project aquila - eng.auburn.edu

PROJECT AQUILA

211 ENGINEERING DRIVE

AUBURN, AL 36849

CRITICAL DESIGN REVIEW REPORT

JANUARY 15, 2016

1

Table of Contents Section 1: Summary of CDR Report .......................................................................................... 12

Team Summary............................................................................................. 12

Launch Vehicle Summary ............................................................................. 12

Payload Summary ......................................................................................... 13

Section 2: Changes Made Since PDR ........................................................................................ 14

Changes Made to Vehicle ............................................................................. 14

Changes Made to Payload ............................................................................. 14

Changes Made to Project Plan ....................................................................... 14

Section 3: Launch Vehicle......................................................................................................... 15

Mission Statement ........................................................................................ 15

Major Milestone Schedule ............................................................................ 15

System-Level Design Review ....................................................................... 17

Structure ................................................................................................... 17

Propulsion ................................................................................................ 20

Aerodynamics........................................................................................... 24

Final Dimensional Drawings ......................................................................... 27

Test Descriptions and Results ....................................................................... 35

Materials Testing ...................................................................................... 35

Wind Tunnel Testing ................................................................................ 38

System Level Functional Requirements ........................................................ 39

Workmanship and Manufacturing ................................................................. 40

Design Integrity ............................................................................................ 41

Fin Shape and Style .................................................................................. 41

Materials in Fins, Bulkheads and Structural Elements ............................... 41

Assembly Procedures................................................................................ 43

Verification Plan ....................................................................................... 44

Mass Statement......................................................................................... 54

2

Mission Performance Criteria........................................................................ 55

Section 4: Subscale Flight Results ............................................................................................. 56

Flight Data .................................................................................................... 56

Predicted vs. Actual Performance.................................................................. 57

Subscale Impact on Full-Scale Design .......................................................... 57

Section 5: Recovery Subsystem ................................................................................................. 58

Recovery System Outline .............................................................................. 58

Subscale Analysis ......................................................................................... 59

Requirement Validation ................................................................................ 60

Parachutes..................................................................................................... 62

Drift .............................................................................................................. 69

Ejection System ............................................................................................ 71

Altimeters ..................................................................................................... 74

Attachment Hardware ................................................................................... 77

Section 6: Aerodynamic Analysis Payload ................................................................................ 79

System Level Design Review ........................................................................ 79

Payload Structure .......................................................................................... 79

Payload Electronics ....................................................................................... 81

Arduino .................................................................................................... 82

Servos....................................................................................................... 82

Accelerometer .......................................................................................... 83

GPS .......................................................................................................... 84

Wiring ...................................................................................................... 84

Design Requirements .................................................................................... 85

Manufacturing and Assembly........................................................................ 87

Risk Mitigation ............................................................................................. 88

Payload Integration ....................................................................................... 88

Payload Concept Features and Definition ...................................................... 89

Science Value ............................................................................................... 90

3

Payload Objectives ................................................................................... 90

Payload Success Criteria ........................................................................... 90

Testing and Simulation ............................................................................. 91

Section 7: Payload Fairing (PLF) ............................................................................................ 111

System Level Design Review ...................................................................... 111

Design Overview .................................................................................... 111

Materials ................................................................................................ 113

Design Requirements .................................................................................. 114

Testing ........................................................................................................ 115

Aerodynamic Design Testing (Completed) ............................................. 115

Charge Deployment Testing ................................................................... 116

Drag Strip Deployment Testing .............................................................. 117

Full Scale Testing ................................................................................... 118

Science Value ............................................................................................. 118

Section 8: Safety ..................................................................................................................... 119

Checklists ................................................................................................... 119

Final Assembly Checklist ....................................................................... 119

Launch Procedures Checklist .................................................................. 121

Safety Officer ............................................................................................. 124

Airframe Hazard Analysis ........................................................................... 124

Airframe Failure Modes .......................................................................... 124

Environmental Effects ............................................................................ 132

Airframe Risk Mitigation – Testing Systems........................................... 133

Scientific Payloads Hazard Analysis ........................................................... 135

Scientific Payload Risk Mitigation – Payload Fairing ............................. 136

Scientific Payload Risk Mitigation – WAFLE ......................................... 142

Recovery Hazard Analysis .......................................................................... 152

4

Recovery Risk Mitigation - Materials ..................................................... 157

Recovery Risk Mitigation - Construction ................................................ 161

Outreach Hazard Analysis ........................................................................... 165

Environmental Effects................................................................................. 168

Vehicle Effects on Environment ............................................................. 168

Environmental Effects on the Vehicle ..................................................... 169

Section 9: Project Plan ............................................................................................................ 170

Budget Plan ................................................................................................ 170

Funding Plan ............................................................................................... 171

Timeline ..................................................................................................... 172

Section 10: Educational Engagement....................................................................................... 175

Drake Middle School 7th Grade Rocket Week ........................................... 175

Rocket Week Plan of Action ................................................................. 176

Rocket Week Launch Day .................................................................... 177

Rocket Week Learning Objectives ........................................................ 178

Gauging Success ................................................................................... 179

Samuel Ginn College of Engineering E-Day ............................................... 179

Boys Scouts of America Space Exploration Merit Badge ............................ 179

Space Exploration Merit Badge Requirements ...................................... 180

Boy Scouts of America - AUSL Requirements...................................... 181

Boy Scouts of America - Plan of Action ............................................... 182

Boy Scouts of America: Goals .............................................................. 183

Girl Scouts of the USA - Space Badge ........................................................ 183

Rocket Day ................................................................................................. 183

Rocket Day – Outline ........................................................................... 184

Rocket Day – Safety ............................................................................. 184

Auburn Junior High School Engineering Day ............................................. 185

5

Section 11: Conclusion............................................................................................................ 186

6

List of Figures Figure 3.1: Full Rocket Rendering ............................................................................................. 15

Figure 3.2: 5 Inch Braided Isogrid ............................................................................................. 19

Figure 3.3: Motor Tube Rendering ............................................................................................ 21

Figure 3.4: Aerotech L1520T Thrust Curve ............................................................................... 21

Figure 3.5: Aeropack Motor Retention ...................................................................................... 23

Figure 3.6: Fin Rendering.......................................................................................................... 24

Figure 3.7: Upper Section Dimensions ...................................................................................... 27

Figure 3.8: Lower Section Dimensions ...................................................................................... 28

Figure 3.9: Booster Tube ........................................................................................................... 29

Figure 3.10: Upper Body Tube .................................................................................................. 30

Figure 3.11: Fins ....................................................................................................................... 31

Figure 3.12: Filament Wound Body Tube .................................................................................. 32

Figure 3.13: Bulkhead ............................................................................................................... 33

Figure 3.14: Motor Tube Bulkhead............................................................................................ 34

Figure 3.15: Carbon Fiber Test Data ......................................................................................... 36

Figure 3.16: Carbon Fiber Test Results ..................................................................................... 36

Figure 3.17: HIPs Data .............................................................................................................. 37

Figure 3.18: HIPS Test Results ................................................................................................. 37

Figure 3.19: Three Point Bending Test ...................................................................................... 38

Figure 3.20: FPS vs Lb Force .................................................................................................... 38

Figure 3.21: Wind Tunnel Test .................................................................................................. 39

Figure 3.22: Fin Shapes ............................................................................................................. 41

Figure 3.23: Patran Tube Model ................................................................................................ 42

Figure 4.1: Subscale Open Rocket Model .................................................................................. 56

Figure 5.1: Parachute Configuration .......................................................................................... 58

Figure 5.2: Subscale Parachute Configuration ........................................................................... 59

Figure 5.3: Parachute Shape Parameters .................................................................................... 63

Figure 5.4: Main Parachute Visualization .................................................................................. 66

Figure 5.5: Pictures of Tender Descender in Undeployed and Deployed Configurations ............ 68

Figure 5.6: Solenoid Circuit ...................................................................................................... 72

7

Figure 5.7: Custom CO2 System Housing ................................................................................. 73

Figure 5.8: Custom CO2 System Assembly ............................................................................... 73

Figure 5.9: Altus Metrum TeleMega Altimeter .......................................................................... 75

Figure 5.10: Altus Metrum TeleMetrum Altimeter .................................................................... 75

Figure 5.11: Taoglas FXP240 433 MHz ISM Antenna .............................................................. 76

Figure 6.1: WAFLE system ....................................................................................................... 79

Figure 6.2: Grid Fin Fairing ...................................................................................................... 80

Figure 6.3: Aerodynamic Grid fin ............................................................................................. 81

Figure 6.4: Arduino Uno ........................................................................................................... 82

Figure 6.5: HiTec HS-5685MH Digital Super Torque Servo ..................................................... 83

Figure 6.6: ADXL335 Triple-axis Accelerometer ...................................................................... 84

Figure 6.7: WAFLE Electronics Schematic ............................................................................... 85

Figure 6.8: Starting point of the flow. ........................................................................................ 94

Figure 6.9: Flow directed over the grid fins. .............................................................................. 95

Figure 6.10: 0.2 Mach flow over a grid fin. ............................................................................... 95

Figure 6.11: Total Fin Axial Force Coefficient versus Angle of Attack Mach 8 Low Pressure, Low

Temperature .............................................................................................................................. 99

Figure 6.12: Fin Normal Force Coefficient versus Angle of Attack Mach 8 High Pressure, High

Temperature .............................................................................................................................. 99

Figure 6.13: Fin Normal Force Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low

Temperature ............................................................................................................................ 100

Figure 6.14: Fin Normal Force Coefficient versus Angle of Attack Mach 0.1 High Pressure, High

Temperature ............................................................................................................................ 100

Figure 6.15: Fin Moment Coefficient versus Angle of Attack Mach 8 Low Pressure, Low

Temperature ............................................................................................................................ 101

Figure 6.16: Fin Moment Coefficient versus Angle of Attack Mach 8 High Pressure, High

Temperature ............................................................................................................................ 101

Figure 6.17: Fin Moment Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low

Temperature ............................................................................................................................ 102

Figure 6.18: Fin Moment Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low

Temperature ............................................................................................................................ 102

8

Figure 6.19: Vortex Shedding Testing Visualization................................................................ 109

Figure 7.1: PLF Curvature ....................................................................................................... 111

Figure 7.2: Overall PLF Structure ........................................................................................... 112

Figure 7.3: Overall PLF Assembly .......................................................................................... 112

Figure 7.4: PLF Half ............................................................................................................... 113

Figure 7.5: Partially Deployed PLF ......................................................................................... 113

Figure 7.6: PLF Half ............................................................................................................... 113

Figure 7.7: Charge Required for a Given Force (assume L = 1in) ............................................ 116

Figure 11.1: Picture from Rocket Week 2014 .......................................................................... 176

Figure 11.2: A photo taken from DMS 7th Grade Rocket Week in April 2014 ........................ 178

Figure 11.3: Space Exploration Merit Badge ........................................................................... 182

Figure 11.4: A Photo taken from Auburn Junior High School Engineering Day....................... 185

9

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: Major Milestone Schedule ........................................................................................ 15

Table 3.2: Vehicle Length ......................................................................................................... 17

Table 3.3: Aerotech L1520T Motor Specifications .................................................................... 22

Table 3.4: Fin Dimensions ........................................................................................................ 25

Table 3.5: Manufacturing Phase Schedule ................................................................................. 43

Table 3.6: Verification Plan....................................................................................................... 44

Table 3.7: Mass Estimates and Growth Allowance .................................................................... 54

Table 4.1: Simulation Data vs. Flight Data ................................................................................ 57

Table 5.1: Recovery Requirement Validation ............................................................................ 60

Table 5.2: Parachute Shape Pugh Chart ..................................................................................... 64

Table 5.3: Main Parachute Dimensions ..................................................................................... 65

Table 5.4: Kinetic Energy Calculations ..................................................................................... 67

Table 5.5: Drift Calculations ..................................................................................................... 71

Table 6.1: Aerodynamic Analysis Payload Design Requirements .............................................. 85

Table 6.2: Aerodynamic Analysis Payload Risk Mitigation ....................................................... 88

Table 6.3: Aerodynamic Payload Success Criteria ..................................................................... 90

Table 6.4: Aerodynamic Payload Simulations and Tests ............................................................ 91

Table 6.5: SolidWorks Simulation Run Cases ........................................................................... 96

Table 6.6: Aerodynamic Payload Fortran- Flight and Dynamic model ....................................... 97

Table 6.7: Sample Data Mach=0.8 Low Pressure, Low Temperature ....................................... 103

Table 6.8: Sample Data at Mach=0.8 High Pressure, High Temperature .................................. 104

Table 6.9: Sample Data at Mach 0.1 Low Pressure, Low Temperature .................................... 105

Table 6.10: Sample Data at Mach 0.1 High Pressure. High Temperature ................................. 106

Table 7.1: PLF Design Requirements ...................................................................................... 114

Table 7.2: Drag Strip Deployment Test Matrix ........................................................................ 117

Table 9.1: Risk Mitigation Table - Airframe ............................................................................ 127

Table 9.2: Risk Mitigation Table - Autoclave .......................................................................... 128

10

Table 9.3: Risk Mitigation Table - Filament Winder................................................................ 129

Table 9.4: Risk Mitigation Table - Carbon Fiber ..................................................................... 130

Table 9.5: Risk Mitigation Tables - Epoxy .............................................................................. 131

Table 9.6: Risk Mitigation Tables - Airframe Environment Effects ......................................... 133

Table 9.7: Risk Mitigation Tables – Wind Tunnel Testing ....................................................... 133

Table 9.8: Risk Mitigation Tables – Tensile Test Rig .............................................................. 134

Table 9.9: Risk Mitigation Table - Operations ......................................................................... 136

Table 9.10: Risk Mitigation Table – Payload Fairing Testing .................................................. 138

Table 9.11: Risk Mitigation Table – Payload Fairing Construction .......................................... 141

Table 9.12: Risk Mitigation Table – Operations ...................................................................... 142

Table 9.13: Risk Mitigation Table – WAFLE Testing ............................................................. 147

Table 9.14: Risk Mitigation Table – WAFLE Construction ..................................................... 148

Table 9.15: Risk Mitigation Table – WAFLE Materials .......................................................... 150

Table 9.16: Risk Mitigation Table - Flight Recovery Operations ............................................. 152

Table 9.17: Risk Mitigation Tables - Wind Tunnel Testing ..................................................... 154

Table 9.18: Risk Mitigation Table - Tensile Test Rig .............................................................. 155

Table 9.19: Risk Mitigation Tables - Shear Pin Test Rig ......................................................... 156

Table 9.20: Risk Mitigation Table - Kevlar ............................................................................. 157

Table 9.21: Risk Mitigation Tables - Nylon ............................................................................. 158

Table 9.22: Risk Mitigation Tables - Carbon Dioxide .............................................................. 158

Table 9.23: Risk Mitigation Table - Black Powder .................................................................. 160

Table 9.24: Risk Mitigation Table - Fiberglass ........................................................................ 161

Table 9.25: Risk Mitigation Table - Orbital Sander ................................................................. 161

Table 9.26: Risk Mitigation Table - Sewing Machine .............................................................. 163

Table 9.27: Risk Mitigation Table - Hand Tools ...................................................................... 164

Table 9.28: Risk Mitigation Table - Outreach Operations ........................................................ 166

Table 9.29: Risk Mitigation Table - Outreach Construction ..................................................... 167

Table 9.30: Risk Mitigation Table - Outreach Materials .......................................................... 168

Table 10.1: Initial Budget Estimates ........................................................................................ 170

Table 10.2: Funding Sources ................................................................................................... 171

Table 10.3: Launches and Vehicle Timeline ............................................................................ 173

11

Table 10.4: Subsystem Timeline ............................................................................................. 173

Table 10.5: Competition Timeline ........................................................................................... 174

12

Section 1: Summary of CDR Report

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

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.

