes. It was to the booster section main
parachute for stability, since it is falling separately from the rest of the rocket. A spill hole was
added to the payload main parachute to accommodate the Tender Descender. This spill hole is
necessary with our configuration of dual-deploying from the same compartment at the top of the
rocket. Shock cord runs through this spill hole to keep the Tender Descender and drogue
parachute attached to the rocket after main parachute deployment. In accordance with the general
rule of thumb, the spill hole is close to 20% of the total base diameter of the chute. The 20%
diameter of the spill hole is chosen because it only reduces the area of the parachute by about
4%. This allows enough air to go through the spill hole to stabilize the booster section without
drastically altering the descent rate.
Figure 4.8: Parachute Shape Parameters
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Figure 5.4: Parachute Gore Parameters
A Pugh chart was created to determine the best choice of parachute shape. This Pugh chart is
shown in Table 5.2. Table 4.1: 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
Parachute areas for hemispherical shaped chutes are determined using the following equation:
2
2
: force: density of air
: drag coefficient: descent velocity
D
D
FAC V
F
CV
ρ
ρ
⋅=
⋅ ⋅
Example calculation for a section of rocket weighing 10 lbm at a descent rate of 16 ft/s:
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𝐴𝐴 =2 ∗ 10𝑙𝑙𝑙𝑙𝑚𝑚 ∗ 32.2 𝑓𝑓𝑡𝑡𝑠𝑠2
0.076474 lb𝑚𝑚𝑓𝑓𝑡𝑡3 ∗ 1.5 ∗ �16𝑓𝑓𝑡𝑡𝑠𝑠 �
2 = 21.9 𝑓𝑓𝑡𝑡2
Table 4.2: Main Parachute Dimensions
Booster (Bottom) Main Payload (Top) Main
Area of chute 17.27 ft2 30.17 ft2
Diameter of chute 39.84 in 52.56 in
Diameter of spill hole 7.92 in 10.56 in
Height of each gore 31.29 in 41.3 in
Width of each gore 20.94 in 27.5 in
Number of gores 6 6
Figure 4.9: Main Parachute Visualization
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The recovery team has designed deployment to ensure that the 75 ft-lb kinetic energy limit is not
reached. Since the rocket is recovered in two separate pieces, the team simply had to calculate
descent rates for each section, and then use this descent velocity to calculate kinetic energy.
2
: mass: descent
12
velocity
KE
V
V
m
m= ⋅
Example calculation for a section of rocket weighing 10 lbm at a descent rate of 16 ft/s,
2
2
101 * 162 32.2
39.75m ft lblb ftKE ft ss
= ⋅ =
⋅
Table 4.3: Kinetic Energy Calculations
Section Mass (lbm) Kinetic Energy (ft-lb)
Payload Section Recovery (2 Parachutes)
Avionics Fairing
Structure
12.783 50.81
Booster Section Recovery(1 Parachute)
Grid Fins and Electronics Motor (After Burnout)
Structure
7.292 28.99
Section 4.4.2: Manufacturing
The parachutes are manufactured at Auburn University by recovery team members. Once they
have been designed, 1:1 gore templates are made using SolidWorks. These templates include an
extra inch on each side for sewing and hemming purposes. Then, these templates are used for
cutting out all of the gores from the team's rolls of orange and blue ripstop nylon. The parachutes
this year are six gores each, and once all gores have been cut, the sewing process can begin.
First, two gores are pinned together down one side and sewn, using strong, nylon thread and a
straight stitch. Now, to ensure strength in our parachute seams, we "butterfly" sew, which means
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we invert the seam that was just sewn and sew a new seam that encases the first one. This
process is repeated until all seams are sewn and reinforced in that way. After all main seams are
sewn, the spill hole and the bottom of the cute are hemmed. After hemming, the paracord shroud
lines are added. Because of the thickness of the paracord, a zigzag stitch is used to continually go
back and forth between the chute and the paracord. This ensures that the connection is secure.
The parachutes are complete after this step and are then inspected by the team to ensure there are
no flaws in production.
Section 4.4.3: Deployment Process
The AUSL team is utilizing the Tender Descender in our recovery systems to enable us to deploy
both a drogue and main parachute simultaneously in a single separation. The Tender Descender
is shown in Figure 5.6. This system deploys more parachutes with fewer separations, reducing
the chance of failure of the recovery portion of flight.
Figure 4.10: Pictures of Tender Descender in Undeployed and Deployed Configurations
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The Tender Descender system works by attaching the drogue lines to a bag containing the main
parachute and the Tender Descender system itself, while the Tender Descender is then attached
directly to shock cord that is anchored to an U-bolt within the fairings. This allows the main
parachute to remain undeployed in its bag. Then at 1000 ft. altitude, the team's altimeters fire an
e-match, igniting a small black powder charge within the Tender Descender that separates its two
connections. This releases the attachment to the shock cord allowing the drogue lines to pull the
bag off the main parachute, thus deploying the main chute just below the drogue.
During the team’s testing of the Tender Descender system, several problems were encountered
that led to improvements on the original implementation of the device. First, the recommended
Tinder Rocketry configuration of the Tender Descender has drogue parachute and Tender
Descender separating completely from the rocket and being recovered separately. This creates
the possibility of losing the drogue parachute with each launch. To prevent this, another shock
cord through the main parachute’s spill hole was attached to the Tender Descender to keep the
drogue attached to the upper section. This prevents the loss of the drogue and allows it to
contribute a small amount of additional drag along with the main parachute. Retaining the
drogue parachute is also useful in reducing the kinetic energy of impact in the event that the
main parachute fails to deploy.
The team also chose to sew a custom bag to hold the main parachute before the Tender
Descender deploys. Made of rip-stop nylon, the bag provides needed strength while also being
incredibly light and compact. The team ran into tangling issues housing both a drogue parachute
and the main parachute in a single compartment, and the thin rip-stop nylon bag alleviated those
troubles substantially.
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The Tender Descender device itself relies on an e-match to be fired in order to separate from the
shock cord tethering it to the rocket and allow the bag to be pulled off the main parachute. The
team decided to set the device with two e-matches for redundancy. This requires several wires to
extend from the altimeter bay to the Tender Descender, which is located several feet above the
rocket while the drogue is fully deployed. To lessen the chance of entanglement or damage to the
wires, the wires are fed through a plastic casing that is then heated to shrink it. This condenses all
four separately insulated wires into a single tube and prevents tangling while the drogue is
already deployed. Additional fasteners that attach the tube of wires to the shock cord connecting
the Tender Descender prevents the tube from flailing about or being tugged on while the system
falls under drogue.
The Tender Descender L2 model that the team will use is rated to withstand a maximum of 2000
pounds of shock, 500 pounds of release weight, and 75 pounds of rocket weight. These values
are well above the expected loading conditions the device will experience during flight and
recovery of Auburn's rocket.
With these alterations and the validation of a successful subscale rocket recovery, the team is
confident in the ability of the Tender Descender system to recover our rocket safely.
Section 4.4.4: Drift
The distance the rocket will drift during descent can be estimated with the following equation.
𝐷𝐷𝐷𝐷𝑖𝑖𝑓𝑓𝑡𝑡 = 𝑊𝑊𝑖𝑖𝑖𝑖𝑑𝑑 𝑆𝑆𝑆𝑆𝑆𝑆𝑆𝑆𝑑𝑑 ∗ 𝐴𝐴𝑙𝑙𝑡𝑡𝑖𝑖𝑡𝑡𝐴𝐴𝑑𝑑𝑆𝑆 𝐶𝐶ℎ𝑎𝑎𝑖𝑖𝑎𝑎𝑆𝑆𝐷𝐷𝑆𝑆𝑠𝑠𝑐𝑐𝑆𝑆𝑖𝑖𝑡𝑡 𝑉𝑉𝑆𝑆𝑙𝑙𝑉𝑉𝑐𝑐𝑖𝑖𝑡𝑡𝑉𝑉
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 are two stages of descent. First, the rocket descends under the drogue parachute from an
altitude of 5280 ft. to 1000 ft. Then the rocket separates and both the booster section and the
payload section descend to the ground under their respective main parachutes at a velocity of 16
ft/s.
The rate of descent under drogue is calculated with the following equation:
𝐷𝐷𝑆𝑆𝑠𝑠𝑐𝑐𝑆𝑆𝑖𝑖𝑡𝑡 𝑉𝑉𝑆𝑆𝑙𝑙𝑉𝑉𝑐𝑐𝑖𝑖𝑡𝑡𝑉𝑉 = �2 ∗ 𝐹𝐹𝑉𝑉𝐷𝐷𝑐𝑐𝑆𝑆
𝐴𝐴𝑖𝑖𝐷𝐷 𝐷𝐷𝑆𝑆𝑖𝑖𝑠𝑠𝑖𝑖𝑡𝑡𝑉𝑉 ∗ 𝐷𝐷𝐷𝐷𝑎𝑎𝑎𝑎 𝐶𝐶𝑉𝑉𝑆𝑆𝑓𝑓𝑓𝑓𝑖𝑖𝑐𝑐𝑖𝑖𝑆𝑆𝑖𝑖𝑡𝑡 ∗ 𝑃𝑃𝑎𝑎𝐷𝐷𝑎𝑎𝑐𝑐ℎ𝐴𝐴𝑡𝑡𝑆𝑆 𝐴𝐴𝐷𝐷𝑆𝑆𝑎𝑎
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With a total rocket weight of 20.075 lbm after burnout and a drogue diameter of 22.11 inches
(which corresponds to an area of 2.67ft2):
𝐷𝐷𝑆𝑆𝑠𝑠𝑐𝑐𝑆𝑆𝑖𝑖𝑡𝑡 𝑉𝑉𝑆𝑆𝑙𝑙𝑉𝑉𝑐𝑐𝑖𝑖𝑡𝑡𝑉𝑉 = �2 ∗ 20.075𝑙𝑙𝑙𝑙𝑚𝑚 ∗ 32.2𝑓𝑓𝑡𝑡𝑠𝑠2
0.076474 𝑙𝑙𝑙𝑙𝑚𝑚𝑓𝑓𝑡𝑡3 ∗ 1.5 ∗ 2.67𝑓𝑓𝑡𝑡2= 64.97
𝑓𝑓𝑡𝑡𝑠𝑠
This yields a descent velocity of 64.97 ft/s under drogue. The estimated drift distances for a
variety of wind speeds are shown in Table 5.5 below.
Table 4.4: Drift Calculations
Section 4.5: Testing Section 4.5.1: Subscale Testing
The subscale rocket was 3/5 the scale of the full scale rocket, giving it a diameter of 3 inches.
This reduced diameter presented an interesting challenge to the Auburn recovery team, as our
custom CO2 recovery system is limited to use in a 5 inch diameter rocket and must be modified
or replaced to fit in the smaller rocket. The subscale used a static nosecone instead of fairings, so
the team ejected this nosecone to deploy parachutes. The subscale rocket’s recovery
configuration is illustrated in Figure 4.3.
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Figure 4.11: Subscale Recovery Configuration
A drogue parachute and main parachute were dual-deployed out of the top of the rocket by
ejecting the nose cone. This was done using the Tender Descender, which allows dual-
deployment from the same compartment. The main parachute was given a spill hole for this
configuration, to keep the Tender Descender and drogue parachute attached after main parachute
deployment. This was a 3/5 subscale, which reduced the size of parachutes needed. Our
subscale drogue parachute was 22 inches in diameter and our subscale main parachute was 34
inches in diameter.
Black powder ejection was used for the subscale flight. Three-gram charges were made using
black powder, electric matches and plastic tubing. The avionics bay board was modified to fit
into the 2.75” interior diameter of the avionics bay, and employed a different assembly of
altimeters and batteries than the full scale rocket. This setup allowed the board to be only 5.5”
long and 2.6” wide and still have all necessary electronics properly mounted to it. The board had
one battery and altimeter mounted on each side of the avionics bay board which saved space and
reduced interference between the altimeters. The team used an Altus Metrum TeleMetrum
Altimeter as the primary altimeter and a PerfectFlite MAWD as the secondary altimeter for
redundancy. Each altimeter was wired with electric matches and black powder charges. For the
ejection of the Tender Descender, electric matches from the “main” port of each altimeter were
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placed into the Tender Descender, which was filled with black powder. Because the Tender
Descender is located far from the avionics bay, the electric matches were connected via eighteen
feet of wire that were shrink-wrapped then secured along the shock cord.
The nose cone was ejected at apogee, releasing the drogue. This is the only section separation
that occurred in the subscale flight. The Tender Descender separated at 750 ft. for main
parachute deployment; this pulled the main parachute out of the bag in which it was contained.
Auburn’s subscale rocket flight was successful and the rocket was capable of being launched
again the same day. All recovery systems worked as expected, validating our use of the Tender
Descender dual deploy system.
Section 4.6: Requirement Verification The Auburn Student Launch team has developed a strategy for meeting all requirements outlined
in the 2015-2016 Student Launch Handbook. The intended method of validation for all recovery
requirements is outlined in the table below.
Table 4.5: Recovery Requirement Validation
Requirement Number Requirement Validation Method
2.1 Deployment of Recovery Devices
Ground tests of recovery system using controlled deployment. Separation was achieved using official recovery hardware and was confirmed in multiple tests.
2.2 Ground Ejection for Drogue & Main Parachute
Ground tests of recovery system using controlled deployment. Separation was achieved using official recovery hardware and was confirmed in multiple tests.
2.3 At Landing, Max KE of 75ft- lbf for each Independent Section
Calculation and subscale testing
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2.4 Recovery system Electrical Circuits Independent of Payload Electrical Circuits
The recovery electronics are completely independent of all other electronics in the rocket. Recovery electronics have independent power supplies.
2.5
Recovery System Must Contain a Redundant, Commercially Available Altimeter
The recovery system uses and Altus Metrum TeleMega as the primary altimeter and an Altus Metrum TeleMetrum altimeter as a redundant backup.
2.6 Exterior Arming Switch for each Altimeter
Each altimeter is armed with an external key switch.
2.7 Dedicated Power Supply for each Altimeter
Each altimeter has a separate battery power supply.
2.8 Arming Switch Capable of being Locked in the ON Position
Key switches lock in armed and unarmed positions and can only be changed with key.
2.9 Removable Shear Pins used for Main & Drogue Parachute Compartment
Nylon machine screws used at all rocket separation sections.
2.10
Electronic tracking Device Installed in Rocket to Transmit the Location of the Tethered Vehicle or any Independent Section to a Ground Receiver
Rocket falls in two independent sections. The upper section with the avionics bay has GPS tracking from the recovery system altimeters. The lower section has GPS tracking from the WAFLE payload. Additionally, each section has an RF tracker.
2.10.1
An Active Electronic Tracking Device shall be connected to any Independent Rocket Section or Payload Component
Each section has an RF tracker.
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2.10.2
The Electronic Tracking Device shall be fully Functional during Official Flight at Competition Launch Site
RF trackers have been ground tested in lab and confirmed through full scale flights. Altimeter tracking capabilities have also been ground tested and confirmed through flights.
2.11
Recovery System Electronics shall not be affected by other On-Board Electronics during Flight
The addition of patch antennae amplifies the tracking capabilities of the recovery electronics. Flight testing has confirmed the lack of interference between on-board electronics.
2.11.1
Recovery System Electronics must be placed in a Separate Compartment away from any other Radio Frequency/ Magnetic Wave Producing Device
Recovery system is housed in a compartment of the rocket separate from all other electronics and is blocked from other transmitters by carbon fiber bulk plates. Patch antennae ensure that signals can be transmitted out of the rocket without interference.
2.11.2 Recovery System Electronics Shielded from all On-Board Transmitting Devices
Recovery system is housed in a compartment of the rocket separate from all other electronics and is blocked from other transmitters by carbon fiber bulk plates. Patch antennae ensure that signals can be transmitted out of the rocket without interference.
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2.11.3
Producing Magnetic Waves Recovery System Electronics Shielded from any other On- Board Transmitting Devices
Recovery system is housed in a compartment of the rocket separate from all other electronics and is blocked from other transmitters by carbon fiber bulk plates. Patch antennae ensure that signals can be transmitted out of the rocket without interference.
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Section 5: Full Scale Results
The Auburn Student Launch team has attempted four full scale launches as of this report.
Unfortunately, due to a series of failures, a very limited amount of data has been collected from
these flights.
The team will be launching a fifth full scale flight on April 1, 2016 to validate the design.
Section 5.1: Project Aquila Test Launch 1 Launch Date: January 16, 2016
Launch Location: Samson, AL
This launch was a static aerodynamic test of the vehicle design. The grid fins were attached but
not activated. The fairing was replaced with a static nose cone with the same aerodynamic
properties.
The upper main parachute came out at apogee and resulted in high drift. However, this flight was
able to confirm stability on ascent.
As a result of this launch, the parachute bag was redesigned and ground tested to eliminate the
risk of premature deployment.
Section 5.2: Project Aquila Test Launch 2: Launch Date: January 30, 2016
Launch Location: Samson, AL
This launch was intended to be a fully operational test of all launch systems and payloads. Both
the WAFLE system and the payload fairing were armed. However, the motor exploded shortly
after the rocket left the rail. After evaluation by RSOs and Aerotech, this failure was determined
to be the result of a faulty motor supplied by the manufacturer.
Section 5.3: Project Aquila Test Launch 3: Launch Date: February 20, 2016
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Launch Location: Sylacauga, AL
This launch was a second attempt at a fully operational launch. All systems were armed and
active. Shortly before motor burnout the fairing prematurely split open. This caused catastrophic
damage to the entire rocket. The drogue parachute came out when the fairing split and burned up
in the motor. The rest of the rocket came down ballistic as the lack of drogue prevented the upper
main from being pulled out of the bag and the lower section was falling too fast for the bottom
main to properly deploy.
Section 5.4: Project Aquila Test Launch 4: Launch Date: March 5, 2016
Launch Location: Manchester, TN
For a third time, the Auburn team attempted to launch a fully operational full scale rocket. Again
the motor exploded and again this was determined to be the result of an error by the
manufacturer. The isogrid structure and the recovery BAE survived both the explosion and the
ballistic impact but all other systems were irreparably damaged. This failure encouraged the team
to switch motors from the Aerotech L1520T to the Loki L1482.
82
Section 6: Payload Fairing
Section 6.1: Design Overview Traditionally, a payload fairing (PLF) is used to protect a scientific payload during the launch
process. However, for Project Aquila, the PLF will house the drogue and one of the main
parachutes. Prior to deployment, the PLF will act as the aerodynamic nose cone. A low-drag
elliptical design was chosen do to the low-altitude, low speed nature of the competition. In order
to line up flush with the rocket main body, the wall thickness of the fairing was chosen to be
0.125 inches. A plot showing the curvature can be seen in Figure 7.1.1.
Figure 7.1.1: PLF Curvature
The overall height of the fairing is 13 inches. Figure 7.1.2 is an overview of the complete
assembly with current dimensions. The two fairings (Section A) are attached to the nose cone
base (Section B) via hinges (Section D). These hinges will allow the fairing to separate while
still retaining the two individual fairing halves (NASA 3.2.5.1). The overall assembly will be
mated to the rocket main body via a sleeve (Section C). This sleeve will be inserted at the top of
the main body and will be permanently affixed. The main and drogue parachutes will be placed
inside the fairings. The base and the sleeve will hold the shock chord which will be attached to a
bulk plate at the base of the sleeve.
83
Figure 7.1.2: Overall PLF Assembly
Figure 7.1.3 shows the PLF in a partially deployed configuration. The charge bay is situated at
the highest point of the PLF system. This will allow the maximum moment to be created during
charge detonation. Side A of the charge bay will contain the black powder charge. Once shut,
recovery wadding will be placed in the PLF to ensure the protection of the payload and other
vital components (PLF.REQ.2). This side of the charge bay will also be lined with fiber glass to
ensure structural integrity of the charge bay during detonation (PLF.REQ.1). Ribs (Figure D)
have also been added to aid in the overall structural integrity of the fairing halves (PLF.REQ.4).
To ensure a proper seal, a plug will be place into Side B of the charge bay.
84
Figure 7.1.3: Partially Deployed PLF
As a result of system testing, several design changes have been incorporated in order to ensure
the overall success of the design. Previously, a single inner lip (side A), which mates with a
recess on side B, was being used to prevent air from entering the PLF during flight. However,
this did not result in a proper an aerodynamic seal. To mitigate this issue, a 1 in. outer lip was
added to side A, and was tapered to fit the profile of side B. As an added measure of safety,
paraffin wax was used to act as an aerodynamic seal on all seams. These changes can be seen in
detail in Figure 7.1.4.
