PROJECT AQUILA 211 ENGINEERING DRIVE AUBURN, AL 36849

Transcription

1 PROJECT AQUILA 2 ENGINEERING DRIVE AUBURN, AL CRITICAL DESIGN REVIEW REPORT JANUARY 5, 206

2 Table of Contents Section : Summary of CDR Report... 2 Team Summary... 2 Launch Vehicle Summary... 2 Payload Summary... 3 Section 2: Changes Made Since PDR... 4 Changes Made to Vehicle... 4 Changes Made to Payload... 4 Changes Made to Project Plan... 4 Section 3: Launch Vehicle... 5 Mission Statement... 5 Major Milestone Schedule... 5 System-Level Design Review... 7 Structure... 7 Propulsion Aerodynamics Final Dimensional Drawings Test Descriptions and Results Materials Testing Wind Tunnel Testing System Level Functional Requirements Workmanship and Manufacturing Design Integrity... 4 Fin Shape and Style... 4 Materials in Fins, Bulkheads and Structural Elements... 4 Assembly Procedures Verification Plan Mass Statement... 54

3 Mission Performance Criteria Section 4: Subscale Flight Results Flight Data Predicted vs. Actual Performance Subscale Impact on Full-Scale Design Section 5: Recovery Subsystem Recovery System Outline Subscale Analysis Requirement Validation Parachutes Drift Ejection System... 7 Altimeters Attachment Hardware Section 6: Aerodynamic Analysis Payload System Level Design Review Payload Structure Payload Electronics... 8 Arduino Servos Accelerometer GPS Wiring Design Requirements Manufacturing and Assembly Risk Mitigation Payload Integration Payload Concept Features and Definition Science Value

4 Payload Objectives Payload Success Criteria Testing and Simulation... 9 Section 7: Payload Fairing (PLF)... System Level Design Review... Design Overview... Materials... 3 Design Requirements... 4 Testing... 5 Aerodynamic Design Testing (Completed)... 5 Charge Deployment Testing... 6 Drag Strip Deployment Testing... 7 Full Scale Testing... 8 Science Value... 8 Section 8: Safety... 9 Checklists... 9 Final Assembly Checklist... 9 Launch Procedures Checklist... 2 Safety Officer Airframe Hazard Analysis Airframe Failure Modes Environmental Effects Airframe Risk Mitigation Testing Systems Scientific Payloads Hazard Analysis Scientific Payload Risk Mitigation Payload Fairing Scientific Payload Risk Mitigation WAFLE Recovery Hazard Analysis

5 Recovery Risk Mitigation - Materials Recovery Risk Mitigation - Construction... 6 Outreach Hazard Analysis Environmental Effects Vehicle Effects on Environment Environmental Effects on the Vehicle Section 9: Project Plan Budget Plan Funding Plan... 7 Timeline Section 0: Educational Engagement Drake Middle School 7th Grade Rocket Week Rocket Week Plan of Action Rocket Week Launch Day Rocket Week Learning Objectives Gauging Success Samuel Ginn College of Engineering E-Day Boys Scouts of America Space Exploration Merit Badge Space Exploration Merit Badge Requirements Boy Scouts of America - AUSL Requirements... 8 Boy Scouts of America - Plan of Action Boy Scouts of America: Goals Girl Scouts of the USA - Space Badge Rocket Day Rocket Day Outline Rocket Day Safety Auburn Junior High School Engineering Day

6 Section : Conclusion

7 List of Figures Figure 3.: Full Rocket Rendering... 5 Figure 3.2: 5 Inch Braided Isogrid... 9 Figure 3.3: Motor Tube Rendering... 2 Figure 3.4: Aerotech L520T Thrust Curve... 2 Figure 3.5: Aeropack Motor Retention Figure 3.6: Fin Rendering Figure 3.7: Upper Section Dimensions Figure 3.8: Lower Section Dimensions Figure 3.9: Booster Tube Figure 3.0: Upper Body Tube Figure 3.: Fins... 3 Figure 3.2: Filament Wound Body Tube Figure 3.3: Bulkhead Figure 3.4: Motor Tube Bulkhead Figure 3.5: Carbon Fiber Test Data Figure 3.6: Carbon Fiber Test Results Figure 3.7: HIPs Data Figure 3.8: HIPS Test Results Figure 3.9: Three Point Bending Test Figure 3.20: FPS vs Lb Force Figure 3.2: Wind Tunnel Test Figure 3.22: Fin Shapes... 4 Figure 3.23: Patran Tube Model Figure 4.: Subscale Open Rocket Model Figure 5.: Parachute Configuration Figure 5.2: Subscale Parachute Configuration Figure 5.3: Parachute Shape Parameters Figure 5.4: Main Parachute Visualization Figure 5.5: Pictures of Tender Descender in Undeployed and Deployed Configurations Figure 5.6: Solenoid Circuit

8 Figure 5.7: Custom CO2 System Housing Figure 5.8: Custom CO2 System Assembly Figure 5.9: Altus Metrum TeleMega Altimeter Figure 5.0: Altus Metrum TeleMetrum Altimeter Figure 5.: Taoglas FXP MHz ISM Antenna Figure 6.: WAFLE system Figure 6.2: Grid Fin Fairing Figure 6.3: Aerodynamic Grid fin... 8 Figure 6.4: Arduino Uno Figure 6.5: HiTec HS-5685MH Digital Super Torque Servo Figure 6.6: ADXL335 Triple-axis Accelerometer Figure 6.7: WAFLE Electronics Schematic Figure 6.8: Starting point of the flow Figure 6.9: Flow directed over the grid fins Figure 6.0: 0.2 Mach flow over a grid fin Figure 6.: Total Fin Axial Force Coefficient versus Angle of Attack Mach 8 Low Pressure, Low Temperature Figure 6.2: Fin Normal Force Coefficient versus Angle of Attack Mach 8 High Pressure, High Temperature Figure 6.3: Fin Normal Force Coefficient versus Angle of Attack Mach 0. Low Pressure, Low Temperature Figure 6.4: Fin Normal Force Coefficient versus Angle of Attack Mach 0. High Pressure, High Temperature Figure 6.5: Fin Moment Coefficient versus Angle of Attack Mach 8 Low Pressure, Low Temperature... 0 Figure 6.6: Fin Moment Coefficient versus Angle of Attack Mach 8 High Pressure, High Temperature... 0 Figure 6.7: Fin Moment Coefficient versus Angle of Attack Mach 0. Low Pressure, Low Temperature Figure 6.8: Fin Moment Coefficient versus Angle of Attack Mach 0. Low Pressure, Low Temperature