13

Table 1.3: Launch Vehicle Summary

Total Length 73.125 inches

Final Mass Estimate 26.8 lbs

Motor Selection Aerotech L1520T

Payload Summary

The Auburn Student Launch Team will be completing a Payload Fairing and an Aerodynamic

Analysis Payload. The payload fairing will serve as the nose cone of the launch vehicle and the

main and drogue parachutes will be ejected from the fairing via the Tender Descender system The

Payload Fairing is detailed in Section 7. The aerodynamic analysis payload, dubbed WAFLE,

consists of grid fins that will be protected by fairings and actuated by servos. This payload will be

used as an airbrake on ascent of the vehicle. This payload is detailed in Section 6 of this report.

14

Section 2: Changes Made Since PDR

Changes Made to Vehicle

The following is a list of changes that have occurred to the vehicle body:

• The fuselage tubes will no longer be wrapped in carbon fiber from in-house filament

winder. Instead, the carbon fiber meshes will be covered with a Kevlar sock.

• The motor retention system will no longer be made in-house. The team has chosen to use

a commercial Aeropack motor retention system.

Changes Made to Payload

The following is a list of the changes that have been made to the structure of subsystems of the

Wall Armed Fin-Lattice Elevator (WAFLE):

• The design of the outer fairing has changed. The fairing has transformed from a hollow

hemispherical fairing to a solid ogive fairing.

• The outer fairing constructing material has changed from a filament wound carbon fiber to

a printed plastic.

• The actuation system has changed from a gear directly attached to the center of the grid

fin, to an actuation system with the gear set within a U-brace integrated into the grid fin.

• The grid fin dimensions have been changed. The length and chord of the fin has increased,

however the lattice thickness has decreased. The mounting structure on the fin has changed

from an A-frame design to a U-brace design to accommodate the actuation system.

• The servos have been repositioned and placed within a slot in the airframe instead of stored

under the hemispherical fairing.

Due to multiple changes to the design of the rocket, the hinge line of the grid fins have moved

relative to the center of gravity of the rocket. The hinge line moved from 0.293 inches forward of

the wet CG to 1.054 aft of the wet CG. The distance between the hinge line and the dry CG

increased from 7.717 inches to 8.436 inches, with the CG forward of the hinge line. Likewise, the

hinge line moved aft of the main separation by 1.22 inches.

Changes Made to Project Plan

15

Section 3: Launch Vehicle

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. AUSL requires

the highest quality of all components in order to reach the goals set by NASA in this year’s

competition.

Major Milestone Schedule

This schedule includes the major competition deadlines set forth in the NASA Student Launch

Handbook as well as dates for construction and test launches. A more detailed version of the

manufacturing, testing, and launch schedules can be found in Section 11: Project Plan .

Table 3.1: Major Milestone Schedule

Date Event Description

Figure 3.1: Full Rocket Rendering

16

10/24/2015 Subscale Construction Begin construction of 3/5

subscale

11/01/2015 Subscale Completion Complete construction of

subscale

11/07/2015 Subscale Launch First launch day for

subscale.

11/21/2015 Subscale Launch Second launch day for

subscale.

12/20/2015 Full Scale Design Completed

Complete full scale design from subscale results.

01/05/2016 Begin Construction of Full Scale

Begin constructing two full scale rockets

01/14/2016 Complete Critical Design Review

Paper and Presentation completed

01/15/2016 Complete Construction of Full Scale

Have a completed full scale test rocket

01/16/2016 Full Scale Launch First full scale launch

testing avionics and PLF

01/25/2016 CDR Presentation Complete the CDR

presentation

01/30/2016 Full Scale Launch Second full scale launch testing WAFLE and RS

02/13/2016 Full Scale Launch Third full scale launch

testing all systems together

02/14/2016 Complete Construction of Full Scale

Have two competition-ready full scale rockets

02/20/2016 Full Scale Launch Fourth full scale launch with competition rocket

03/14/2016 Flight Readiness Review Complete FRR paper and

presentation

17

03/17/2016 – 03/30/2016 FRR Teleconference Complete FRR

presentation

04/13/2016 Launch Readiness Review Complete LRR in

Huntsville, AL

04/16/2016 Competition Launch Day Launch full scale rocket in

competition

04/29/2016 Post Launch Assessment Review

Complete PLAR

System-Level Design Review

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. 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.

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 3.2: Vehicle Length

Component Length (Inches)

18

Nose Cone (Fairing) 13.125

Upper Tube 22

Aerodynamics Payload Section 5

Booster Section 33

Total 73.125

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 altitude.

The body tubes will be constructed using carbon fiber braiding, a process that involves taking

individual strands of carbon fiber and stitching them into a tightly-wound braid. The carbon fiber

braids that are produced will be formed into an isogrid structure around a 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

this skin, the result is a lightweight, aerodynamic body. Using this wrapped isogrid method, the

mass of the body tubes will be decreased by approximately 20 to 30 percent less than if the tubes

were constructed using only filament wound carbon fiber, while also maintaining the same

compressive strength properties as a carbon fiber tube. This mass reduction was confirmed using

tube samples constructed by team members using final production methods. An image of a sample

of the braided isogrid structure without the aerodynamic skin can be seen in Figure 3.2.

19

Figure 3.2: 5 Inch Braided Isogrid

Couplers:

The couplers serve as a joint between two body tube sections. The couplers are designed to separate

during the recovery phase of the flight. To accomplish this, the lower body tube is attached to the

coupler using four 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 couplers will be constructed using high impact polystyrene (HIPS) on a TAZ4 3D printer.

Designs printed from HIPS are very accurate; using this material, the printer’s accuracy is 50

microns. This allows for the manufacturing of high-quality, accurate parts. Couplers have very

thin walls relative to their length and this can cause the HIPS material to be prone to cracking when

placed under stress. To correct this, two layers of carbon fiber infused with resin will be epoxied

to the inside wall of these couplers. By re-enforcing the plastic with the composite material, the

structure becomes 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

20

and launch day conditions. The tank will be placed forward of the grid fin section, near the CG

location, and is secured to the launch vehicle body by two aluminum pins. As the tank will not be

subjected to a large force, the team is confident that the pins will hold the tank securely without

fear of a shear failure. The tank will be constructed using high impact polystyrene (HIPS) on a

TAZ4 3D printer.

Bulkheads:

Bulkheads are typically flat plates used to increase the structural strength of a rocket. They are also

used to create airtight spaces and to divide the body into separate compartments. In rockets, they

are commonly used to separate payload bays and to mount equipment for avionics and payloads.

For rockets similar in size to the Project 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. This was chosen due to the simplicity of manufacturing with pre-impregnated carbon fiber.

The interior diameter for the circular cross-sectional rocket will be 5 inches and the bulkheads are

designed to fit perfectly into this size. All bulkheads for this rocket will be 0.25 inches thick.

Centering Rings:

The purpose of the centering rings is to center a smaller cylindrical body or tube inside a tube of a

larger diameter. In the case of high powered model rocketry, centering rings can be used as an

engine block in motor mounts. The Project 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 attach to the

motor retention. The centering rings are made of carbon fiber and manufactured using the

Computer Numerical Control (CNC) machine at Auburn University Aerospace Design Lab due to

the availability and the teams experience with using carbon fiber. The centering rings have an outer

diameter of 5 inches with an inner diameter of 3 inches. The thickness of each ring is approximately

0.25 inch. The centering rings have a mass of 2.16 oz., determined from sample pieces.

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.

21

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

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.

22

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 3.3: 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

Motor Tube:

To contain the motor on the rocket, a carbon fiber motor tube is being used. The motor tube will

be made by braiding carbon fiber strands and then filament wound around a mandrel that is the

same diameter of the motor. The 3D braided carbon fiber material was chosen for its strength

relative to its weight when compared to a solid tube. Basalt fiber was considered to be used for the

23

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.

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

24

Aerodynamics

The aerodynamics system requires the rocket remain stable during flight. The placement and

design of the aerodynamic surfaces determines the center of pressure along the length of the rocket.

Fins:

The stability of the rocket is controlled by the fins. The primary purpose of the fins is to keep the

center of pressure aft of the center of gravity. The greater drag on the fins will keep them behind

the upper segments of the vehicle, thus allowing the rocket to fly straight along the intended flight

path. They are also helpful in minimizing the chances of weather-cocking. Fins serve as an ideal

addition to the vehicle body as they are lightweight and easy to manufacture using the CNC

machine. A clipped delta planform has been selected for the fins. Four fins will be machined from

0.2 inch thick carbon fiber flat plates. A rendering of the fin design is shown in Figure 3.6.

Figure 3.6: Fin Rendering

When attached, the trailing edge of each fin will be located slightly forward of the end of the body

tube. This design feature will theoretically provide some impact protection for the fins when the

rocket hits ground. Carbon fiber of 1.03 oz/in3 density has been selected as the material due to its

stiffness, strength, and light weight. Each fin will have a surface area of 54 in2 (summing both

25

sides), making the fin surface area total equal to 216 in2. The total component mass is 13.5 ounces.

These dimensions provide the vehicle with a projected stability of 2.25 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 cosΓ𝑐𝑐

)^2

Table 3.4: 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 0.2 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

26

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

The flutter velocity is directly reflective of the aero-elastic conditions of the structure/fin system.

The catastrophic flutter phenomenon results from coupling of aerodynamic forces creating a

positive feedback loop. The increase in either torsion or bending drives an infinitely looped

increase in the other motion. Since it is assumed that the fins are rigidly fixed and cantilevered to

an infinitely stiff rocket body, the fin twist (torsion) and fin plunge (bending) are the only two

degrees of freedom.

Once this flutter velocity is exceeded, the air, inversely, amplifies the oscillations and significantly

increases the energy within the respective fin. As velocity increases, the fin twist and plunge are

no longer damped. At this velocity, known as the divergent speed, one degree of freedom usually

diverges while the other remains neutral. Structural failure usually occurs at or just above this

velocity. Due to certain failure of the structure associated with potential aero-elastic flutter, the

flutter velocity is applied to the design as a “never-to-exceed” parameter.

There are various ways to minimize the chances of experiencing fin flutter. Increasing fin retention

by strengthening the joints between the fins and rocket body is one way to supplement system

stability. Furthermore, additional layers of carbon fiber and epoxy applied to portions of the fins

as well as the joints should provide extra defense against aero-elastic flutter. The Finite Element

Method (FEM) will be implemented via FORTRAN and PATRAN to optimize the aero-elastic

fin/body combination.

27

Final Dimensional Drawings

Figure 3.7: Upper Section Dimensions

28

Figure 3.8: Lower Section Dimensions

29

Figure 3.9: Booster Tube

30

Figure 3.10: Upper Body Tube

31

Figure 3.11: Fins

32

Figure 3.12: Filament Wound Body Tube

33

Figure 3.13: Bulkhead

34

Figure 3.14: Motor Tube Bulkhead

35

Test Descriptions and Results

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. On

36

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.

The next planned test is a tensile testing test, which is scheduled for the third week of January

2016. The samples used for this testing were acquired from the same batch as the previous testing

to eliminate a potential source of error.

Figure 3.15: Carbon Fiber Test Data

Figure 3.16: Carbon Fiber Test Results

37

Figure 3.17: HIPs Data

Figure 3.18: HIPS Test Results

38

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 3.20: 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 3.19: Three Point Bending Test

39

System Level Functional Requirements

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:

1. The deployment charge shall induce separation without harming the structural integrity of

the PLF.

Figure 3.21: Wind Tunnel Test

40

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.

Workmanship and Manufacturing

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 will be tested for strength to make sure all components are capable of handling

the expected loads. Finite element analysis has been used in the software Patran in order to

understand the structural stress involved in the rockets flight.

The rockets flight has been simulated on a simulation software OpenRocket. To verify the results

of the simulation, wind tunnel testing, Computational fluid dynamics and calculations by hand

have all been performed.

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 will use former launch vehicle data, design faults and

failures as an example to anticipate and mitigate any future 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

41

member has any question or doubt about any reason an answer is sought from either the team’s

faculty advisors, safety advisors or any other reliable source before any proceeding of activities.

Design Integrity

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.

Figure 3.22: Fin Shapes

The fins will be manufactured from the same carbon fiber plates as the bulkheads and centering

rings. The same data used to verify the bulkheads and centering rings will be used to ensure the

fins are capable of withstanding any inflight or landing forces.

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.

Materials in Fins, Bulkheads and Structural Elements

Body Tubes:

42

The structural tubes of the launch vehicle are going to be constructed using a 3D braided carbon

fiber isogrid structure. As this is something the team has not done in past years, structural data will

need to be collected for this structure. To do this, using the same material and manufacturing

method, a test sample will be made consisting of an equal diameter of the tubes that will be used

on the launch vehicle. This sample will then be placed into a load cell to determine the maximum

load of the structure. This will allow us to determine that the structure is capable of safely

completing the mission. The structure will experience a maximum of 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.

Figure 3.23: 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 will be made of. The samples were placed in a three point

bending test as well as a tensile stress test.

Coupler:

To verify the coupler functions correctly, ground tests of the separation will be performed. Once

proven on the ground, a subscale flight test using this coupler component will be used.

43

Ballast Tank:

By running simulations, the team is able to determine where the center of gravity is located. Once

the launch vehicle is manufactured a final simulation will be run using real component weights. If

the center of gravity is not where initially predicted, the ballast tank will be used to correct the

location. Throughout the project, this will be re-examined to ensure stable flight.

Assembly Procedures

Manufacturing of the vehicle generally takes two weeks to produce and assemble the components.

To account for this the team plans to start manufacturing three weeks prior to any scheduled

launches. This allows for one extra week if any problems arise during the manufacturing process.

The typical manufacturing schedule can be seen in Table 3.5.

Table 3.5: Manufacturing Phase Schedule

Week Events Percentage of completion

1 Manufacture major components. Such as body tubes, nose cone, fins. 50%

2 Begin assembly of subsystems.

Such as booster sections, fin assemblies.

90%

3 Assemble completed rocket 100%

Manufacturing body tubes using braided structures is a very time consuming process. The first

four weeks of January will be used to manufacture the tube structures, both braided and

nonbraided. These tubes are the most time consuming component to manufacture and the event of

a crash would have negative effects on the team’s timeline. To mitigate the effects of a total loss

crash, six tube sections and three motor tube sections will be produced during this time, which will

allow for the construction of three full scale rockets, two with braided body tubes and one with

nonbraided body tubes.

Several flat plates of carbon fiber have been produced at various thicknesses. The plates will be

placed in a CNC router to be shaped into flat components. These components include fins, bulk

plates and centering rings.

44

Verification Plan

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 3.6: Verification Plan

Team Requirement

NASA Requirement Section/Number

Requirement Statement

Verification Method

Execution of Method

AU – 1 Vehicle 1.1

The vehicle shall deliver the payload to an apogee altitude of 5,280 feet above ground level (AGL).

Analysis Demonstration Testing

Launch vehicle and check altimeters

AU – 2 Vehicle 1.2

The vehicle shall carry one commercially available, barometric altimeter for recording the official altitude used in the competition scoring.

Inspection Demonstration

Purchase and calibrate one commercially available altimeter

AU – 3 Vehicle 1.2.1

The official scoring altimeter shall report the official competition altitude via a series of beeps to be checked after the competition flight.

Inspection

Testing

Test the altimeter to verify it creates audible beeps

AU – 4 Vehicle 1.2.2

Teams may have additional altimeters to control vehicle electronics and payload experiment(s).

Demonstration

The team may use additional altimeters.

45

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


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.


Ensure beeps are audible, launch successfully


At the launch field, to aid in determination of the vehicle’s apogee, all audible electronics, except for the official altitude-determining altimeter shall be capable of being turned off.

Inspection Demonstration Testing

Ensure all electronics can be turned off and back on

AU – 8 Vehicle 1.2.3.1 The official, marked altimeter will not be damaged

Inspection

Analysis Testing

Design the electronics housing to prevent damage to altimeter

AU – 9 Vehicle 1.2.3.2 The team will report to the NASA official designated to record the altitude with their official, marked altimeter on the day of the launch.