Figure 7.4.1: PLF Design Improvements
85
To prevent premature separation, the two sides of the PLF were previously connected by
horizontal pins. They have since been replaced by four vertical sheer brackets which can be seen
in Figure 7.1.5. These four brackets will contain vertically oriented sheer pins that will sheer
upon detonation. It has been determined that the top two sheer pins will be 25-lb sheer pins,
whereas the bottom two sheer pins will be 10-lb sheer pins. This configuration ensure that the
fairing halves will not separate prematurely (PLF REQ.4).
Figure 7.1.5: Shear Bracket
Section 6.2: Payload Fairing Materials The fairing halves, the charge bay, and the sheer brackets will be additively manufactured using
Acrylonitrile Butadiene Styrene (ABS) thermoplastic. This material was chosen because of its
toughness and ability to withstand significant impacts. ABS is also easily manipulated and
repaired after initial production. The base and the sleeve will also be additively manufactured;
however, High Impact Polystyrene (HIPS) will be utilized for production. This material performs
well when impacting and when subjected to bending.
Nylon sheer pins will be used to ensure the fairing does not separate during flight. The force
produced by the charge will snap the sheer pins and allow the fairing to separate. The metal
hinge will rotate on a small metal pin.
Section 6.3: Payload Fairing Testing Section 6.3.1: Aerodynamic Design Testing
The overall aerodynamic design of the fairing was tested at various scales. A 1:5 subscale model
(Figure 7.3.1.1) of the entire rocket was tested in a wind tunnel. This test included a static
86
version of the PLF system. From this test, drag and vibrational data as collected and evaluated.
The collected data showed that design was sound and testing could continue.
A 3:5 subscale rocket (Figure 7.3.1.2) was launched. Again, this test used a static version of the
PLF. The goal of this flight was to prove that this elliptical design would perform well in
transonic conditions. During the flight, the overall rocket appeared to be extremely stable.
Therefore, the flight was deemed a success and the aerodynamic design of the fairing was
finalized.
Section 6.3.2: Charge Chamber Strength Testing
Test article: Charge Chamber
Test Description:
The deployment of the Payload Fairing is induced by a black powder charge. The payload
fairing is set to electronically deploy at apogee. The force of the detonation will separate the
fairing and break the sheer pins. To ensure that the chare chamber would remain structurally
intact, a series of test detonations of the black powder charge in the chamber was completed
(PLF.REQ.1, PLF.REQ.2).
This test only included the charge bay portion of the Payload Fairing system. The test rig can be
seen in Figure 7.3.2.1. The goal of this test was to determine the “breaking point” of the
chamber. Various black powder charges were tested in order to determine which charge size
Figure 7.3.1.1: 1:5 Subscale
Figure 7.3.1.2: 3:5 Subscale
87
would most efficiently induce separation, while not damaging the chamber. The maximum black
powder charge was limited by the capacity of the charge chamber. Below is a table that displays
the details of each individual test.
Test Black Powder
(grams)
Open? Visible Damage
1 0.1 Yes None
2 0.2 Yes None
3 0.3 Yes None
4 0.4 Yes None
5 0.5 Yes None Table 7.3.2.1: Charge Chamber Test Results
Figure 7.3.2.1: Charge Chamber Test Rig
While testing, the team prioritized safety. Safety equipment was brought to the testing area, as
well as first aid supplies. The team was properly distanced from the charge before detonation,
and the black powder charge was safely detonated by triggering the electronic match remotely.
A fire extinguisher was also brought to the test site.
Conclusion:
Even when filled to the maximum capacity, the detonation of the gun powder did not result in
structural damage to the charge chamber. The optimal amount of gun powder was chosen to be
0.3 grams.
88
Section 6.3.3: Ground Testing
Test Articles: PLF v.1, PLF v.2, PLF v.3
Test Description:
Extensive ground testing of the PLF was performed prior to integration on to the rocket.
Throughout the testing process, several version of the PLF were developed. Table 7.3.3.1
describes each version of the PLF and the results of each individual ground test.
Test
Article
Description of Test Article Results Conclusion
PLF v.1 - 4 Horizontal 10-lb sheer
pins
- Inner seal only
- 0.3 grams of black powder
Successful separation,
no structural damage
Ready for integration on
to the rocket
PLF v.2 - 2 10-lb vertical sheer pins
- Inner seal
- 0.5-in outer seal
- 0.3 grams of black powder
Successful separation,
broken sheer pin
bracket
Bracket not properly
secured
PLF v.2 - 2 10-lb vertical sheer pins
- Inner seal
- 0.5-in outer seal
- 0.3 grams of black powder
Successful separation,
no structural damage
Ready for integration on
to the rocket
PLF v.3 - 2 10-lb vertical sheer pins
- 2 25-lb vertical sheet pins
- inner seal
- 1.0-in outer seal
- Paraffin wax seal on all
seams
- 0.4 grams of black powder
Successful separation,
no structural damage
Ready for integration on
to the rocket
Table 7.3.3.1: Ground Test Results
89
Figure 7.3.3.1: Still from ground test of PLF v.1
While performing ground tests, the team prioritized safety. Safety equipment was brought to the
testing area, as well as first aid supplies. The team was properly distanced from the charge
before detonation, and the black powder charge was safely detonated by triggering the electronic
match remotely. A fire extinguisher was also brought to the test site.
Section 6.3.4: Water Seal Test
Test Article: PLF v.3
Test Description:
Due to the PLF v.2 failure on the Aquila III Launch (discussed in next section), it was concluded
that the PLF system must be completely air tight. For PLF v.3, the outer seal was extended to
1.0-in and the seams were sealed with paraffin wax. Prior to ground testing, the entire PLF was
filled with water and checked for leaks.
Conclusion:
No leaks were observed. Since the PLF is water tight, it can reasonably be assumed that the PLF
is also air tight.
Section 6.3.5: Full Scale Testing
Test Articles: PLF v.1, PLF v.2, PLF v.3, Static Nose Cone
Test Description:
After ground testing was complete, a series of full scale tests were attempted. Table 7.3.5.1
summarizes the results and conclusion of each full-scale test.
90
Launch Test Article Description of Test Article Results Conclusion
Aquila I Static Nose
Cone
Aerodynamically similar full
scale nose cone constructed
from hard foam and fiber glass
The rocket remained
stable throughout the
flight
The aerodynamic
design of the PLF
performs well in
transonic conditions.
Aquila II PLF v.1 - 4 Horizontal 10-lb sheer pins
- Inner seal only
- 0.3 grams of black powder
Motor CATO, no
useable data
None
Aquila
III
PLF v.2 - 2 10-lb vertical sheer pins
- Inner seal
- 0.5-in outer seal
- 0.3 grams of black powder
PLF deployed
prematurely at Mach
0.6.
Air broke through the
outer/inner seals at the
stagnation point forcing
the fairings to deploy.
Aquila
IV
PLF v.3 - 2 10-lb vertical sheer pins
- 2 25-lb vertical sheet pins
- inner seal
- 1.0-in outer seal
- Paraffin wax seal on all
seams
- 0.4 grams of black powder
Motor CATO, no
useable data
None
Table 7.3.5.1: Full Scale Test Results
Section 6.4: Payload Fairing Requirements NASA
3.2.5.1
The fairings and payload must be tethered to
the main body to prevent small objects from
getting lost in the field.
Each half of the PLF will
retained to the main body of
the rocket via hinges.
PLF.REQ.1 The deployment charge shall induce
separation without harming the structural
integrity of the PLF.
Extensive testing will be done
to determine the optimal
charge size. The charge bay
will be lined with carbon
fiber to maintain structural
integrity.
91
PLF.REQ.2 The deployment charge shall not harm the
recovery payload contained within the PLF.
Recovery wadding will be
placed in the PLF to protect
payloads and rocket
components
PLF.REQ.3 The entire PLF system shall remain
structurally intact during the following phases:
launch, separation, drift, and landing.
Ribs have been integrated
into each half of the PLF to
prevent flexing.
PLF.REQ.4 Premature separation of the PLF system shall
not occur.
A lip has been integrated to
prevent air from entering into
the system during flight. Pin
connectors will hold the side
92
Section 7: Aerodynamic Analysis Payload - WAFLE
Section 7.1: Experiment Concept Grid fins are an exciting new concept within the aerospace world. Grid fins were first used as
aerodynamic control surfaces in 1970 in Russia. The Russian grid fin was used as an
aerodynamic stabilizer for missiles on fighter jets and a stabilizer on their launch vehicle ejection
pods. The United States replicated the Russian design on a few bombs and missiles. Only a few
companies have implemented grid fins as drag control surfaces and most information regarding
these applications is proprietary. Therefore, there is little data available on the characteristics of
grid fins and how they integrate with other systems on a high-powered rocket.
The intent of this payload is to answer the questions still held about grid fins and how they react
in flight on a rocket.
Section 7.2: Science Value Section 7.2.1: Payload Objectives
The overall objective for the aerodynamic analysis payload is to obtain accurate aerodynamic
data for an aerodynamic protuberance. The protuberance chosen is the grid fin. Due to the scares
data for the grid fin, many tests and simulations will be performed to acquire the data.
The secondary objective is for the aerodynamic payload to provide drag to the rocket to insure
that the rocket completes Vehicle Requirement 1.1. The grid fins will deploy gradually to
increase the drag until the acceleration of the rocket reaches the desired acceleration for the
rocket to reach the mile high requirement. The Arduino will be used to command the servos that
turn the grid fin. The Arduino will be instructed by the accelerometer when to deploy the fins.
Section 7.2.2: Mission Success Criteria
Table 7.1: Aerodynamic Payload Success Criteria
Criteria
Number Criteria Method of Validation
93
AU1
All Aerodynamic data must
be validated through
analytical and experimental
testing.
A full aerodynamic analysis of the grid
fins will be conducted through
computational fluid dynamics (CFD),
subsystem wind tunnel testing, and in-
flight sensors.
AU2
Grid fins must stay stowed
until boost phase is
complete.
A delay is written into the code to delay
the actuation of the grid fins until the end
of the burn time.
AU3
Electronics must stay
stationary throughout the
flight
The electronics are mounted to the
Electronics Sled and secured between
threaded rods on the LANTERN.
AU4
Servos must remain in
direct contact with the
gears of the grid fins
throughout the flight.
The gear of the servos are secured to the
grid fin by means of a plastic bracket
screwed into the side of the grid fins and
incasing the gear of the servo. The servos
are actuated multiple times to insure that
the fins are secured properly.
AU5
Arduino must accurately
predict the flight path of
the vehicle.
Visual confirmation is done with the
code to insure that it is written correctly.
Changing the value of the final height to
one foot is preformed and the sensor is
lifted to the corresponding height to
achieve visual confirmation of the codes
implementation.
AU6
Grid fins must be deployed
with precision to correct
the vehicle’s trajectory.
Low tolerances between parts insures
that the parts move in sequence and in
time. The code will print the angles of
the servos during flight. Post flight
inspection will review these angles.
94
AU7
Grid fins must stay
deployed under the force
applied by the flow.
Wind tunnel testing and structure testing
will insure stationary deployment.
Section 7.3: Scientific Experiment Grid fins are a new type of control surface in the realm of aerospace, therefore there is minimal
public data on how the control surface reacts in flight to an external flow. Research was performed
and general ideas and parameters were determined to obtain a general idea of how they perform.
A design for the grid fins were decided upon, as illustrated in the previous grid fins subsystem
section.
Within this section list and describes the simulations that are planned and that have been performed
to validate the theory and researched values. Following is a chart of theses simulations and test:
Table 7.2: Aerodynamic Payload Simulation List
Simulations Intent
Computational Fluid
Dynamics (CFD)
More accurate models and data can be
obtained through this method of
investigation. This method will provide
the most accurate simulation of the flow
through and around the fin.
SolidWorks Flow
SolidWorks has a simulation tool
available to provide a visual and
approximated data for geometries within
a fluid. This program is used to provide
rough estimates of the characteristics of
the fin in flight.
95
Fortran- Flight and
Dynamic model
A Fortran simulation has been created to
model how the fin will react when
attached to an airframe under flight
conditions.
Drag Profile
A Matlab simulation was created to
provide a profile of the drag parameters
and the trajectory of the rocket. This will
allow for rough magnitudes to be
determined and assist with input data for
other simulations.
Aerodynamic Load
Testing
Wind tunnel test will be performed to
experimentally validate research and
simulation data for the forces that the fin
will experience. Different angles will be
investigated to acquire an overview of
the characteristics of the fin.
Vortex Shedding Testing
Water tunnel experiments will be
performed to investigate the vortex
shedding of the grid fin. Flow
visualization will also be performed on
the fairing and fin at different angles of
attack.
1:5 Scale Test
A 1:5 aerodynamic scale model of the
Aquila rocket and WAFLE was built and
tested in a subsonic wind tunnel.
Aerodynamic data was collected about
the aerodynamic subscale rocket and
WAFLE.
96
3:5 Scale Test
A 3:5 aerodynamic scale model of the
Aquila rocket and WAFLE was built and
launched. Data was collected and
observed about the subscale aerodynamic
model and WAFLE.
Full Scale Test
A full scale rocket with working
payloads will be built and launched. The
payload systems will be validated.
Computational Fluid Dynamics
The Computational Fluid Dynamics is a branch of fluid mechanics that uses numerical analysis
using Navier Stokes equations to solve and analyze problems that involve fluid flows. A geometry
is imported into Pointwise meshing software that allows the parameters of the flow to be defined
as well as how it interacts with the geometry. Assumptions are made about the flow. The algorithm
is implemented and beings to try to converge the Navier Stokes equations. This method is the most
accurate way to develop characters about the aerodynamic parameters of the grid fin and the rocket.
Before testing though the rocket was meshed in the program Pointwise on a Linux computer. The
process began by uploading a CAD drawing of the rocket to the program Pointwise. The parts of
the rockets were then each separated and connectors were placed on the edges of each part. The
amount of points in each connector was based on how the flow came into contact with the part.
97
Figure 7.1: Connectors and Points on a Fin in Pointwise
Once the points were properly spaced on each object the domains were then created for each piece
on the rocket. Unstructured domains were created in each section of connectors individually, this
was so the grid fins would have the appropriate holes where needed.
Figure 7.2: Domain on a Fin in Pointwise
98
Figure 7.3: Domain on a Grid Fin in Pointwise
Once the domain was created for each piece the mesh was then created for all of the objects. An
anisotropic tetrahedral extrusion method, also known as T-Rex, was then performed in order to
create unstructured boundary layer meshes on the grid fins and other pieces of the rocket. Once
done holes were created where the grid fins and fins were in order to reconnect all pieces of the
rocket. The entire mesh around the rocket was watertight. After the meshing was complete
boundaries were then created in order to run the watertight meshed rocket in fluent, which is a
powerful CFD software tool.
99
Figure 7.4: Rocket with Mesh and All Domains in Pointwise
Figure 7.5: Final Mesh and Boundaries on Rocket in Pointwise
100
SolidWorks Flow Simulation
In order to execute the testing, simulation and inspection was conducted. The team did SolidWorks
fluid flow simulation on a 3-D CAD model of a grid fin with ideal dimensions. SolidWorks Flow
Simulation is an intuitive CFD (computational fluid dynamics) tools that enables the user to
simulate liquid or gas flow in real world conditions. This program also runs “what if” scenarios
and efficiently analyzes the effects of a fluid flow. Also, the team did a visual inspection of the
flow of a grid fin within a controlled environment.
The logic behind flow simulation is to virtually see what happens aerodynamically to a grid fin
under certain flight parameters. In order to see this, the team created a 3D CAD model of an HIPS
grid fin. SolidWorks program has a fluid flow simulation that allows the user to place a virtual
model within a controlled environment.
In addition to visualizing through simulation, the need to visualize generally what happens in a
real world scenario as flow moves through the lattice on the grid fin. The best way to accomplish
this is through testing in a water tunnel where colored dyes can be added that follow the flow
through and around the grid fin and its attached fairing.
The team took measurements of the pressure created over the surface of the grid fin. The pressure
is caused by drag. The simulation allowed the team to change certain variables, such as the
dimensions of the grid fins. Also, the team was allowed to control the environment in which the
grid fin was set in. The team had control over the temperature, speed of flow, and direction of
flow.
The first test on the grid fin was the 0.1 mach. The flow of air is coming from the positive Y going
into the top face of the grid fin, meaning that the velocity of the flow is going in the negative Y
direction. Once the environment properties were set a flow trajectory was placed. The starting
point of the flow trajectory was placed an offset of 2 inches away from the face of the grid fin.
Various starting points were placed over the face to represent the start of the flow. Figure 7.6
shows an example of placing starting points of the flow over the face of the grid fin.
101
Figure 7.6: Starting points created over the top face of the grid fin to demonstrate the starting point of the flow.
After the starting points of the flow trajectory were placed, the appearance of the flow was
represented in lines and arrows. The lines and arrows represented pressure due to the flow. In
order to get accurate data, the number of iterations that the program was allowed to run was 175.
In Figure 12, the result were that the incoming flow was at 14.74097 lbf/in^2(lime green). Once
the flow passes directly over the face of the grid fin, the pressure slightly increased in certain areas
to 14.83414 lbf/in^2(yellow). The area most affected by this higher pressure is at the base of the
grid fin. Then, the pressure decreased to 14.69439 lbf/in^2 (turquoise) once the flow passed
through the lattices.
102
Figure 7.7: Flow is directed over the grid fins and represented by arrows and lines. The lines and arrows are representatives of pressure due to flow.
Similar to the first run, the second had starting points to represent the beginning of the flow. The
starting points were placed at two inches from the top face of the grid fin. The difference between
the second test and the first is the Mach number. The Mach number in which the grid fin was
placed perpendicular to is .2 Mach. In Figure 7.8, it represents the flow simulation of the grid fin
under a .2 Mach flow trajectory. The simulation was allowed to run at 200 iterations to give a
more precise result.
103
Figure 7.8: A flow representation of 0.2 Mach flow over a grid fin. The arrows and lines represent the pressure over the grid fin.
The result of the flow simulation shows that at a 0.2 Mach the incoming flow is at 14.83976
lbf/in^2. Once the flow comes in contact with the grid fin, the higher pressure is at the base of the
grid fin at 15.30267 lbf/in^2(yellow-orange). Finally, when the flow passes through the lattices
the pressure is decreased to 14.22253 lbf/in^2(dark blue) to 14.37684 lbf/in^2(light blue).
In conclusion, the in an ideal like state environment the grid fin will receive a large amount of
pressure on the face perpendicular to the flow in both 0.1 and 0.2 Mach. Once the flow passed
through the lattices it decreased then increased once completely passed through.
Table 7.3: SolidWorks simulation run cases
Mach P1 P2 |P1-P2| P0 |P0-P2|
0.1 14.74097 14.69439 0.04658 14.83414 0.13975
0.2 14.83976 14.37684 0.46292 15.30267 0.92583
The pressures over a grid fin under a 0.1 and 0.2 Mach flow varied throughout the surface of the
object. The pressure increased once in contact with the surface of the grid fin, then decreased as
104
it passed through the lattices. The data is not accurate however due to certain entities missing that
are in a life-like scenario, such as change in acceleration of the rocket. The data is precise however,
because it helps explain how the pressure from the flow will act once the rocket is launched. The
flow visualization data from the water tunnel gives a rough understanding of how the air will
interact with the lattice structure, but due to the fact that the testing environment differs from the
launch environment the final vehicle will encounter it is not to be considered as a precise test.
Fortran- Flight and Dynamic model
A code written in Microsoft Visual FORTRAN was used to analyze the aerodynamics of the
subsonic grid fin design. A goal is to obtain a working value for the coefficient of drag to estimate
the drag force on the fins.