9 Figure 6.9: Vortex Shedding Testing Visualization Figure 7.: PLF Curvature... Figure 7.2: Overall PLF Structure... 2 Figure 7.3: Overall PLF Assembly... 2 Figure 7.4: PLF Half... 3 Figure 7.5: Partially Deployed PLF... 3 Figure 7.6: PLF Half... 3 Figure 7.7: Charge Required for a Given Force (assume L = in)... 6 Figure.: Picture from Rocket Week Figure.2: A photo taken from DMS 7th Grade Rocket Week in April Figure.3: Space Exploration Merit Badge Figure.4: A Photo taken from Auburn Junior High School Engineering Day

10 List of Tables Table.: General Team Information... 2 Table.2: Team Leadership... 2 Table.3: Launch Vehicle Summary... 3 Table 3.: Major Milestone Schedule... 5 Table 3.2: Vehicle Length... 7 Table 3.3: Aerotech L520T Motor Specifications Table 3.4: Fin Dimensions Table 3.5: Manufacturing Phase Schedule Table 3.6: Verification Plan Table 3.7: Mass Estimates and Growth Allowance Table 4.: Simulation Data vs. Flight Data Table 5.: Recovery Requirement Validation Table 5.2: Parachute Shape Pugh Chart Table 5.3: Main Parachute Dimensions Table 5.4: Kinetic Energy Calculations Table 5.5: Drift Calculations... 7 Table 6.: Aerodynamic Analysis Payload Design Requirements Table 6.2: Aerodynamic Analysis Payload Risk Mitigation Table 6.3: Aerodynamic Payload Success Criteria Table 6.4: Aerodynamic Payload Simulations and Tests... 9 Table 6.5: SolidWorks Simulation Run Cases Table 6.6: Aerodynamic Payload Fortran- Flight and Dynamic model Table 6.7: Sample Data Mach=0.8 Low Pressure, Low Temperature Table 6.8: Sample Data at Mach=0.8 High Pressure, High Temperature Table 6.9: Sample Data at Mach 0. Low Pressure, Low Temperature Table 6.0: Sample Data at Mach 0. High Pressure. High Temperature Table 7.: PLF Design Requirements... 4 Table 7.2: Drag Strip Deployment Test Matrix... 7 Table 9.: Risk Mitigation Table - Airframe Table 9.2: Risk Mitigation Table - Autoclave

11 Table 9.3: Risk Mitigation Table - Filament Winder Table 9.4: Risk Mitigation Table - Carbon Fiber Table 9.5: Risk Mitigation Tables - Epoxy... 3 Table 9.6: Risk Mitigation Tables - Airframe Environment Effects Table 9.7: Risk Mitigation Tables Wind Tunnel Testing Table 9.8: Risk Mitigation Tables Tensile Test Rig Table 9.9: Risk Mitigation Table - Operations Table 9.0: Risk Mitigation Table Payload Fairing Testing Table 9.: Risk Mitigation Table Payload Fairing Construction... 4 Table 9.2: Risk Mitigation Table Operations Table 9.3: Risk Mitigation Table WAFLE Testing Table 9.4: Risk Mitigation Table WAFLE Construction Table 9.5: Risk Mitigation Table WAFLE Materials Table 9.6: Risk Mitigation Table - Flight Recovery Operations Table 9.7: Risk Mitigation Tables - Wind Tunnel Testing Table 9.8: Risk Mitigation Table - Tensile Test Rig Table 9.9: Risk Mitigation Tables - Shear Pin Test Rig Table 9.20: Risk Mitigation Table - Kevlar Table 9.2: Risk Mitigation Tables - Nylon Table 9.22: Risk Mitigation Tables - Carbon Dioxide Table 9.23: Risk Mitigation Table - Black Powder Table 9.24: Risk Mitigation Table - Fiberglass... 6 Table 9.25: Risk Mitigation Table - Orbital Sander... 6 Table 9.26: Risk Mitigation Table - Sewing Machine Table 9.27: Risk Mitigation Table - Hand Tools Table 9.28: Risk Mitigation Table - Outreach Operations Table 9.29: Risk Mitigation Table - Outreach Construction Table 9.30: Risk Mitigation Table - Outreach Materials Table 0.: Initial Budget Estimates Table 0.2: Funding Sources... 7 Table 0.3: Launches and Vehicle Timeline

12 Table 0.4: Subsystem Timeline Table 0.5: Competition Timeline... 74

13 Section : Summary of CDR Report Team Summary Table.: General Team Information Team Affiliation Mailing Address Title of Project Auburn University 2 Engineering Drive Auburn, AL Project Aquila Table.2: Team Leadership Student Team Lead Safety Officer Academic Advisor NAR/Tripoli Advisor Cassandra Seelbach Austin Phillips Dr. Joseph Majdalani Dr. Eldon Triggs Launch Vehicle Summary Table.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. 2

14 Table.3: Launch Vehicle Summary Total Length Final Mass Estimate Motor Selection inches 26.8 lbs Aerotech L520T 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. 3

15 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 inches forward of the wet CG to.054 aft of the wet CG. The distance between the hinge line and the dry CG increased from 7.77 inches to inches, with the CG forward of the hinge line. Likewise, the hinge line moved aft of the main separation by.22 inches. Changes Made to Project Plan 4

16 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. Figure 3.: Full Rocket Rendering 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 : Project Plan. Table 3.: Major Milestone Schedule Date Event Description 5

17 0/24/205 Subscale Construction /0/205 Subscale Completion /07/205 Subscale Launch /2/205 Subscale Launch Begin construction of 3/5 subscale Complete construction of subscale First launch day for subscale. Second launch day for subscale. 2/20/205 0/05/206 0/4/206 0/5/206 Full Scale Design Completed Begin Construction of Full Scale Complete Critical Design Review Complete Construction of Full Scale Complete full scale design from subscale results. Begin constructing two full scale rockets Paper and Presentation completed Have a completed full scale test rocket 0/6/206 Full Scale Launch 0/25/206 CDR Presentation 0/30/206 Full Scale Launch 02/3/206 Full Scale Launch First full scale launch testing avionics and PLF Complete the CDR presentation Second full scale launch testing WAFLE and RS Third full scale launch testing all systems together 02/4/206 Complete Construction of Full Scale Have two competitionready full scale rockets 02/20/206 Full Scale Launch 03/4/206 Flight Readiness Review Fourth full scale launch with competition rocket Complete FRR paper and presentation 6

18 03/7/206 03/30/206 FRR Teleconference 04/3/206 Launch Readiness Review 04/6/206 Competition Launch Day Complete FRR presentation Complete LRR in Huntsville, AL Launch full scale rocket in competition 04/29/206 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 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 inches. Component lengths are shown in Table 3.2. Table 3.2: Vehicle Length Component Length (Inches) 7

19 Nose Cone (Fairing) 3.25 Upper Tube 22 Aerodynamics Payload Section 5 Booster Section 33 Total 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

20 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 9

21 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 2.6 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