Demonstration The team is timely and organized in gathering data and reporting to NASA official

AU – 10 Vehicle 1.2.3.3 The altimeter will not report an apogee altitude over 5,600 feet AGL.

Demonstration Testing

Design and test launch vehicle to meet altitude requirement

46

AU – 11 Vehicle 1.2.3.4 The rocket will be flown at the competition launch site.

Demonstration Team will launch the rocket at the appropriate site on launch day

AU – 12 Vehicle 1.3 The launch vehicle shall be designed to be recoverable and reusable. Reusable is defined as being able to launch again on the same day without repairs or modifications

Testing Analysis Demonstration Inspection

Trajectory simulations and testing will ensure the launch vehicle is recoverable and reusable

AU – 13 Vehicle 1.4 The launch vehicle shall have a maximum of four (4) independent sections. An independent section is defined as a section that is either tethered to the main vehicle or is recovered separately from the main vehicle using its own parachute

Demonstration Team will design and build launch vehicle that can have, but does not require, four independent sections

AU – 14 Vehicle 1.5 The launch vehicle shall be limited to a single stage

Demonstration Team will design and build a single-stage launch vehicle

AU – 15 Vehicle 1.6 The launch vehicle shall be capable of being prepared for flight at the launch site within 2 hours, from the time the Federal Aviation Administration flight waiver opens.

Demonstration Team will be timely and organized to ensure vehicle is prepared on time

47

AU – 16 Vehicle 1.7 The launch vehicle shall be capable of remaining in launch-ready configuration at the pad for a minimum of 1 hour without losing the functionality of any critical on-board component.

Testing Batteries shall be tested with full electronics to verify their life

AU – 17 Vehicle 1.8 The launch vehicle shall be capable of being launched by a standard 12 volt direct current firing system. The firing system will be provided by the NASA-designated Range Services Provider

Demonstration Vehicle will be designed and tested to be launched by the standard 12 volt DC system

AU – 18 Vehicle 1.9 The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA), and/or the Canadian Association of Rocketry (CAR).

Demonstration Vehicle will be designed around commercially available, certified motors

AU – 19 Vehicle 1.9.1 Final motor choices must be made by the Critical Design Review (CDR).

Demonstration CDR will determine which motor the team will

48

use for competition

AU – 20 Vehicle 1.9.2 Any motor changes after CDR must be approved by the NASA Range Safety Officer (RSO), and will only be approved if the change is for the sole purpose of increasing the safety margin.

Demonstration If the change is made to increase safety margin, NASA RSO will allow the change

AU – 21 Vehicle 1.10 The total impulse provided by a launch vehicle shall not exceed 5,120 Newton-seconds (L-class).

Demonstration Launch vehicle impulse will be designed to not exceed 5,120 Newton-seconds.

AU – 22 Vehicle 1.11 Pressure vessels on the vehicle shall be approved by the RSO

Analysis Testing

Inspection of pressure vessel by RSO standards by testing.

AU – 23 Vehicle 1.11.1 The minimum factor of safety (Burst or Ultimate pressure versus Max Expected Operating Pressure) shall be 4:1 with supporting design documentation included in all milestone reviews

Inspection

Analysis Testing

Testing of the low-cycle fatigue.

AU – 24 Vehicle 1.11.2 Each pressure vessel shall include a pressure relief valve that sees the full pressure of the tank.

Inspection

Analysis Testing

Inspection of each pressure vessel and testing of the pressure relief valve to see does it work as inspected.

49

AU – 25 Vehicle 1.11.3 Full pedigree of the tank shall be described, including the application for which the tank was designed, and the history of the tank, including the number of pressure cycles put on the tank, by whom, and when.


The team will inspect the tank along with documentation of testing and history.

AU – 26 Vehicle 1.12 All teams shall successfully launch and recover a subscale model of their full-scale rocket prior to CDR. The subscale model should resemble and perform as similarly as possible to the full-scale model, however, the full-scale shall not be used as the subscale model.


A subscale and full scale launch will be completed.

AU – 27 Vehicle 1.13 All teams shall successfully launch and recover their full-scale rocket prior to FRR in its final flight configuration. The rocket flown at FRR must be the same rocket to be flown on launch day.

Testing Demonstration Testing

A test of the rocket will be exhibit demonstration all hardware functions properly.

AU – 28 Vehicle 1.13.1 The vehicle and recovery system shall have functioned as designed.

Testing Testing of vehicle will show how recovery system functions.

50

AU – 29 Vehicle 1.13.2.1 If the payload is not flown, mass simulators shall be used to simulate the payload mass.


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.

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 of the adaptability of the systems notice to payload changes of the external surfaces through testing.

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


Inspection of the motor will be done by the team to ensure it is flown through full-scale testing.

51

acceleration of the competition flight.

AU – 33 Vehicle 1.13.4 The vehicle shall be flown in its fully ballasted configuration during the full-scale test flight. Fully ballasted refers to the same amount of ballast that will be flown during the competition flight.

Testing Demonstration

Testing of the vehicle will demonstrate it being fully ballasted.

AU – 34 Vehicle 1.13.5 After successfully completing the full-scale demonstration flight, the launch vehicle or any of its components shall not be modified without the concurrence of the NASA Range Safety Officer (RSO).

Demonstration The team will demonstrate that it did not alter any components or vehicle after demonstration flight.

AU – 35 Vehicle 1.14 Each team will have a maximum budget of $7,500 they may spend on the rocket and its payload(s).

Demonstration The team will demonstrate its budget of the competition rocket to validate its cost.

AU – 36 Vehicle 1.15.1 The launch vehicle shall not utilize forward canards.

Demonstration The team will demonstrate how the launch vehicle does not utilize canards.

AU – 37 Vehicle 1.15.2 The launch vehicle shall not utilize forward firing motors.

Demonstration A demonstration of the launch vehicle will demonstrate it not utilizing

52

forward firing motors.

AU – 38 Vehicle 1.15.3 The launch vehicle shall not utilize motors that expel titanium sponges (Sparky, Skidmark, MetalStorm, etc.)

Demonstration The team will demonstrate that the motor does not expel titanium sponges through test flight.

AU – 39 Vehicle 1.15.4 The launch vehicle shall not utilize hybrid motors.

Demonstration The team will exhibit how the launch vehicle does not utilize hybrid motors.

AU – 40 Vehicle 1.15.5 The launch vehicle shall not utilize a cluster of motors.

Demonstration A demonstration and inspection of 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

53

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 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.

54

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.

Mass Statement

The mass of the rocket and all of its subsystems was calculated using optimal mass calculations

from OpenRocket. In addition to using final masses from last year as a basis, a brick sample of

carbon fiber was created to have an accurate density measurement since most of the parts will be

manufactured using carbon fiber. This density test is exceedingly important given the method of

mass estimation. Since construction methods vary drastically from each manufacturer, as well as

different resin and cloth systems varying, it is highly important to get an accurate model of the

density.

Having determined an accurate density for the carbon fiber of the rocket, and the structure of the

rocket being the most significant portion of the weight of the structures of the rocket, the team

used estimates from last year’s rocket to determine the initial size estimate of the rest of the

subsystem components. The team believes that this model presents an estimate that is sufficient.

As the program develops, the model will attain a higher and higher accuracy in its simulation.

Table 3.7: Mass Estimates and Growth Allowance

Section Mass (lb) Percentage

Structure 10.8 40.28%

Recovery 4.51 16.82%

Grid Fins 2.00 7.46%

Electronics 1.52 5.66%

Motor 8.00 29.83%

55

Total 26.81 100%

The values in Table 3.7 are based off of simulations ran in OpenRocket and using data from the

3/5th subscale flight to determine accurate mass values.

Mission Performance Criteria

Vehicle:

1. The vehicle must have an apogee of 5280 feet AGL

2. The vehicle must be recoverable and reusable

Grid Fins:

1. All Aerodynamic data must be validated through analytical and experimental testing.

2. Charging line must disengage upon liftoff.

3. Grid fins must stay stowed until boost phase is complete.

4. Electronics must stay stationary throughout the flight

5. All electronics must come online once initiated after boost and stay online throughout the

flight.

6. Servos must remain in direct contact with the gears of the grid fins throughout the flight.

7. Arduino must accurately predict the flight path of the vehicle.

8. Grid fins must be deployed with precision to correct the vehicle’s trajectory.

9. Grid fins must stay deployed under the force applied by the flow.

Fairing:

1. The PLF tether line shall retain the PLF to the rocket main body post fairing separation.

2. The PLF system will be considered a success if the parachutes are successfully deployed,

and if the fairings remain structurally intact following landing.

56

Section 4: Subscale Flight Results

Flight Data

For the subscale flight, a scaling factor of 3/5th was applied to the full-scale rocket in order to

correlate the modeling to real-world results. The scaling factor was chosen in order to closely

resemble the full-scale flight. At 3/5 scale, the forces are realistic to what the full-scale would

experience. Unfortunately, the team was not able to use the CO2 system in the sub-scale. Thus, the

black-powder backup was utilized. In addition, the rocket was flown with a 3/5th scale aerodynamic

model to test how the system will work on the full-scale rocket.

The OpenRocket Design, shown in Figure 4.1, was as follows:

Simulations provided a predicted apogee of 5236 feet, and a static stability margin of 2.6 calibers.

The subscale weighed 8.25 pounds.

The first launch of the sub-scale took place on 7 November 2015 in Samson, AL under SoAR

(Southern Area Rocketry). This launch was categorized as a failure because the recovery system

failed to deploy resulting in the rocket nose diving into the ground. The team determined that the

recovery system failed because the vehicle did not have sufficient ejection force and material was

overly compressed into the parachute bay.

The second launch was on 21 November 2015 at the Phoenix Missile Works launch area. The

Aerotech K-805 motor originally selected was not available; instead an Aerotech K-1100 motor

was used. The Aerotech K-1100 has a faster burn time and higher thrust, giving the subscale a

higher maximum velocity. The maximum velocity reached a reported 1060 ft/s and an apogee of

6279 feet.

Figure 4.1: Subscale Open Rocket Model

57

Because a different motor was used during the flight test from that used in simulations, a second

simulation was made for the new motor. It has been determined that the grid fins in an undeployed

position are not a source of aerodynamic instability. This result shows that the grid fins will not

destabilize the rocket during its ascent, and will continue to be a viable and safe scientific payload.

The successful recovery of the rocket also proved that deploying a parachute using a tender

descender is a reliable means of recovery.

Predicted vs. Actual Performance

Table 4.1: Simulation Data vs. Flight Data

OpenRocket Simulation

Recorded Flight Data Percent Error

Apogee (Feet) 6158 6279 1.92%

Max Velocity (Ft/s) 1306 1060 23.2%

A reason for the discrepancy in the velocity can be attributed to the team being unable to simulate

the effects of drag caused by the grid fin payload in OpenRocket. There were also some anomalies

in the recorded flight data, resulting in unreliable and possibly inaccurate data. Several full scale

test launches will be performed to verify the operation of all altimeters and to gather more accurate

flight data

Subscale Impact on Full-Scale Design

The subscale flight tests determined that the design is very stable and will perform effectively.

Because of the complex aerodynamic shape of the rocket, simulations are not very accurate and

more tests will need to be run. The design of the rocket will not change as a result of this subscale

flight.

58

Section 5: Recovery Subsystem

Recovery System Outline

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 750

ft. At the second event (at 750 ft.) the booster section will deploy its own main parachute and

separate completely from the payload section, which will also deploy its own main parachute. The

rocket will be recovered in two sections. The payload and booster sections are not tethered together

and descend independently. The avionics bay is located in the bottom of the payload section. Both

the drogue and payload main parachutes are deployed out of the nosecone fairings, making use of

the Tinder Rocketry Tender Descender Dual Deploy parachute system. These parachutes are

deployed along with the Tender Descender via the opening of the nosecone fairings. The booster

section descends under its own main parachute, deployed from the separation at the second event

via Auburn’s custom CO2 ejection system. The parachute deployment is shown in Figure 5.1.

Figure 5.1: Parachute Configuration

59

Subscale Analysis

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

recovery systems are all designed for a 5 inch diameter rocket and must be modified or replaced

to fit in the smaller rocket. The subscale used a standard nosecone instead of fairings, so the team

ejected this nosecone to deploy parachutes. The subscale rocket’s recovery configuration is

illustrated in Figure 5.2.

Figure 5.2: Subscale Parachute 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

60

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 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.

Requirement Validation

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 5.1: Recovery Requirement Validation

Requirement Number Requirement Validation Method

2.1 Deployment of Recovery Devices Test Deployment System

2.2 Ground Ejection for Drogue & Main Parachute Test Deployment System

2.3 At Landing, Max KE of 75ft- lbf for each Independent Section

Calculation and subscale testing

61

2.4 Recovery system Electrical Circuits Independent of Payload Electrical Circuits

Make Separate Electrical Circuits for Recovery and Payload

2.5

Recovery System Must Contain a Redundant, Commercially Available Altimeter

Add redundant altimeter to recovery system

2.6 Exterior Arming Switch for each Altimeter

Add Exterior Arming Switch for each Altimeter

2.7 Dedicated Power Supply for each Altimeter

Put Separate Dedicated Power supply for each Altimeter

2.8 Arming Switch Capable of being Locked in the ON Position

Make Sure Locking Mechanism Locks the Switch when Turned ON

2.9 Removable Shear Pins used for Main & Drogue Parachute Compartment

Put Removable Shear Pins for Main and Drogue Parachute

2.10

Electronic tracking Device Installed in Rocket to Transmit the Location of the Tethered Vehicle or any Independent Section to a Ground Receiver

Place Electronic Tracking Device in the Tethered Vehicle and to any Independent Section and Test the Signal Location

2.10.1

An Active Electronic Tracking Device shall be connected to any Independent Rocket Section or Payload Component

Attach Electronic Tracking Device to any Independent Section & Payload Section and Test the Signal Location

2.10.2

The Electronic Tracking Device shall be fully Functional during Official Flight at Competition Launch Site

Test and Bring extra Electronic Tracking Devices at the Official Flight

2.11 Recovery System Electronics shall not be affected by other

Test Recovery System Electronics along with the

62

On-Board Electronics during Flight

other On-Board Electronics to Ensure Signal Strength

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 Electronics will be placed in a separate compartment away from any other Radio Frequency/ Magnetic Wave Producing Device

And Tested to Ensure Signal Strength

2.11.2 Recovery System Electronics Shielded from all On-Board Transmitting Devices

The Recovery System is designed be shielded by the avionics bay.

2.11.3 Recovery System Electronics Shielded from all On-Board Transmitting Devices

The Recovery System is designed be shielded by the avionics bay.

2.11.4

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

The Recovery System is designed to be shielded by the avionics bay.

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 drogue parachute will be a small, circular parachute constructed of rip-stop nylon with 0.5

inch tubular Kevlar shroud lines. At apogee, the drogue will be deployed from the top of the rocket,

out of the payload fairing. This will stabilize 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

⋅ ⋅=

63

The team’s rocket has a length of 73.125 in and a diameter of 5.25 in:

𝑑𝑑𝐷𝐷𝑅𝑅𝐷𝐷𝐺𝐺𝐷𝐷𝐷𝐷 = �4 ∙ 73.125 𝑖𝑖𝑖𝑖 ∙ 5.25 𝑖𝑖𝑖𝑖

𝜋𝜋 = 22.11 𝑖𝑖𝑖𝑖

The recovery system involves two main parachutes. Each main parachute will be constructed of

rip-stop nylon with 0.5 inch tubular Kevlar shroud lines.

Both main parachutes will be 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 will be deployed

from the top of that section. The other main parachute will deploy through the top of the rocket,

following the drogue. A spill hole will be added to both main parachutes. It will be added to the

booster section main parachute for stability, since it is falling separately from the rest of the rocket.