The required design parameters, in English units, were obtained to input into the program. The
outputs are as follows:
Table 7.4: Aerodynamic Payload Fortran- Flight and Dynamic model
Mach Number (0.1-0.8)
Atmosphere Temperature (ºR) (511.650 - 456.894 )
Atmospheric Pressure (lb/in^2) (13.6802 lb/in^2- 7.54617 lb/ in^2)
Reference Length (66.174 in)
Reference Area (21.55 in^2)
Nose Length (9.126 in)
Nose-Center body Length (66.174 in)
Total Body Length (69.3 in)
Maximum Body Radius (5.24 in)
Radius Body at Tail (5.24 in)
Nose to Fin Hinge Line (39.118 in)
105
Nose to Moment Center (43.7 in)
Nose Type (0)
Body CL to Base of Grid Fin (4.12 in)
Min Radius for grid points (0.5 in)
Body CL to Grid fin tip (9.12 in)
Height of fin support base (2.5 in)
Span of fin support base (1.5 in)
Total height of fin (0.5 in)
Chord length of fin (2 in)
Average fin element thickness (0.125 in)
Fin base corner type number cells in base
corner (1)
Fin tip corner type number cells in tip
corner (1)
Number cells in spanwise direction (5)
Number cells in vertical direction (2)
Number vortices per element chordwise (1)
Number vortices per element spanwise (1)
Fin “stall” angle (alpha max) (deg) (20)
Fin “stall” angle (delta max) (deg) (20)
Total number of fins (4)
Roll angle for configuration (15)
106
The axial force coefficient, moment coefficient, and normal force coefficient are outputs of the
program. For reference, low pressure is a pressure of 7.54617 pounds per square inch and high
pressure is 13.6802 pounds per square inch. Low temperature is 456.894 degrees Rankine and high
temperature is 511.650 degrees Rankine. The reference pressure and temperature correspond to an
altitude of 600 feet for low and 5860 feet for high. The total axial force remains constant as seen
in the plot below:
Figure 7.9: Total Fin Axial Force Coefficient versus Angle of Attack Mach 8 Low Pressure, Low Temperature
0
0.5
1
1.5
2
2.5
3
3.5
-20 -15 -10 -5 0 5 10 15 20
Fin
Axia
l For
ce
Angle of Attack (degrees)
107
Figure 7.10: Fin Normal Force Coefficient versus Angle of Attack Mach 8 High Pressure, High Temperature
Figure 7.11: Fin Normal Force Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low Temperature
-2
-1.5
-1
-0.5
0
0.5
1
1.5
2
-20 -15 -10 -5 0 5 10 15 20
Fin
Nor
mal
For
ce C
oeffi
cien
t
Angle of Attack (degrees)
-150
-100
-50
0
50
100
150
-20 -15 -10 -5 0 5 10 15 20
Fin
Nor
mal
For
ce C
oeffi
cien
t
Angle of Attack (degrees)
108
Figure 7.12: Fin Normal Force Coefficient versus Angle of Attack Mach 0.1 High Pressure, High
Temperature
The behavior of the fin moment coefficient is plotted below.
Figure 7.13: Fin Moment Coefficient versus Angle of Attack Mach 8 Low Pressure, Low Temperature
-150
-100
-50
0
50
100
150
-20 -15 -10 -5 0 5 10 15 20
Fin
Nor
mal
For
ce C
oeffi
cien
t
Angle of Attack (degrees)
-0.05
-0.04
-0.03
-0.02
-0.01
0
0.01
0.02
0.03
0.04
0.05
-20 -15 -10 -5 0 5 10 15 20
Fin
Mom
ent C
oeffi
cien
t
Angle of Attack (degrees)
109
Figure 7.14: Fin Moment Coefficient versus Angle of Attack Mach 8 High Pressure, High Temperature
Figure 7.15: Fin Moment Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low Temperature
-0.15
-0.1
-0.05
0
0.05
0.1
0.15
-20 -15 -10 -5 0 5 10 15 20
Fin
Mom
ent C
oeffi
cien
t
Angle of Attack (degrees)
-10
-8
-6
-4
-2
0
2
4
6
8
10
-20 -15 -10 -5 0 5 10 15 20
Fin
Mom
ent C
oeffi
cien
t
Angle of Attack (degrees)
110
Figure 7.16: Fin Moment Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low Temperature
The drag is calculated using the following equation:
Sample calculations of the drag for varying degrees of alpha is shown below.
Table 7.5: Sample Data Mach=0.8 Low Pressure, Low Temperature
Alpha (degrees) Fin 1 Drag
(lbf) Fin 2 Drag (lbf)
Fin 3 Drag
(lbf)
Fin 4 Drag
(lbf)
-15 -7526.98803 -7018.445745 -7514.453538 -7035.456842
-14 1460.738409 2298.157817 1482.560413 2270.880311
-13 9104.463021 9490.315161 9114.297333 9477.588404
-12 8364.231765 7826.41637 8350.195375 7844.146547
-10
-8
-6
-4
-2
0
2
4
6
8
10
-20 -15 -10 -5 0 5 10 15 20
Fin
Mom
ent C
oeffi
cien
t
Angle of Attack (degrees)
212 DD C A Vρ=
111
-11 -81.09937339 -1182.534279 -110.0120397 -1145.360851
-10 -8453.448184 -9116.323926 -8470.675464 -9093.853561
-9 -9033.726299 -8472.56012 -9018.9737 -8491.284573
-8 -1279.76665 240.3988596 -1240.264141 189.999107
-7 7659.479773 8817.295868 7689.329719 8778.400484
-6 9521.522032 8946.008968 9506.518684 8965.24403
-5 2562.923166 194.387911 2500.871239 273.6031369
-4 -6791.03415 -9117.75553 -6852.510443 -9039.6077
-3 -9819.595401 -9248.562304 -9804.634684 -9267.797512
-2 -3546.625268 1750.278581 -3407.661352 1572.504923
-1 6337.759414 14359.51281 6554.407 14089.57223
0 9984.605207 9984.605207 9984.605207 9984.605207
1 6337.759414 14359.51281 6554.407 14089.57223
2 -3546.625268 1750.278581 -3407.661352 1572.504923
3 -9819.595401 -9248.562304 -9804.634684 -9267.797512
4 -6791.03415 -9117.75553 -6852.510443 -9039.6077
5 2562.923166 194.387911 2500.871239 273.6031369
6 9521.522032 8946.008968 9506.518684 8965.24403
7 7659.479773 8817.295868 7689.329719 8778.400484
8 -1279.76665 240.3988596 -1240.264141 189.999107
9 -9033.726299 -8472.56012 -9018.9737 -8491.284573
10 -8453.448184 -9116.323926 -8470.675464 -9093.853561
11 -81.09937339 -1182.534279 -110.0120397 -1145.360851
112
12 8364.231765 7826.41637 8350.195375 7844.146547
13 9104.463021 9490.315161 9114.297333 9477.588404
14 1460.738409 2298.157817 1482.560413 2270.880311
15 -7526.98803 -7018.445745 -7514.453538 -7035.456842
16 -9585.635074 -9794.927587 -9591.184497 -9787.792615
Table 7.6: Sample Data at Mach=0.8 High Pressure, High Temperature
Alpha (degrees) Fin 1 Drag
(lbf) Fin 2 Drag (lbf)
Fin 3 Drag
(lbf)
Fin 4 Drag
(lbf)
-15 -7492.321265 -6985.569621 -7478.891451 -7002.580718
-14 1452.973667 2287.665324 1474.795671 2259.023944
-13 9061.40564 9445.522313 9071.818441 9432.795556
-12 8325.468502 7789.869379 8311.432111 7808.338314
-11 -79.92975306 -1175.857485 -108.8424193 -1138.684057
-10 -8413.421024 -9072.551705 -8430.648303 -9050.830353
-9 -8991.074978 -8433.313245 -8976.889786 -8452.037698
-8 -1274.317739 239.0369931 -1234.81523 188.6372406
-7 7623.283977 8774.768266 7653.133924 8735.872882
-6 9476.959778 8904.524323 9461.956429 8923.759386
-5 2550.964784 194.3118127 2488.912857 273.5270386
-4 -6759.394179 -9073.611906 -6819.8285 -8995.464076
-3 -9773.44672 -9205.716638 -9758.680298 -9224.757551
-2 -3530.900616 1738.460835 -3391.9367 1561.939104
113
-1 6307.832911 14286.72021 6523.321954 14017.93817
0 9937.793765 9937.793765 9937.793765 9937.793765
1 6307.832911 14286.72021 6523.321954 14017.93817
2 -3530.900616 1738.460835 -3391.9367 1561.939104
3 -9773.44672 -9205.716638 -9758.680298 -9224.757551
4 -6759.394179 -9073.611906 -6819.8285 -8995.464076
5 2550.964784 194.3118127 2488.912857 273.5270386
6 9476.959778 8904.524323 9461.956429 8923.759386
7 7623.283977 8774.768266 7653.133924 8735.872882
8 -1274.317739 239.0369931 -1234.81523 188.6372406
9 -8991.074978 -8433.313245 -8976.889786 -8452.037698
10 -8413.421024 -9072.551705 -8430.648303 -9050.830353
11 -79.92975306 -1175.857485 -108.8424193 -1138.684057
12 8325.468502 7789.869379 8311.432111 7808.338314
13 9061.40564 9445.522313 9071.818441 9432.795556
14 1452.973667 2287.665324 1474.795671 2259.023944
15 -7492.321265 -6985.569621 -7478.891451 -7002.580718
16 -9540.805653 -9748.909004 -9545.958688 -9741.774032
Table 7.7: Sample Data at Mach 0.1 Low Pressure, Low Temperature
Alpha (degrees) Fin 1 Drag
(lbf) Fin 2 Drag (lbf)
Fin 3 Drag
(lbf)
Fin 4 Drag
(lbf)
-15 -299921.8777 -274851.1012 -300151.0799 -275119.6974
114
-14 65846.45778 107128.234 65470.02821 106686.3385
-13 370939.6334 389949.3593 370765.5082 389745.7312
-12 333499.9794 307004.4448 333742.2918 307288.8664
-11 -12266.47141 -66494.24222 -11769.44891 -65911.85851
-10 -346894.8859 -379538.3328 -346596.0301 -379187.7952
-9 -360342.8125 -332726.5141 -360595.3089 -333022.7009
-8 -39273.3124 35620.7199 -39958.4766 34817.04817
-7 318864.0683 375876.561 318343.051 375264.1848
-6 379931.5922 351598.3454 380190.8808 351902.6441
-5 84149.62723 -32463.10685 85217.71253 -31211.50628
-4 -293399.6384 -407924.7617 -292349.3316 -406695.2358
-3 -391991.6522 -363895.8135 -392249.8703 -364197.5537
-2 -99566.99976 160955.2881 -101975.7076 158153.4751
-1 323876.0069 717055.4604 320137.3878 712802.4482
0 403345.2824 403345.2824 403345.2824 403345.2824
1 323876.0069 717055.4604 320137.3878 712802.4482
2 -99566.99976 160955.2881 -101975.7076 158153.4751
3 -391991.6522 -363895.8135 -392249.8703 -364197.5537
4 -293399.6384 -407924.7617 -292349.3316 -406695.2358
5 84149.62723 -32463.10685 85217.71253 -31211.50628
6 379931.5922 351598.3454 380190.8808 351902.6441
7 318864.0683 375876.561 318343.051 375264.1848
8 -39273.3124 35620.7199 -39958.4766 34817.04817
115
9 -360342.8125 -332726.5141 -360595.3089 -333022.7009
10 -346894.8859 -379538.3328 -346596.0301 -379187.7952
11 -12266.47141 -66494.24222 -11769.44891 -65911.85851
12 333499.9794 307004.4448 333742.2918 307288.8664
13 370939.6334 389949.3593 370765.5082 389745.7312
14 65846.45778 107128.234 65470.02821 106686.3385
15 -299921.8777 -274851.1012 -300151.0799 -275119.6974
16 -388937.4986 -399245.1549 -388843.5548 -399134.5628
Table 7.8: Sample Data at Mach 0.1 High Pressure. High Temperature
Alpha (degrees) Fin 1 Drag
(lbf) Fin 2 Drag (lbf)
Fin 3 Drag
(lbf)
Fin 4 Drag
(lbf)
-15 -302215.0139 -276773.5744 -302446.902 -277046.6473
-14 66434.65492 108326.0834 66051.40597 107876.0046
-13 373865.5265 393156.3981 373689.6659 392949.299
-12 336052.376 309165.2999 336297.6434 309454.1541
-11 -12459.07328 -67488.13798 -11956.54361 -66897.49352
-10 -349657.7893 -382783.6 -349355.1884 -382427.8193
-9 -363102.4399 -335078.7428 -363358.9082 -335378.9015
-8 -39444.80315 36555.2994 -40139.50244 35740.73042
-7 321453.7096 379306.5235 320923.6469 378686.0065
-6 382842.4843 354091.0672 383105.6199 354399.5976
-5 84595.39335 -33738.95163 85679.32169 -32468.86751
116
-4 -295893.7739 -412108.9742 -294828.8795 -410861.7349
-3 -394996.976 -366486.1231 -395258.8857 -366792.3321
-2 -99867.82516 164500.383 -102312.8389 161656.0045
-1 327139.0897 726122.8449 323344.8606 721806.1128
0 406489.9098 406489.9098 406489.9098 406489.9098
1 327139.0897 726122.8449 323344.8606 721806.1128
2 -99867.82516 164500.383 -102312.8389 161656.0045
3 -394996.976 -366486.1231 -395258.8857 -366792.3321
4 -295893.7739 -412108.9742 -294828.8795 -410861.7349
5 84595.39335 -33738.95163 85679.32169 -32468.86751
6 382842.4843 354091.0672 383105.6199 354399.5976
7 321453.7096 379306.5235 320923.6469 378686.0065
8 -39444.80315 36555.2994 -40139.50244 35740.73042
9 -363102.4399 -335078.7428 -363358.9082 -335378.9015
10 -349657.7893 -382783.6 -349355.1884 -382427.8193
11 -12459.07328 -67488.13798 -11956.54361 -66897.49352
12 336052.376 309165.2999 336297.6434 309454.1541
13 373865.5265 393156.3981 373689.6659 392949.299
14 66434.65492 108326.0834 66051.40597 107876.0046
15 -302215.0139 -276773.5744 -302446.902 -277046.6473
16 -391988.6196 -402448.0922 -391893.0903 -402335.9146
117
The calculated drag force coefficient is useful for predicting drag on the grid fins. Calculating the
drag force on the grid fins is important for insuring the structural integrity of the grid fins and of
the related components. Moreover, a calculated drag force is invaluable for investigating the
possibility of maneuvering the rocket for a safe and quick recovery. A percent error of 10-15 %
is expected when comparing data obtained in FORTRAN to experimental data.
Drag Profile
A drag simulation was performed in MATLAB to see what forces are going to act on the grid fins
through various velocities and altitudes during flight. This simulation was necessary to acquire
rough estimates for the other simulations. The altitude and velocities were determined through an
Open Rocket simulation, and the area for the grid fin was iterated three times at angles of thirty,
sixty, and ninety. This figure compares the design trajectory (the values designed for when the
rocket) velocity vs altitude graph and the optimal (trajectory to complete the mission) velocity and
altitude graph. Using the Drag equation various drag forces were obtained and plotted against
altitude and velocity.
Table 7.9: Calculated Drag and Acceleration Values
Drag Estimate of Fins at Max Velocity, 45 degrees 53.1661 pound-force
Drag Estimate of Fins at Max Velocity, 90 degrees 13.4517 pound-force
Drag Estimate of Rocket at Max Velocity 96.9410 pound-force
Drag Estimate of Rocket and Fins at Max Velocity, 45
degrees 150.1071 pound-force
Drag Estimate of Rocket and Fins at Max Velocity, 90
degrees 110.3927 pound-force
Max Acceleration at Max Velocity with Fins, 45
degrees -185.1896 feet per square second
Max Acceleration at Max Velocity with Fins, 90
degrees -136.1933 feet per square second
118
Vortex Shedding Testing
A 3-D printed, full-scale grid fin and fairing was placed into a water tunnel for observational data
to be acquired. The model is set up using a test rig to allow for the grid fin to be placed at all of
the different angles it will experience during flight. Dye was inserted to the flow upstream of the
model to allow for visualization of the vortices and any other adverse flow effects that could
negatively impact the performance of the grid fin during flight. Pictures and video are to be taken
to allow for future analysis and increased understanding of the system being tested.
During the test a 3/8” rod was screwed into the faring and grid fin system. Next, connect the 3/8”
rod was attached to the adjustable angle arm in the water tunnel. The adjustable arm is adjusted
to where the grid fin system is perpendicular to the flow of the water. While holding the dye port,
the water tunnel ran through a range of hertz. The dye port was turned on after the water tunnel
speed was reached. The dye from the port flew through the grid fin showing the vortices of the
flow.
Runs at both low and high speeds with dye injected upstream of the fin increased turbulence of the
flow. This validates the hypothesis established during the design phase of the grid fin. When
laminar flow enters the grid fin the flow transitions into turbulent and creates vortices downstream.
They are more pronounced in the high speed flow tests due to the higher Reynolds number
associated with it. Vortices are also present in the low speed flows, but their size is not as large.
The transition to turbulent flow and the vortices created indicated a large increase in pressure drag
by the grid fins, which is their primary purpose. The test also shows that the flow remains turbulent
for a short distance downstream. Therefore, the visualization indicates that the flow will be laminar
when interacting with the main fins of the rocket. No numerical data was gathered, as this was
only a visualization test.
119
Figure 7.17: Vortex Shedding Testing visualization
1:5 Scale Test
A 1:5 scale model was built of the rocket for wind tunnel testing. A 1:5 scale WAFLE section was
built for the model. The WAFLE section was inactive for the test. The actuating system for the
fins was not scalable for a test one-fifth the scale. Therefore, the WAFLE section was an in active
aerodynamic version of the section.
One-fifth scale grid fins and fairings were printed using HIPS. The fins and fairing were epoxied
to the body of the rocket in the stored position. The epoxied fins and fairing were located in the
same position on the rocket as the full scale.
The fins and fairing were tested before being placed in the wind tunnel. The structure was deemed
secure and safe for the wind tunnel. Once inside the tunnel, aerodynamic data was gathered and
recorded at subsonic speeds. The fairing and fin remained secure to the body of the model
throughout the test. Thus the fairing and fin was structurally and aerodynamically certified at the
1:5 scale level.
3:5 Scale Test
120
After the 1:5 scale test, a 3:5 scale launch was performed. A 3:5 scale model of the Aquila rocket
was built and 3:5 scale models of all payload systems were designed and integrated within the
model. The WAFLE actuation system was deemed to be non-scalable to the 3:5 scale level.
Therefore, the WAFLE segment was built as an aerodynamic model and did not actuate throughout
the flight.
The process for manufacturing the fins and faring remained the same as the 1:5 scale test. Both
subsystems were printed using additive manufacturing and used the HIPS material. The fins and
fairings were then epoxied to the rocket body in a stored position. After applying a load to the fins
in the axial direction to insure full adhesion, the fins and fairing were deemed worthy to fly.
The rocket was transported to a launch site and then launched. Once retrieved after touch down,
the WAFLE segment was inspected. The fins and fairing for the segment remained secured to the
body of the rocket throughout the flight. Since the rocket traveled at approximately Mach 0.8, the
fairing was safely assumed to break Mach 1. With that assumption and the fins and fairing
remaining secured to the rocket, the WAFLE segment was certified at the 3:5 scale level.
Full Scale Test
The Full Scale Test will be the final certification for the WAFLE system. The Full Scale Test will
have a full scale WAFLE system designed and built for it. The WAFLE system will join the other
payload system in the Full Scale Test. The Full Scale test will validate that the rocket can achieve
the mile height requirement and that the WAFLE system operates as desired.
During the first Full Scale Test, the WAFLE system recorded data. Unfortunately this data was
not reflective of the desired flight due to the rocket reaching a higher altitude than expected.
However, since the electronics survived the flight and proved that they can record the data for the
flight, the WAFLE system was validated to continue with minimal changes.
During the second Full Scale flights, the failure destroyed the WAFLE second. Therefore, no data
was collected during the flight.
The failure in the third Full Scale flight resulted the destruction of the WAFLE system, therefore
no data was retrieved from this flight. However, since the grid fins and the fairings were survived
121
the impact with the ground, it was verified that they were structurally sound in their design and
configuration.
The motor failure in the fourth Full Scale test resulted in another failure to collect data.