22 Figure 3.3: Motor Tube Rendering Motor: The motor selected for the competition is the Aerotech L520T. 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 L520T Thrust Curve This motor was chosen based on OpenRocket simulations, as it provides the roughly 3-to- thrustto-weight ratio desired for stable and predictable flight. 2

23 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 3-to- 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/s 2 which delivers a max velocity of 857 ft/s or close to Mach = Table 3.3: Aerotech L520T Motor Specifications Motor Specifications Manufacturer Motor Designation Diameter Length Impulse Total Motor Weight Propellant Weight Average Thrust Maximum Thrust Burn Time Aerotech L520T 2.95 in 20.9 in 3769 N-s 28 oz 62.8 oz 340 lbs 382 lbs 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 22

24 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. inch thick and is designed to fit around an Aerotech L520T-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 23

25 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.03 oz/in 3 density has been selected as the material due to its stiffness, strength, and light weight. Each fin will have a surface area of 54 in 2 (summing both 24

26 sides), making the fin surface area total equal to 26 in 2. The total component mass is 3.5 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 Single fin CC NNαα at subsonic speeds: XX = nn XX ii(cc NNαα ) ii ii= nn ii= (CC NNαα ) ii 2ππ ss2 AA rrrrrr (CC NNαα ) = ββss + + ( 2 AA ffffff cos Γ )^2 cc Table 3.4: Fin Dimensions Trapezoidal Fin Dimensions Root Chord Tip Chord 6.25 in 2.5 in Height Sweep 6 in 3.68 in Sweep Angle 3.5 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 25

27 the wings. The equation below represents the NACA flutter boundary equation with thin plate theory included. GG.337AARR VV ff = aa 3 PP(λλ + ) 2(AAAA + 2)( tt cc )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. 26

28 Final Dimensional Drawings Figure 3.7: Upper Section Dimensions 27

29 Figure 3.8: Lower Section Dimensions 28

30 Figure 3.9: Booster Tube 29

31 Figure 3.0: Upper Body Tube 30

32 Figure 3.: Fins 3

33 Figure 3.2: Filament Wound Body Tube 32

34 Figure 3.3: Bulkhead 33

35 Figure 3.4: Motor Tube Bulkhead 34

36 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, 205. 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 lb, while the 50% infill had an average maximum load of only lb. The solid infill test pieces had an average weight of lb, while the 50% infill had an average weight of 0.07 lb. This meant that the 23.% increase in weight caused by increasing the infill from 50% to a full 00% was responsible for only a 5.% 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 35

37 average, the 6 layer samples of carbon fiber weighed lb, while the 0 layer samples weighed lb. The 6 layer samples recorded an average yield force of 44.2 lb, while the 0 layer samples recorded an average yield force of lb. By increasing the number of layers of carbon fiber from 6 to 0, a 04.8% increase in performance was recorded, at the expense of only a 0.8% increase in weight. The next planned test is a tensile testing test, which is scheduled for the third week of January 206. The samples used for this testing were acquired from the same batch as the previous testing to eliminate a potential source of error. Figure 3.5: Carbon Fiber Test Data Figure 3.6: Carbon Fiber Test Results 36

38 Figure 3.7: HIPs Data Figure 3.8: HIPS Test Results 37

39 Figure 3.9: Three Point Bending Test 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 3.2. Feet Per Second Lb Force Drag Figure 3.20: FPS vs Lb Force 38

40 Figure 3.2: Wind Tunnel Test System Level Functional Requirements Vehicle:. The vehicle must maintain stability of 2 or more calibers. 2. The vehicle must have a factor of safety of at least Structural components must remain attached to launch vehicle. Grid Fins:. 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:. The deployment charge shall induce separation without harming the structural integrity of the PLF. 39

41 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 40

42 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 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. Body Tubes: Materials in Fins, Bulkheads and Structural Elements 4

43 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 75 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 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. 42

44 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 2 Events Manufacture major components. Such as body tubes, nose cone, fins. Begin assembly of subsystems. Such as booster sections, fin assemblies. Percentage of completion 50% 90% 3 Assemble completed rocket 00% 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. 43

45 Verification Plan The team has developed a set of requirements that covers all points addressed in the 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 Method of AU Vehicle. 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.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.2. 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.2.2 Teams may have additional altimeters to control vehicle electronics and payload experiment(s). Demonstration The team may use additional altimeters. 44

46 AU 5 Vehicle.2.2. At the Launch Readiness Review, a NASA official will mark the altimeter that will be used for the official scoring Inspection Demonstration Complete safety check at LRR AU 6 Vehicle At the launch field, a NASA official will obtain the altitude by listening to the audible beeps reported by the official competition, marked altimeter. Inspection Demonstration Ensure beeps are audible, launch successfully AU 7 Vehicle At the launch field, to aid in determination of the vehicle s apogee, all audible electronics, except for the official altitudedetermining altimeter shall be capable of being turned off. Inspection Demonstration Testing Ensure all electronics can be turned off and back on AU 8 Vehicle.2.3. The official, marked altimeter will not be damaged Inspection Analysis Testing Design electronics housing prevent damage altimeter the to to AU 9 Vehicle 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 0 Vehicle The altimeter will not report an apogee altitude over 5,600 feet AGL. Demonstration Testing Design and test launch vehicle to meet altitude requirement 45

47 AU Vehicle The rocket will be flown at the competition launch site. Demonstration Team will launch the rocket at the appropriate site on launch day AU 2 Vehicle.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 AU 3 Vehicle.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 Testing Analysis Demonstration Inspection Demonstration Trajectory simulations and testing will ensure the launch vehicle is recoverable and reusable Team will design and build launch vehicle that can have, but does not require, four independent sections AU 4 Vehicle.5 The launch vehicle shall be limited to a single stage Demonstration Team will design and build a singlestage launch vehicle AU 5 Vehicle.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 46

48 AU 6 Vehicle.7 The launch vehicle shall be capable of remaining in launchready configuration at the pad for a minimum of hour without losing the functionality of any critical on-board component. Testing Batteries shall be tested with full electronics to verify their life AU 7 Vehicle.8 The launch vehicle shall be capable of being launched by a standard 2 volt direct current firing system. The firing system will be provided by the NASA-designated Range Services Provider AU 8 Vehicle.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 Demonstration Vehicle will be designed and tested to be launched by the standard 2 volt DC system Vehicle will be designed around commercially available, certified motors AU 9 Vehicle.9. Final motor choices must be made by the Critical Design Review (CDR). Demonstration CDR will determine which motor the team will 47