A spill hole will also be 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 will run 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 will be 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 5.3: Parachute Shape Parameters

64

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 5.2: 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:

65

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:

𝐴𝐴 =2 ∗ 10𝑙𝑙𝑙𝑙𝑚𝑚 ∗ 32.2 𝑓𝑓𝑡𝑡𝑠𝑠2

0.076474 lb𝑚𝑚𝑓𝑓𝑡𝑡3 ∗ 1.5 ∗ �16𝑓𝑓𝑡𝑡𝑠𝑠 �

2 = 21.9 𝑓𝑓𝑡𝑡2

Table 5.3: 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

66

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

= ⋅ =

⋅

Figure 5.4: Main Parachute Visualization

67

Table 5.4: 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

Rip-stop nylon is a well-known, preferred material for parachutes. The way the fabric is woven

makes it more resistant to tearing. This is desirable because the team can trust that a small tear will

not spread and ruin a whole parachute. Not only is it strong, but it is also a thin and lightweight

fabric. This will keep the rocket’s weight low and allow it to be easily stored inside the rocket

body.

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.

68

Figure 5.5: Pictures of Tender Descender in Undeployed and Deployed Configurations

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 750 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

69

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.

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

ensure a factor of safety for the device in loading conditions it will experience during flight and

recovery.

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.

Drift

The distance the rocket will drift during descent can be estimated with the following equation.

70

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 will descend under the drogue parachute from

an altitude of 5280 ft. to 750 ft. Then the rocket will separate and both the booster section and the

payload section will descend to the ground under their respective main parachutes at a velocity of

16 ft/s.

The rate of descent under drogue can be calculated with the following equation:

𝐷𝐷𝑆𝑆𝑠𝑠𝑐𝑐𝑆𝑆𝑖𝑖𝑡𝑡 𝑉𝑉𝑆𝑆𝑙𝑙𝑉𝑉𝑐𝑐𝑖𝑖𝑡𝑡𝑉𝑉 = �2 ∗ 𝐹𝐹𝑉𝑉𝐷𝐷𝑐𝑐𝑆𝑆

𝐴𝐴𝑖𝑖𝐷𝐷 𝐷𝐷𝑆𝑆𝑖𝑖𝑠𝑠𝑖𝑖𝑡𝑡𝑉𝑉 ∗ 𝐷𝐷𝐷𝐷𝑎𝑎𝑎𝑎 𝐶𝐶𝑉𝑉𝑆𝑆𝑓𝑓𝑓𝑓𝑖𝑖𝑐𝑐𝑖𝑖𝑆𝑆𝑖𝑖𝑡𝑡 ∗ 𝑃𝑃𝑎𝑎𝐷𝐷𝑎𝑎𝑐𝑐ℎ𝐴𝐴𝑡𝑡𝑆𝑆 𝐴𝐴𝐷𝐷𝑆𝑆𝑎𝑎

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.

71

Table 5.5: Drift Calculations

Ejection System

The Auburn team is utilizing a custom ejection system employing 12g CO2 cartridges rather than

a typical black powder charge found in most rockets of this size. The team has deemed ejection

via CO2 safer, more reliable, and less detrimental to the rockets and its contents if there were a

malfunction during flight.

The team seeks to replace the large black powder charges that separate the rocket via explosive

force, which is the common method of separation in amateur rocketry. Obviously, black power

introduces risks since it’s a highly explosive material. A misfire or electrical-magnetic interference

could inadvertently ignite the black powder and cause damage to the rocket or anyone working

with it. Due to these risks, the team made the goal of minimizing the amount of black powder used

in recovery, and switched to CO2 ejection systems last year. However, these CO2 ejection systems

were still activated by a small amount of black powder.

This year’s team came up with a new idea to completely remove black powder from the custom

CO2 system. Utilizing a magnetically driven solenoid, the CO2 cartridges can be punctured via

electromotive force, instead of the explosive force from black powder. The solenoid ejection

system is activated by a custom built, altimeter driven electronic circuit, as seen in Figure 5.8.

Wind Speed (mph) Wind Speed (ft/s) Drift Under Drogue (ft) Drift Under Main (ft) Total Drift (ft)5 7.33 511.31 343.75 855.06

7.5 11.00 766.97 515.63 1282.5910 14.67 1022.63 687.50 1710.13

12.5 18.33 1278.28 859.38 2137.6615 22.00 1533.94 1031.25 2565.19

17.5 25.67 1789.60 1203.13 2992.7220 29.33 2045.25 1375.00 3420.25

72

Figure 5.6: Solenoid Circuit

While the team has developed a working prototype of this system, the obstacle of integrating the

solenoid circuit into the constrained space of the rocket still remains. As such, the team decided

to leave the solenoid circuit as a project to be continued in next year’s Student Launch competition,

and instead implement an updated version of last year’s custom black powder driven CO2 ejection

system.

Auburn’s custom CO2 ejection system is activated when the altimeters fire an e-match that ignites

a small (0.15g) black powder charge, propelling a 12g CO2 cartridge into a pin. This punctures

the cartridge and releases the CO2, quickly pressurizing the inside of the rocket and causing the

separation and deployment of the parachute. All charges will be activated by the altimeters at an

altitude of 750ft. Since the inner components are equipped with rubber O-rings, the black powder

charges are completely sealed within each chamber and the combustion is effectively contained.

This system is mounted on the bulk plate just below the avionics bay, in the same compartment as

the booster main parachute.

This year’s edition of the custom CO2 system makes several improvements over the previous

model. The chambers are arranged in a pyramid configuration, keeping them closer to the center

of the bulk plate and increasing the amount of length available. The system is also assembled in a

way that is much more user friendly and reduces the chances of broken parts or premature

deployment. The casing is constructed out of high density polyethylene and the inner components

73

are aluminum. All parts are fabricated in house by our machinist. Auburn’s custom CO2 system

is shown in Figure 5.9 and Figure 5.10.

Figure 5.7: Custom CO2 System Housing

Figure 5.8: Custom CO2 System Assembly

Black powder analysis was performed to provide a backup means of separation should there be

unforeseen difficulties with the CO2 system. The following equation was used to estimate the

grams of black powder needed to pressurize the inside of the rocket body tube to between 5 psi

and 20 psi.

𝐵𝐵𝑙𝑙𝑎𝑎𝑐𝑐𝐵𝐵 𝑃𝑃𝑉𝑉𝑃𝑃𝑑𝑑𝑆𝑆𝐷𝐷 𝑊𝑊𝑆𝑆𝑖𝑖𝑎𝑎ℎ𝑡𝑡 (𝑎𝑎𝐷𝐷𝑎𝑎𝑔𝑔𝑠𝑠) = 𝐶𝐶 ∗ 𝐷𝐷2 ∗ 𝐿𝐿

In this equation, C is a constant related the desired pressure in PSI, D is the inside diameter of the

rocket body in inches, and L is the length of the rocket body tube. For a 5 inch diameter and 32

inch long section, between 1.6 g and 6.4 g will be needed. Electronic matches will be used to ignite

74

the black powder charge. To hold the black powder charge, charge cups were designed and

fabricated using a 3D printer. The charge cups were designed such that the head of an e-match can

be retained while allowing the wires of the electronic match to connect to the altimeters. The

charge cups are sealed with adhesive tape to contain the black powder. The black powder system

would incorporate a primary and secondary charge, the secondary providing redundancy should

the primary fail. The primary altimeter will be set to ignite the primary charge at the main deploy

height of 750 feet. The other will be programmed to ignite the secondary charge 650 feet to ensure

that both charges do not combust simultaneously.

Black powder ejection is designated as a backup because it’s been successfully utilized in past

launches, but the requirement of explosive material is preferably avoided. The black power could

be ignited by electromagnetic interference (EMI) or any sparks at any point, posing a risk for those

near the rocket and proper operation of the rocket.

Altimeters

The avionics bay will house two altimeters to satisfy redundant system requirements. Both

altimeters will fire the fairing charge at the apogee height of 1 mile (5280 feet) to deploy the

fairings and thus the drogue parachute. Then both altimeters will fire the main deployment at an

altitude of 750 ft.

The team will be using one Altus Metrum TeleMega as the primary and one Altus Metrum

TeleMetrum as the secondary altimeters. Previous rockets used the TeleMetrum as the primary

altimeter and the PerfectFlite Stratologger as backup. The altimeters have been switched out due

to the expandability of the TeleMega. The TeleMega has 4 additional sets of pyro connectors,

allowing for future expansion if necessary. It can also have a second battery easily installed into

dedicated screw terminals for additional power for pyro ignition purposes. The TeleMega also has

a more advanced accelerometer for more detailed flight data acquisition. The previously used

altimeters required a standard 9V battery, which is larger and heavier than the battery used for the

Altus Metrum altimeters. Additionally, using two Altus Metrum altimeters will make

programming quicker and easier, as they share an interface program. This makes any last minute

or on site adjustments across both boards simpler. Should one of the Altus Metrum altimeters fail,

the PerfectFlite Mawd or Stratologger can be used as additional backup. All altimeters are capable

75

of tracking in flight data, apogee and main ignition, GPS tracking, and accurate altitude

measurement up to a maximum of 25,000 feet.

Figure 5.10: Altus Metrum TeleMetrum Altimeter

Another reason the Altus Metrum altimeters are preferred are their radio frequency (RF)

communication abilities. Both TeleMega and TeleMetrum are capable of communicating with a

Yagi-Uda antenna operated by the team at a safe distance at any point during the launch. It can be

monitored while idle on the ground or while in flight. While on the ground, referred to as “idle

mode”, the team can use the computer interface to ensure that all ejection charges are making

proper connections. Via the RF link, the main and apogee charges can be fired to verify

functionality, which was used to perform ground testing. The voltage level of the battery can also

Figure 5.9: Altus Metrum TeleMega Altimeter

76

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. The altimeter can even be rebooted. While in flight, referred to as “flight

mode”, the team can be constantly updated on the status of the rocket via the RF transceiver. It

will report altitude, battery voltage, igniter status, and GPS status. However, in flight mode,

settings can’t be configured and the communication is one way from the altimeter to the RF

receiver.

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, a new antenna will be used.

The Altus Metrum altimeters can have their whip antennas replaced with any antenna desired, so

an SMA cable will be connected to the board and run to the outside of the rocket. On the outside

the team will attach three flexible patch-antennae. The Taoglas FXP240 433 MHz ISM Antenna

is the team's selection and can be seen in Figure 5.13. The advantage of this antenna is it will

conform to the shape of the rocket and have a negligible effect on the aerodynamics of the rocket.

Figure 5.12: Taoglas FXP240 433 MHz ISM Antenna

77

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.

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 will be done wherever possible. Should a wire happen to be a resonant length and is unable

to be shortened or lengthened, a low-pass filter can be implemented to block the high frequency

noise.

The avionics bay will be approximately 10 inches in length, and have an inner diameter of 4.75

inches. This segment of the rocket will be the coupling link between payload and booster sections

of the rocket, so the length and diameter are both fixed. Within this bay, there will be a carbon

fiber board that slides into a set of rails, and on this board all the altimeters and batteries will be

attached. The altimeters and batteries will be mounted on opposing sides of the board, with one

battery and altimeter per side. Since carbon fiber is an effective shielding material (50dB

attenuation), this board will act as shielding between the two altimeters and minimize cross-talk

as well as near-field coupling. This board will also be easily removable for connecting the

altimeters to computers for configuration and for charging the altimeters’ batteries.

Attachment Hardware

The parachutes are attached to their respective bulk plates by a system of shock cord, quick links

and U-bolts.

The Auburn recovery team considered two different materials to be used for shock cord. Kevlar

was considered for shock cord and shroud lines because of its strength. Compared to another

material, such as nylon, Kevlar has a much higher strength to weight ratio— 2514 KN*m/kg

compared to 69 KN*m/kg. However, the team decided to use tubular nylon as the material for

78

shock cord. Tubular nylon consists of a nylon tube which is made from exceptionally high strength

material which is both light and strong. Tubular nylon is easy to handle and also cost efficient. The

wrap around webbing increases the overall strength per inch. Tubular nylon is highly flexible and

pliable. Due to its pliability, it tends to glide better over rough or jagged surfaces preventing the

wear and tear that occurs more with flat webbing. One inch width of tubular nylon webbing can

withstand about 4,000 pounds of pressure. By using a material known to be strong, the team

ensures failure is less likely to happen in this component.

The team will be using U-bolts to attach the parachutes to the bulk plates. U-bolts have proven to

be more reliable in the teams past experience because the shock cord is less likely to tangle on the

bolt. An eye-bolt is more susceptible to failure for this reason, as the shape allows cord to wrap

around it. Additionally, a U-bolt provides two points of attachment between the shock cord and

the bulk plate (where an eye-bolt has one), effectively halving the resultant force on each bulk

plate connection.

79

Section 6: Aerodynamic Analysis Payload

System Level Design Review

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, GPS, Accelerometer, Servos, and an Arduino.

Figure 6.1: 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 an inner bulk plate that allows them to protrude out

from 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.

Payload Structure

Fairing:

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

80

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 6.2: Grid Fin Fairing

Grid Fin:

The grid fins are lattice shape control surfaces. An illustration can be viewed in Figure 6.3. 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.

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

81

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 6.3: Aerodynamic Grid fin

Payload Electronics

The electronic subsystems for the WAFLE are the Arduino, Servos, Accelerometer, and GPS. The

Servos are located 4.1 inches aft of the top of the WAFLE. The servos are set within a cradle and

protrude from the airframe. The batteries for the system are forward of the servos and secured to

a bulk plate. The Arduino, Accelerometer, and GPS are aft of the servos secured to a bulk plate.

Rods run through the three bulk plates sandwiching them together and are permanently fixed to

the permanent bulk plate on the top of the WAFLE.

82

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 6.4: Arduino Uno

Servos

HiTec HS-5685MH Digital Super Torque Servo is a high torque servo connected to the Arduino

Uno. The servo is used to orient an attached object (grid fin) to a specific angle based on given

inputs. 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. The drag force created by the grid fins will create a moment on the

grid fins that needs to be countered by a large amount of torque by the servo. The calculated

maximum amount of torque needed will be 13.8 kg/cm for the entire flight. This servo had the best

trade-off between size (4.1 x 2 x 3.8 cm) and torque (11.3-12.9 kg/cm) and the calculated

maximum amount of torque was a generous estimate. There is also a large probability that the

maximum amount of torque needed will not be necessary at the specific flight conditions, therefore

83

the maximum amount of torque is overestimated. In the event of a miscalculated for the maximum

amount of torque necessary, a failure in the stability of the grid fin will not cause an overall failure

in the system. The grid fin will be able to readjust and slow down the rocket as before, but in an

extended period of time.

Figure 6.5: HiTec HS-5685MH Digital Super Torque Servo

Accelerometer

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. The ADXL335 is used to measure the acceleration of an object. This triple-axis sensor

allows the ADXL335 to record acceleration in the x, y, and z directions of the chip. This will allow

the sensor to instruct the Arduino when it will need to tell the servos to pitch. The ADXL335 is a

commercially available accelerometer. It also is compatible with the Arduino Uno. The team has

also had experience with the ADXL335. With a low operating voltage of 5 volts and a small size

of 1.9 x 1.9 x0.314 cm, the ADXL335 would be an ideal sensor for the team.

84

Figure 6.6: ADXL335 Triple-axis Accelerometer

GPS

A commercially available GPS tracker will be used within the WAFLE to validate acceleration of

the rocket and the height that Arduino will calculate. The other function for the GPS will be to

broadcast the location of the booster section. This will allow the booster section to be obtained

after touchdown.

Selection of the optimal GPS will be forthcoming. Research and testing will be performed to find

a GPS that functions ideally with the WAFLE system.

Wiring

The wiring for the WAFLE system is illustrated in the schematic. The voltage source supplies

power to the servos directly. It also powers the Arduino, which in turn powers the accelerometer.

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.

85

Figure 6.7: WAFLE Electronics Schematic

Design Requirements

Design requirements for the aerodynamic analysis payload set forth by team leadership and by

NASA are outlined in Table 6.1.