Section 7.4: Flight Performance Predictions An OpenRocket model and a Matlab simulation were created to evaluate the performance of
WAFLE at 45 degrees and 90 degrees of leading edge sweep. The OpenRocket model of the
Project Acquila rocket modeled the WAFLE as a transition, which provided important constants
for calculating drag force and acceleration, such as: max altitude, density of air at max altitude,
max velocity in transition, simulated mass in transition, and the axial force coefficient of the
rocket. Further, a Matlab simulation was created to calculate the drag and acceleration forces
acting on the WAFLE. The following two tables present the WAFLE and vehicle constants used
in OpenRocket model and Matlab simulation:
Table 7.10: WAFLE Constants
Height of Fins 5.1 inches
Span of Fins 2 inches
Area of a Hole in the Fin 0.4356 square inches
Area of a Fin 3.666 square inches
Table 7.11: Vehicle Constants
Diameter of Vehicle 5 inches
Reference Area of Vehicle 19.6350 inches
The following estimations were obtained from a paper titled Curvature and Leading Edge Sweep
Back Effects on Grid Fin Aerodynamics Characteristics by Mark S. Miller:
122
Table 7.12: Estimations from Miller's Document
Axial Force Coefficient for Fins at 45
degrees 0.83
Axial Force Coefficient for Fins at 90
degrees 0.21
The following table lists values obtained from an OpenRocket model of the Project Aquila rocket
with the WAFLE modeled as a transition:
Table 7.13: Constants from Transition Open Rocket
Altitude 1659.3 feet
Density at Altitude 0.00226361 Slugs per cubed foot
Max Velocity in Transition 745.5 feet per second
Simulated Mass in Transition 26.1 pound-mass
Axial Force Coefficient of Body 0.78374
The following table presents the drag and acceleration values calculated in the Matlab simulation
for the WAFLE:
Table 7.14: Calculated Drag and Acceleration Values
Drag Estimate of Fins at Max Velocity, 45 degrees 53.1661 pound-force
Drag Estimate of Fins at Max Velocity, 90 degrees 13.4517 pound-force
Drag Estimate of Rocket at Max Velocity 96.9410 pound-force
Drag Estimate of Rocket and Fins at Max Velocity, 45
degrees 150.1071 pound-force
Drag Estimate of Rocket and Fins at Max Velocity, 90
degrees 110.3927 pound-force
123
Max Acceleration at Max Velocity with Fins, 45
degrees -185.1896 feet per square second
Max Acceleration at Max Velocity with Fins, 90
degrees -136.1933 feet per square second
The data obtained indicates that a 45 degree sweep is optimal for the WAFLE system, because
the drag is maximized with a drag force of 53.1661 pound-force. At max velocity, the max drag
force at a 45 degree sweep reaches 150.1071 pound-force. Maximum drag enables WAFLE to
effectively act as air breaks, so that our rocket does not overreach the maximum altitude.
Section 7.5: Payload Design Wall Armed Fin-Lattice Elevator
The Wall Armed Fin-Lattice Elevator (WAFLE) is the primary aerodynamic payload system. This
system will be integrated into the rocket 43.125 inches aft of the fairing tip. The overall length of
the WAFLE is 8.85 inches. The system is composed of multiple subsystems including: Grid fins,
Outer Fairings, 10 DOD IMU, RF Tracker, Servos, and an Arduino.
Figure 7.18: WAFLE system
A fairing is located at the tip of the WAFLE. The fairing extends 4.10 inches aft of the rocket.
Four fairings are mounted on the rocket; oriented 90 degrees from one another. Aft of each fairing
are the servos. The servos are mounted on the Servo plate that allows them to protrude out from
124
the airframe and remain flush with the outer face of the fairing. The servos are the point of rotation
for each grid fin, so the servo gear is embedded within the grid fin base. The grid fin extends 5.10
inches aft of the servos; terminating 1.16 inches aft of the waffle.
The internal subsystems for the WAFLE are the Arduino, Servos, 10 DOF IMU, and RF Tracker.
All electronic of the WAFLE are located on the LANTERN. The breadboard holding the 10 DOF
IMU is located on one side of the electronics sled, while the Arduino is located on the other side
of the sled. A battery is located below the sled on the bottom bulk plate of the LANTERN. The U-
bolt that holds the parachute for the booster section is located on the top plate of the LANTERN.
Threaded rods hold the LANTERN plates together as well as hold the electronics sled in place.
Arduino
The Arduino Uno is a single-board microcontroller that provides digital I/O pins of 14/6 and analog
I/O pins of 6/0. The pins can be used to send and receive signals shared with connected devices
such as the servos and the sensors. The primary use of the Arduino is to send commands to the
servos and receive data from the sensor telling it when to actuate. The Arduino Uno will read input
data from an accelerometer and a GPS and use those inputs to output a rotation angle for the servos
to pitch the grid fins in order to reach a specific altitude. A rechargeable battery source will power
the Arduino, which will supply the necessary power for all inputs and outputs.
Figure 7.19: Arduino Uno
Servos
125
Savox SV-1270TG High Voltage Monster Torque Servos is a very high torque servo which is
connected to the Arduino Uno. Using an input the Servo orients an attached object (grid fin) to a
specific angle. The torque produced by the Servo locks the grid fin into place in order to counteract
the forces on the grid fin. The Savox SV-1270TG Servo is powerful producing a torque of 35.07
kg/cm at 7 volts. The torque produced is very strong compared to the dimension and weight of the
Servo which is 40.3 x 20.2 x 37.2 cm and 56 g.
The HiTec HS-5685MH Servo was replaced by the Savox SV-1270TG Servo because the Savox
Servo provided a much higher torque. When testing how strong the HiTec Servo was the grid fins
could be moved by the force created by a person’s hand while at full deployment. This was opposed
to the Savox Servo which did not budge at full deployment when a force was created by a person’s
hand.
Figure 7.20: Savox SV-1270TG
10-DOF IMU Breakout
The 10-DOF IMU Breakout replaced the ADXL335 Triple-axis Accelerometer due to the
capabilities of the IMU Breakout. The IMU Breakout has a 3-axis accelerometer, gyroscope and
compass as well as a barometric pressure/temperature sensor. This is much more efficient that
126
using three different devices to measure the required values to calculate a final height in order to
actuate the grid fins. The IMU Breakout has a temperature range of -40 to 85o C, a pressure
range between 300 and 1100 hPa, is capable of detecting an acceleration up to +/- 16 g’s and has
a gyroscopic scale of +/- 2000 degree-per-second. The IMU Breakout can run on either 3 or 5V
and is accurate up to +/- 3 feet for both altitude and acceleration readings.
Figure 7.21: 10-DOF IMU
RF TRACKER
A 30 milliwatt RC-HP Transmitter has been chosen to broadcast the location of the booster
section up to a 10 mile range. This transmitter was chosen due to the unnecessary size and
additional capabilities of the high-tech GPS system that was originally chosen. Initially, a GPS
system would calculate the position, velocity and acceleration of the booster section in order to
actuate the grid fins. A 10-DOF system was determined to be more accurate than a GPS for
calculating velocity and a final position. The RC-HP Transmitter is now solely used for
broadcasting a location for retrieval after descent. A CR2032 battery is used inside the
transmitter and has a 1 week battery life. The frequency of the RC-HP Transmitter is 222.450
MHz.
Fairing
127
The fairing will allow the WAFLE section to obtain a more aerodynamic form and reduce the
stress formed within the servos and grid fins. The fairing will be made of High Impact Polystyrene
(HIPS) and printed by means of additive manufacturing. The ease of manufacturing, low cost, and
high impact strength made HIPS the obvious choice of materials to make the redesigned fairing
from. The fairing will be 4.10 inches in length and 2 inches in width.
The fairing is configured with an ogive-like shape. This shape will allow for the local flow velocity
on the fairing to remain close to freestream velocity. The attempt is to prevent the flow over the
fairing from breaking Mach 1. This would impede the flow through the grid fins and reduce the
overall drag on the fins.
Figure 7.22: Grid Fin Fairing
Grid Fin
The grid fins are lattice shape control surfaces. An illustration can be viewed in Figure 9. The
lattice shape allows flow to pass the fin but will still impair the flow on the lattice surfaces. This
will provide some drag but will allow the root chord moment to be small. A small root chord will
mean that the torque required for the fin to actuate is also small. This reason is why grid fins are
an ideal chose for use in control surfaces on rockets, and subsequently this mission.
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The grid fins are one of the main payloads on the rocket. Since the grid fins create drag but are
still practical to actuate, they are used to correct the trajectory of the vehicle. The grid fins are
deployed perpendicularly to the direction of flow to create the drag. The grid fins will deploy
during flight and use drag to control the rocket’s target apogee. The intent is to accurately complete
the Vehicle Requirement 1.1.
In order to evaluate how the grid fins will interact once deployed, the team will construct visual
testing of the fluid flow through the lattices of the grid fins. Therefore, a basic lattice fin has been
designed and implemented to act as the primary grid fin. The lattice was designed to be easy to
model and manufacture, and still obtain adequate drag characteristics. The length of the grid fin is
5.91 inches, span of 2 inches, and height of 0.77 inches. The holes are 0.66 x 0.66 inches, making
the lattice thickness 0.05 inches. The fins are printed with HIPS through a process of additive
manufacturing. This material, like the fairing, will withstand the high strain induced by the external
flow.
Figure 7.23: Aerodynamic Grid fin
The wiring for the WAFLE system is illustrated in the schematic. The 7.4 voltage source supplies
power to the servos directly. It also powers the Arduino, which in turn powers the 10 DOF IMU.
Signal lines run from the servos to the Arduino in order to communicate when it needs to actuate.
The acceleration in each axis is output from the accelerometer to the Arduino.
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Figure 7.24: WAFLE electronics schematic
Section 7.6: Requirement Verification Table 7.15: Aerodynamic Payload System Validation Table
Requirement
Number Requirement Method of Validation
3.2.6 An aerodynamic analysis of
structural protuberances
A full aerodynamic analysis of the
grid fins is conducted through
computational fluid dynamics
(CFD), subsystem wind tunnel
testing, and in-flight sensors.
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3.2.6.1
Grid Fin payload is self-
contained within a separate
segment of the rocket.
The WAFLE system is built to be a
self-contained and is removable
from the rest of the booster section.
3.2.6.2
Aerodynamic fairing is
firmly adhered to the gird
fin segment.
The fairing contains screw holes that
allow the fairing to be hard mounted
to the airframe.
3.2.6.3
Bulk heads sealing the ends
of the segment are
stationary throughout flight
A permanent bulk plate will seal the
top section of the rocket. The bottom
on the segment will be secured with
pins to insure that the WAFLE
segment does not separate from the
booster segment.
3.2.6.4
Grid fins must stay
deployed during the decent
phase of the trajectory.
When the Arduino detects that
apogee has occurred, the fins will be
deployed.
3.2.6.5 Grid fins must stow away at
100 feet.
Arduino will be informed from
sensors that 100 feet is reached and
will implement the storing sequence.
AU1
All Aerodynamic data must
be validated through
analytical and experimental
testing.
A full aerodynamic analysis of the
grid fins is conducted through
computational fluid dynamics
(CFD), subsystem wind tunnel
testing, and in-flight sensors.
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AU2
Grid fins must stay stowed
until boost phase is
complete.
Redundant timer will be
implemented into the system to
insure that the code iteration does
not engage. This pause timer will
wait until the acceleration of the
rocket is within a safe range before
starting the Arduino calculations.
AU3
Electronics must stay
stationary throughout the
flight
The electronics will be adhered to a
stationary plate within the airframe.
This plate and mounting bolds will
be secured to a stationary plate
within the rocket.
AU4
Servos must remain in
direct contact with the gears
of the grid fins throughout
the flight.
The gears of the servos will be
imbedded into the U-bracket base of
the grid fin by means of a metal bar.
Do to the high strength of the metal
bar and HIPS, the fin will stay
attached.
AU5
Arduino must accurately
predict the flight path of the
vehicle.
Testing and accurate simulation
modeling will insure accurate
prediction.
AU6
Grid fins must be deploy
with precision to correct the
vehicle’s trajectory.
The Arduino will tell the servos to
rotate a specific degree. Since the
grid fins are directly attached to the
servos, the fins will see the same
rotation.
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AU7
Grid fins must stay
deployed under the force
applied by the flow.
The Arduino will not be actuated
until the flow force is under the
maximum torque provided by the
servos.
Section 7.7: Payload Integration Purpose
The purpose of the document is to describe the concept of the primary aerodynamic payload
system known as the “Wall Armed Fin-Lattice Elevator (WAFLE). Also, the document will
explain how the system is integrated into the rocket. As well as what the system will do for the
complete the mission of the team.
Scope
The primary aerodynamic payload system is what the team calls the “Wall Armed Fin-Lattice
Elevator. This system is integrated into the rocket 43.125 inches aft of the fairing tip. The
system is composed of various systems including: Grid fins, Outer Fairings, GPS,
Accelerometer, Servos and an Arduino. The WAFLE system is used to create drag and actuate,
to allow the vehicle to slow down and fly the correct trajectory. In creating this system, various
challenges and deployment complexities have come across the team. Many challenges like size
and shape of the grid fins were an issue to the team along with the type of servos to use to
actuate the grid fins.
Integration Strategy
The integration is used to create a working system to actuate and create drag for the rocket.
Previous plans were to connect the system permanently, however this idea had some flaws. For
example, if the rocket were damaged in a field test, then the system cannot be salvaged. Thus,
the idea of a better system that allowed one to remove it from a damaged rocket was created.
The team decided to create what is known as a “LANTERN.” The LANTERN is a system built
to hold batteries, Arduino, and 10 DOF sensor. The team has also created a Servo Plate which
mounts under the LANTERN. The Servo Plate will secure and house the servos.
The two systems were made for easy installation and replacement. In case the system is
destroyed or damaged in a field test, the system is easily accessible and removable. Also, the
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other way around if the section of the rocket that contained the system is damaged, then the
system could be removed.
Phase Integrations
Phase 1: General Unit
• LANTERN • Arduino • Breadboard • Four Servos • Servo plate • Key switches • Key switch key • Battery
Phase 2: Carbon Fiber Tube and Servo plate
Slots cut into the tube 11.1 inches from the top of the tube and should be width and height of the
servos.
The bulk plate with three angle brackets attached serves as a Servo Plate.
Phase 3: LANTERN body
The LANTERN has two bulk plates connected by two threaded rods with nuts.
The LANTERN body has an Arduino/Breadboard sled to cradle the equipment.
Four Angle brackets and one U-bolt with washers.
Phase 4: Breadboard Specific
The breadboard should contain one 10 DOF IMU sensor.
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Figure 7.25: LANTERN Configuration
WAFLE
The design of the LANTERN system and servo plate utilize a design based around the complete
system being removable from the WAFLE, as well as robustness. The servo plate is designed to
be a permanent aspect of the rocket body, whereas the LANTERN needs to be removed prior to
and after launches to access the Arduino and servo leads. The servo plate has three 0.75 inch L-
brackets that are aligned with three holes in the isogrid tube and secured with #10 bolts. Bondo
All-Purpose putty was added in each grid section that the bolts pass through for structural
integrity and to reduce the likelihood of a tear-out failure. The bolts are secured to the L-
brackets on the interior of the rocket with standard locknuts. The LANTERN is secured to the
isogrid tube in the same manner, with the exception being that the L-brackets have threaded
holes tapped into them in place of locknuts on the interior. This method was chosen due to lack
of accessibility once the LANTERN is inserted into the rocket. The LANTERN has four
connection points, as opposed to the previously used three. This is mainly to increase stability of
the section and its sensitive electronics since three connection points has been proven to be
sufficiently secure.
LANTERN
The LANTERN houses all of the grid fin payload’s electronics in a secure, removable unit. It is
compromised to two 0.25 inch thick, carbon fiber bulkplates that are connected via two 0.625
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inch diameter threaded rods. They allow for a two-sided sled made of balsa wood and fiberglass
to hold the Arduino and 10-DOF in a vertically oriented position. The sled is epoxied onto two
carbon fiber tubes that line up and slide around the threaded rods. This allows for the electronics
to be removed from the LANTERN for post flight analysis without needing to take the entire
LANTERN system. The battery for the system lays on the bottom bulkplate and is secured via
Velcro strips. This layout was chosen so that the leads from the servos could easily be threaded
up through a hole in the bottom bulkplate of the LANTERN to be attached to the breadboard.
The design driver for this section of the rocket was removability due to the complexity of the
electronics.
Servo Plate
The servos are connected to the servo plate with 0.032 inch thick 6061-T6 aluminum sheet
brackets that were manufactured in house. The brackets contact each side of the servo and are
secured with epoxy. The tail of the bracket extends into the rocket body through rectangular
slots in the isogrid tube. All four of the servo brackets’ holes line up at the center of the rocket
and a #10 bolt passes through them, where it is threaded into a T-nut that is epoxied into a
wooden block. These brackets were designed knowing that the servos would need to be
removable but provide the necessary strength to hold the servo/grid fin assembly in place during
flight.
Figure 7.26: Servo Bracket
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Section 8: Safety
Section 8.1: Safety Officer Team member Austin Phillips is the ideal choice for a safety officer. He is an aerospace
engineering graduate at Auburn University and now a senior in polymer and fiber engineering at
Auburn University. Austin is a fully trained and certified EMT and firefighter in the state of
Alabama. Working full-time as a firefighter for the City of Auburn as well as being a student at
Auburn, Austin is well versed in crisis-management and safety practices. His extensive training
makes him an invaluable resource towards maintaining safety throughout the competition. In
addition, having a High Powered Level 1 and 2 certification, and very close to completing his
level 3, Austin is well versed in the challenges and safety hazards that are associated with the
construction of a high-powered rocket.
The safety officer is responsible for producing the main checklists for the vehicle, watching over
construction of the different vehicle elements, among other definable responsibilities. Austin will
produce the main check-lists that will be used for checking the different parts of construction,
payload integration, and flight readiness. He will be involved in the construction of the different
vehicle elements to ensure that all components of the vehicle are built to a certain standard that is
ensures complete safety during flight. Austin will provide any immediate medical care that could
be required if a team member is hurt or ill while in the lab or if a team member or bystander is
injured at a launch. He will be responsible for inspecting the different vehicle components at the
end of their construction and for the final vehicle inspection before the rocket has its final
inspection by the RSO.
Section 8.2: Airframe Hazard Analysis Safety is taken into consideration for every part of building the rocket. There are steps that will
be taken by the airframe team to ensure the safety of the members while they construct the
airframe for the rocket. There are three different areas that we will look at while considering
failure modes for safety protocols for airframe: operations, materials, and construction.
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Section 8.2.1: Airframe Failure Modes
All of these failure modes have been taken into consideration and the proper mitigations have
been put into effect to ensure the safety of team members and the environment. Mitigation tables
for failure modes within airframe are listed in the following section.
Operations Failure Modes:
• Transport
o Not properly transported
o Airframe damaged it Transportation
• Storage
o Stored in wet area
o Stored in dirty area
• Ground Operations
o Cracks in the carbon fiber
o Gaps between different parts
o Excess epoxy
o Lack of epoxy
• Launch
o Cracks in Airframe
o Airframe breaking apart
Construction Failure Modes:
• Autoclave
o left in Autoclave by Previous user
o Drain strainer not properly cleaned
o Explosive breakage of glass vessels
o Burns to hands and other body parts
o Lacerations to hands and other body parts
o Trauma to users eyes
o Materials catching on fire
o Breathing toxic fumes
139
o Autoclave not set on correct setting
• Aluminum mandrel
o Hands caught in mandrel
o Burns from touching mandrel after it comes out of autoclave
o Injury due to torque of mandrel while wrapping material
• Filament Winder
o Fingers caught in moving parts
o Exposure to epoxy and carbon fiber
o Loose clothing and/or hair caught in winder
Materials Failure Modes:
• Carbon Fiber
o Allergic dermatitis from coming in contact with carbon fiber
o Skin irritation from coming in contact with carbon fiber
o Respiratory irritation from breathing in particles
o Trauma to users eyes from fragments of carbon fiber
o Carbon fiber should be kept away from electrical equipment
• Epoxy
o Trauma to eyes from epoxy coming in contact with eyes
o Setting up before work is completed
o Mixing too much epoxy
o Heating up and melting through container
o Improper disposal
Personal hazards that could occur during the construction of the airframe and during the launch
have been assess to ensure the safety of team members and people in the area around the launch
site. Mitigation tables have been put in place to make team members aware of these hazards to
minimize the risk of them occurring, these mitigation tables are listed below. Along with the
mitigation tables team members are required to read over the MSDS sheets that pertain to the
material or machine that they are working with. To prevent personal hazards while operating the
autoclave each team member should be knowledgeable about how the autoclave operates by
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reading over the operator’s manual for the autoclave, alone with looking over the mitigation
table that has been put in place. Table 8.1: Risk Mitigation Table – Airframe
Many problems can occur as a result of improper care of the rocket airframe. Table 9.1
summarizes potential problems, their potential effects, what we are doing to prevent these
problems from ever occurring, and what we will do in the case that they do.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Airframe not
properly
transported (1)
Damage to
airframe 4 3
Custom made boxes
with foam inserts
have been created in
which the airframe is
transported to and
from events to
protect it from
damaging vibrations
or slipping
1
Airframe not
properly stored (2)
Damage to
airframe 4 3
Airframe is stored in
locked aerospace lab
or custom shipping
boxes when not
being constructed or
tested
1
141
Cracks in Airframe
(3)
Breaks on launch
injuring team
members or
bystanders
5
3
Airframe is
inspected at every
stage of construction
and pre-launch
inspections with the
use of a checklist are
conducted before
each launch to
confirm structural
integrity
1
Gaps between
airframe and other
parts of the rocket
(4)
Failure during
launch or early
separation
resulting in high
velocity projectiles
causing injuries to
team members or
spectators
5 3
Airframe
components are
constructed using
specialized tools to
ensure exact
dimensions and a
prelaunch inspection
is conducted before
each launch to ensure
that there are no gaps
between parts
1
Lack of Epoxy (5)
Airframe breaks
apart during launch
causing pieces to
fall on spectators
5 3
Epoxy is mixed in
3:1 ratio to ensure
maximum bonding.