49 AU 20 Vehicle.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. AU 2 Vehicle.0 The total impulse provided by a launch vehicle shall not exceed 5,20 Newton-seconds (Lclass). AU 22 Vehicle. Pressure vessels on the vehicle shall be approved by the RSO AU 23 Vehicle.. The minimum factor of safety (Burst or Ultimate pressure versus Max Expected Operating Pressure) shall be 4: with supporting design documentation included in all milestone reviews AU 24 Vehicle..2 Each pressure vessel shall include a pressure relief valve that sees the full pressure of the tank. Demonstration Demonstration Analysis Testing Inspection Analysis Testing Inspection Analysis Testing use for competition If the change is made to increase safety margin, NASA RSO will allow the change Launch vehicle impulse will be designed to not exceed 5,20 Newtonseconds. Inspection of pressure vessel by RSO standards by testing. Testing of the low-cycle fatigue. Inspection of each pressure vessel and testing of the pressure relief valve to see does it work as inspected. 48

50 AU 25 Vehicle..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. AU 26 Vehicle.2 All teams shall successfully launch and recover a subscale model of their fullscale rocket prior to CDR. The subscale model should resemble and perform as similarly as possible to the fullscale model, however, the full-scale shall not be used as the subscale model. AU 27 Vehicle.3 All teams shall successfully launch and recover their fullscale 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. Inspection Demonstration Demonstration Testing Testing Demonstration Testing The team will inspect the tank along with documentation of testing and history. A subscale and full scale launch will be completed. A test of the rocket will be exhibit demonstration all hardware functions properly. AU 28 Vehicle.3. The vehicle and recovery system shall have functioned as designed. Testing Testing of vehicle will show how recovery system functions. 49

51 AU 29 Vehicle.3.2. If the payload is not flown, mass simulators shall be used to simulate the payload mass. Inspection Demonstration Payload will be flown. AU 30 Vehicle 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 3 Vehicle 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 fullscale demonstration flight AU The full-scale motor does not have to be flown during the fullscale 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 fullscale flight, it is desired that the motor simulate, as closely as possible, the predicted maximum velocity and maximum Demonstration Testing Inspection Demonstration Demonstration of the adaptability of the systems notice to payload changes of the external surfaces through testing. Inspection of the motor will be done by the team to ensure it is flown through fullscale testing. 50

52 acceleration of the competition flight. AU 33 Vehicle.3.4 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. AU 34 Vehicle.3.5 After successfully completing the fullscale demonstration flight, the launch vehicle or any of its components shall not be modified without the concurrence of the NASA Range Safety Officer (RSO). AU 35 Vehicle.4 Each team will have a maximum budget of $7,500 they may spend on the rocket and its payload(s). AU 36 Vehicle.5. The launch vehicle shall not utilize forward canards. AU 37 Vehicle.5.2 The launch vehicle shall not utilize forward firing motors. Testing Demonstration Demonstration Demonstration Demonstration Demonstration Testing of the vehicle will demonstrate it being fully ballasted. The team will demonstrate that it did not alter any components or vehicle after demonstration flight. The team will demonstrate its budget of the competition rocket to validate its cost. The team will demonstrate how the launch vehicle does not utilize canards. A demonstration of the launch vehicle will demonstrate it not utilizing 5

53 AU 38 Vehicle.5.3 The launch vehicle shall not utilize motors that expel titanium sponges (Sparky, Skidmark, MetalStorm, etc.) AU 39 Vehicle.5.4 The launch vehicle shall not utilize hybrid motors. AU 40 Vehicle.5.5 The launch vehicle shall not utilize a cluster of motors. Demonstration Demonstration Demonstration forward firing motors. The team will demonstrate that the motor does not expel titanium sponges through test flight. The team will exhibit how the launch vehicle does not utilize hybrid motors. A demonstration and inspection of the launch vehicle to validate it does not use a cluster of motors. To ensure compliance with requirement AU-, 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 52

54 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-0, 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-2, 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-6, 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

55 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 fullscale 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 % Recovery % Grid Fins % Electronics % Motor % 54

56 Total % The values in Table 3.7 are based off of simulations ran in OpenRocket and using data from the 3/5 th subscale flight to determine accurate mass values. Mission Performance Criteria Vehicle:. The vehicle must have an apogee of 5280 feet AGL 2. The vehicle must be recoverable and reusable Grid Fins:. 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:. 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. 55

57 Section 4: Subscale Flight Results Flight Data For the subscale flight, a scaling factor of 3/5 th 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 3/5 th scale aerodynamic model to test how the system will work on the full-scale rocket. The OpenRocket Design, shown in Figure 4., was as follows: Figure 4.: Subscale Open Rocket Model 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 205 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 2 November 205 at the Phoenix Missile Works launch area. The Aerotech K-805 motor originally selected was not available; instead an Aerotech K-00 motor was used. The Aerotech K-00 has a faster burn time and higher thrust, giving the subscale a higher maximum velocity. The maximum velocity reached a reported 060 ft/s and an apogee of 6279 feet. 56

58 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.: Simulation Data vs. Flight Data OpenRocket Simulation Recorded Flight Data Percent Error Apogee (Feet) % Max Velocity (Ft/s) % 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. 57

59 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.. Figure 5.: Parachute Configuration 58

60 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 59

61 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 shrinkwrapped 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 Student Launch Handbook. The intended method of validation for all recovery requirements is outlined in the table below. Table 5.: Recovery Requirement Validation Requirement Number Requirement Validation Method Deployment of Recovery Devices Ground Ejection for Drogue & Main Parachute At Landing, Max KE of 75ftlbf for each Independent Section Test Deployment System Test Deployment System Calculation and subscale testing 60

62 Recovery system Electrical Circuits Independent of Payload Electrical Circuits Recovery System Must Contain a Redundant, Commercially Available Altimeter Exterior Arming Switch for each Altimeter Dedicated Power Supply for each Altimeter Arming Switch Capable of being Locked in the ON Position Removable Shear Pins used for Main & Drogue Parachute Compartment Electronic tracking Device Installed in Rocket to Transmit the Location of the Tethered Vehicle or any Independent Section to a Ground Receiver An Active Electronic Tracking Device shall be connected to any Independent Rocket Section or Payload Component The Electronic Tracking Device shall be fully Functional during Official Flight at Competition Launch Site Make Separate Electrical Circuits for Recovery and Payload Add redundant altimeter to recovery system Add Exterior Arming Switch for each Altimeter Put Separate Dedicated Power supply for each Altimeter Make Sure Locking Mechanism Locks the Switch when Turned ON Put Removable Shear Pins for Main and Drogue Parachute Place Electronic Tracking Device in the Tethered Vehicle and to any Independent Section and Test the Signal Location Attach Electronic Tracking Device to any Independent Section & Payload Section and Test the Signal Location Test and Bring extra Electronic Tracking Devices at the Official Flight 2. Recovery System Electronics shall not be affected by other Test Recovery System Electronics along with the 6

63 On-Board Electronics during Flight other On-Board Electronics to Ensure Signal Strength Recovery System Electronics must be placed in a Separate Compartment away from any other Radio Frequency/ Magnetic Wave Producing Device Recovery System Electronics Shielded from all On-Board Transmitting Devices Recovery System Electronics Shielded from all On-Board Transmitting Devices Producing Magnetic Waves Recovery System Electronics Shielded from any other On- Board Transmitting Devices 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 The Recovery System is designed be shielded by the avionics bay. The Recovery System is designed be shielded by the avionics bay. 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. d DROGUE = π 4 L TUBE d TUBE 62