Table 6.1: Aerodynamic Analysis Payload Design Requirements

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.

86

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.

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

The gears of the servos will be imbedded into the U-bracket base of the grid fin by means of a metal bar.

87

of the grid fins throughout the flight.

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.

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.

Manufacturing and Assembly

The WAFLE system has very few subsystems that are manufactured. Most of the systems are

electronics were bought from a commercially available distributor. However, there are a few

subsystems and structural components to the WAFLE that will need to be manufactured to move

to the integration and assembly phase.

The grid fin fairings and grid fins will be manufactured from HIPS plastic. Printer time needs to

be allocated for the prints. Extra time needs to be accounted for in the event that anomalies occur

in the printing process.

A carbon fiber tube with an inner diameter of 5 inches and an outer diameter of 5.25 inches will

act as body of the WAFLE. This tube will have the other subsystems mounted to the exterior and

interior. The carbon fiber will be rolled from prepreg carbon fiber and will be baked in an autoclave

until the resin has fully cured. After the tube cures, the tube should then be trimmed to 15.85

inches.

The bulk plates that will hold the servos and other electronics in place will be made in a similar

manner to the carbon fiber tube. Carbon fiber prepreg will be cut into sheets and laid out to make

plates with dimensions 36 x 36 x 0.25 inches. These plates will be placed on the CNC router and

88

circular bulk plates will be machined from them. These bulk plates will be 5 inches in diameter

with a thickness of 0.25 inches.

The electronics will not undergo manufacturing. However, the accelerometer and GPS will have

to be wired into the Arduino. The servos and batteries will be wired into the Arduino once the

WAFLE is assembled.

Risk Mitigation

Table 6.2: Aerodynamic Analysis Payload Risk Mitigation

Failure Event Result Mitigation

Non-symmetrical deployment of grid fins Large change in trajectory

If large off axis acceleration is detected the fins will disengage to a stored

position.

Battery power runs out during flight

Arduino will now function and the grid fins will not be

deployed.

Fully charged batteries will be stored within the rocket

before launch.

Fairing screws fall out and the fairing falls off.

The fins will be at a high risk of destruction

Insure that the screws are secured before flight.

Mounting screws connecting the WAFLE section and the engine

section come out.

The rocket will fall apart and be terminated if the

engine is boosting.

Insure that the screws are mounted correctly.

Accelerometer fails The fins will stop deploying or shut down

Insure that the accelerometer is in working

order and backups are on hand during system checks

Payload Integration

The integration of the subsystems into WAFLE starts with the carbon fiber tube. Slots for the

servos will be cut into the tube. The slots should be 11.1 inches from the top of the tube and should

be the width and height of the servos. The permanent bulk plate for the parachute is then epoxied

89

to the top of the tube. This will have holes for the U-bolt as well as the linking rods that will hold

the bulk plates for the other subsystems. Threaded rods can be slid into the rods once the epoxy

dries. A bulk plate with the battery can be slid onto the linking rods first. Screws forward of the

plate to insure that it remains in the desired position. The servo cradle is slid in aft of the batteries

and will be 11.1 inches aft of the top of the tube. The Arduino bulk plate will follow that servo

cradle into the tube. The bulk plate array will be secure with bolts to insure that the plates do not

move.

The servos will be mounted into the cradle of the bulk plate array. The grid fins will be the next

components to be integrated into the WAFLE. The gear of the servo will slide into the slot of the

grid fin base and secured. The final subsystem that is integrated is the fairing. The fairing will be

placed forward of the servo and will be secured with screws into the airframe.

After the WAFLE is assembled, it can then be tested and certified. After certified, the WAFLE

will join the booster section to complete the lower section of the rocket.

Payload Concept Features and Definition

The concept of gird fins is a very new idea within the aerospace world. The invention of the grid

fin occurred in the 1970s by the Russians. The Russian’s used the grid fin as an aerodynamic

stabilizer for missiles on their fighter jet as well as for stabilizers on their ejection pod of their

launch vehicles. The United States replicated the Russian design on a few bombs and missiles.

The concept of the grid fin has not been implemented to as drag control surfaces by on a few

companies, and fewer companies have implemented them onto rockets. 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.

90

Science Value

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.

Payload Success Criteria

Table 6.3: Aerodynamic Payload Success Criteria

Criteria

Number Criteria Method of Validation

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.

Redundant timer will be implemented into the system to insure that the code

iteration does not engage.

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 tested

and verified for security.

AU4

Servos must remain in direct contact with the gears of the grid fins throughout the flight.

Testing and small tolerances between the gear and servo will insure stability

throughout flight.

91

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.

“Model In-The-Loop” testing and test flights will validate deployment precision

for the full scale.

AU7 Grid fins must stay

deployed under the force applied by the flow.

Wind tunnel testing and structure testing will insure stationary deployment.

Testing and Simulation

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 6.4: Aerodynamic Payload Simulations and Tests

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.

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.

92

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.

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.

Sub-Full Scale Test A full scale aerodynamic model with a working WAFLE system will be launched. This test will

validate the WAFLE system for the Full Scale Test.

Full Scale Test A full scale rocket with working payloads will be built and launched. The payload systems will be validated.

Computational Fluid Dynamics (CFD):

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.

93

SolidWorks Flow Simulation:

In order to execute the testing, simulation and inspection was conducted. The team did SolidWorks

fluid flow simulation on a 3D CAD model of a grid fin with ideal dimensions. SolidWorks Flow

Simulation is an intuitive CFD 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 6.8

shows an example of placing starting points of the flow over the face of the grid fin.

94

Figure 6.8: 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 6.9, the result were that the incoming flow was at 14.74097 lbf/in2 (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/in2 (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/in2 (turquoise) once the flow passed through

the lattices.

95

Figure 6.9: Flow directed over the grid fins.

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

Figure 6.10: 0.2 Mach flow over a grid fin.

96

placed perpendicular to is 0.2 Mach. In Figure 6.10, it represents the flow simulation of the grid

fin under a 0.2 Mach flow trajectory. The simulation was allowed to run at 200 iterations to give

a more precise result.

The result of the flow simulation shows that at a 0.2 Mach the incoming flow is at 14.83976 lbf/in2.

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/in2 (yellow-orange). Finally, when the flow passes through the lattices the pressure

is decreased to 14.22253 lbf/in2 (dark blue) to 14.37684 lbf/in2 (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 6.5: 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 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.

97

The required design parameters, in English units, were obtained to input into the program. The

outputs are as follows:

Table 6.6: Aerodynamic Payload Fortran- Flight and Dynamic model

Mach Number (0.1-0.8)

Atmosphere Temperature (ºR) (511.650 - 456.894 )

Atmospheric Pressure (lb/in2) (13.6802 lb/in2- 7.54617 lb/ in2)

Reference Length (66.174 in)

Reference Area (21.55 in2)

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)

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)

98

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)

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:

99

Figure 6.11: Total Fin Axial Force Coefficient versus Angle of Attack Mach 8 Low Pressure, Low Temperature

Figure 6.12: Fin Normal Force Coefficient versus Angle of Attack Mach 8 High Pressure, High 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)

-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


100

Figure 6.13: Fin Normal Force Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low Temperature

Figure 6.14: 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.

-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


-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


101

Figure 6.15: Fin Moment Coefficient versus Angle of Attack Mach 8 Low Pressure, Low Temperature

Figure 6.16: Fin Moment Coefficient versus Angle of Attack Mach 8 High Pressure, High Temperature

-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


-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


102

Figure 6.17: Fin Moment Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low Temperature

Figure 6.18: Fin Moment Coefficient versus Angle of Attack Mach 0.1 Low Pressure, Low Temperature

The drag is calculated using the following equation:

-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


-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


103

212 DD C A Vρ=

Sample calculations of the drag for varying degrees of alpha is shown below.

Table 6.7: 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

-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

104

10 -8453.448184 -9116.323926 -8470.675464 -9093.853561

11 -81.09937339 -1182.534279 -110.0120397 -1145.360851

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 6.8: Sample Data at Mach=0.8 High Pressure, High Temperature


-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

-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

105

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 6.9: Sample Data at Mach 0.1 Low Pressure, Low Temperature


-15 -299921.8777 -274851.1012 -300151.0799 -275119.6974

-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

106

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

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 6.10: Sample Data at Mach 0.1 High Pressure. High Temperature


-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

107

-5 84595.39335 -33738.95163 85679.32169 -32468.86751

-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

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:

108

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.

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 fairing and grid fin system. Next, 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.

109

Figure 6.19: 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 inactive

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:

110

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 fly separately from the

other payload systems. During this test, the WAFLE system will be tested to fulfill a desired height

requirement. The competition altitude will be chosen as the desired apogee and the system will be

programed to ensure the system does not surpass that programmed altitude. The altimeter will

record the apogee. Acceleration and height will also be recorded with the WAFLE system. If the

system is able to achieve the desired altitude safely, then the system will be certified to be

integrated into the Full Scale rocket along with the other payloads. The final Full Scale test, with

all payloads operation, will validate that the rocket can achieve the mile height requirement and

that the WAFLE system, and all other payloads, operates as desired.

111

Section 7: Payload Fairing (PLF)

System Level Design Review

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.

Figure 7.1: PLF Curvature

Several design changes have been incorporated in order to ensure the overall success of the

design. The overall height of the fairing was increased from 9 inches to 13 inches. This increased

volume allows more room for the recovery system. Figure 8.2 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). To prevent the fairing halves

from colliding with the rocket structure post-deployment, fabric will be attached to the sides of

the fairing allowing the fairing a maximum separation of 180° (PLF.REQ.3). 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

112

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.

Figure 7.2: Overall PLF Structure

Figure 7.3 shows the PLF in a partially deployed configuration, while Figure 7.4 shows half of the

PLF system in the undeployed configuration. The charge bay has been moved to the highest point

Figure 7.3: PLF in Partially Deployed Configuration

113

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 and a small amount of recovery

wadding. The recovery wadding will be used to ensure the protection of the payload and other vital

components (PLF.REQ.2). This side of the charge bay will be lined with carbon fiber to ensure

structural integrity of the charge bay during detonation (PLF.REQ.1). Ribs (Figure 7.4) 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.

To prevent air from entering the PLF during flight, a lip has been integrated onto Side A of the

PLF system. This lip will mate with a recess on Side B, therefore creating an aerodynamic seal.

To prevent premature separation, the two sides of the PLF will be connected by pins (PLF REQ.5).

These pin connections can be seen in Figure 7.4.

Materials

The fairing halves, the base, and the sleeve 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 charge bay and the hinges will also be additively manufactured; however, High

Impact Polystyrene (HIPS) will be utilized for production. This material performs well when

Figure 7.4: PLF Half

114

impacting and when subjected to bending. When compared to ABS, HIPS can with stand higher

temperatures.

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 hinge will

rotate on a small metal pin.

Design Requirements

Design requirements for the PLF set forth by team leadership and by NASA are outlined in Table

7.1.

Table 7.1: PLF Design 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.

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 hinges shall not deploy more than 180°

Fabric strips will be attached to the inside of the PLF preventing the halves from separating more than 180°

PLF.REQ.4 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.5 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 sid

115

Testing

A rigorous course of testing is in progress to ensure the strength of the structure and the reliability

of the operation of the payload fairing.

Aerodynamic Design Testing (Completed)

The overall aerodynamic design of the fairing was tested at various scales. A 1:5 subscale model

(Figure 7.5) of the entire rocket was tested in a wind tunnel. This test included a static 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 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.

Manufacturing of the full scale, fully-functional PLF has begun.

Figure 7.5: 1:5 Subscale Model

Figure 7.6: 3:5 Subscale PLF

116

Charge Deployment Testing

The deployment of the Payload Fairing will be 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 fairing and payload will remain structurally intact, a

series of test detonations of the black powder charge in the chamber will be completed

(PLF.REQ.1, PLF.REQ.2).

The first test will only include the charge bay portion of the Payload Fairing system. The team

will create a fixture for the chamber to be held in place while testing. Initially, ten test runs using

various black powder charge sizes will be done in order to find a maximum charge size the

chamber can withstand without breaking or being compromised. If ten test runs are not sufficient,

more test runs will be completed. Once a charge size has been determined for the chamber, the

team will proceed to conduct tests using a full scale grounded Payload Fairing. Based off of the

chamber testing, various black powder charge sizes will be chosen to be tested to determine which

charge size will most efficiently induce separation. The following equation will be utilized to

determine the charge sizes, where N is the amount of black powder in milligrams, F is the force in

pounds, and L is the length of the chamber (Knowles). A plot was also produced to show the

relationship between the amount of black powder and the forces produced.

𝑁𝑁 = 0.00052𝑥𝑥103(𝐹𝐹𝐿𝐿)

Figure 7.7: Charge Required for a Given Force (assume L = 1in)

01020304050

0 10 20 30 40 50 60 70 80 90 100

Char

ge S

ize

(mg)

Required Force (lb)

Force vs Charge Size

117

While testing, the team will prioritize safety. Safety equipment will be brought to the testing area,

as well as first aid supplies. A blast shield will be placed around the test fixture to protect the test

technicians from potential shrapnel created by the detonation. The team will be properly distanced

from the charge before detonation, and will safely detonate the black powder charge by triggering

the electronic match remotely. A fire extinguisher will also be brought to the test site.

Drag Strip Deployment Testing

This test is to ensure that the payload fairing system and drogue parachute activate properly at

speeds similar to the flight speed at or near apogee. The deployment system is expected to activate

at apogee. However, in the event that the deployment is either premature or late the recovery

system must still successfully deploy. The estimated max speed of the rocket is Mach 0.79, and

this test is designed to ensure the deployment of the fairing and drogue parachute within roughly

20% of that speed to certify that a premature activation will not cause detrimental harm to the

recovery system. The drogue parachute will be the rockets only form of a controlled descent until

the main parachute deploys at 1000 ft.; therefore, it is imperative that the fairing system activates

and the drogue parachute deploys correctly for a nominal descent.

The testing will be done by attaching a to-scale model of the fairing system and avionics section

to a vehicle, driving at various velocities, and activating the payload fairing system to deploy the

parachute system. This will certify that the rocket is ready for flight and the recovery system is

going to be functional. To perform this test an attachment system will be designed in such a way

that the rocket will be securely attached to the vehicle and simulate the forces acting on the rocket

at various points throughout flight. Once the test velocity is reached the fairing system will be

activated by an altimeter. The Drag Strip Deployment Test Matrix is shown below.

Table 7.2: Drag Strip Deployment Test Matrix

Speed (mph) Description

Run 1 10 Runs 1-3 are a proof of concept that the parachute will deploy at low apogee speeds. Successful deployment will result in more testing.

Run 2 10

Run 3 10

Run 4 30 Runs 4 - 12 are to ensure proper deployment of the parachute in the event of a late fairing activation. Run 5 30

118

Run 6 30

Run 7 60

Run 8 60

Run 9 60

Run 10 100

Runs 10 - 12 are to ensure proper deployment of the parachute in the event of a premature fairing activation.

Run 11 100

Run 12 100

Run 13 120

Run 14 120

Run 15 120

The drag strip deployment test will be performed safely, by ensuring that the black powder in the

fairing system is handled correctly, that the vehicle is driven by someone licensed by the

government to do so, and that the rocket is attached securely to the vehicle. Safety equipment will

be worn, road laws obeyed, and the test will be performed far away from any pedestrians,

bystanders, and other vehicles.

Full Scale Testing

After all of the above tests have been completed and all minor design changes have been finalized,

the PLF system will be integrated and launched as part of a full scale rocket.

Science Value

Traditionally, a payload fairing, or PLF, is used to protect a scientific payload from the pressures

and atmospheric heating experienced by the vehicle during launch. Once a launch vehicle has

reached an altitude where atmospheric density is negligible, the PLF is jettisoned and allowed to

fall back to Earth. However, for the purposes of Project Aquila, the PLF will be used to protect a

section of the recovery system from the forces experienced during launch. The PLF will also act

as an aerodynamic nose cone to ensure rocket stability until apogee is reached.