Each part is
inspected to ensure
that is has sufficient
epoxy
1
142
Collision with bird
(6)
Damage to
airframe 4 2
Testing has been
performed on nose
cone and airframe to
ensure strength is
sufficient to
withstand minor
collisions
1
Airframe breaks
apart in flight (7)
High speed objects
falling on
spectators
5 3
Strain and stress tests
have been performed
on sample materials
to confirm integrity
of materials with a
safety factor of at
least 2 times
1
Table 8.2: Risk Mitigation Table – Autoclave
Improper use of the autoclave has potential to cause injury to operators or damage to the
autoclave and lab. Table 9.2 summarizes the risks involved in the operation of the autoclave and
the guidelines that we are following to prevent injury.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Debris flies up into
user’s eyes (1)
Trauma to the
user’s eyes 3 3
Operators are
required to wear
safety glasses or face
shield while
operating autoclave
1
Material left in
autoclave (2)
Damage to
autoclave and
material
4 3
Operators double
check autoclave to
ensure it is empty
before operating
1
143
Door not properly
closed (3)
Damage to material
inside autoclave 2 3
Operators must
double check that
doors are closed and
locked before turning
autoclave on
1
Wrong cycle
selected (4)
Damage to material
inside autoclave 2 3
Autoclave is stored
in a locked lab where
only authorized and
trained users may
operate autoclave.
1
Material
experiences
explosive breakage
when autoclave is
opened (5)
Can cause severe
injuries to users 5 2
Operators must wear
proper PPE and
always keep hands,
head, and face clear
while opening.
1
Touching hot
materials (6)
Severe burns to
users 4 3
Proper PPE such as
heat and cut resistant
gloves are required
before opening
autoclave.
1
Materials catch fire
(7)
Damage to the
autoclave and
materials will
occur. Possible risk
of fire spreading to
the rest of building
and causing harm
to individuals
5 3
A fire extinguisher is
kept in the same
room as the
autoclave and is
easily accessible. If
fire spreads
personnel will
contact 911
immediately.
2
144
Toxic Fumes (8)
Can cause
respiratory
problems
5 5
Respirators are
required when
working with
potentially hazardous
materials. Lab is
properly ventilated at
all times.
1
Unauthorized use
(9)
Damage to
Autoclave,
materials, and to
personnel
5 3
Lab where autoclave
is located is locked
up and can only be
accessed by
authorized personnel
1
Table 8.3: Risk Mitigation Table - Filament Winder
The filament winder is an expensive piece of equipment that be damaged or cause injury if
operated incorrectly. Table 9.3 summarizes common incorrect procedures that can lead to
damage or injury and how to prevent them from occurring.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Mandrel not
secured properly
(1)
Improper
construction of
rocket body tubes
leading to
structural failure
5 3
Only trained team
members have
permission to operate
filament winder.
Operator must check
that the mandrel is
properly secured with
winder clamps.
Equipment must be
supervised while in
use.
1
145
Improper winding
angles for the
specific stresses
occurred during
flight (2)
Improper
construction of
rocket body tubes
leading to
structural failure
4 3
Material testing has
been performed on
samples to ensure
that the winding
angles used on our
rocket tubes will be
strong enough to
withstand expected
forces
1
Winder runs out of
resin when using
dry filaments (3)
Structural integrity
of rocket body
tubes is
compromised
leading to a
structural failure
during flight
4 3
Supervision of
equipment is required
while in operation
1
Filament does not
unroll correctly (4)
Improper
construction of
rocket body tubes
leading to
structural failure or
damage to
equipment
4 3
Supervision of
equipment is required
while in operation
1
Hair gets caught on
mandrel (5)
Hair ripped out,
scalp injuries 4 2
Long hair must be
pulled back while
operating filament
winder
1
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Table 8.4: Risk Mitigation Table - Carbon Fiber
Carbon fiber can be a harmful substance if it is not handled properly, especially when cutting or
sanding it. Table 9.4 summarizes the potential harmful effects of carbon fiber and how to avoid
them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Allergic reaction
from coming in
contact with carbon
fiber (1)
Skin irritation 3 4
Proper PPE must be
worn when handling
carbon fiber
1
Debris flies up into
users eyes (2)
Trauma to the
users eyes 3 3
Safety glasses must
be worn when cutting
or sanding carbon
fiber
1
Toxic particles (3) Respiratory
irritation 3 3
Respirators are
required when
cutting or sanding
carbon fiber
1
Electrical shock (4) Burn or
electrocution 4 2
No exposed electrical
wires may be present
in section of lab
where carbon fiber is
being manufactured
or cut
1
Table 8.5: Risk Mitigation Tables – Epoxy
Epoxy plays an integral part in the construction of our rocket; however, it can have harmful
effects if handled improperly. Table 9.5 summarizes the problems that can arise and what we
have done to prevent them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
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Improper
Ventilation (1)
Vapors can cause
headache, nausea,
and irritate the
respiratory system
4 5
Lab is ventilated at
all times. Proper PPE
is required when
working with epoxy.
2
Skin Contact (2) Can cause skin
irritation 2 5
Disposable gloves are
required when
working with epoxy.
1
Degradation of
Epoxy Resin (3)
Bonds weakly
resulting in parts
that break easily
4 3
Epoxy is stored in an
air conditioned lab
between 40°F and
120°F
1
Spilling and leaking
(4)
Hardens on work
table or lab
equipment
damaging the
equipment
2 4
Personnel using
epoxy must not be
distracted by any
other tasks. In the
case of a spill, paper
towels are used to
clean up and stop
leakage. Warm water
and soap must be
used to clean up
messes immediately
1
148
Fire Hazard (5)
Damage to lab
area, equipment,
and personnel
5 3
Epoxy is not stored
or used near heat
sources. In the case
of a fire, a fire
extinguisher is stored
in an easily
accessible location in
lab
1
Epoxy gets in user’s
eyes (6)
Damage to the
user’s eyes 5 2
Safety glasses must
be worn when using
epoxy
1
Epoxy setting up
before work is
finished (7)
Waste of epoxy
that is not used 2 3
Epoxy is mixed in
small amounts and
only when it will be
applied immediately
1
Epoxy burning
through container
(8)
Potential fire
hazard and damage
to lab
2 3
Mixed epoxy must be
supervised and the
user must be aware of
how hot the epoxy is
as it starts to set
1
Epoxy not properly
disposed (9)
Potential fire
hazard and damage
to lab
2 3
All wasted epoxy
will be cured and
allowed to cool
before disposal
1
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Section 8.2.2: Airframe Risk Mitigation – Testing Systems
Table 9.6: Risk Mitigation Tables – Wind Tunnel Testing
The wind tunnel is a sophisticated piece of equipment that requires trained personnel to operate
safely and efficiently. Table 9.6 summarizes the problems that would result from improper
operation of the wind tunnel and the guidelines that are followed to prevent them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Debris in the wind
tunnel (1)
Damage to wind
tunnel, object
being tested, or
personnel
4 3
Test objects are
inspected before
testing to ensure they
will not break. Wind
tunnel is inspected
for loose debris
before each use
1
Open test section
(2)
Incorrect results
calculated from the
wind tunnel that
can have
potentially
damaging effects
on the rocket in the
future
5 2
Operators double
check that doors are
shut and locked
before turning wind
tunnel power on
1
Inexperienced
personnel (3)
Damage to project
and equipment due
to incorrect
operation of the
wind tunnel or
personnel injury
5 3
Wind tunnel is stored
in a locked lab to
prevent any
unauthorized use
1
150
Running the wind
tunnel too high (4)
Can cause
structural damage
within the wind
tunnel, hurt the
intended test
object, and hurt the
engine running the
wind tunnel
5 3
Wind tunnel must be
supervised by trained
operator while in use.
Wind speed is
limited to less than
160 feet per second.
1
Overusing Motor
(5)
Engine becomes
damaged and
would cost large
amounts of money
to repair or replace
5 3
Use of wind tunnel
must be scheduled in
advance. Periodic
checks of the system
are performed to
keep engine running
properly
1
Section 8.3: Scientific Payloads Hazard Analysis During the process of building a rocket, safety is constantly kept in mind. Safety is even more
critical in this year’s competition due to the complexities introduced in payload integration.
Guidelines have been implemented to ensure the safety of the members of the scientific payload
team during the construction and testing phases of Project Aquila. There are three different
sections that are being looked at while considering failure modes for safety protocols for the
scientific payloads: operations, construction, and materials.
Operations Failure Modes:
• Mission Processes
• Testing
• Personnel Risks (Operator and Observers)
• Environmental Risks (Macro and Micro)
• Vehicle Risks (Launch, Flight, and Recovery)
151
• Controller Risks (Electrical and Mechanical)
Construction Failure Modes:
• Hand Tools
• Soldering Equipment
• Drill Press
• Band Saw
• Autoclave
• Personnel Risks
• Environmental Risks
• Vehicle Risks
Materials Failure Modes:
• Carbon Fiber
• Aluminum
• Epoxy
• Electric Servos
• Copper Wires
• Flux and Soldering Materials
• Personnel Risks
• Environmental Risks
Section 8.3.1: Scientific Payload Risk Mitigation – Payload Fairing
Table 9.7: Risk Mitigation Table – Operations
Many risks are present during the launch of our vehicle. To avoid any risks to the C5
mission, vehicle, operators, or spectators any such risks must be predetermined and accounted
for. Because of the numerous risks involved, Table 9.7 details the potential hazards specifically
related to the payload fairings during the preparation, launch, and recovery of the vehicle as well
as guidelines to address these hazards and reduce or remove their impact.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
152
Premature
charge ignition
on the ground
(1)
Premature fairing
separation,
destroyed clips and
pins, potentially
scrubbed launch.
Remote chance of
harm to attending
members.
5 2
The black powder is
stored in a safe, closed
container which can
only be interacted with
via an electronic
ignition that is
connected to the
altimeter.
1
Premature
charge ignition
on ascent (2)
Premature fairing
separation,
compromised and
uncontrolled flight.
5 2
The black powder is
stored in a safe, closed
container which can
only be interacted with
via an electronic
ignition that can only
be fired when the
altimeter has registered
apogee.
1
Black powder
fails to ignite
(3)
No fairing
separation, failure
to deploy
parachute,
uncontrolled
descent.
5 3
Two electronic matches
are rigged for ignition,
a primary and
secondary for backup.
1
PLF hinges
break (4)
PLF falls away
from rocket, is
potentially lost in
the launch field
below.
3 2
The hinges are made of
steel and placed to
reduce unnecessary
stress.
1
153
PLF is
damaged
during flight
or on landing
(5)
Destruction of
nosecone which
can lead to failure
of other
subsystems such as
the recovery
system
4 2
The PLF has been
tested to ensure it will
withstand any forces it
will encounter. PLF is
reinforced with
fiberglass and ribbing.
1
The
deployment
charge
damages the
structural
integrity of the
PLF (6)
The PLF requires
repair or
replacement,
violating a critical
mission
requirement
5 1
PLF separation testing
has been performed
confirming that the
charge does not
damage the PLF in any
way
1
The
deployment
charge
damages the
recovery
payload within
the PLF (7)
A critical mission
requirement is
compromised,
repairs or
replacements may
need to be made
before reuse
4 1
Testing has been
performed to ensure
that deployment charge
separates the PLF but
does not damage any
other components.
1
PLF
structurally
compromised
by
aerodynamic
forces in flight
(8)
Operation of PLF
is compromised,
flight of the rocket
may also be
compromised
4 2
The aerodynamic
forces have been
simulated through
testing and through
full-scale launch
testing.
1
154
Table 9.8: Risk Mitigation Table – Payload Fairing Testing
The payload fairings serve a critical role in the rocket as both a structural component and
scientific priority. We have performed extensive testing to ensure that the fairings will function
correctly. Table 9.8 catalogs these tests and any associated risks to the component being tested,
the overall system, or team members.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Three Point
Bending Test
specimen partially
shatters (1)
Sharp debris would
be left around the
testing area.
2 1
The test specimen
was manufactured
and loaded to ensure
that it would bend
and fracture but not
shatter.
1
Three Point
Bending Test
specimen
improperly loaded
(2)
Test results may
not be accurate and
may interfere with
further testing.
2 1
Team members have
been trained about
qualities of the test
specimen and
procedure prior to
conducting the test.
1
Improper choice of
Three Point
Bending test
specimen (3)
Test results will
not be accurate and
likely will need to
be repeated with a
proper specimen.
3 2
The team has
manufactured a
representative
specimen that
reflects the properties
of the rocket.
1
155
Rocket model
comes loose from
vehicle (4)
The vehicle or
rocket model could
become damaged,
and potentially the
operation of the
vehicle
compromised.
3 1
The rocket model
and its mount are
checked several
times prior to testing
1
Vehicle breaks
during testing (5)
Delays to the
testing, potentially
harm to the vehicle
operator, rocket
model, or mount.
2 1
The vehicle has been
maintained and
inspected prior to
testing to ensure
proper operation.
1
Obstacle exists in
testing area (6)
The obstacle could
cause unexpected
changes to testing,
or if impacted
during testing
could cause
complications to
the vehicle, rocket
model, or operator.
2 1
A suitable testing
area has been
determined in
advance of testing
that has no obstacles.
This is verified again
on the day of the test
and testing is
postponed if
potential obstacles
cannot be removed
1
156
PLF activation or
testing damages
recovery system (7)
Repair or
replacement of the
recovery system
will be required,
potentially
delaying further
testing
3 2
The PLF and its
deployment charge
have been tested to
ensure that they will
not affect the
recovery system
1
Charge
Deployment
Testing throws
shrapnel (8)
Shrapnel can injure
nearby testers or
damage elements
of the test
4 2
Team members must
stay a safe distance
away and a blast
shield will be utilized
for protection
1
Ignition of black
powder during
handling or setup
(9)
Injury can occur to
team members or
to elements of the
test nearby
3 3
Black powder is
always handled with
extreme caution
1
Black powder fails
to ignite during
testing (10)
Team members
must remove the
unignited black
powder, exposing
them to risk if there
is a delay in the
electric signal
2 1
If charges do not
ignite, members will
not approach test
section for at least 2
minutes. Face and
hand protection will
be worn when
dealing with live
charges
1
157
Fumes from black
powder charge
testing are inhaled
(11)
Team members
may experience
adverse health
effects of inhaling
fumes and
particles.
2 4 Testing area will be
properly ventilated 1
PLF breaks into
several pieces from
charge testing (12)
Fragments of the
PLF could cause
cuts or pierce shoes
of tester during
clean up and repair
2 4
All team members
must wear thick,
close-toed shoes and
be very observant
when approaching
the testing model. In
the case of
fragmentation,
testing area is cleared
of any fragments
immediately after it
is safe to approach.
2
Table 9.9: Risk Mitigation Table – Payload Fairing Construction
The construction of the payload fairing is performed mainly on a 3D printer which can easily be
damaged or cause injury if operated incorrectly. Table 9.9 summarizes the risks of operating the
3D printer and how we are working to prevent them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
158
Hair or items
become entangled
with 3D printer
machinery (1)
If interacted with
in close quarters,
hair, clothing, or
other items could
become entangled
with the 3D
Printer.
3 2
3D printer is
operated inside an
enclosure. Care is
exercised by
operators when
setting up prints.
Long hair is tied
back and loose items
are removed.
1
Interaction with
hot 3D printer
machinery causes
burns (2)
Interaction with the
extruder, extruded
plastic, or other
high temperature
machinery can
cause burns on
hands or body.
3 2
Team members must
wait a sufficient time
after the printing
process finishes.
Operators may not
interact with the
machinery at all
when it is in
operation.
1
Interaction during
operation results in
jammed or injured
fingers or other
appendages (3)
Interaction with the
3D printer during
operation could
easily result in
appendages
becoming caught in
machinery that
continues to
operate. This can
result in damage or
harm to these
appendages.
2 4
Operators may not
interact directly with
the 3D printer during
its operation. Any
alterations to its
process during
operation must be
made with software.
Printer must be
allowed to cool down
before touching it or
the print
1
159
3D Printer
produces fumes as
a byproduct of
construction (4)
Nearby team
members may be
adversely affected
by fumes if inhaled
in a large quantity
in a short period of
time, such as when
working or
monitoring the 3D
printer in the
immediate vicinity
of it.
2 4
Team members
monitor the 3D
printer and its
progress periodically
rather than
continuously. The
immediate area is
properly ventilated
any other work or
construction occurs
at a distance from the
3D printer at which
the fumes are
dispersed and not
dangerous.
2
Shock due to
physical alterations
to 3D Printer (5)
If a team member
contacts the inner
electronics of the
3D printer while it
is drawing power
they risk shock or
electrocution if
improperly
handled.
1 4
Team members are
not allowed to alter
or tamper with the
inner electronics of
the 3D Printer and
are only serviced by
knowledgeable
members if it is
unplugged and
properly grounded.
1
160
Fire (6)
The presence of
foreign objects
and/or poor
maintenance of
wiring and parts
could result in a
fire, damaging the
3D Printer, its
contents, and
potentially
5 2
The 3D Printer is
inspected before and
after each use and the
inside cleared of any
foreign objects.
Operators are
knowledgeable of the
location and use of
fire extinguishers and
safety procedures
prior to operation
1
In any risk considerations, environmental risks to and by the vehicle should be kept in mind as
well. Fragments of plastic from the payload fairings and the use of black powder could prove a
risk to the environment if either malfunctions. It is most likely that the PLF will fail during
ascent in which case it threatens to spread plastic over a large area. The team will be careful to
collect as much as they can, but the large area makes it difficult to be certain that it is all
recovered. Any remaining plastic will not biodegrade and could adversely affect the
environment for hundreds of years. Also, malfunctioning black powder could ignite either
before takeoff or after descent and risks provoking a fire in the immediate area. Team members
will be prepared for a fire and have access to a fire extinguisher. They will also vacate the area if
the fire spreads or threatens any team members or spectators.
Section 8.3.2: Scientific Payload Risk Mitigation – WAFLE
Table 9.10: Risk Mitigation Table – Operations
The Wall Armed Fin-Lattice Elevator (WAFLE) represents a component of dual importance as
both scientific payload and tool to shed velocity to reach the target altitude. Because of its
significance, any risks or concerns that may threaten the mission, vehicle, spectators or operators
must be predetermined and tempered or prevented. Table 9.10 catalogs the risks that could
impact the WAFLE system during preparation, launch, and recovery of the vehicle.
161
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Grid fins do not
deploy in any
capacity (1)
No data is
gathered on the
grid fins and
stability is
slightly
compromised
2 3
A full aerodynamic
analysis of the grid
fins is conducted
through
computational fluid
dynamics (CFD),
subsystem wind
tunnel testing, and in-
flight sensors. If the
Arduino detects that
apogee has occurred
but the fins have not,
the fins will be
deployed for a more
stable descent.
1
162
All 4 grid fins do
not deploy
simultaneously (2)
Significant
instability due
to unbalanced
aerodynamic
forces
5 1
A full aerodynamic
analysis of the grid
fins is conducted
through
computational fluid
dynamics (CFD),
subsystem wind
tunnel testing, and in-
flight sensors. When
the Arduino detects
that apogee has
occurred, the fins will
be deployed. Servos
will be checked for
redundancy and
proper function. All
wiring will be
checked for security
and proper
connections.
1
163
Grid fins
structurally unable
to handle
magnitude of
aerodynamic
forces (3)
Significant
instability due
to unbalanced
aerodynamic
forces and
potential for
damage and to
body and/or
fins of the
vehicle
3 2
In order to evaluate
how the grid fins will
interact once
deployed, the team
will construct visual
testing of the fluid
flow through the
lattices of the grid
fins. Therefore, a
basic lattice fin has
been designed and
implemented to act as
the primary grid fin.
The fairing contains
screw holes that allow
the fairing to be hard
mounted to the
airframe.
1
164
Grid fins deploy
prematurely (4)
Significant
stress on grid
fins which may
lead to damage
to fuselage
and/or fins,
catastrophic
failure, or
significant
instability due
to unbalanced
aerodynamic
forces
4 2
A redundant timer
will be implemented
into the system to
ensure that the code
iteration does not
engage until after
boost phase and
correct altitude. A full
aerodynamic analysis
of the grid fins is
conducted through
computational fluid
dynamics (CFD),
subsystem wind
tunnel testing, and in-
flight sensors.