64 The team s rocket has a length of in and a diameter of 5.25 in: iiii 5.25 iiii dd DDDDDDDDDDDD = = 22. iiii ππ 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 63

65 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 2 Ease of Manufacturing 2 2 Stability 2 Total Parachute areas for hemispherical shaped chutes are determined using the following equation: 64

66 2 F A = ρ C V D F : force ρ : density of air C D : drag coefficient V : descent velocity 2 Example calculation for a section of rocket weighing 0 lbm at a descent rate of 6 ft/s: 2 0llll mm 32.2 ffff AA = ss lb 2 = 2.9 ffff2 mm ffff ffff3.5 6 ss Table 5.3: Main Parachute Dimensions Booster (Bottom) Main Payload (Top) Main Area of chute 7.27 ft ft 2 Diameter of chute in in Diameter of spill hole 7.92 in 0.56 in Height of each gore 3.29 in 4.3 in Width of each gore in 27.5 in Number of gores

67 Figure 5.4: Main Parachute Visualization 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. KE = m V 2 m : mass V : descent velocity Example calculation for a section of rocket weighing 0 lbm at a descent rate of 6 ft/s, 0lbm ft KE = * ft lb 2 ft = 32.2 s 2 s

68 Table 5.4: Kinetic Energy Calculations Section Mass (lbm) Kinetic Energy (ft-lb) Payload Section Recovery (2 Parachutes) Avionics Fairing Structure Booster Section Recovery( Parachute) Grid Fins and Electronics Motor (After Burnout) Structure 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. 67

69 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 68

70 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. 69

71 DDDDDDDDDD = WWWWWWWW SSSSSSSSSS AAAAAAAAAAAAAAAA CChaaaaaaaa DDDDDDDDDDDDDD VVVVVVVVVVVVVVVV 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 6 ft/s. The rate of descent under drogue can be calculated with the following equation: 2 FFFFFFFFFF DDDDDDDDDDDDDD VVVVVVVVVVVVVVVV = AAAAAA DDDDDDDDDDDDDD DDDDDDDD CCCCCCCCCCCCCCCCCCCCCC PPPPPPPPPPhuuuuuu AAAAAAAA With a total rocket weight of lbm after burnout and a drogue diameter of 22. inches (which corresponds to an area of 2.67ft 2 ): DDDDDDDDDDDDDD VVVVVVVVVVVVVVVV = llll mm 32.2 ffff ss llll mm ffff ffff 2 = ffff ss This yields a descent velocity of ft/s under drogue. The estimated drift distances for a variety of wind speeds are shown in Table 5.5 below. 70

72 Table 5.5: Drift Calculations Wind Speed (mph) Wind Speed (ft/s) Drift Under Drogue (ft) Drift Under Main (ft) Total Drift (ft) Ejection System The Auburn team is utilizing a custom ejection system employing 2g 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

73 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 (0.5g) black powder charge, propelling a 2g 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 72

74 are aluminum. All parts are fabricated in house by our machinist. Auburn s custom CO2 system is shown in Figure 5.9 and Figure 5.0. 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. BBBBBBBBBB PPPPPPPPPPPP WWWWWWWWhtt (gggggggggg) = 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.6 g and 6.4 g will be needed. Electronic matches will be used to ignite 73

75 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 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 74

76 of tracking in flight data, apogee and main ignition, GPS tracking, and accurate altitude measurement up to a maximum of 25,000 feet. Figure 5.9: Altus Metrum TeleMega Altimeter Figure 5.0: 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 75

77 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 Figure 5.2: Taoglas FXP MHz ISM Antenna 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 FXP MHz ISM Antenna is the team's selection and can be seen in Figure 5.3. 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. 76

78 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 (/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 0 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 254 KN*m/kg compared to 69 KN*m/kg. However, the team decided to use tubular nylon as the material for 77

79 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. 78

80 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 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.: WAFLE system A fairing is located at the tip of the WAFLE. The fairing extends 4.0 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 5.0 inches aft of the servos; terminating.6 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 79

81 high impact strength made HIPS the obvious choice of materials to make the redesigned fairing from. The fairing will be 4.0 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. 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 80

82 during flight and use drag to control the rocket s target apogee. The intent is to accurately complete the Vehicle Requirement.. 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 5.9 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. 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. 8

83 Arduino The Arduino Uno is a single-board microcontroller that provides digital I/O pins of 4/6 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 3.8 kg/cm for the entire flight. This servo had the best trade-off between size (4. x 2 x 3.8 cm) and torque ( 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 82

84 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.9 x.9 x0.34 cm, the ADXL335 would be an ideal sensor for the team. 83

85 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. 84

86 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.. Table 6.: Aerodynamic Analysis Payload Design Requirements Requirement Number Requirement An aerodynamic analysis of structural protuberances Method of Validation A full aerodynamic analysis of the grid fins is conducted through computational fluid dynamics (CFD), subsystem wind tunnel testing, and in-flight sensors. 85

87 AU AU2 AU3 AU4 Grid Fin payload is selfcontained within a separate segment of the rocket. Aerodynamic fairing is firmly adhered to the gird fin segment. Bulk heads sealing the ends of the segment are stationary throughout flight Grid fins must stay deployed during the decent phase of the trajectory. Grid fins must stow away at 00 feet. All Aerodynamic data must be validated through analytical and experimental testing. Grid fins must stay stowed until boost phase is complete. Electronics must stay stationary throughout the flight Servos must remain in direct contact with the gears The WAFLE system is built to be a self-contained and is removable from the rest of the booster section. The fairing contains screw holes that allow the fairing to be hard mounted to the airframe. 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. When the Arduino detects that apogee has occurred, the fins will be deployed. Arduino will be informed from sensors that 00 feet is reached and will implement the storing sequence. A full aerodynamic analysis of the grid fins is conducted through computational fluid dynamics (CFD), subsystem wind tunnel testing, and in-flight sensors. 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. 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. The gears of the servos will be imbedded into the U-bracket base of the grid fin by means of a metal bar. 86

88 AU5 AU6 AU7 of the grid fins throughout the flight. Arduino must accurately predict the flight path of the vehicle. Grid fins must be deploy with precision to correct the vehicle s trajectory. Grid fins must stay deployed under the force applied by the flow. Do to the high strength of the metal bar and HIPS, the fin will stay attached. Testing and accurate simulation modeling will insure accurate prediction. 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. 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 5.85 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 87