119

Section 8: Safety

Checklists

Final Assembly Checklist

Final Rocket Assembly

Initial Check-off Points

Check rocket tube for any structural imperfections acquired during transport.

Check rocket tube for structural integrity and flight readiness.

Check payload fairing and grid fins for any structural imperfections acquired during transport.

Check payload fairing and grid fins for proper functioning and mission readiness.

Check parachutes for any imperfections that could be a problem during recovery operations.

Check parachutes and parachute bags for flight readiness.

Check avionics for proper functioning.

Check carbon dioxide expulsion system for flight readiness and ensure all parts are functioning correctly.

Check motor casing for any structural imperfections acquired during transport.

Check motor mount, motor casing, and thrust plate for flight readiness.

Check the couplers for structural integrity and flight readiness.

Check primary fins for structural integrity and flight readiness.

Check shock cords for flight readiness.

Check the assembled fin section with motor mount, motor casing, and other motor mounting items.

Assemble carbon dioxide expulsion system and pack drogue chute.

Pack main chute and insert into position.

Assemble avionics bay and attach shock cord to avionics bay.

120

Attach avionics bay using proper fasteners (i.e. bolts or shear pins).

Attach payload bay and nose cone section to the rest of the rocket.

Check all connections and assemblies on the rocket.

Insert rocket motor into motor casing.

Complete final check of the assembled rocket.

AUSL Safety Officer Signature

X

AUSL President Signature

X

121

Launch Procedures Checklist

Final Construction Check:


Check for proper connections between all the sections.

Check the main body tube for final flight readiness.

Check launch lugs for proper operation.

Check fins and fin connections for final flight readiness.

Check engine mount for final flight readiness.

Overall rocket construction readiness check.

Final Scientific Payloads Check:


Check to ensure all grid fins are undamaged and ready for flight.

Check payload fairing to ensure it is ready for flight.

Check any cameras to ensure they are on and ready for the flight.

Check charge status of electrical system for the launch-pad.

Overall launch-pad readiness check.

Final Overall Systems Check:


Final overall check of rocket construction.

Final overall check of launch-pad and rail.

Final overall check of ground support systems readiness.

122

Final overall check of personnel and observers readiness.

Final overall launch readiness check.


X


X

123

Launch Procedures Check:


Place rocket on the launch rails ready for mission process.

Have unnecessary personnel move to safe location for launch process.

Attach rocket umbilical for launch sequence.

Have qualified personnel place electronic igniter on insertion/ ignition device.

Remove mechanical system safeties.

Have all personnel move to proper launch operations locations.

Ensure launch area is cleared for system operation.

Initialize mission process.

Receive proper feedback from the rocket.

Receive proper feedback on rocket reaching proper launch angle.

Check with range officer to ensure range is all clear and ready for launch.

Receive final all clear for launch readiness.

Initiate motor ignition.

Check for proper ignition.


X


X

124

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 check lists 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.

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.

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.

125

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

o Autoclave not set on correct setting

• Aluminum mandrel

o Hands caught in mandrel

126

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

reading over the operator’s manual for the autoclave, alone with looking over the mitigation table

that has been put in place.

127

Table 9.1: Risk Mitigation Table - Airframe

Potential Risk Potential Effect Impact Risk1 Mitigation Risk2

Airframe not properly

transported (1)

Damage to airframe 4 3

Airframe will be transported in a custom made

shipping container to protect it from

damaging vibrations or slipping

1

Airframe not

properly stored (2)


Airframe will be stored in a specified container in a dry

and cool room when not being constructed

or tested

1

Cracks in Airframe (3)

Breaks on launch injuring team members or bystanders

5

3

Airframe will be inspected at every

stage of construction and a pre-launch

inspection with the use of a checklist

will be conducted 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 will be constructed using

specialized tools to ensure exact

dimensions and a prelaunch inspection will be conducted 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 will be mixed in correct proportions and airframe will be

inspected during construction and

1

128

before launch for lack of epoxy

Collision with bird (6)


Testing will be performed on nose

cone and airframe to ensure strength is

sufficient to withstand collisions and during all flights

the sky will be checked for nearby

birds in flight

1

Airframe breaks apart in flight (7)

High speed objects falling on spectators

5 3

Strain and stress tests will be performed on sample materials to confirm integrity of

materials under more than expected

conditions

Table 9.2: Risk Mitigation Table - Autoclave


Debris flies up into user’s eyes (1)

Trauma to the user’s eyes 3 3

Students will wear safety glasses or face

shield while operating autoclave

1

Material left in autoclave (2)

Damage to autoclave and

material 4 3

Users will ascertain that the autoclave is

empty before operating it.

1

Door not properly closed (3)

Damage to material inside autoclave 2 3 Make sure the door is

fully closed. 1

Wrong cycle selected (4)

Damage to material inside autoclave 2 3

Only authorized and trained users may operate autoclave.

1

Material experiences

Can cause severe injuries to users 5 2 Wear proper PPE and

always keep hands, 1

129

explosive breakage when autoclave is

opened (5)

head, and face clear while opening.

Touching hot materials (6)

Severe burns to users 4 3

Wear proper PPE such as heat and cut

resistant gloves, leather apron, and

leather sleeve protector.

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

In the case of a fire a fire extinguisher

must be kept in the same room as the autoclave and be

easily accessible. If fire spreads contact 911 immediately.

2

Toxic Fumes (8) Can cause respiratory problems

5 5

Respirators will be required when working with

potentially hazardous materials. Proper

ventilation of the lab at all times.

1

Unauthorized use (9)

Damage to Autoclave,

materials, and to personnel

5 3

Lab where autoclave is located is locked up by authorized

personnel. Autoclave is also locked to

prevent unauthorized use.

1

Table 9.3: Risk Mitigation Table - Filament Winder


130

Mandrel not secured properly

(1)

Improper construction of

rocket body tubes leading to

structural failure

5 3

Only trained team members will use the

filament winder, making sure the

mandrel is properly secured with winder grippers and a pin

through the mandrel. Constant supervision of equipment while

in use.

1

Improper winding angles for the

specific stresses occurred during

flight (2)



structural failure

4 3

Material testing will be performed on samples to ensure

that 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 Constant supervision of mandrel while it is

operating 1

Filament does not unroll correctly (4)



structural failure or damage to equipment

4 3

Winder will be watched closely during winding

process

1

Table 9.4: Risk Mitigation Table - Carbon Fiber


Allergic reaction from coming in

contact with carbon fiber (1)

Skin irritation 3 4 Wear proper PPE

when handling carbon fiber

1

131

Debris flies up into users eyes (2)

Trauma to the users eyes 3 3

Wear safety glasses when working with

carbon fiber 1

Toxic particles (3) Respiratory irritation 3 3

Wear proper breathing apparatus when working with

carbon fiber

1

Electrical shock (4) Burn or electrocution 4 2

Carbon fiber is electrically

conductive so it should be kept away

from electrical equipment or

machinery

1

Table 9.5: Risk Mitigation Tables - Epoxy


Improper Ventilation (1)

Vapors can cause headache, nausea,

and irritate the respiratory system

4 5

Keep lab ventilated at all times when

working with epoxy. Also wear respiratory PPE when working

with epoxy.

2

Skin Contact (2) Can cause skin irritation 2 5

Wear proper lab clothing when

working with epoxy. If epoxy gets on skin wash off with soap

and water

1

Degradation of Epoxy Resin (3)

Bonds weakly resulting in parts that break easily

4 3

Epoxy will be stored in an air conditioned lab between 40°F and

120°F

1

Spilling and leaking (4)

Hardens on work table or lab equipment

2 4 Handle Epoxy

carefully. If spilled, use paper towel to clean up and stop

1

132

damaging the equipment

leakage. Use warm water and soap to clean up messes

immediately

Fire Hazard (5) Damage to lab

area, equipment, and personnel

5 3

Keep epoxy away from high heat

sources. If fire starts use Foam or carbon dioxide to put out.

Fire extinguisher will be stored in an easily accessible location in

lab

1

Epoxy gets in user’s eyes (6)

Damage to the user’s eyes 5 2 Wear safety glasses

while using epoxy 1

Epoxy setting up before work is

finished (7)

Waist of epoxy that is not used 2 3

Epoxy will be mixed in small amounts

when needed 1

Epoxy burning through container

(8)

Potential fire hazard and damage

to lab 2 3

Never mix epoxy and leave it unattended

and be aware of how hot the epoxy is as it

starts to set

1

Epoxy not properly disposed (9)


to lab 2 3

Cure epoxy and let it cool before disposing

to prevent possible fire

1

Environmental Effects

When constructing the airframe there are environmental concerns that will be addressed. These

concerns include how the airframe affects the environment and also how the environment affects

the airframe. A risk mitigation table has been put in place for airframe environment effects to make

team members aware of the impact they can have on each other. This mitigation table has been

listed below.

133

Table 9.6: Risk Mitigation Tables - Airframe Environment Effects


Harmful toxic fumes released into

environment by autoclave (1)

Damage to environment and toxic air supply

4 3

Make sure that autoclave is always properly ventilated before operating

1

Epoxy not properly disposed (2)


to lab 2 3

Cure epoxy and let it cool before disposing

to prevent possible fire

1

Airframe not recovered after

launch (3)

Hazard to the environment from carbon fiber and

epoxy

3 2

Airframe will be tracked using a GPS during the launch to ensure that it will not

be lost

1

Airframe Risk Mitigation – Testing Systems

Table 9.7: Risk Mitigation Tables – Wind Tunnel Testing


Debris in the wind tunnel (1)

Damage to wind tunnel, object

being tested, or personnel

4 3

Inspect test object to ensure it will not

break. Inspect wind tunnel before use for

loose debris

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 Check that doors are

securely locked before each test

1

Inexperienced personnel (3)

Damage to project and equipment due

to incorrect operation of the

5 3

Lab with wind tunnels will be

locked to prevent any unauthorized use

1

134

wind tunnel or personnel injury

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 speed will be limited to less than 160 ft per second. Only authorized personnel can

operate the wind tunnel

1

Overusing Motor (5)

Engine becomes damaged and

would cost large amounts of money to repair or replace

5 3

Scheduling for use of the wind tunnel will

be necessary. Periodic checks of the system will be performed to keep

engine running properly

1

Table 9.8: Risk Mitigation Tables – Tensile Test Rig


Object being tested is improperly

aligned (1)

Results acquired from tests are

incorrect and result in a weaker rocket

in the future

4 4

Operation will be supervised by a

trained member of the faculty at all

times

1

Fractured particles during test (2)

Irritation to eyes or injury from dust or high speed particles

4 4

All personnel must stay a safe distance away from tensile test rig while in

operation. Goggles are required while

equipment is running

1

Large forces generated with

incorrect operation (3)

Bodily 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

1

135


Damage to machine,

personnel, and projects

5 2

Machine will be 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

Scientific Payloads Hazard Analysis

During the process of building a rocket, safety is constantly kept in mind. With the design concept

for this year’s payload integration being, it is being thought of even more so. There will be

guidelines implemented to ensure the safety of the members of the scientific payload team while

the construction and testing of the system is occurring. There are three different sections that are

being looked at while considering failure modes for safety protocols for the scientific payloads:

operations, materials, and construction processes.


• Mission Processes

• Testing

• Personnel Risks (Operator and Observers)

• Environmental Risks (Macro and Micro)

• Vehicle Risks (Launch, Flight, and Recovery)

• Controller Risks (Electrical and Mechanical)

Construction Failure Modes:

• Hand Tools

• Soldering Equipment

• Drill Press

• Band Saw

• Autoclave

136

• 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

Scientific Payload Risk Mitigation – Payload Fairing

Table 9.9: Risk Mitigation Table - Operations


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 will be stored in a safe container and only be interacted

with via an electronic ignition

that will be connected to the

altimeter.

1

Premature charge ignition on ascent

(2)

Premature fairing

separation, compromised

and

5 2

The black powder will only be

interacted with via an electronic

ignition that will be

1

137

uncontrolled flight.

connected to the altimeter. Drag strip

deployment testing will be performed to

ensure premature separation does not jeopardize recovery.

Black powder fails to ignite (3)

No fairing separation, failure to deploy

parachute, uncontrolled

descent.

5 3

Two electronic matches will be

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 will be made of a sturdy

material and placed to reduce

unnecessary stress. Fabric strips will be placed to ensure the

hinges to not overextend beyond

180 degrees.

1

PLF is damaged during flight or on

landing (5) 2 2

The material and design of the PLF will be tested to

ensure it will withstand any forces it will

encounter. Carbon fiber ribs will structurally

reinforce the PLF.

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

Extensive testing will be done to

determine that the materials and

structure designed for the PLF will withstand any

1

138

forces directed at the PLF. Placement

and size of the charges will also be considered, tested,

and modified to avoid damage.

Carbon fiber ribs will reinforce the PLF internally.

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

Extensive testing will be done to determine the

optimal charge size. If deemed necessary,

reinforcement may be added to the charge bay to

ensure structural integrity.

1

PLF structurally compromised by

aerodynamic forces in flight (8)

Operation of PLF is

compromised, flight of the

rocket may be compromised

4 2

The aerodynamic forces will be

simulated and their danger mitigated in wind tunnel testing

and subscale launches prior to

competition.

1

Table 9.10: Risk Mitigation Table – Payload Fairing Testing


Three Point Bending Test

specimen partially shatters (1)

Sharp debris would be left around the

testing area. 2 1

The test specimen will be carefully

manufactured and loaded to ensure that

it will bend and fracture but not

shatter.

1

139

Three Point Bending Test

specimen improperly loaded

(2)

Test results may not be accurate and may interfere with

further testing.

2 1

Team members will be knowledgeable

about 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 1

The team will be certain to

manufacture a representative specimen that

accurately reflects the properties of the

rocket.

1

PLF activation damages vehicle during drag strip

testing (4)

The vehicle’s operator could be

harmed. The vehicle or rocket mount could be made unsuitable

for further testing and require repairs.

3 1

Caution and creativity will be exercised when

designing a method of mounting the

rocket model that will avoid contact

with the vehicle and maintain a safe

distance from the walls of the vehicle.

1

Rocket model comes loose from

vehicle (5)

The vehicle or rocket model could become damaged, and potentially the

operation of the vehicle

compromised.

3 1

The rocket model and its mount will be

checked several times prior to testing and no unnecessary

force or use of it will occur.

1

Vehicle breaks during testing (6)

Delays to the testing, potentially harm to the vehicle

operator, rocket model, or mount.

2 1

The vehicle will be well maintained and

inspected prior to testing to ensure proper operation.

1

Obstacle exists in testing area (7)

The obstacle could cause unexpected changes to testing,

or if impacted

2 1 A suitable testing

area will be determined in

advance of testing

1

140

during testing could cause

complications to the vehicle, rocket model, or operator.

that has no obstacles. This will be verified again on the day of the test and the test will be postponed if potential obstacles cannot be removed.

PLF activation or testing damages

recovery system (8)

Repair or replacement of the recovery system will be required,

potentially delaying further

testing

3 2

The PLF will be carefully designed to

minimize any potential for damage

to the system it is supposed to protect.

1

Charge Deployment

Testing throws shrapnel (9)

Shrapnel can injure nearby testers or damage elements

of the test

4 2

Team members will be kept at a safe

distance and a blast shield will be utilized

for protection

1

Ignition of black powder during

handling or setup (10)

Injury can occur to team members or to elements of the

test nearby

3 3

Team members will handle the black

powder with extreme caution and minimize

time handling the powder

1

Black powder fails to ignite during

testing (11)

Team members must remove the unignited black

powder, exposing them to risk if there

is a delay in the electric signal

2 1

Team members will wait an extended portion of time

before approaching the test model and

will interact with the model and black

powder carefully and with hand and eye

protection.

1

Fumes from black powder charge

testing are inhaled (12)

Team members may experience adverse health

effects of inhaling fumes and particles.