1
Electronics detach
or become loose
during flight (5)
Center of
gravity will
change causing
slight to
significant
instability
which may lead
to undesirable
flight path
and/or
malfunction of
grid fins
4 2
Careful and extensive
measures will be
taken to insure all
electronics are
securely attached to a
stationary plate within
the airframe during
assembly. The plate as
well as all mounting
bolts have been
extensively tested for
security and ability to
handle all stresses.
1
165
Electronics fail to
come online after
boost (6)
No data is
gathered on the
grid fins and
stability is
slightly
compromised
2 1
During assembly, all
electronics will be
checked for proper
connectivity and
security, and tested to
insure the Arduino is
receiving power.
Testing will verify the
time delay during the
startup of the Arduino
and the security of the
startup. The
electronics will be
adhered to a
stationary plate within
the airframe. This
plate and mounting
bolds will be secured
to a stationary plate
within the rocket.
1
166
Servos lose ability
to deploy grid fins
(7)
One or more
grid fins will
fail to deploy
causing
significant
instability and
potentially
shifting the
center of
gravity
5 2
A full aerodynamic
analysis of the grid
fins is conducted
through
computational fluid
dynamics (CFD),
subsystem wind
tunnel testing, and in-
flight sensors. The
gears of the servos
have been imbedded
into the U-bracket
base of the grid fin by
means of a metal bar.
Due to the high
strength of the metal
bar and HIPS, the fin
will stay attached. If
large off axis
acceleration is
detected the fins will
disengage to a stored
position.
1
167
Malfunction with
WAFLE system (8)
Vehicle
trajectory will
be altered
resulting
undesirable
flight path and
potentially
collateral
damage and/or
loss of asset
5 1
The team will ensure
that all electronic
systems are in
working order and
backups are on hand
during system checks.
If large off axis
acceleration is
detected the fins will
disengage to a stored
position.
1
Flaws or
weaknesses in grid
fins (9)
Instability of
flight or heavy
vibration
causing
undesirable
trajectory or
debris to fall
back to the
Earth
2 1
In order to evaluate
how the grid fins will
interact once
deployed, the team
has constructed visual
testing of the fluid
flow through the
lattices of the grid
fins. A basic lattice
fin has been designed
and implemented to
act as the primary grid
fin.
1
168
Improper battery
power or voltage
(10)
Electronics will
fail. No data is
gathered on the
grid fins and
stability is
slightly
compromised
2 1
A new and correct
voltage battery will be
used and tested to
ensure all electronics
will have optimal
power and voltage,
and function properly.
Fully charged
batteries will be
stored within the
rocket before launch.
1
Improper range of
motion and angle
in servos (11)
May cause one
or more grid
fins to extend
too far or not
far enough
causing slight
instability
during flight
2 2
A full aerodynamic
analysis of the grid
fins is conducted
through
computational fluid
dynamics (CFD),
subsystem wind
tunnel testing, and in-
flight sensors. If large
off axis acceleration is
detected the fins will
disengage to a stored
position.
1
169
Servos do not
operate at same
speed (12)
May cause one
or more grid
fins to extend
at different
speeds than the
others causing
slight
instability
during flight
2 2
A full aerodynamic
analysis of the grid
fins is conducted
through
computational fluid
dynamics (CFD),
subsystem wind
tunnel testing, and in-
flight sensors. If large
off axis acceleration is
detected the fins will
disengage to a stored
position.
1
170
Table 9.11: Risk Mitigation Table – WAFLE Testing The WAFLE system’s criticality to mission success obliges diligence to ensure that it will
function properly in competition conditions. We have performed comprehensive testing to
ensure the system will function as intended during flight. Table 9.11 specifies the risks present
in this testing and instructions to mitigate or avoid suck risks.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Grid fins impacts
hard surface or
tools (1)
The grid fins are
structurally and
aerodynamically
compromised and
will need to be
remanufactured
3 2
Great care is taken in
the handling and
transportation of the
grid fins at all times
as well as
constructing,
mounting and
working on them.
Keen observation
and testing is
conducted on all
components.
1
171
Grid fins motion
improperly while
they are being
worked on by hand
(2)
Pinching of the
fingers or hand
working on the fins
and momentary
discomfort
1 2
Hand work on the
grid fins is always
short and focused.
This is to avoid
extended contact, or
injury due to
complacency. Any
personnel working
on grid fins must
have proper
knowledge of grid fin
operation to be aware
of pinch points.
1
Grid fins are
structurally
compromised by
aerodynamic loads
in wind tunnel
testing (3)
The grid fins will
need to be
remanufactured
and potentially
strengthened for
future testing.
4 1
The grid fins are be
well-manufactured
and the test well-
monitored to ensure
that the aerodynamic
loads applied are
proper for testing and
do not exceed the
grid fin’s limitations.
1
Grid fins deploy
unsymmetrically in
wind tunnel
testing(4)
Potential
movement in the
test model the grid
fins are attached to.
3 1
The test model must
be secured in such a
way that movement
or uneven forces will
be measured but
contained and will
not damage or affect
the testing area.
1
172
Unintended items,
such as screws,
bolts, or small
tools, enter the
testing area during
wind or water
tunnel tests (5)
Unpredictable
interaction and
potential harm to
the grid fins, test
model, wind/water
tunnel, or team
members.
5 2
Team members take
great precaution to
check over the
testing area prior to
installing the fins and
test model and again
before initiation of
the tests. Any tools
used are accounted
for before beginning
the tests.
2
Table 9.12: Risk Mitigation Table – WAFLE Construction
Risk is present during all phases of Project Aquila, including during the manufacture and
assembly of components and parts. Uncoordinated construction can result in a chaotic
environment conducive to accidents and contributes to risk. Table 9.12 catalogs the risks
involved in construction of the WAFLE system and protocols to reduce or eliminate hazards to
the system, vehicle, or team members.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Improper material
used in the
manufacturing of
the grid fins (1)
The grid fins are
unstable and may
compromise flight
dynamics.
1 3
Thorough research
and testing has been
done on the material
and the general
design throughout to
ensure the fins meet
the expected
requirements.
1
173
Improper material
used in the
manufacturing of
the payload fairings
(2)
The payload
fairings may cause
unexpected
instability or
complications.
2 3
Extensive research
and comprehensive
testing has taken
place to verify that
the chosen material
meets the
requirements and
functions properly.
1
Servos are
improperly
installed (3)
The grid fins may
deploy improperly
and could damage
the fins or
adversely affect
flight dynamics.
2 4
The servos are
checked after
construction and
tested thoroughly
before any launches
occur.
1
Payload fairings
are improperly
manufactured (4)
Payload fairings
will not function as
designed and could
damage other
elements of the
rocket and would
certainly affect
flight dynamics
1 4
Close attention is
given to the
manufacturing
process and
extensive testing has
confirmed that the
properties of the
manufactured part
match that of the
design.
1
174
Improper tools are
used to
manufacture parts
of the grid fins (5)
The grid fins may
not be
manufactured to
proper
specifications and
may require
additional
modification or
remanufacture.
4 1
Exact methods of
manufacture
including tools large
and small are
determined prior to
the beginning of
construction.
1
Grid fins are
improperly stored
during or after
manufacture (6)
The grid fins may
not properly set or
cure or could
become damaged
due to poor
environment or
contact, which
could result in
improper shape or
other
specifications.
3 1
The grid fins are
stored in a sizeable
space that is dry,
room temperature,
and not crowded by
tools or other
materials or project
pieces.
1
175
Table 9.13: Risk Mitigation Table – WAFLE Materials Risks to the mission, vehicle, spectators or operators are inherent to every facet of the project and
the rocket, including the choices of materials used for the final system. To minimize these risks,
excellent care was taken to select the most appropriate and safest materials and not to
compromise the integrity of the materials at any point in time. Table 9.13 catalogs the liabilities
concerning material selection and application as well as justification for the selections made.
Material Potential Effect Impact Risk1 Mitigation Risk2
Batteries (1) Insufficient
power 2 2
A new battery is being
used and tested with
an electronic multi-
tester to ensure proper
function. Fully
charged batteries are
be stored within the
rocket before launch.
1
Accelerometer (2)
Receiving false
or inaccurate
data, causing
the Arduino to
make improper
course
corrections
2 1
ADXL335 Triple-axis
Accelerometer was
chosen as the
temporary
accelerometer for the
mission and WAFLE.
Validation of the
accelerometer is being
conducted and a final
selection process will
occur.
1
176
HIPS – High
Impact
Polystyrene (3)
Structural
Failure 2 1
In order to evaluate
how the grid fins will
interact once
deployed, the team
will construct visual
testing of the fluid
flow through the
lattices of the grid
fins. Therefore, a
basic lattice fin has
been designed and
implemented to act as
the primary grid fin.
1
Aluminum (4) Structural
Failure 4 1
Material was chosen
for its light weight
and structural
integrity
1
177
Arduino Uno (5)
Electronic
failure or
undesirable
grid fin
deployment
3 2
All electronics and
computing has been
extensively tested to
ensure reliability and
redundancy of
accurate flight path
corrections. A
redundant timer will
be implemented into
the system to insure
that the code iteration
does not engage. This
pause timer will wait
until the acceleration
of the rocket is within
a safe range before
starting the Arduino
calculations.
1
Carbon Fiber (6) Structural
Failure 4 1
Material was chosen
for its light weight
and structural
integrity
1
Copper Wires (7)
Electric
connections
fail
3 2
The electronics adhere
to a stationary plate
within the airframe.
This plate and
mounting bolds are
secured to a stationary
plate within the
rocket.
1
178
Electric Servos (8)
Electric
connections
fail, servos do
not
3 2
The servo provides
enough torque to lock
the secondary object
in place in order to
counteract opposing
forces on the object.
The HS-5685MH
servo was chosen due
to the high amount of
torque provided.
1
Adhesives (Epoxy,
Flux and Soldering
Materials (9)
Copper wires
may detach. 2 1
Adhesives will be
tested with the full
aerodynamic analysis
of the grid fins.
Conducted through
computational fluid
dynamics (CFD),
subsystem wind
tunnel testing, and in-
flight sensors.
1
Environmental concerns should be considered in any mission or launch event. The WAFLE
system introduces environmental risks during the 3D printing of the grid fins and during
recovery operations.
In the former instance, 3D printing of the grid fins can result in significant amounts of waste
plastic due to inferior or damaged parts or unused redundancies. Disposal of this waste plastic
increases the accumulation of plastic in landfills and other improper locations in the
environment. Because HIPS plastic is not biodegradable, it will continue to accumulate and
never cease to impact the environment. To mitigate this impact, waste HIPS plastic has been
179
collected and stored to be recycled at the conclusion of Project Aquila or reused internally for
future projects. During descent, the WAFLE system is at risk when the rocket impacts the
ground or objects near the surface. If impact is severe, the grid fins or fairings may fracture and
leave fragments in the general area surrounding the landing site. HIPS plastic is not
biodegradable and thus will remain in place for an indefinite period of time, likely leaving a
negative impact on the ecosystem. To mitigate this, the grid fins will be automatically stowed
away once the system reaches 100 feet in altitude to minimize effects of the impact.
Furthermore, the team will immediately inspect the grid fins upon recovery and search the
surrounding area for shards or fragments and collect them.
Section 8.4: Recovery Hazard Analysis Safety is taken into consideration for every part of building the rocket. There are steps that will
be taken by the recovery team to ensure the safety of the members while they construct the
recovery system for the rocket. There are three different areas that we will look at while
considering failure modes for safety protocols for recovery: operations, materials, and
construction.
The recovery system is one of the most important systems on our rocket since it provides for the
safe descent of our airframe. Table 9.14 addresses the problems that could arise in our recovery
system and the guidelines which we are following to prevent them from occurring. Table 8.6: Risk Mitigation Table - Flight Recovery Operations
Potential
Failure Potential Effect Impact Risk1 Risk Mitigation Risk2
The
parachute(s) is
not packed
properly. (1)
The parachute does
not fully deploy
causing rocket to
fall in an
uncontrolled
manner.
5 4
Strict packing
instructions are
followed by the team
members when packing
the parachutes.
1
180
Parachute tears
(2)
The parachute
fabric material is
torn causing the
rocket to fall in an
uncontrolled
manner
5 3
Parachutes have been
tested out on all full-
scale flights. Container
in which the parachute
has been smoothed to
not contain any sharp
edges. Parachutes have
been reinforced at any
potential tear locations
1
Parachute fails
to deploy (3)
Parachute fails to
deploy causing the
rocket to fall in an
uncontrolled
manner
5 4
Ground testing and
flight testing has been
performed to ensure
that the parachute will
deploy. On the day of
launch, systems will be
checked to ensure the
parachute will deploy at
the proper time.
2
The shock
cords break
after
deployment of
parachutes. (4)
Uncontrolled
descent of the
rocket with
potential crowd
endangerment.
5 3
Shock cords have been
tested on all sub-scale
and full-scale launches
of our rocket
1
181
Winds blow
rocket off
course. (5)
Rocket could
become lost,
damaged, or could
endanger observers.
5 3
The rocket will not be
launched if weather
conditions are not
suitable. All parts of the
rocket have a GPS
locater device securely
attached.
1
The parachute
deploys at the
incorrect time.
(6)
Structural damage
to rocket causing
unsafe descent or
location of descent
potentially
endangering
observers.
5 4
Recovery system has
been tested on our full-
scale launches to ensure
that it will deploy at the
correct time. Altimeter
data is checked before
each launch to ensure
that it is responding
correctly
2
The altimeter
fails. (7)
The parachute
deploys at incorrect
time or not at all
resulting in
structural damage
or uncontrolled
descent. Potentially
endangering
observers.
5 3
Our rocket has a backup
altimeter in the case
that one fails. All
altimeter data is
checked prelaunch to
determine if they are
responding correctly
1
182
The drogue
parachute fails
to deploy. (8)
Uncontrolled
descent until main
parachute opening
then resulting in
structural damage
with potential
endangerment of
observers.
5 4
Drogue deployment
systems have been
tested on our full-scale
launch to ensure that
they respond correctly
2
Table 9.15: Risk Mitigation Table - Tensile Test Rig
The tensile test rig is a large piece of equipment that produces large forces. As a result, it can be
dangerous if not operated properly. Table 9.15 summarizes possible risks and the guidelines that
we follow to avoid them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Object being tested
is improperly
aligned (1)
Results acquired
from tests are
incorrect and result
in a weaker rocket
in the future
4 4
Equipment must be
supervised by a
trained member of
the faculty at all
times while in
operation
1
Fractured particles
during test (2)
Irritation to eyes or
injury from dust or
high speed particles
4 4
All personnel must
stay a safe distance
away from tensile
test rig while in
operation. Goggles
are required while
equipment is running
1
183
Heavy weights and
high forces
generated (3)
Body damage,
specifically crushed
body extremities,
from misuse of
machine while
testing
5 2
While machine is in
operation, people
may not approach
within five feet of the
machine, marked by
tape on the floor
1
Unauthorized use
(4)
Damage to
machine,
personnel, and
projects
5 2
Machine is kept
powered off in a
locked lab when not
in use
1
Improper testing
material (5)
Unneeded use of
machine, possible
damage to
machine, and waste
of material
3 3
All workers must
check with
authorized personnel
before testing
materials
1
Table 8.7: Risk Mitigation Tables - Shear Pin Test Rig
The shear pin test rig is useful in testing our equipment but can lead to injury if operated
incorrectly. Table 9.16 summarizes the risks involved in operating the shear pin test rig and how
we are addressing them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Shear pin being
tested is improperly
aligned (1)
Results acquired
from tests are
incorrect and have
a damaging effect
on the rocket in the
future
4 4
Only trained
operators may use the
shear pin test rig.
Shear pin alignment
is always by operator
checked immediately
before testing
1
184
Fractured particles
during test (2)
Damage to eyes
and body
extremities when
the item being
tested fractures
4 4
All personnel must
stay a safe distance
away from tensile
test rig while
performing test.
Safety eyewear must
also be worn along
with proper clothing
covering body
extremities
1
Heavy weights and
high forces
generated (3)
Body damage,
specifically crushed
body extremities,
from misuse of
machine while
testing
5 2
All personnel are
kept at a safe distance
whenever equipment
is in operation
1
Unauthorized use
(4)
Damage to
machine,
personnel, and
shear pin
5 2
Shear pin test rig is
locked up and
powered off when
not in use
1
Improper testing
material (5)
Unneeded use of
machine, possible
damage to
machine, and waste
of material
3 3
Authorized personnel
must be present
during operation to
ensure that proper
procedures are
followed
1
Section 8.4.1: Recovery Risk Mitigation – Materials
185
Table 9.17: Risk Mitigation Tables - Nylon
Nylon is a strong material which we are using as shock cord for our parachutes and recovery
system. This material can cause minor injury and irritation if not handled with proper care. Table
9.17 summarizes the problems working with nylon can cause and how we will prevent these
problems from occurring.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Breathing in fiber
dust (1)
Respiratory
Problems 4 4
Respirators are
required when cutting
Nylon
1
Fiber dust in eyes
and on skin (2)
Can cause irritation
to eyes and skin 3 4
Eye protection is
required in lab when
people are working
with Nylon. If dust
gets in eyes rinse out
immediately with
water
1
Nylon catches fire
(3)
Nylon will melt
and cause severe
burns if it comes
into contact with
skin
5 3
Nylon is kept away
from sparks and open
flames. A fire
extinguisher is
always easily
accessible in lab. If
skin is exposed to hot
nylon submerge area
in cold running water
and immediately seek
medical attention
1
186
Table 9.18: Risk Mitigation Tables - Carbon Dioxide
The carbon dioxide ejection system allows us to reduce the amount of black powder that is
necessary on our rocket. However, working with this system presents its own risks. Table 9.18
summarizes these risks and how we are working to prevent them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Improper
Ventilation (1)
CO2 gas can cause
headaches, nausea,
and loss of
consciousness in
high doses
5 4
Lab is ventilated at
all times. If working
with large amounts of
CO2 the tests are
performed outside.
1
Explosion of
canisters containing
CO2 (2)
Canister shrapnel
can cause serious
cuts to the body
5 3
Cylinders are stored
upright in a proper
storage device, in a
well-ventilated and
secure area, protected
from the weather.
Storage area
temperatures never
exceed 100 °F
1
Broken O-Ring (3) CO2 can leak into
the surrounding air 3 4
O-rings are inspected
before every use and
all faulty O-rings are
replaced immediately
1
Over pressurizing
rocket (4)
Over pressurization
can cause problems
with deployment of
the parachute and
damage the rocket
4 3
Tests have been
performed to confirm
the amount of CO2
that is needed to
pressurize the rocket
to cause separation.
1
187
Under pressurizing
rocket (5)
The parachute
doesn’t come out at
all resulting in the
rocket becoming a
high speed
projectile
4 4
Ground tests have
been performed
before full scale use.
System will be
checked beforehand
to ensure that it will
function correctly
1
Table 9.19: Risk Mitigation Table - Black Powder
Black powder is necessary in the use of our recovery system; however, it can be extremely
dangerous if not handled with care. Table 9.19 summarizes the dangers of working with black
powder and the rules that have put into place to prevent injury form its use.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Improper
Ventilation (1)
Black powder is
hazardous to the
respiratory system
when inhaled. Also
particles may form
explosive mixtures
in air.
5 4
Lab is ventilated at
all times. Ventilation
masks are required
when working with
black powder
1
188
Powder comes into
contact with the
body (2)
Can irritate skin
and eyes 4 3
Proper PPE is
required when
working with black
powder. In the case
of skin contact,
personnel will wash
area with soap and
water. In the case of
eye contact,
personnel will flush
large amounts of
water into eyes and
seek immediate
medical attention
1
Highly Reactive
Substance (3)
Can cause fires
resulting in human
injury or
destruction of
equipment, and in
large amounts it
can cause
explosions causing
injuries due to heat
or flying shrapnel
5 4
Black powder is
stored in a marked
container and kept
away from heat,
sparks, and open
flames. Extreme care
is required while
using black powder
in order to avoid
impact or friction.
Fire extinguisher is
easily accessible at
all times.
1
189
Improper storage
(4)
Degrades material
and possible
combustion
resulting in injuries
and loss of
equipment
5 4
Black powder is
stored between 40°F
to 120°F in a cool dry
place in a tightly
sealed container. It is
also stored separate
from all the other
flammables
1
Improper
measuring of black
powder for rocket
use (5)
If measured
amount is too
small, the
parachute will not
eject resulting in
the rocket
becoming a high
speed projectile
5 4
Testing has been
carried out to confirm
that calculations for
the amount of
powder needed are
correct
1
Table 9.20: Risk Mitigation Table – Fiberglass
Fiberglass is a durable material that we are using to reinforce our airframe. Working with
fiberglass can lead to skin irritation and injury if the proper safety equipment is not used when
handling it. Table 9.20 summarizes potential risks associated with working with fiberglass and
what we are doing to minimize them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Ventilation issues
(1)
Can cause
respiratory
problems
4 4
Lab is properly
ventilated and
respirators are
required when
working with
fiberglass
190
Eye and Skin
contact (2)
Can cause irritation
with skin and eyes 3 5
Proper clothing and
eye protection is
required when
working with
fiberglass
Section 8.4.2: Recovery Risk Mitigation - Construction
Table 9.21: Risk Mitigation Table - Orbital Sander
The orbital sander has fast moving parts and can cause moderate to severe injury if operated
incorrectly. Table 9.21 presents the risks of operating the orbital sander and what guidelines have
been put into place to ensure safe operation.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Injuries to hands
and fingers from
moving parts (1)
Injury or loss of
extremities 5 4
Gloves are required
to operate sander.