89 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 Battery power runs out during flight Fairing screws fall out and the fairing falls off. Mounting screws connecting the WAFLE section and the engine section come out. Accelerometer fails Large change in trajectory Arduino will now function and the grid fins will not be deployed. The fins will be at a high risk of destruction The rocket will fall apart and be terminated if the engine is boosting. The fins will stop deploying or shut down If large off axis acceleration is detected the fins will disengage to a stored position. Fully charged batteries will be stored within the rocket before launch. Insure that the screws are secured before flight. Insure that the screws are mounted correctly. 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. 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 88

90 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. 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 970s 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. 89

91 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.. 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 AU AU2 AU3 AU4 Criteria All Aerodynamic data must be validated through analytical and experimental testing. Grid fins must stay stowed until boost phase is complete. Electronics must stay stationary throughout the flight Servos must remain in direct contact with the gears of the grid fins throughout the flight. Method of Validation A full aerodynamic analysis of the grid fins will be conducted through computational fluid dynamics (CFD), subsystem wind tunnel testing, and inflight sensors. Redundant timer will be implemented into the system to insure that the code iteration does not engage. The electronics will be adhered to a stationary plate within the airframe. This plate and mounting bolds will be tested and verified for security. Testing and small tolerances between the gear and servo will insure stability throughout flight. 90

92 AU5 AU6 AU7 Arduino must accurately predict the flight path of the vehicle. Grid fins must be deploy with precision to correct the vehicle s trajectory. Grid fins must stay deployed under the force applied by the flow. Testing and accurate simulation modeling will insure accurate prediction. Model In-The-Loop testing and test flights will validate deployment precision for the full scale. 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 Computational Fluid Dynamics (CFD) SolidWorks Flow Fortran- Flight and Dynamic model Intent 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 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. A Fortran simulation has been created to model how the fin will react when attached to an airframe under flight conditions. 9

93 Drag Profile Aerodynamic Load Testing Vortex Shedding Testing :5 Scale Test 3:5 Scale Test Sub-Full Scale Test Full Scale Test 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. 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. 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. A :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. 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. 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. 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. 92

94 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. 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. 93

95 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 75. In Figure 6.9, the result were that the incoming flow was at lbf/in 2 (lime green). Once the flow passes directly over the face of the grid fin, the pressure slightly increased in certain areas to lbf/in 2 (yellow). The area most affected by this higher pressure is at the base of the grid fin. Then, the pressure decreased to lbf/in 2 (turquoise) once the flow passed through the lattices. 94

96 Figure 6.9: Flow directed over the grid fins. Figure 6.0: 0.2 Mach flow over a grid fin. 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 95

97 placed perpendicular to is 0.2 Mach. In Figure 6.0, 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 lbf/in 2. Once the flow comes in contact with the grid fin, the higher pressure is at the base of the grid fin at lbf/in 2 (yellow-orange). Finally, when the flow passes through the lattices the pressure is decreased to lbf/in 2 (dark blue) to lbf/in 2 (light blue). In conclusion, the in an ideal like state environment the grid fin will receive a large amount of pressure on the face perpendicular to the flow in both 0. 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 P P2 P-P2 P0 P0-P The pressures over a grid fin under a 0. 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. 96

98 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.-0.8) Atmosphere Temperature (ºR) ( ) Atmospheric Pressure (lb/in 2 ) ( lb/in lb/ in 2 ) Reference Length (66.74 in) Reference Area (2.55 in 2 ) Nose Length Nose-Center body Length Total Body Length Maximum Body Radius Radius Body at Tail Nose to Fin Hinge Line Nose to Moment Center (9.26 in) (66.74 in) (69.3 in) (5.24 in) (5.24 in) (39.8 in) (43.7 in) Nose Type (0) Body CL to Base of Grid Fin Min Radius for grid points Body CL to Grid fin tip Height of fin support base Span of fin support base Total height of fin Chord length of fin (4.2 in) (0.5 in) (9.2 in) (2.5 in) (.5 in) (0.5 in) (2 in) 97

99 Average fin element thickness Fin base corner type number cells in base corner Fin tip corner type number cells in tip corner (0.25 in) () () Number cells in spanwise direction (5) Number cells in vertical direction (2) Number vortices per element chordwise () Number vortices per element spanwise () Fin stall angle (alpha max) (deg) (20) Fin stall angle (delta max) (deg) (20) Total number of fins (4) Roll angle for configuration (5) The axial force coefficient, moment coefficient, and normal force coefficient are outputs of the program. For reference, low pressure is a pressure of pounds per square inch and high pressure is pounds per square inch. Low temperature is degrees Rankine and high temperature is 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: 98

100 Fin Axial Force Angle of Attack (degrees) Figure 6.: Total Fin Axial Force Coefficient versus Angle of Attack Mach 8 Low Pressure, Low Temperature 2 Fin Normal Force Coefficient Angle of Attack (degrees) Figure 6.2: Fin Normal Force Coefficient versus Angle of Attack Mach 8 High Pressure, High Temperature 99

101 50 Fin Normal Force Coefficient Angle of Attack (degrees) Figure 6.3: Fin Normal Force Coefficient versus Angle of Attack Mach 0. Low Pressure, Low Temperature 50 Fin Normal Force Coefficient Angle of Attack (degrees) Figure 6.4: Fin Normal Force Coefficient versus Angle of Attack Mach 0. High Pressure, High Temperature The behavior of the fin moment coefficient is plotted below. 00

102 Fin Moment Coefficient Angle of Attack (degrees) Figure 6.5: Fin Moment Coefficient versus Angle of Attack Mach 8 Low Pressure, Low Temperature Fin Moment Coefficient Angle of Attack (degrees) Figure 6.6: Fin Moment Coefficient versus Angle of Attack Mach 8 High Pressure, High Temperature 0

103 Fin Moment Coefficient Angle of Attack (degrees) Figure 6.7: Fin Moment Coefficient versus Angle of Attack Mach 0. Low Pressure, Low Temperature Fin Moment Coefficient Angle of Attack (degrees) Figure 6.8: Fin Moment Coefficient versus Angle of Attack Mach 0. Low Pressure, Low Temperature The drag is calculated using the following equation: 02

104 D= C AρV 2 D Sample calculations of the drag for varying degrees of alpha is shown below. 2 Alpha (degrees) Table 6.7: Sample Data Mach=0.8 Low Pressure, Low Temperature Fin Drag (lbf) Fin 2 Drag (lbf) Fin 3 Drag (lbf) Fin 4 Drag (lbf)

105 Alpha (degrees) Table 6.8: Sample Data at Mach=0.8 High Pressure, High Temperature Fin Drag (lbf) Fin 2 Drag (lbf) Fin 3 Drag (lbf) Fin 4 Drag (lbf)

106 Alpha (degrees) Table 6.9: Sample Data at Mach 0. Low Pressure, Low Temperature Fin Drag (lbf) Fin 2 Drag (lbf) Fin 3 Drag (lbf) Fin 4 Drag (lbf)