2 4 Care will be taken to properly ventilate the

testing area. 1

141

PLF breaks into several pieces from charge testing (13)

Fragments of the PLF could cause

cuts or pierce shoes of tester during

clean up and repair

2 4

All team members will wear thick,

close-toed shoes and be very observant when approaching the testing model.

The testing area will be cleared of any

fragments immediately after it is safe to approach.

2

Table 9.11: Risk Mitigation Table – Payload Fairing Construction


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

Team members will take care to minimize interaction with the

3D Printer and ensure no

unnecessary items of clothing will be worn

and hair will be secured.

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 will wait a sufficient time after the conclusion

of the printing process and

confirmation via software to interact

with the final product and will 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

2 4

Team members will not operate directly with the 3D printer during its operation. Any alterations to its

1

142

becoming caught in machinery that

continues to operate. This can

result in damage or harm to these appendages.

process during operation will be

made with software and will wait a

sufficient amount of time to allow

operations to finish before interaction.

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 will monitor the 3D printer and its

progress periodically rather than

continuously and that any other work or construction will

occur 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 will not be allowed to

alter or tamper with the inner electronics of the 3D Printer and will only be serviced

by knowledgeable members if it is unplugged and

properly grounded.

1

Scientific Payload Risk Mitigation – WAFLE

Table 9.12: Risk Mitigation Table – Operations


Grid fins do not deploy in any capacity (1)

No data is gathered on the

grid fins and stability is

2 3 A full aerodynamic analysis of the grid fins is conducted

through

1

143

slightly compromised

computational fluid dynamics (CFD), subsystem wind

tunnel testing, and in-flight sensors. When the Arduino detects

that apogee has occurred, the fins will

be deployed.

All 4 grid fins do not deploy

simultaneously (2)

Significant instability due to unbalanced aerodynamic

forces

5 1

A full aerodynamic analysis of the grid fins is conducted

through computational fluid dynamics (CFD), subsystem wind

tunnel testing, and in-flight sensors. When the Arduino detects

that apogee has occurred, the fins will be deployed. Servos will be checked for

redundancy and proper function. All

wiring will be checked for security

and proper connections.

1

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

1

144

the fairing to be hard mounted to the

airframe.

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 insure that the code iteration does not engage until after boost phase and

correct altitude. A full aerodynamic analysis

of the grid fins is conducted through computational fluid dynamics (CFD), subsystem wind

tunnel testing, and in-flight sensors.

1

Electronics detach or become loose during flight (5)

Center of gravity will

change causing slight to

significant instability

which may lead to undesirable

flight path and/or

malfunction of grid fins

4 2

Careful and extensive measures will be taken to insure all

electronics are securely attached to a stationary plate within

the airframe. The plate as well as all

mounting bolts will be extensively tested for security and ability to

handle all stresses.

1

Electronics fail to come online after

boost (6)

No data is gathered on the



2 1

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

1

145

adhered to a stationary plate within

the airframe. This plate and mounting

bolds will be secured to a stationary plate within the rocket.

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



tunnel testing, and in-flight sensors. 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. If large off axis acceleration is

detected the fins will disengage to a stored

position.

1

Malfunction with WAFLE system (8)

Vehicle trajectory will

be altered resulting

undesirable flight path and

potentially collateral

damage and/or loss of asset

2 1

The team insures 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

2 1 In order to evaluate

how the grid fins will interact once

deployed, the team

1

146

undesirable trajectory or debris to fall back to the

Earth




implemented to act as the primary grid fin.

Improper battery power or voltage

(10)

Electronics will fail. No data is gathered on the



1 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



tunnel testing, and in-flight sensors. If large off axis acceleration is detected the fins will disengage to a stored

position.

1

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



tunnel testing, and in-flight sensors. If large off axis acceleration is

1

147


position.

Table 9.13: Risk Mitigation Table – WAFLE Testing


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 will be 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 will be conducted on all

components.

1

Payload Fairings impact hard

surface or tools (2)

The fairings are damaged and

require re-manufacture to

assure they fit and function properly

3 2

Great care will be taken in the handling

of the payload fairings at all times

1

Grid fins motion improperly while

they are being worked on by hand

(3)

Pinching of the fingers or hand

working on the fins and momentary

discomfort

1 2

Hand work on the grid fins will be short and focused. This is to avoid extended

contact, or injury due to complacency. Any

personnel working on grid fins will have proper knowledge of grid fin operation to be aware of pinch

points.

1

Grid fins are structurally

compromised by aerodynamic loads

The grid fins will need to be

remanufactured and potentially

4 1 The grid fins will be well-manufactured and the test well-

monitored to ensure

1

148

in wind tunnel testing (4)

strengthened for future testing.

that the aerodynamic loads applied are

proper for testing and do not exceed the

grid fin’s limitations.

Grid fins deploy unsymmetrically in wind tunnel testing

(5)

Potential movement in the

test model the grid fins are attached to.

3 1

The test model will 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

Unintended items, such as screws, bolts, or small tools, enter the

testing area during wind or water tunnel tests (6)

Unpredictable interaction and

potential harm to the grid fins, test

model, wind/water tunnel, or team

members.

5 2

Team members will 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 will be accounted for before beginning the tests.

2

Table 9.14: Risk Mitigation Table – WAFLE Construction


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 will be

done on the material and the general

design throughout to ensure the fins meet

the expected requirements.

1

Improper material used in the

manufacturing of

The payload fairings may cause

unexpected 2 3

Extensive research and comprehensive

testing will take place to verify that the chosen material

1

149

the payload fairings (2)

instability or complications.

will meet the requirements and

will function properly.

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 will be 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 will be given to the manufacturing

process and extensive testing will

confirm that the properties of the

manufactured part match that of the

design.

1

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 will be

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 conditions for storage will be

reviewed prior to beginning of

construction and will be immediately

available afterwards. Included in these are a sizeable space that

is dry, room temperature, and not crowded by tools or

1

150

other materials or project pieces.

Table 9.15: Risk Mitigation Table – WAFLE Materials

Material Potential Effect Impact Risk1 Mitigation Risk2

Batteries (1) Insufficient power 2 2

A new battery will be used and tested with an electronic multi-

tester to ensure proper function. Fully

charged batteries will 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

HIPS – High Impact

Polystyrene (3)

Structural Failure 2 1

In order to evaluate how the grid fins will

interact once deployed, the team




implemented to act as the primary grid fin.

1

151

Aluminum (4) Structural Failure 4 1

Material was chosen for its light weight

and structural integrity

1

Arduino Uno (5)

Electronic failure or

undesirable grid fin

deployment

3 2

All electronics and computing will be

extensively tested to ensure reliability and

redundancy of accurate flight path

corrections. 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 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

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

1

152

to the high amount of torque provided.

Adhesives (Epoxy, Flux and Soldering

Materials (9)

Copper wires may detach. 2 1

Adhesives will be tested with the full

aerodynamic analysis of the grid fins.

Conducted through computational fluid dynamics (CFD), subsystem wind

tunnel testing, and in-flight sensors.

1

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.

Table 9.16: 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 will be

followed by the team members to ensure a proper packing of the

parachutes. A checklist will be filled out during the packing process and signed off by both the

recovery team lead and safety officer

1

Parachute tears (2)

The parachute fabric material is torn causing the

rocket to fall in an

5 3

Fabric material of the parachute will be

strength tested before actual use. Container in which the parachute is kept will not contain

1

153

uncontrolled manner

any sharp edges. Parachute will be reinforced at any

potential tear locations

Parachute fails to deploy (3)

Parachute fails to deploy causing the rocket to fall in an

uncontrolled manner

5 4

Multiple tests will occur with the parachute to ensure the parachute

will deploy. On the day of launch systems will be checked to ensure

the parachute will deploy at the proper time. Payload fairing

has testing requirements that must be obtained to ensure proper parachute

deployment.

2

The shock cords break

after deployment of parachutes. (4)

Uncontrolled descent of the

rocket with potential crowd endangerment.

5 3

Shock cords will be subjected to tensile testing to ensure the

strength capabilities of the chosen cord

material.

1

Tensile strength test of shock chord back lashes (5)

Damage to body parts of the

workers involved 4 3

All workers when performing the tensile

test must stay an appropriate distance

away from the testing area. People performing the tensile test must also

be wearing safety glasses and appropriate

lab clothing.

1

Winds blow rocket off course. (6)

Rocket could become lost,

damaged, or could endanger observers.

5 3

The rocket will not be launched in improper

weather conditions. All parts of the rocket will

have a GPS locater device securely

attached.

1

The parachute deploys at the

Structural damage to rocket causing

5 4 Recovery systems will be thoroughly tested

2

154

incorrect time. (7)

unsafe descent or location of descent

potentially endangering observers.

prior to flight operations, and

checklists will be completed and signed

off by the recovery team lead and the safety

officer.

The altimeter fails. (8)

The parachute deploys at

incorrect time or not at all resulting

in structural damage or

uncontrolled descent.

Potentially endangering observers.

5 3

Extensive testing will be performed on flight

computer and associated electronics

ensuring proper functioning. During testing and prior to

launch, checklists will be filled out and signed by proper supervisors.

1

The drogue parachute fails to deploy. (9)

Uncontrolled descent until main parachute opening then resulting in

structural damage with potential

endangerment of observers.

5 4

Drogue deployment systems will be

thoroughly tested, checked off, and signed

off prior to launch operations.

2

Table 9.17: Risk Mitigation Tables - Wind Tunnel Testing


Debris in the wind tunnel (1)

Damage to wind tunnel or object

being tested 4 3

Inspect test object to ensure it will not

break. Inspect wind tunnel before use for

loose debris

1

Open test section (2)

Incorrect results calculated from the

wind tunnel that can have

potentially damaging effects

5 2

Check that doors are securely locked

before each test and all test equipment is properly calibrated

1

155


Inexperienced personnel (3)

Damage to project and equipment due

to incorrect operation of the

wind tunnel

5 3

Lab with wind tunnels will be

locked to prevent any unauthorized use

1

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 speed will be limited to less than 160 ft per second. Only authorized personnel can

operate the wind tunnel

1

Overusing Motor (5)

Engine becomes damaged and

would cost large amounts of money to repair or replace

5 3

Scheduling for use of the wind tunnel will

be necessary. Periodic checks of the system will be performed to keep

engine running properly

1

Table 9.18: Risk Mitigation Table - Tensile Test Rig


Object being tested is improperly

aligned (1)


incorrect and result in a weaker rocket

in the future

4 4

Operation will be supervised by a

trained member of the faculty at all

times

1


Irritation to eyes or injury from dust or high speed particles

4 4

All personnel must stay a safe distance away from tensile test rig while in

operation. Goggles are required while

equipment is running

1

156

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

1


Damage to machine,

personnel, and projects

5 2

Machine will be kept powered off in a

locked lab when not in use

1




of material

3 3


authorized personnel before testing

materials

1

Table 9.19: Risk Mitigation Tables - Shear Pin Test Rig


Shear pin being tested is improperly

aligned (1)


incorrect and have a damaging effect


4 4

Shear pin is carefully measured by an authorized team

member and double checked by a second

authorized team member to ensure proper alignment

1


Damage to eyes and body

extremities when the item being tested fractures

4 4

All personnel must stay a safe distance away from tensile

test rig while performing test.

Safety eyewear must also be worn along

with proper clothing covering body

extremities

1

157

Heavy weights and high forces

generated (3)

Body damage, specifically crushed body extremities, from misuse of machine while

testing

5 2

Don’t touch object being tested when

machine is active and stay a safe distance

away

1


Damage to machine,

personnel, and shear pin

5 2

Have machine locked up by an authorized

worker and keep power off

1




of material

3 3


authorized personnel to make sure they

have the authorization to test a

shear pin

1

Recovery Risk Mitigation - Materials

Table 9.20: Risk Mitigation Table - Kevlar


Breathing in Fiber Dust (1)

Severe respiratory problems 4 5

Respirators will be required when

working with Kevlar 1

Fiber dust in eyes and on skin (2)

Can cause irritation to both eyes and

skin 3 4

Eye protection will be required in lab when people are

working with Kevlar. If dust gets in eyes

rinse out immediately with water

1

Contact with moving Kevlar

fiber (3)

Minor to severe lacerations 5 3

Only trained personnel will be

allowed to work with Kevlar. Work area

will be kept clear of people except

operator. Call 911 in

1

158

the case of any serious lacerations

Leaving in direct sunlight (4)

Discoloration of Kevlar 1 1

Kevlar will be stored in closed containers

in lab 1

Table 9.21: Risk Mitigation Tables - Nylon


Breathing in fiber dust (1)

Respiratory Problems 4 4

Respirators will be required when

working with Nylon 1

Fiber dust in eyes and on skin (2)

Can cause irritation to eyes and skin 3 4

Eye protection will be 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

Keep nylon away from sparks and open

flames. Keep fire extinguisher in the same room when

working with nylon. If skin is exposed to hot nylon submerge area in cold running

water and immediately seek medical attention

1

Table 9.22: Risk Mitigation Tables - Carbon Dioxide



CO2 gas can cause headaches, nausea,

and loss of 5 4

Lab will be ventilated at all times when

working with CO2. If 1

159

consciousness in high doses

working with large amounts of CO2 the

test will be performed outside.

Explosion of canisters containing

CO2 (2)

Canister shrapnel can cause serious cuts to the body

5 3

Cylinders will be stored upright or in a

proper storage device, in a well-

ventilated and secure area, protected from the weather. Storage

area temperatures will not exceed 100

°F

1

Broken O-Ring (3) CO2 can leak into the surrounding air 3 4

Periodic checks of O rings on canisters

will be implemented. All faulty O-rings will be replaced

immediately

1

Over pressurizing rocket (4)

Over pressurization can cause problems with deployment of the parachute and damage the rocket

4 3

Trained members of the recovery team

will determine appropriate amount

of CO2 pressure. Tests will be

performed to confirm calculations before

full scale use.

1

Under pressurizing rocket (5)

The parachute doesn’t come out at all resulting in the rocket becoming a

high speed projectile

4 4

Trained members of the recovery team

will determine appropriate amount

of CO2 pressure. Grounded tests will be performed before

full scale use

1

160

Table 9.23: Risk Mitigation Table - Black Powder



Black powder is hazardous to

respiratory system when inhaled. Also particles may form explosive mixtures

in air.

5 4

Lab will be kept ventilated at all times when working with

black powder. Ventilation masks

will be required when working with black

powder

1

Powder comes into contact with the

body (2)

Can irritate skin and eyes 4 3

Proper PPE will be required when

working with black powder. Wash skin if contact is made and flush large amounts of water into eyes if

contact is made there. Afterwards 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. Care will be practiced in order to

avoid impact or friction. Fire

extinguisher will be available at all times.

1

Improper storage (4)

Degrades material and possible combustion

resulting in injuries and loss of equipment

5 4

Black powder will be stored between 40°F

to 120°F in a cool dry place in a tightly

sealed container. It will not be stored

with any other flammables

1

161

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

Extensive testing will be done to ensure that the proper amount of

powder is used

1

Table 9.24: Risk Mitigation Table - Fiberglass


Ventilation issues (1)

Can cause respiratory problems

4 4

Lab will be properly ventilated and

respirators will be required when working with

fiberglass

Eye and Skin contact (2)

Can cause irritation with skin and eyes 3 5

Proper clothing and eye protection must

be worn when working with

fiberglass

Recovery Risk Mitigation - Construction

Table 9.25: Risk Mitigation Table - Orbital Sander


Injuries to hands and fingers from moving parts (1)

Injury or loss of extremities 5 4

Using thick gloves to operate the sander. Turn the sander off

when 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

162

Electric Shock (3) Electrocution 5 3

Sander will be stored and operated in a dry

lab and inspected regularly to ensure

that there is no exposed wiring

1

Unintentional Starting (4)

Damage to equipment,

projects, or bodily harm

5 4

Sander will be turned off and unplugged before it is moved.