Sander will be
powered off if not in
use
1
Eye Damage (2)
Wood chip, metal
particles, or other
debris hitting eyes
and damaging them
5 4
Safety glasses are
required when
operating the orbital
sander
1
Electric Shock (3) Electrocution 5 3
Sander is stored and
operated in a dry lab
and inspected
regularly to ensure
that there is no
exposed wiring
1
191
Unintentional
Starting (4)
Damage to
equipment,
projects, or bodily
harm
5 4
Before moving
sander will be turned
off and unplugged
Operator must check
that switch is in the
off position before
connecting it to a
power source
1
Improper Tool
Storage (5)
Misuse of tool by
unauthorized
personnel or
damage to
equipment due to
environment
5 3
Sander is stored in
dry lab which is
locked at all times
1
Hazardous Work
Environment (6)
Damage to body,
work area, or
project from debris
in work area
5 4
Sander and work area
must be cleaned
before and after
every use of the
orbital sander
1
Improper Work
Attire (7) Damage to body 5 4
Proper clothing and
PPE are required to
operate the orbital
sander
1
Dust, carbon fiber
and metal shards,
and air quality (8)
Damage to throat
and lungs 5 5
Respirators are
required for everyone
in lab while using the
orbital sander on
hazardous materials
1
192
Project is not
secured down (9)
Damage to project
and damage to
hands from high
speed objects
4 3
Operator must check
that project is
securely clamped
down before turning
on sander
1
Over-reaching (10) Severe cuts to
body 5 2
Operators must not
do any other task
while sanding.
Operator must be
aware of moving
parts
1
Improper Tool
Maintenance (11)
Dull or ineffective
tool that causes
unsafe handling
and damage to
body or project
5 3
Sander is cleaned
before and after each
use. Sand paper is
replaced periodically.
1
Over Exerting Tool
(12)
Causes damage to
project due to
excessive force
applied to tool
3 3
Personnel must be
trained to operate
orbital sander
1
Improper Tool
Replacement Parts
(13)
Tool becomes
unusable 3 3
Only replacement
parts intended for the
orbital sander will be
used to fix it
1
Table 9.22: Risk Mitigation Table - Sewing Machine
The sewing machine is a necessary piece of equipment since we have elected to create our own
parachutes. However, improper operation can result in minor to moderate injuries. Table 9.22
193
summarizes the risks that are present when operating the sewing machine and how we are
addressing them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Sewing over fingers
(1)
Hurting fingers and
causing irreparable
damage to the
equipment
4 3
Operators must be
trained before using
machine. Operator
must be aware of
machine at all times
while in use
1
Pin misuse (2)
Damage to body
from the pins and
damage to project
3 3
Proper training is
required to use the
sewing materials
1
Improper machine
use (3)
Inexperienced
personnel can
damage material
and damage self
5 3
Personnel must be
trained before they
can use sewing
machine
1
Cord can fray (4) Can cause a fire 5 3
Any broken or worn
out parts will be
placed immediately.
Operator must
inspect machine
before using.
1
Cord can be a
tripping hazard (5)
Can cause people
to trip and injure
themselves
3 3
Machine must be
plugged in close to
wall so that the
chord is not extended
over any walkways
1
194
Table 9.23: Risk Mitigation Table - Hand Tools
Hand tools such as utility knives, scalpels, and screwdrivers are necessary in the creation of the
airframe and the assembly of the rocket. Table 9.23 summarizes risks associated with them and
how we are handling these risks.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Improper use (1)
Irreparable bodily
harm can occur.
Damage to project
5 4
Tools may only be
operated by
authorized personnel.
Team leads will
advise to make sure
the hand tool in use
is appropriate for the
specific project job
1
Body damage from
tools (2)
Severe cuts and
tetanus can
possibly infect
wound
5 4
Proper clothing must
be worn at all times
to prevent damage to
body. If damage does
occur clean wound
and provide first aid.
Visit a doctor if
wound doesn’t heal
properly and
infection is seen
1
195
Improper tool
maintenance (3)
Damage to project
or body from tools
breaking or not
working as
designed
5 4
All tools have
maintenance
regularly. Any tools
deemed beyond
repair are disposed of
and replaced
immediately.
1
Flying Debris (4)
Debris may cause
eye and/or bodily
damage
5 3
Proper clothing and
eye protection are
required to operate
tools.
1
Insecure
workbench or
project (5)
Materials or tools
slip and can cause
injury to operator
5 4
Project must be
secured properly by
straps, clamps, or
through help by a
work partner before
any hand tool use.
1
Improper tool
storage (6)
Damage to tools
and potential for
unauthorized use
5 4
All hand tools have a
designated place to
be stored. All tools
will be kept under
lock
1
Section 8.5: Outreach Hazard Analysis Safety is the primary concern in every aspect of the AUSL rocket program, especially when
young children are involved. There are steps that will be taken during the outreach program to
ensure safety to the children in the community and will allow them the most amount of
enjoyment while learning about rockets. The three primary safety concerns are: Operations,
Construction, and Materials.
196
Operations Failure Modes:
• Transportation to outreach site
o Car accident
• Introduction/help students design their rockets
o Children jam fingers
o Children hurt by tools
• Multiple rocket launchings
• Rocket stands fall
• Rockets have mid-air collisions
• Rockets land in the woods
Construction:
• Tools for rocket kits
o Children incapable of using tools
• Model rocket motor
o Children accidentally ignite motor during time other than directed
Materials:
• Model rocket kits
o Children break rocket model
o Hard pieces may hurt children
Table 8.8: Risk Mitigation Table - Outreach Operations
During our outreach programs potential risks to people and the environment will always exist.
Table 9.24 summarizes the most common risks and what we are doing to prevent them from
occurring.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
197
Car Accident (1) Ranges from minor
injuries to death 5 3
All participants are
required to wear
seatbelts and only
licensed drivers may
operate motor
vehicles. Anyone
being transported by
team members must
sign waivers
releasing the team
from liability in the
event of an accident.
1
Children jam
fingers (2)
Children experience
minor pain 2 2
USLI team
demonstrates how to
perform all tasks for
rocket completion
and help the children
when needed. All
minors are supervised
at all times.
1
Children
accidentally hurt by
tools (3)
Children could
experience trauma
to numerous body
areas
3 2
All tools that could
prove dangerous to
children are operated
by USLI team
members while
wearing necessary
protective equipment.
1
198
Mid-air rocket
collisions (4)
Rockets would not
reach highest
altitude due to mid-
air collision
1 2
Students’ rockets are
launched from
significant distances
from each other.
Rockets are launched
one at a time
1
Rocket stands fall
(5)
Failure of rocket
launch 2 2
All equipment is
examined prior to
departing for the
outreach event. Any
non-functioning
equipment must be
fixed or replaced.
1
Rockets fall in the
woods (6)
Slight
environmental
contamination.
2 2
All rockets are not
designed to achieve
significant distance
and all must be
recovered.
1
Table 8.9: Risk Mitigation Table - Outreach Construction
When working with children around tools at our outreach events, the potential for them to harm
themselves always exists. Table 9.25 common risks are evaluated along with how we are
working to prevent them from occurring.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
199
Children ignite
motor at time other
than directed (1)
Trauma to hands,
eyes, ears, nose, 5 2
Children must be
under constant
supervision and any
potentially
dangerous materials
are handled by the
USLI outreach team
1
Children incapable
of using tools (2)
Danger to child,
and other
children’s face,
hands, and body
3 2
Children are under
constant supervision
and any potentially
dangerous use of
tools will result in
apprehension of the
tool. The task will
then be completed by
the USLI outreach
team for the child
1
Table 9.26: Risk Mitigation Table - Outreach Materials
Rockets materials can be dangerous if not treated with proper care. Table 9.26 details the risks
involved and what we are doing to prevent any injury from their use.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Children Break
Rocket model (1)
Student will not be
able to launch a
rocket or participate
in the primary
outreach activity
2 2
Students are under
constant supervision
and any misbehavior
will be handled
appropriately
1
200
Hard pieces may
hurt children (2)
Trauma to children
hands, eyes, nose,
mouth, ears
2 2
Students are under
constant supervision
and any misbehavior
will be handled
appropriately
1
Section 8.6: Environmental Effects Section 8.6.1: Vehicle Effects on Environment
Rockets have many diverse effects on the environment both in their operation and their
construction. The most significant environmental effects that will be part of Auburn University’s
“Project Aquila” will result from use of epoxy, carbon fiber, carbon dioxide, and 3D-printed
HIPS plastic. During construction, the use and curing of epoxy releases volatile organic
compounds along with other unhealthy gases and chemicals. Furthermore, additional unused but
cured epoxy is common after construction. Waste epoxy is contained in epoxy cups that are
thrown away and placed in landfills where they add to large amounts of non-biodegradable trash
and leak hazardous chemicals into ground around and below the landfill site. Additionally during
construction the carbon fiber, when machined, releases tiny dust particles into the air that are
extremely small and are difficult to filter out of the air. People that breathe in this dust could
experience lung, eye, and skin irritation. Also, carbon dioxide is a dangerous gas for humans
breathe and could displace oxygen in the lungs resulting in symptoms of hypoxia. Construction
will also feature the use of a 3D printer, which are capable of producing ultrafine particles during
the printing process which can settle in the lungs or the bloodstream and cause adverse effects.
Furthermore, the material used for 3D printed products will be high-impact polystyrene (HIPS)
plastic, which is non-biodegradable and rarely recycled. It is common for extra 3D printed parts
to be manufactured for redundancy, demonstration, or testing purposes, and thus some waste
HIPS is to be expected.
During rocket launch, when the rocket motor is ignited, exhaust from the motor will burn
anything immediately near the exhaust. This could potentially set fire to the fields where the
rocket will be launched or the surroundings where it will land. The ignition also releases
201
additional carbon dioxide and hydrogen chloride, which can cause internal and external irritation
to anyone that comes in contact with it. Table 9.27: Risk Mitigation Table – Environmental Effects
The construction and operation of our rocket can have potentially harmful effects to the
environment. Table 9.27 presents these possible risks, their effects, and what we are doing to
mitigate them.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Harmful toxic
fumes released into
environment by
autoclave (1)
Damage to
environment and
toxic air supply
4 3
Autoclave and lab is
properly ventilated at
all times
1
Epoxy not properly
disposed (2)
Potential fire
hazard and damage
to lab
2 3
All wasted epoxy
will be cured and
allowed to cool
before disposal
1
3D printer
malfunctions
creating waste
material (3)
Material is non-
biodegradable so it
must be disposed
of in a landfill
2 5
3D prints are
supervised by trained
operators to prevent
printing areas and the
creation of waste
material
1
Airframe not
recovered after
launch (4)
Hazard to the
environment from
carbon fiber and
epoxy
3 2
Airframe is tracked
using a GPS during
the launch to ensure
that it will not be lost
1
202
Electronics and
battery cells are
lost during launch
(5)
Battery acid is
corrosive and can
damage the
environment
4 3
Airframe is tracked
using a GPS during
the launch. In the
case of an accident,
the launch area will
be searched to ensure
that all electronics
are recovered
1
Rocket motor sets
grass on fire (6)
Damage to grass
and potential to
spread to damage
more property
3 4
Fire extinguishers are
kept near launch sites
to extinguish any
grass fires
immediately
1
Rocket lands in
major waterway (7)
Contamination of
water from the
motor ash and the
composites used to
build the rocket
3 2
Rockets are not
launched near major
waterways such as
streams or rivers
1
Section 8.6.2: Environmental Effects on the Vehicle
The environment can also effect the integrity and flight of the rocket, most significantly through
humidity, wind current, thermal fluctuations, and visibility. Exposure to humidity can cause
corrosion in the different metals and materials used in the structure as well as damage on-board
electronics and launch-electronics. Wind currents are both a danger during transport, on the
launch-pad before launch, and most critically during flight where wind can cause recovery to
become unpredictable and extremely difficult to track. This can also cause additional problems if
the rocket lands somewhere particularly hazardous or vulnerable. Additionally, thermal
fluctuations can cause different materials to behave differently than intended, flex and become
structurally deficient, or damage relevant electronics or cause thermal noise to occur in the
electronics. Visibility is also a concern during launch and operation. The launch of a rocket in
203
the midst of mild fog or low-hanging clouds can result in the rocket becoming difficult to track
or lost altogether. Table 9.28: Risk Mitigation Table – Environmental Effects on Rocket
The environment can have harmful effects on the operation and integrity of our rocket. Table
9.28 presents these effects and what we have done to minimize their effect.
Potential Risk Potential Effect Impact Risk1 Mitigation Risk2
Rocket materials
corrode if stored
improperly (1)
Rocket is no longer
useable 5 2
Rocket will be stored
in a dry lab or in its
custom made
transportation boxes.
In the case of rain, all
launch will be
postponed
1
High humidity can
damage electronics
causing them to
malfunction (2)
Electrical systems
fail and the
parachutes do not
deploy resulting in
our rocket
becoming a high
speed projectile
that could injure
onlookers or
personnel
5 2
Rocket will be stored
in a dry lab and all
systems must be
checked before
launch to ensure that
they are working
properly
1
High wind causes
rocket to drift a
long way (3)
Rocket could
damage buildings
if it lands on them
or may be lost
3 3
Launch is delayed or
postponed in the case
of high winds
1
204
Low cloud ceiling
(4)
Rocket may be lost
and unrecoverable 3 2
Launch is delayed or
postponed in the case
of a low cloud
ceiling
205
Section 9: Launch Operations Procedures
Section 9.1: Parts Checklists Grid Fins Parts Check
Initial Check-off Points
Lantern
2 Bulk plates
1 Arduino/Breadboard sled
14 5/16 nuts
2 Angle Brackets
1 U-Bolt (with washer)
6 #10-32 bolts
10 #10-32 nuts
7 #4-40 bolts
9 #4-40 nuts
Arduino
Breadboard
1 10 DOF IMU sensor
Servos
Servo platform
1 Bulk plates
3 Angle Brackets
5 #10-32 bolts
5 #10-32 nuts
9 #10-32 bolts
5 #10-32 nuts
206
1 Key switches
1 Key switch key
1 Battery
Payload Fairing
Initial Check-off Points
Male half of PLF
Female half of PLF
Shoulder coupler assembly
2 Hinges
8 Hinge bolts
8 Hinge nuts
Black Powder
2 Ematches
6 Shoulder attachment bolts
6 Shoulder attachment nuts
Wax
Small Drill Bit (for extracting pins)
Spray Paint
Masking Tape
Scissors
Gloves
Eye Protection
Paint Brushes
Sand Paper
Cups
Wooden Compressors
Masks
Pick
207
File
Screw Driver
Avionics Parts Checklist
Initial Check-off Points
1 Altimeter Bay
1 Altus Telemega
1 Perfectflite Stratologger
2 Standoff sets
2 Key switches
1 Altimeter board
10 ematches
1 Patch antennae
Slide avionics board into avionics bay
Make sure ejection charge wires protrude from proper ends of BAE
Section 9.2: Final Assembly Checklists General setup
Initial Check-off Points
Set up tables.
