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 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

https://tigermailauburn-my.sharepoint.com/personal/cas0058_tigermail_auburn_edu/Documents/Auburn%20Student%20Launch/SL%202015%20-%202016/FRR/FRR.docx#_Toc445705648

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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


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


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


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


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


The official, marked

altimeter will not be

damaged

Inspection

Analysis

Testing

Design the

electronics

housing to

prevent

damage to

altimeter


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


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


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


The launch vehicle

shall be limited to a

single stage

Demonstration

Team will

design and

build a single-

stage launch

vehicle


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


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


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


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


Final motor choices

must be made by the

Critical Design

Review (CDR).

Demonstration

CDR will

determine

which motor

the team will

use for

competition


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


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.


Pressure vessels on

the vehicle shall be

approved by the RSO

Analysis

Testing

Inspection of

pressure vessel

by RSO

standards by

testing.

50


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.


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.


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


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.


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


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


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.


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.


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


The launch vehicle

shall not utilize

forward canards.

Demonstration

The team will

demonstrate

how the launch

vehicle does

not utilize

canards.


The launch vehicle

shall not utilize

forward firing

motors.

Demonstration

A

demonstration

of the launch

vehicle will

demonstrate it

not utilizing

forward firing

motors.


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.


The launch vehicle

shall not utilize

hybrid motors.

Demonstration

The team will

exhibit how the

launch vehicle

does not utilize

hybrid motors.

57


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.

59

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

66

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:

69

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

76

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


79

2.11.3

Producing Magnetic Waves Recovery System Electronics Shielded from any other On- Board Transmitting Devices


80

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



broken sheer pin

bracket

Bracket not properly

secured


- Inner seal

- 0.5-in outer seal





to the rocket


- 2 25-lb vertical sheet pins

- inner seal

- 1.0-in outer seal

- Paraffin wax seal on all

seams





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


Motor CATO, no

useable data

None

Aquila

III


- Inner seal

- 0.5-in outer seal


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


- 2 25-lb vertical sheet pins

- inner seal

- 1.0-in outer seal

- Paraffin wax seal on all

seams


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


-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


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


-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


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


-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


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


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



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



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



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 Rocket at Max Velocity 96.9410 pound-force

Drag Estimate of Rocket and Fins at Max Velocity, 45

degrees 150.1071 pound-force



Max Acceleration at Max Velocity with Fins, 45

degrees -185.1896 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 Rocket at Max Velocity 96.9410 pound-force





123





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.

128

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.

129

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.

130

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.

131

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.

132

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

133

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.

134

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

135

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

137

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.

138

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

140

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.


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.


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

146

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.


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.


147

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

149

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.


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.


• 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.


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.


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.


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


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


fins is conducted

through

computational fluid

dynamics (CFD),

subsystem wind


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


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


fins is conducted

through

computational fluid

dynamics (CFD),

subsystem wind


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



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



interact once

deployed, the team

has constructed visual


flow through the


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


fins is conducted

through

computational fluid

dynamics (CFD),

subsystem wind


flight sensors. If large

off axis acceleration is



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


fins is conducted

through

computational fluid

dynamics (CFD),

subsystem wind


flight sensors. If large

off axis acceleration is



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.


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.


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



interact once

deployed, the team

will construct visual


flow through the


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


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.


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


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.


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


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.


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.


Improper

Ventilation (1)

CO2 gas can cause

headaches, nausea,

and loss of

consciousness in

high doses

5 4


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.


Improper

Ventilation (1)

Black powder is

hazardous to the

respiratory system

when inhaled. Also

particles may form

explosive mixtures

in air.

5 4


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.


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.


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.


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.


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


• 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.


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.


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.


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


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


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.


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.


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


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


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


Set up tables.

Unload all boxes and equipment

Unpack booster section

Inspect fins and tube for damage caused during travel

208

Grid Fins Section


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


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

209


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)


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

210

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

211

**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


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:


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

212

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



X X

Section 9.4: Setup for Launch/ Igniter Installation Setup on Launcher and Igniter Installation Check:


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

213

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



X X

Section 9.5: Launch Procedures Final Construction Check:


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:


214

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



X X

Launch Procedures Check:


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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X X

Section 9.6: Troubleshooting mall cracks in material


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


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


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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Check that hinges are attached correctly

If problem remains, loosen hinge screws, align halves and retighten screws

Grid fins do not respond


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


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


Screw launch lugs in farther

If launch lugs are still not secure, drill new holes and screw them in there

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X X

Section 9.7: Post-Flight Inspection Post Flight Inspection Check:


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



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.