Switch will be in 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 will be 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

Clean all work areas before and after every use of the orbital sander

1

Improper Work Attire (7) Damage to body 5 4

Proper clothing and PPE will be 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

Project is not secured down (9)

Damage to project and damage to

hands from high speed objects

4 3

Properly secure project with clamps before turning on

sander

1

Over-reaching (10) Severe cuts to body 5 2

Ensure proper footing and balance

while operating 1

163

sander. Always turn off sander before

performing another task

Improper Tool Maintenance (11)

Dull or ineffective tool that causes unsafe handling and damage to body or project

5 3

Sander must remain clean, sand paper

replaced periodically, and inspections made

on wires

1

Over Exerting Tool (12)

Causes damage to project due to

excessive force applied to tool

3 3 Operators will be trained in proper

operation of sander 1

Improper Tool Replacement Parts

(13)

Tool becomes unusable 3 3

Only use replacement parts intended for the orbital sander

1

Table 9.26: Risk Mitigation Table - Sewing Machine


Sewing over fingers (1)

Hurting fingers and causing irreparable

damage to the equipment

4 3

Operators will 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 Regular maintenance

of machine will occur before and

after use. Machine’s

1

164

chords will be looked over regularly.

Cord can be a tripping hazard (5)

Can cause people to trip and injure

themselves 3 3

Machine will be plugged in close to wall and the chord

will not be extended over any walkways

1

Table 9.27: Risk Mitigation Table - Hand Tools


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 will 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

Improper tool maintenance (3)

Damage to project or body from tools

breaking or not working as designed

5 4

Regular scheduled maintenance will

occur for all tools. Tools beyond repair will be disposed of.

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

165

Insecure workbench or

project (5)

Materials or tools slip and can cause injury to operator

5 4

Project will be secured properly by straps, clamps, or through help by a

work partners before any hand tool use.

1

Improper tool storage (6)

Damage to tools and potential for unauthorized use

5 4

All hand tools will have a designated place to be stored.

All tools will be kept under lock

1

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.


• 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

166

• 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 9.28: Risk Mitigation Table - Outreach Operations


Car Accident (1) Ranges from minor injuries to death 5 3

All participants will wear seatbelts and

only licensed drivers will operate motor vehicles. Anyone

being transported by team members will

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 will demonstrate how to perform all tasks for rocket completion

and help the children when needed. All

minors will be 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 will be operated by USLI

team members while wearing necessary

protective equipment.

1

167

Mid Air rocket collisions (4)

Rockets would not reach highest

altitude due to mid-air collision

1 2

Students’ rockets will be launched from

significant distances from each other. Rockets will be

launched one at a time

1

Rocket stands fall (5)

Failure of rocket launch 2 2

All equipment will be examined prior to departing for the

outreach event. Any non-functioning

equipment will be fixed or replaced.

1

Rockets fall in the woods (6)

Slight environmental contamination.

2 2

All rockets will not be designed to

achieve significant distance and all will

be recovered.

1

Table 9.29: Risk Mitigation Table - Outreach Construction


Children ignite motor at time other

than directed (1)

Trauma to hands, eyes, ears, nose, 5 2

Children will be under constant

supervision and any potentially

dangerous materials will be 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 will be under constant

supervision and any potentially

dangerous use of tools will result a

removal of the tool. The task will then be

completed by the

1

168

USLI outreach team for the child

Table 9.30: Risk Mitigation Table - Outreach Materials


Children Break Rocket model (1)

Student will not be able to launch a

rocket or participate in the primary

outreach activity

2 2

Students will be under constant

supervision and any misbehavior will be

handled appropriately

1

Hard pieces may hurt children (2)

Trauma to children hands, eyes, nose,

mouth, ears 2 2

Students will be under constant

supervision and any misbehavior will be

handled appropriately

1

Environmental Effects

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 nondegradable 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

169

used for 3D printed products will be high-impact polystyrene (HIPS) plastic, which is

nondegradable 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 additional carbon

dioxide and hydrogen chloride, which can cause internal and external irritation to anyone that

comes in contact with it.

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 the

midst of mild fog or low-hanging clouds can result in the rocket becoming difficult to track or lost

altogether.

170

Section 9: Project Plan

Budget Plan

The budgets displayed in Table 10.1 are an initial approximation of the expenditures required for

the overall project. The approximations are conservative, assuming excess quantities of materials

and no price breaks. These approximations put the rocket on the pad for $2274.10, putting us well

below the maximum budget of $7,500 outlined in requirement 1.14 of the NASA Student Launch

Handbook. We hope to bring a large squad of team members to the competition this year, as travel

from Auburn to Huntsville is relatively cheap. Assuming we travel with 30 team members, the

lodging costs will be approximately $4000. Assuming $2,274.10 for the rocket on the pad and

$4,000 for travel, this leaves $16,725.90 for overhead costs, test motors, educational engagement,

and any other testing and development costs, based on the $23,000 amount for total funding

presented in Table 10.1.

Table 10.1: Initial Budget Estimates

Item Price Source

Rocket on the Pad

Carbon Fiber $222.50 US Composites

Resin $75.50 US Composites

Paint $70.00 Eastwood

Aerotech L1520T

Motor $169.99 Sirius Rocketry

RMS 75/3840 Motor

Case and Associated

Hardware

$385.20 Apogee Rockets

Rail buttons $6.00 Chris's Rocket Supplies

http://www.uscomposites.com/carbonpage.html

http://www.uscomposites.com/epoxy.html

http://www.eastwood.com/paints/primers.html

https://www.siriusrocketry.biz/ishop/aerotech-high-power-l1520t-ps-reload-kit-hazmat-910.html

https://www.apogeerockets.com/Rocket_Motors/Rouse-Tech_Casings/75mm_Casings/RMS-75_3840_Casing

http://www.csrocketry.com/store/

171

Ripstop Nylon $200.00 Rockywoods

Tubular Nylon $90.00 Chris's Rocket Supplies

Altus Metrum (x2) $600.00 Chris's Rocket Supplies

CO2 System $180.00 Tinder Rocketry

ABS Material $300.00

HDPE $70.00 McMaster-Carr

Tender Descender $85.00 Apogee Components

Total $2,274.10

Travel and Lodging

Team Travel to

Competition in

Huntsville

$4,000.00 *assuming 30 members

will be traveling

Funding Plan

The team has secured funding from the sources presented in Table 10.2. This money will cover

the cost of the rocket on the pad, the purchase of capital equipment as needed, the cost of subscale

and full scale test launch motors, programming and materials for our educational engagement

events, travel and housing for the team at the competition in Huntsville, Alabama, and any other

costs associated with designing, building, and launching our competition rocket.

Table 10.2: Funding Sources

Source Amount

Alabama Space Consortium $13,000

http://www.rockywoods.com/Fabrics-Kits/All-Ripstop-Nylon-Fabrics



http://www.tinderrocketry.com/#!peregrine-8--12g/c23p3

https://www.apogeerockets.com/Building_Supplies/Parachutes_Recovery_Equipment/Tender_Descender

172

Auburn University Organization Board $5,000

Auburn University College of Engineering $5,000

Total Funding $23,000

Timeline

As CDR wraps up, the timeline for the rest of the competition is primarily focused on the

production of the full scale rocket and the multiple test launches the team plans to complete by

FRR. The team plans to complete three full scale launches: one to verify the PLF and recovery

electronics, one to verify the WAFLE system, and one to verify the complete, full functioning

rocket. Should these be successful, the team plans to build an additional complete rocket and

complete and additional launch of this second rocket.

The full timeline can be found in Appendix C. 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 is on schedule for building the full scale rocket. As of the time of this report, the rocket

for the full scale test of the PLF and recovery system is almost completely finished. There are two

launches close to Auburn in January and the team will be attending one to complete the first full

scale launch.

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.

Table 10.2 shows the overall schedule of the rocket build superimposed with launch dates available

to the Auburn team. This shows the three planned full scale launches and available launch dates

around the time in which these builds should be complete.

173

Table 10.3: Launches and Vehicle Timeline

Table 10.2 outlines the basic timeline of the payload subsystems and the recovery subsystem.

These graphics show that subsystem testing and manufacturing will be completed in late January.

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.

174

Table 10.5: Competition Timeline

175

Section 10: 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.

Drake Middle School 7th Grade Rocket Week

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 700 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 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,

176

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

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 rockets. AUSL team

members and certified professionals will take care of this portion at the launch event. The students

177

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.

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.

178

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

179

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.

Samuel Ginn College of Engineering E-Day

Event Date: February 27, 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.

Boys Scouts of America Space Exploration Merit 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 of the

180

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.

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

181

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.

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.

182

• 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.

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

183

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.

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.

Rocket Day

Event Date: March 2016

The Auburn Student Launch team is pleased to announce its new big educational outreach event

dubbed Rocket Day. As its name suggests there will a day of learning and exploration about a

variety of rockets. AUSL will host this educational outreach event to spread rocketry to the

community of Auburn and Opelika. This is a family event so everyone is invited and encouraged

to join. Programming will be available for children from Kindergarten to High School. For younger

children, water bottle rockets will be available, as well as a gallery of inert rockets to look at and

explore. Older children will have the opportunity to launch rockets with up to I-class motors,

depending on maturity and skill level of the children. A variety of prepackaged beginner kits will

be available as well as more advanced kits that involve more freedom to design. All rockets will

require AUSL members to assist and supervise. The goal of this event is engage the community in

STEM fields in a fun hands-on event designed to get parents and children of all ages to design,

build and launch rockets. Many members of the Auburn Student Launch team have fond memories

of building and launching rockets as children and cite this as their inspiration for entering STEM

fields and getting involved in the NASA Student Launch competition. We hope to spread this

inspiration to the next generation of engineers and scientists.

184

Rocket Day – Outline

AUSL has defined the following outline for approaching the community and coordinating this

huge event that will require lots of attention and cooperation.

1. Budget must first be defined for this event to make sure funds are available.

2. Approval from Auburn/Opelika City Project Management to conduct Rocket Day.

3. Safety Handbook for Rocket Day will be completed by CDR for NASA and

Auburn/Opelika city approval.

4. A large location that meets guidelines and requirements set by the NAR, TRA, FAA, and

local city rules must be acquired by CDR.

5. Notify local hospital and Fire Station to have EMTs and Fire Fighters on standby to be

ready for any cautionary event that could take place.

6. Approach Auburn/Opelika City schools to promote Rocket Day.

7. Acquire rockets for Rocket Day

8. Acquire facilities such as tents, tables, restrooms, trashcans, food, water, and concession

stands for this event.

9. Rocket Day will commence on a Saturday at a time to be announced and end by sundown

during March 2016, just like any rocket launch held by NAR and TRA.

10. Launch Field will be cleaned up the following Sunday to leave no evidence to show that

no event had ever occurred at the launch location.

Rocket Day – Safety

AUSL will be taking extreme caution to ensure safety of all participants. Because of the wide

variety of participants we expect to see, EMTs will be on site in addition to the Auburn Student

Launch safety team. All participants will be required to follow these rules:

1. Everyone will follow NASA, NAR, TRA, and FAA requirements and guidelines for launch

safety.

2. NO ONE will design/build a rocket designed to fail or perform in a potentially dangerous

way.

3. Rocketry certified level one and above TRA members will inspect completed rockets and

certified NAR personnel will conduct the rocket launches on the field.

4. Launch Field safety/rules will be announced to everyone building and launching a rocket.

185

5. Certified EMTs and Firefighters will be on standby at ALL times.

6. Rocket launches will be conducted the same way the NAR and the TRA organize rocket

launches.

7. All purchased rocket motors will be sold by certified NAR prefects.

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

186

Section 11: Conclusion

In conclusion, the team is very excited to move into the final testing and construction phase of the

project. We are excited to see the final product of our ambitious payload and the months of effort

we have put towards it. Subscale and subsystem testing provided promising results for the

upcoming full scale tests. We are also looking forward to the many educational outreach events

we have planned for the spring. We hope to spend a lot of time on the launch fields in the upcoming

months and are looking forward to the competition in April.

187

Appenix A: References

Knowles, Vern. Ejection Charge Sizing. Vern’s Rocketry. 2007. Web. 02 October 2015.

Niskanen, Sampo. "OpenRocket Technical Documentation." OpenRocket Simulator. 10 May

2013. Web. 14 Jan. 2016. <http://openrocket.info/index.html>.

Howard, Zachary. "How to Calculate Fin Flutter Speed." Www.apogeerockets.com. 19 July 2011.

Web. 14 Jan. 2016. <https://www.apogeerockets.com/education/downloads/Newsletter291.pdf>.

188

Appenix B: Risk Mitigation Matrices

Airframe Section

Airframe Risk Matrix

Impact→ 1 2 3 4 5

Probability

↓ Negligible Minor Moderate Significant Severe

Very High

High

Medium 1,2 3,4,5,7

Low 6

Very Low

189

Autoclave Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High 8

High

Medium 3,4 1 2,6,9 7

Low 5

Very Low

Filament Winder Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

190

High

Medium 2,3,4 1

Low

Very Low

Carbon Fiber Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High 1

Medium 2,3

Low 4

Very Low

191

Epoxy Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High 2 1

High 4

Medium 7,8,9 3 5

Low 6

Very Low

Airframe Environment Effects Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

192

High

Medium 2 1

Low 3

Very Low

Testing

Wind Tunnel Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High

Medium 1 3,4,5

Low 2

193

Very Low

Tensile Test Rig Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High 1,2

Medium 5

Low 3,4

Very Low

Scientific Payloads – Payload Fairing

Operations Risk Matrix

Impact→ 1 2 3 4 5

Probability


194

Very High

High

Medium 3

Low 5 4 8 1,2

Very Low 7 6

Testing Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High 12,13

Medium 10

Low 8 9

195

Very Low 1,2,6,7,11 3,4,5

Construction Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High 5 3,4

Medium

Low 1,2

Very Low

Scientific Payload - WAFLE

Operations Risk Matrix

Impact→ 1 2 3 4 5

Probability


196

Very High

High

Medium 1

Low 11,12 3 4,5 7

Very Low 10 6,8,9 2

Testing Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High

Medium

Low 3 1,2 6

197

Very Low 5 4


Impact→ 1 2 3 4 5

Probability


Very High

High 4 3

Medium 1 2

Low

Very Low 6 5

Materials Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

198

High

Medium

Low 1 5,7,8

Very Low 2,3,9 4,6

Recovery Section

Flight Recovery Operations Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High 1,3,7,9

Medium 5 2,4,6,8

Low

199

Very Low

Wind Tunnel Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High

Medium 1 1,3,4,5

Low 2

Very Low

Tensile Test Rig Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

200

High 1,2

Medium 5

Low 3,4

Very Low

Shear Pin Test Rig Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High 1,2

Medium 5

Low 3,4

Very Low

201

Kevlar Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High 1

High 2

Medium 3

Low

Very Low 4

Nylon Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High 2 1

202

Medium 3

Low

Very Low

Carbon Dioxide Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High 3 5 1

Medium 4 2

Low

Very Low

Black Powder Risk Matrix

203

Impact→ 1 2 3 4 5

Probability


Very High

High 1,3,4,5

Medium 2

Low

Very Low

Fiber Glass Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High 2

High 1

204

Medium

Low

Very Low

Orbital Sander Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High 8

High 1,2,4,6,7

Medium 12,13 9 3,5,11

Low 10

Very Low

Sewing Machine Risk Matrix

205

Impact→ 1 2 3 4 5

Probability


Very High

High

Medium 2,5 1 3,4

Low

Very Low

Hand Tools Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High 1,2,3,5,6

206

Medium 4

Low

Very Low

Outreach Section

Outreach Operations Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High

Medium 1

Low 4 2,5,6 3

Very Low

207


Impact→ 1 2 3 4 5

Probability


Very High

High

Medium

Low 2 1

Very Low

Materials Risk Matrix

Impact→ 1 2 3 4 5

Probability


Very High

High

208

Medium

Low 1,2

Very Low

209

Appenix C: Competition Calendar

project aquila - eng.auburn.edu

Documents