107 Alpha (degrees) Table 6.0: Sample Data at Mach 0. High Pressure. High Temperature Fin Drag (lbf) Fin 2 Drag (lbf) Fin 3 Drag (lbf) Fin 4 Drag (lbf)

108 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 0-5 % is expected when comparing data obtained in FORTRAN to experimental data. Drag Profile: 07

109 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. 08

110 Figure 6.9: Vortex Shedding Testing Visualization :5 Scale Test: A :5 scale model was built of the rocket for wind tunnel testing. A :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 :5 scale level. 3:5 Scale Test: 09

111 After the :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 :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. 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. 0

112 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 0.25 inches. A plot showing the curvature can be seen in Figure 7.. Figure 7.: 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 3 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 ). 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 80 (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

113 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 2

114 Figure 7.4: PLF Half 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.). 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 3

115 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.. Table 7.: PLF Design Requirements NASA PLF.REQ. PLF.REQ.2 The fairings and payload must be tethered to the main body to prevent small objects from getting lost in the field. The deployment charge shall induce separation without harming the structural integrity of the PLF. The deployment charge shall not harm the recovery payload contained within the PLF. Each half of the PLF will retained to the main body of the rocket via hinges. Extensive testing will be done to determine the optimal charge size. The charge bay will be lined with carbon fiber to maintain structural integrity. Recovery wadding will be placed in the PLF to protect payloads and rocket components PLF.REQ.3 The hinges shall not deploy more than 80 Fabric strips will be attached to the inside of the PLF preventing the halves from separating more than 80 PLF.REQ.4 PLF.REQ.5 The entire PLF system shall remain structurally intact during the following phases: launch, separation, drift, and landing. Premature separation of the PLF system shall not occur. Ribs have been integrated into each half of the PLF to prevent flexing. A lip has been integrated to prevent air from entering into the system during flight. Pin connectors will hold the sid 4

116 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 :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: :5 Subscale Model Figure 7.6: 3:5 Subscale PLF 5

117 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., 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 = xx0 3 (FFFF) Force vs Charge Size Charge Size (mg) Required Force (lb) Figure 7.7: Charge Required for a Given Force (assume L = in) 6

118 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 000 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) Run 0 Run 2 0 Run 3 0 Description Runs -3 are a proof of concept that the parachute will deploy at low apogee speeds. Successful deployment will result in more testing. Run 4 30 Runs 4-2 are to ensure proper deployment of the parachute Run 5 30 in the event of a late fairing activation. 7

119 Run 6 30 Run 7 60 Run 8 60 Run 9 60 Run 0 00 Run 00 Run 2 00 Run 3 20 Run 4 20 Run 5 20 Runs 0-2 are to ensure proper deployment of the parachute in the event of a premature fairing activation. 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. 8

120 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. 9

121 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 AUSL President Signature X X 20

122 Final Construction Check: Launch Procedures Checklist Initial Check-off Points 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: Initial Check-off Points 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: Initial Check-off Points Final overall check of rocket construction. Final overall check of launch-pad and rail. Final overall check of ground support systems readiness. 2

123 Final overall check of personnel and observers readiness. Final overall launch readiness check. AUSL Safety Officer Signature AUSL President Signature X X 22

124 Launch Procedures Check: Initial Check-off Points 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. AUSL Safety Officer Signature AUSL President Signature X X 23

125 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 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. 24

126 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 25

127 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. 26

128 Table 9.: Risk Mitigation Table - Airframe Potential Risk Potential Effect Impact Risk Mitigation Risk 2 Airframe not properly transported () Damage to airframe 4 3 Airframe will be transported in a custom made shipping container to protect it from damaging vibrations or slipping Airframe not properly stored (2) Damage to airframe 4 3 Airframe will be stored in a specified container in a dry and cool room when not being constructed or tested 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 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 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 27

129 before launch for lack of epoxy Collision with bird (6) Damage to airframe 4 2 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 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 Potential Risk Potential Effect Impact Risk Mitigation Risk2 Debris flies up into user s eyes () Trauma to the user s eyes 3 3 Students will wear safety glasses or face shield while operating autoclave Material left in autoclave (2) Damage to autoclave and material 4 3 Users will ascertain that the autoclave is empty before operating it. Door not properly closed (3) Damage to material inside autoclave 2 3 Make sure the door is fully closed. Wrong cycle selected (4) Damage to material inside autoclave 2 3 Only authorized and trained users may operate autoclave. Material experiences Can cause severe injuries to users 5 2 Wear proper PPE and always keep hands, 28

130 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. 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 9 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. 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. Table 9.3: Risk Mitigation Table - Filament Winder Potential Risk Potential Effect Impact Risk Mitigation Risk2 29

131 Mandrel not secured properly () 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. Improper winding angles for the specific stresses occurred during flight (2) Improper construction of rocket body tubes leading to structural failure 4 3 Material testing will be performed on samples to ensure that rocket tubes will be strong enough to withstand expected forces 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 Filament does not unroll correctly (4) Improper construction of rocket body tubes leading to structural failure or damage to equipment 4 3 Winder will be watched closely during winding process Table 9.4: Risk Mitigation Table - Carbon Fiber Potential Risk Potential Effect Impact Risk Mitigation Risk2 Allergic reaction from coming in contact with carbon fiber () Skin irritation 3 4 Wear proper PPE when handling carbon fiber 30

132 Debris flies up into users eyes (2) Trauma to the users eyes 3 3 Wear safety glasses when working with carbon fiber Toxic particles (3) Respiratory irritation 3 3 Wear proper breathing apparatus when working with carbon fiber Electrical shock (4) Burn or electrocution 4 2 Carbon fiber is electrically conductive so it should be kept away from electrical equipment or machinery Table 9.5: Risk Mitigation Tables - Epoxy Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper Ventilation () 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 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 20 F 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 3

133 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 Epoxy gets in user s eyes (6) Damage to the user s eyes 5 2 Wear safety glasses while using epoxy 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 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 Epoxy not properly disposed (9) Potential fire hazard and damage to lab 2 3 Cure epoxy and let it cool before disposing to prevent possible fire 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. 32

134 Table 9.6: Risk Mitigation Tables - Airframe Environment Effects Potential Risk Potential Effect Impact Risk Mitigation Risk2 Harmful toxic fumes released into environment by autoclave () Damage to environment and toxic air supply 4 3 Make sure that autoclave is always properly ventilated before operating Epoxy not properly disposed (2) Potential fire hazard and damage to lab 2 3 Cure epoxy and let it cool before disposing to prevent possible fire 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 Airframe Risk Mitigation Testing Systems Table 9.7: Risk Mitigation Tables Wind Tunnel Testing Potential Risk Potential Effect Impact Risk Mitigation Risk2 Debris in the wind tunnel () 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 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 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 33