Unload all boxes and equipment
Unpack booster section
Inspect fins and tube for damage caused during travel
208
Grid Fins Section
Initial Check-off Points
Remove bolts for grid fin section
Inspect grid fins
Bolt grid fins into booster section
Inspect grid fin electronic section
Check that all wiring is hooked up correctly and that no wires are loose
Secure grid fin electronic section in booster section
Turn key to power on and make sure all grid fins respond
Remove grid fin electronic section from booster section
Remove SD card and check data to ensure sensors are recording correctly
Replace SD card and secure grid fin electronic section in booster section
Tighten bolts and ensure that all components are secure is secure
Place booster section on cradle
Payload Fairing
Initial Check-off Points
Inspect PLF for any cracks or damage that may have been caused during travel
Prepare e-matches and insert e-matches into charge chamber
Pack charge chamber with clay to ensure no space is wasted in charge chamber
Fill charge chamber with black powder charge
Close the PLF and make sure it fits together properly
Insert shear pins on inside of PLF once it is closed and sealed properly
Make sure all pins are still intact and in position
Seal edges with wax to ensure that no air will compromise the PLF
Custom Carbon Dioxide Ejection System
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Initial Check-off Points
Insert e-match heads into three charge cups and seal with epoxy
When epoxy is dry, pour 0.15 grams of black powder into each charge cup
Lubricate outer surfaces of lift pistons
Slide lift pistons and charge cups into charge containment cylinders
Insert charge containment cylinders into charge half of system casing taking care to
run e-match wires through wire holes
Place one CO2 cartridge into each cylinder
Into each chamber of pin side of system casing, place one pin base, one spring, and
one alignment collar
Assemble two halves of casing together making sure that all chambers are aligned
properly
Secure halves together with three bolts
Attach system onto the BAE bottom bulk plate using two U-bolts
Run e-match wires through the wire hole in bottom bulk plate
Attach e-match wires to proper connections
Attach bottom bulk plate to the BAE
Rocket Assembly and Parachute Packing (Carbon Dioxide Ejection System)
Initial Check-off Points
Remove upper sections from shipping boxes
Inspect for any damage caused by transportation
Insert ballast tank into upper section above avionics bay
Insert screws and inspect to ensure it is secure and does not move
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Secure lower part of upper section to booster section with two #2 shear pins
Fold booster main parachute
Attach booster main parachute quick link to booster section U-bolt
Pack booster main parachute into coupler to booster section
Slide the BAE, with the CO2 system attached, into lower part of upper section and
secure with four bolts
Run wires for fairing deployment charges through ballast tank
Slide upper section onto the BAE and attach with four bolts
Run two e-matches through hole in the base of the Tender Descender
Fill Tender Descender with one gram of black powder
Lock two halves of Tender Descender together ensuring that both quick links are in
their proper position
Fold main parachute and pack into deployment bag, running slip hole shock cord
through the bag
Assemble drogue parachute, main parachute, and Tender Descender with shock cord
in proper configuration
Add small handful BARF into top of payload fairing
Pack parachute configuration into payload fairing
Attach bottom of parachute configuration quick link to U-bolt on top of the ballast
tank
Slide payload faring into upper section and secure with four bolts
Motor construction
Insert motor into rocket
Secure motor in place by screwing on motor retention
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**In the case of the use of a black powder ejection system follow these steps after the booster section assembly instead of the CO2 assembly
Initial Check-off Points
Add handful of BARF (Black powder Assurance Recovery Fiber)
Run e-match wires through holes in base of two ejection charge cups
Fill charge cups with six grams of black powder each
Seal charge cups with tape
Connect e-match wires to main ejection charge wires on bottom of the BAE
Attach charge cups to bottom bulk plate of the BAE using tape
Slide the BAE into lower part of upper section and secure with four bolts
AUSL Safety Officer Signature
AUSL President Signature
X X
Section 9.3: Motor Preparation Motor Preparation Check:
Initial Check-off Points
Unpack Loki motor case
Locate clip ring pliers
Undo Clip ring from upper and lower end of motor case
Remove top motor enclosure
Remove nozzle from bottom with nozzle retention ring
Remove old O-rings
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Grease new O-rings and insert into O-ring slots on top and bottom enclosure
Insert composite grain sleeve into motor case
Insert the three motor grains
Undo smoke grain slot snap ring
Insert smoke grain
Replace smoke grain slot snap ring
Insert top motor enclosure into top of case
Replace top enclosure snap ring
Insert motor nozzle and motor nozzle retention ring
Replace bottom enclosure snap ring
Ensure bolt on top motor enclosure is tightly fastened
Motor preparation is complete
AUSL Safety Officer Signature
AUSL President Signature
X X
Section 9.4: Setup for Launch/ Igniter Installation Setup on Launcher and Igniter Installation Check:
Initial Check-off Points
Set the launch box to safe before approaching launch rail
Inspect launch rail for any issues
Lower rail to height for safe rocket insertion
Insert launch lugs on rocket into launch rail
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Lower till completely inserted into launch rail
Raise launch rail to launch position
All launch procedures occur now
All members besides level 2 certified personnel leave launch pad
Ignitor has ends stripped
Ignitor is fully inserted into motor
Ignitor leads are connected to launch controller
Ignitor connection is tested with secondary launch controller
Arm secondary launch controller
Member retreats to safe launch zone
AUSL Safety Officer Signature
AUSL President Signature
X X
Section 9.5: Launch Procedures Final Construction Check:
Initial Check-off Points
Inspect entire outer tube for cracks
Check fins for cracks or any movement
Ensure that launch lugs are secure
Inspect all joints for wiggling
Final Overall Systems Check:
Initial Check-off Points
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Inspect that PLF is secure
Test that grid fins respond
Test that avionics respond with correct beeps and wireless connection
Inspect launch rail for imperfections that could cause problems at launch
AUSL Safety Officer Signature
AUSL President Signature
X X
Launch Procedures Check:
Initial Check-off Points
Mount rocket on launch rail
All team members except safety officer, grid fins lead, and recovery lead, and a level
2 certified member move back to a safe distance
Test avionics and ensure tracking connection is enabled and all altimeters are
responding
Power on grid fins and ensure that they respond
Everyone but a certified level 2 personnel move back to a safe distance
Certified personnel inserts ignitor into rocket engine (See Setup on Launcher and
Ignitor Installation checklist)
Everyone retreats to a safe distance
Ensure launch area is clear of personnel
Check with range RSO to ensure range is all clear and ready for launch
Receive final all clear from RSO
Initiate motor ignition
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AUSL Safety Officer Signature
AUSL President Signature
X X
Section 9.6: Troubleshooting mall cracks in material
Initial Check-off Points
Replace if spare parts are available
Epoxy if part is not a load bearing section
If a major crack is found in airframe, postpone launch until replacement of part
Rocket coupler does not fit well
Initial Check-off Points
If too large, sand down coupler until it fits
If too small, add layers of tape until it fits snugly with no wiggling
PLF does not fit well
Initial Check-off Points
If too large, sand down coupler
If too small, add layers of tape to ensure a tight fit
PLF sections do not align properly
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Initial Check-off Points
Check that hinges are attached correctly
If problem remains, loosen hinge screws, align halves and retighten screws
Grid fins do not respond
Initial Check-off Points
Check that battery connection is not loose
Check that battery is not dead
Check wiring to switch
Check moving parts for anything blocking them
Altimeters do not respond
Initial Check-off Points
Check wiring to batteries
Check that battery is not dead
Check wiring to switch
Check for any crossed wires
If problem continues, replace altimeter
Launch lugs are not secured well
Initial Check-off Points
Screw launch lugs in farther
If launch lugs are still not secure, drill new holes and screw them in there
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AUSL Safety Officer Signature
AUSL President Signature
X X
Section 9.7: Post-Flight Inspection Post Flight Inspection Check:
Initial Check-off Points
Wait until the rocket is on the ground in a safe location
If the rocket is not in a safe location, or if it is out of reach, have the correct personnel
retrieve the rocket
Safety officer will approach the vehicle first to ensure it is safe to retrieve
Retrieve rocket and take back for the RSO and team leads to inspect
Safety officer removes motor case as soon as case is cool enough to handle
Safety officer inspects motor case and immediately cleans case
Airframe team lead inspects all components of the airframe (i.e. fins, body tubes,
couplers, main structure)
Grid fins team lead inspects all components of grid fins
Payload Fairing team lead inspects all components of payload fairing
Recovery team lead inspects all components of recovery section
Safety officer does final all over inspection
Receive all good from RSO
AUSL Safety Officer Signature
AUSL President Signature
X X
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Section 10: Project Plan
Section 10.1: Budget The budgets displayed in Table 10.1 are an initial approximation of the expenditures required for
the overall project. Out final cost for the rocket on the pad is $4,198, putting us well below the
maximum budget of $7,500 outlined in requirement 1.14 of the NASA Student Launch
Handbook. Table 10.1: Final Budget
Vehicle
Cost Per Unit Unit Number Total
Isogrid Tubes (carbon fiber and resin) $284 per tube 2 $568 Fiberglass Sleeve for Isogrid $33 per tube 2 $66 dry weave fiberglass $7 per yard 20 $144 fiberglass resin $15 per unit 4 $60 Loki L1482 Motor $145 per unit 1 $145 RMS 75/3840 Motor Case and Associated Hardware $385 per unit 1 $385 Pre-preg carbon fiber $118 per yard 10 $1,180 Rail Buttons $3 per unit 2 $6 Misc Hardware $100 per unit 1 $100
Vehicle Total $2,654
Recovery Ripstop Nylon $8 per yard 25 $200 Nylon thread $8 per spool 3 $24 Tubular Kevlar $1 per foot 50 $50 Paracord $5 per roll 1 $5 Telemetrum $200 per unit 1 $200 Telemega $300 per unit 1 $300 Tender Descender $85 per unit 1 $85 CO2 $180 per unit 1 $180
Recovery Total $1,044
Payloads HIPPS $30 per roll 4 $120 Servos $90 per unit 4 $360
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Arduino $20 per unit 1 $20
Payloads Total $500
GRAND TOTAL ROCKET ON THE PAD $4,198
Section 10.2: Funding Plan The team has secured funding from the sources presented in Table 10.2. This money covers 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. Table 10.2: Funding Sources
Source Amount
Alabama Space Consortium $13,000
Auburn University Organization Board $5,000
Auburn University College of Engineering $5,000
Total Funding $23,000
Section 10.3: Timeline The timeline is organized around completion of testing and manufacturing of the payloads before
their respective full scale tests. After the full scale launches, the team’s priority will be writing
FRR and created a final, polished competition presentation.
The team followed this schedule quite closely but had to completely rebuild the rocket multiple
times due to motor CATOs. As a result of these failures, the team is now scheduled to launch a
final full scale on April 2, 2016 at Georgia Rockets In the Sky (GRITS).
The full GANTT chart for the competition does not translate well to documentation due to its
size; therefore the events on the GAANT chart were subdivided to provide clarity.
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Table 10.3 shows the overall schedule of the rocket build superimposed with launch dates
available to the Auburn team.
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Table 10.3: Launches and Vehicle Timeline
Table 10.4 outlines the basic timeline of the payload subsystems and the recovery subsystem..
This allows plenty of time for the full scale test of each system and the eventual compilation of
all systems. Table 10.4: Subsystem Timeline
Table 10.3 shows the competition milestones set forth by NASA in the 2015-2016 NASA
Student Launch Handbook. This timeline also shows the team’s timeline for completing the FRR
milestone and the team’s preparation for traveling to Huntsville for the competition.
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Table 10.5: Competition Timeline
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Section 11: Educational Engagement
The Auburn University Student Launch team (AUSL), along with the Department of Aerospace
Engineering at Auburn University, is entering an exciting new era of growth, influence and
leadership, as a devotion for the future advancement of aeronautical and astronautical
engineering swells throughout the department. Just as NASA and the USLI competition has
instilled the spirit of rocketry in AUSL’s team members, AUSL truly aspires to encourage
interest in STEM fields in young students throughout the state of Alabama. Statistical studies
show that more and more young people are losing interest in STEM careers every year.
There are many middle school, high school, and college students that possess talents in math and
science, and they have aspirations to pursue STEM careers in their futures. AUSL plans to use its
influence to enrich the young minds of young students in Auburn and to promote the importance
of STEM careers and aerospace interests throughout the community.
Section 11.1: Drake Middle School 7th Grade Rocket Week Event Date: March 2, 2016-March 4, 2016
This year, AUSL’s primary plans begin with its venture in engaging young students by bringing
a hands-on learning experience for the seventh grade class of J.F. Drake Middle School (DMS).
The program is entitled DMS 7th Grade Rocket Week, and the goal of the program is to instill
interests in math, science, engineering, technology and rocketry through an interactive three-day
teaching curriculum that will reach more than 600 middle school students.. In general, many
students do not know much about rocketry or any relevant interdisciplinary applications that
space exploration entails. The seventh grade science curriculum at DMS focuses on life science
for the year. Therefore, the rocketry unit curriculum will include lessons about g-forces and how
they affect the human body. Also, most students have certainly never built their own rockets. So
additionally, the students will be divided into teams of 2-3 and provided a small alpha rocket to
construct and launch under the supervision of AUSL and certified professionals. This program
was successfully implemented during the 2013-2014 school year, and the school has requested
that we return to repeat the program with the new seventh grade class. A summarized plan of
action is written below, and it detail will be added as more formal pending agreements are made
between the school and the team. Once all formal decisions are made final for the year, a fully
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detailed program handbook will be printed for the teachers and all other administration involved.
The handbook will include specific details regarding the plan of action, the launch, scheduling
outlines, procedures, worksheets, teaching materials, lesson plans, feedback forms, etc. A rough
draft plan of action, an ideal launch plan, and the learning objectives for the outreach program
are provided in the following section.
Figure 11.1: Picture from Rocket Week 2014
Section 11.1.1: Rocket Week Plan of Action
Day 1: The students will participate in an engaging in-class lesson presented by AUSL members.
The lesson will first teach the students about g-forces through a presentation and demonstration.
Secondly, students will learn how the human body reacts under stress in high and low g-force
environments via a presentation and a video. This part of the lesson will be both educational and
highly engaging. A curriculum guide will be provided for the teacher, along with all presentation
materials that are to be utilized. A worksheet will be distributed to the students for them to fill
out key concepts as they follow the lesson.
Day 2: The students will be split into teams of 2-3 and given a small alpha rocket assembly kit
and the required materials to build and decorate the rocket. The teachers will need to divide the
students into teams since the teachers can more appropriately handle their students. AUSL team
members will lead and guide the students and faculty in every step of assembly in a very
organized and well-prepared fashion. At no point will the students be given the motors for their
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rockets. AUSL team members and certified professionals will take care of this portion at the
launch event. The students and faculty will sand, glue, assemble and paint their own rockets as
AUSL team members instruct them to do so.
Day 3: All science classes will head to the P.E. field on DMS’s campus during each period
throughout the day. Students will also be informed of all safety and launch procedures for the
event when they first arrive on the field. A summary of what will take place at the launch and a
launching order will be announced on this day.
Section 11.1.2: Rocket Week Launch Day
The launch day will be held on the DMS P.E. field on the third and final day of the program.
Each period of the school day, four or five science classes will proceed to the launch field. There
will be multiple launch rails set up in sanctioned safe zones in different parts of the field,
meeting all NAR Safety Guidelines for launching model rockets. Each class will be assigned to a
launch rail, and instructions will be delivered by an AUSL member. In the order that they are
called, students will have their rockets prepped for launch by AUSL team members. One
designated 7th grade student from each team will be given a launch controller for the team’s
rocket. At the end of a cued countdown, students will fire their rockets and recover them once
the field has been cleared by the range safety officer. At the end of the period, students return to
their classrooms and continue the day.
A permission slip will require parental permissions for students to launch rockets. AUSL plans to
invite the Southeast Alabama Rocketry Association to supervise the launch site to ensure that all
aspects of the launch are safe and successful.
Additionally, AUSL plans to invite all parents, administrators, local newspaper outlets, etc. to the
event in order to celebrate and promote the students’ work at the launch event. The Auburn
community will be able to see and appreciate the results of what its young student body has
accomplished and learned. The media attention will also recognize AUSL’s goals and efforts to
inspire and communicate the importance of STEM fields, aerospace engineering, and rocketry to
both the students and the greater Auburn community, just as NASA and its Student Launch
competition has inspired AUSL to engineer a launch vehicle.
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Section 11.1.3: Rocket Week Learning Objectives
The learning objectives for the entire outreach program are outlined below:
• Students will learn about the basics about gravity and g-forces.
• Students will learn the basic fundamentals of Newton’s Laws of Motion.
• Students will learn how high and low gravity environments affect the circulatory system,
cognitive processes, and muscle performance in humans.
• Students will learn some specific terms related to rockets and Newton’s Laws of Motion.
• Students will gain an idea of what engineering is and why math and science are so
important.
• Students will learn basic values of teamwork and why communication is important.
• Through the rocket construction and launch event, students will hopefully gain a sense of
accomplishment and confidence in their abilities to work with others to complete projects
that they may have never thought they would get a chance to do.
• Finally, AUSL secretly plans to have at least one student realize that all he or she wants to
do is become a rocket scientist. Although truthfully, the team will be glad to have sparked
any and all interests in math, science, engineering and/or technology in students’ minds
throughout the experience.
Figure 11.2: A photo taken from DMS 7th Grade Rocket Week in April 2014
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Section 11.1.4: Gauging Success
Finally, AUSL will measure the success of the outreach program by utilizing brief feedback
questionnaires. The forms will ask for feedback on different aspects of the program. One form
will be made for teachers to complete. Teachers will be able to express what they liked, what
they disliked, make suggestions for improvements, etc.
Secondly, the students will be assessed by filling out a brief worksheet that will cover some basic
highlights of what they learned from the program based on the learning objectives.
Finally, AUSL will complete a group self-assessment in writing that will highlight program
aspects that were favored, successful, needed improvement, and aspects that were not favored.
AUSL will utilize all of these forms of feedback in order to learn and plan for better ways to
execute student engagement activities in the future.
Section 11.2: Samuel Ginn College of Engineering E-Day Event Date: February 26, 2016
E-Day is an annual open house event during which middle and high school students and teachers
from all over the southeast are invited to tour Auburn University’s campus and learn about the
programs and opportunities that the college of engineering offers. Students will be able to
explore all of the labs and facilities housed in the Samuel Ginn College of Engineering, which
includes the Aerospace Engineering labs and competition team project facilities. They will also
be able to speak with faculty, advisors, organizations, competition teams and Auburn student
engineers while visiting. AUSL will be participating in the event to promote STEM fields,
rocketry, and the NASA Student Launch competition. Students will be informed of AUSL’s
current activities and will learn how they can join organizations like AUSL while attending
school at Auburn. In 2014 and 2015, more than 3,000 students and teachers attended E-Day.
More than half of the attendees were exposed to the work and activities that AUSL had
completed and learned about the Auburn rocket team’s accomplishments in the NASA Student
Launch competition. We hope to see even greater success this year as interest in STEM fields
continues to grow.
Section 11.3: Boy Scouts of America: Space Exploration Badge Through AUSL, boy scouts from Boy Scouts of America can receive the Space Exploration
Badge. The Space Exploration Badge is meant to persuade young scouts to explore the mysteries
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of the universe and build rockets. The boy scouts will be led by students in AUSL who have at
minimum earned a level one rocket certification through either the Tripoli Rocketry Association
or the National Rocketry Association. The Boy Scouts of America have set guidelines as to how
the scouts can receive the Space Exploration Badge. AUSL will follow these requirements to
ensure full completion defined by the Boy Scouts of America.
Section 11.3.1: Space Exploration Merit Badge Requirements
The following are defined guidelines set by the Boy Scouts of America to receive the Space
Exploration Badge.
• Tell the purpose of space exploration and include the following:
1. Historical reasons
2. Immediate goals in terms of specific knowledge
3. Benefits related to Earth resources, technology, and new products
4. International relations and cooperation
• Design a collector's card, with a picture on the front and information on the back, about
your favorite space pioneer. Share your card and discuss four other space pioneers with
your counselor.
• Build, launch, and recover a model rocket.
1. Make a second launch to accomplish a specific objective. Launch to accomplish a
specific objective.
2. If local laws prohibit launching model rockets, do the following activity: Make a
model of a NASA rocket. Explain the functions of the parts.
3. Rocket must be built to meet the safety code of the National Association of
Rocketry.
• Identify and explain the following rocket parts: Body tube; Engine mount; Fins; Igniter;
Launch lug; Nose cone; Payload; Recovery system; Rocket engine.
• Give the history of the rocket.
• Discuss and demonstrate each of the following:
1. The law of action-reaction
2. How rocket engines work
3. How satellites stay in orbit
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4. How satellite pictures of Earth and pictures of other planets are made and
transmitted.
• Do TWO of the following:
1. Discuss with your counselor a robotic space exploration mission and a historic
crewed mission. Tell about each mission’s major discoveries, its importance, and
what was learned from it about the planets, moons, or regions of space explored.
2. Using magazine photographs, news clippings, and electronic articles (such as from
the Internet), make a scrapbook about a current planetary mission.
3. Design a robotic mission to another planet or moon that will return samples of its
surface to Earth. Name the planet or moon your spacecraft will visit. Show how
your design will cope with the conditions of the planet's or moon's environment.
• Describe the purpose, operation, and components of ONE of the following:
1. Space shuttle or any other crewed orbital vehicle, whether government-owned
(U.S. or foreign) or commercial
2. International Space Station
• Design an inhabited base located within our solar system, such as Titan, asteroids, or other
locations that humans might want to explore in person. Make drawings or a model of your
base. In your design, consider and plan for the following:
1. Source of energy
2. How it will be constructed
3. Life-support system
4. Purpose and function
• Discuss with your counselor two possible careers in space exploration that interest you.
Find out the qualifications, education, and preparation required and discuss the major
responsibilities of those positions.
• Failure, by any boy scout, to complete any of the above requirements will disqualify him
from receiving the Space Exploration Merit Badge.
Section 11.3.2: Boy Scouts of America - AUSL Requirements
In addition to the guidelines set by the Boy Scouts of America, AUSL has set requirements that
the Boy Scouts will also follow to receive the Space Exploration Badge.
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• All boy scouts will follow rules/regulations set by the NAR and TRA, just like AUSL.
• All boy scouts will follow safety guidelines set forth the by the AUSL designated safety
officer.
• Boy Scouts will not tamper with their rocket in such a way as to cause the rocket to have
instabilities or incomplete recovery.
• All Boy Scouts will complete the required lesson plan.
• Failure by any Boy Scout to complete any of the above requirements will disqualify him
from receiving the Space Exploration Merit Badge.
Section 11.3.3: Boy Scouts of America - Plan of Action
In February 2016, Boy Scouts will assemble in the Haley Center at Auburn University in the
morning to sign in for the day’s activities. AUSL members will greet the Boy Scouts and their
chaperones. The Boy Scouts will be escorted to the assigned classroom for their merit badge
activities. After lunch, AUSL members will explain the safety rules for building the rockets and
will distribute Alpha rocket kits to the Scouts. Once the scouts have completed their rockets,
everyone will travel to the designated launch site AUSL has acquired, which meets all NAR,
FAA, and Auburn City requirements. While AUSL members setup launch, a designated safety
officer will explain all launch rules and precautions associated with rocketry. Rocket launches
will then commence. All launches will take place in the presence of a registered NAR/TRA
official. After successfully completing their launches, the scouts will be presented with the Space
Exploration Merit Badge, shown in Figure 11.3.
Figure 11.3: Space Exploration Merit Badge
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Section 11.3.4: Boy Scouts of America: Goals
It is intended for every Boy Scout to receive the Space Exploration Badge. AUSL wishes for the
boy scouts to enjoy their learning experience about space and rocketry. AUSL also hopes to
inspire the scouts to pursue a career in engineering.
Section 11.4: Girl Scouts of the USA - Space Badge The Auburn Student Launch team will be conducting an event similar to the Boy Scout Space
Exploration Merit Badge for local area Girl Scout troops. The event will follow all standards and
guidelines set by Girl Scouts of the USA, NASA Student Launch, Tripoli/NAR, Auburn
University and any other relevant parties. Girl Scouts will learn the basics of rocketry and build
and launch their own rockets. Girl Scouts currently does not have a badge equivalent to the
Space Exploration merit badge, but we will be working with the involved troops to develop a
custom badge for this event.
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Section 11.5: Auburn Junior High School Engineering Day
Event Date: October 19, 2015
Auburn Junior High School hosted its first Engineering Day to spur student interest in
engineering and to create an atmosphere where students can gain firsthand experience as to what
it is like to be an engineer. All engineering majors were invited to present their major, clubs, and
teams to encourage students to become engineers. AUSL participated in the event to promote
aerospace engineering, rocketry, and most importantly, the NASA Student Launch Competition.
AUSL talked about rocketry and its components where students were also able to view and hold
some of AUSL’s rockets because for many students they have never seen a rocket or even
touched carbon fiber. AUSL presented to 1,000 students that day in the hopes that at least one
student becomes an aerospace engineer; although, we had plenty of students who said they were
very interested in aerospace engineering because they wanted to build rockets.
Figure 11.4: A Photo taken from Auburn Junior High School Engineering Day
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Section 12: Conclusion
In conclusion, the Auburn team would like to thank the competition organizers for working with
us through our motor failures. We look forward to our launch on April 2nd and feel extremely
confident that it will qualify us to compete in Huntsville. Our rebuild from the last motor failure
is almost complete and a rigorous check of all systems for both safety and functionality will be
completed before the launch date.