135 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 60 ft per second. Only authorized personnel can operate the wind tunnel 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 Table 9.8: Risk Mitigation Tables Tensile Test Rig Potential Risk Potential Effect Impact Risk Mitigation Risk2 Object being tested is improperly aligned () 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 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 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 34

136 Unauthorized use (4) Damage to machine, personnel, and projects 5 2 Machine will be kept powered off in a locked lab when not in use 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 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. Operations Failure Modes: 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 35

137 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 Potential Risk Potential Effect Impact Risk Mitigation Risk2 Premature charge ignition on the ground () 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. 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 36

138 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. 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 80 degrees. 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. The deployment charge damages the structural integrity of the PLF (6) The PLF requires repair or replacement, violating a critical mission requirement 5 Extensive testing will be done to determine that the materials and structure designed for the PLF will withstand any 37

139 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 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. 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. Table 9.0: Risk Mitigation Table Payload Fairing Testing Potential Risk Potential Effect Impact Risk Mitigation Risk2 Three Point Bending Test specimen partially shatters () Sharp debris would be left around the testing area. 2 The test specimen will be carefully manufactured and loaded to ensure that it will bend and fracture but not shatter. 38

140 Three Point Bending Test specimen improperly loaded (2) Test results may not be accurate and may interfere with further testing. 2 Team members will be knowledgeable about the test specimen and procedure prior to conducting the test. 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 The team will be certain to manufacture a representative specimen that accurately reflects the properties of the rocket. 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 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. Rocket model comes loose from vehicle (5) The vehicle or rocket model could become damaged, and potentially the operation of the vehicle compromised. 3 The rocket model and its mount will be checked several times prior to testing and no unnecessary force or use of it will occur. Vehicle breaks during testing (6) Delays to the testing, potentially harm to the vehicle operator, rocket model, or mount. 2 The vehicle will be well maintained and inspected prior to testing to ensure proper operation. Obstacle exists in testing area (7) The obstacle could cause unexpected changes to testing, or if impacted 2 A suitable testing area will be determined in advance of testing 39

141 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. 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 Ignition of black powder during handling or setup (0) 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 Black powder fails to ignite during testing () Team members must remove the unignited black powder, exposing them to risk if there is a delay in the electric signal 2 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. Fumes from black powder charge testing are inhaled (2) Team members may experience adverse health effects of inhaling fumes and particles. 2 4 Care will be taken to properly ventilate the testing area. 40

142 PLF breaks into several pieces from charge testing (3) 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.: Risk Mitigation Table Payload Fairing Construction Potential Risk Potential Effect Impact Risk Mitigation Risk2 Hair or items become entangled with 3D printer machinery () 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. 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. 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 4

143 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. 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. Scientific Payload Risk Mitigation WAFLE Table 9.2: Risk Mitigation Table Operations Potential Risk Potential Effect Impact Risk Mitigation Risk2 Grid fins do not deploy in any capacity () No data is gathered on the grid fins and stability is 2 3 A full aerodynamic analysis of the grid fins is conducted through 42

144 slightly compromised computational fluid dynamics (CFD), subsystem wind tunnel testing, and inflight 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 A full aerodynamic analysis of the grid fins is conducted through computational fluid dynamics (CFD), subsystem wind tunnel testing, and inflight 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. 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 43

145 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 inflight sensors. 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. Electronics fail to come online after boost (6) No data is gathered on the grid fins and stability is slightly compromised 2 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 44

146 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 A full aerodynamic analysis of the grid fins is conducted through computational fluid dynamics (CFD), subsystem wind tunnel testing, and inflight 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. Malfunction with WAFLE system (8) Vehicle trajectory will be altered resulting undesirable flight path and potentially collateral damage and/or loss of asset 2 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 detected the fins will disengage to a stored position. Flaws or weaknesses in grid fins (9) Instability of flight or heavy vibration causing 2 In order to evaluate how the grid fins will interact once deployed, the team 45

147 undesirable trajectory or debris to fall back to the Earth 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. Improper battery power or voltage (0) Electronics will fail. No data is gathered on the grid fins and stability is slightly compromised 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. Improper range of motion and angle in servos () May cause one or more grid fins to extend too far or not far enough causing slight instability during flight 2 2 A full aerodynamic analysis of the grid fins is conducted through computational fluid dynamics (CFD), subsystem wind tunnel testing, and inflight sensors. If large off axis acceleration is detected the fins will disengage to a stored position. Servos do not operate at same speed (2) May cause one or more grid fins to extend at different speeds than the others causing slight instability during flight 2 2 A full aerodynamic analysis of the grid fins is conducted through computational fluid dynamics (CFD), subsystem wind tunnel testing, and inflight sensors. If large off axis acceleration is 46

148 detected the fins will disengage to a stored position. Table 9.3: Risk Mitigation Table WAFLE Testing Potential Risk Potential Effect Impact Risk Mitigation Risk2 Grid fins impacts hard surface or tools () 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. Payload Fairings impact hard surface or tools (2) The fairings are damaged and require remanufacture to assure they fit and function properly 3 2 Great care will be taken in the handling of the payload fairings at all times 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 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. Grid fins are structurally compromised by aerodynamic loads The grid fins will need to be remanufactured and potentially 4 The grid fins will be well-manufactured and the test wellmonitored to ensure 47

149 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 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. 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 9.4: Risk Mitigation Table WAFLE Construction Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper material used in the manufacturing of the grid fins () The grid fins are unstable and may compromise flight dynamics. 3 Thorough research and testing will be done on the material and the general design throughout to ensure the fins meet the expected requirements. 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 48

150 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. 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 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. 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 Exact methods of manufacture including tools large and small will be determined prior to the beginning of construction. 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 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 49

151 other materials or project pieces. Table 9.5: Risk Mitigation Table WAFLE Materials Material Potential Effect Impact Risk Mitigation Risk2 Batteries () Insufficient power 2 2 A new battery will be used and tested with an electronic multitester to ensure proper function. Fully charged batteries will be stored within the rocket before launch. Accelerometer (2) Receiving false or inaccurate data, causing the Arduino to make improper course corrections 2 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. HIPS High Impact Polystyrene (3) Structural Failure 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. 50

152 Aluminum (4) Structural Failure 4 Material was chosen for its light weight and structural integrity 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. Carbon Fiber (6) Structural Failure 4 Material was chosen for its light weight and structural integrity 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. 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 5

153 to the high amount of torque provided. Adhesives (Epoxy, Flux and Soldering Materials (9) Copper wires may detach. 2 Adhesives will be tested with the full aerodynamic analysis of the grid fins. Conducted through computational fluid dynamics (CFD), subsystem wind tunnel testing, and inflight sensors. 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 9.6: Risk Mitigation Table - Flight Recovery Operations Potential Failure Potential Effect Impact Risk Risk Mitigation Risk2 The parachute(s) is not packed properly. () 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 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 52

154 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. 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. 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. The parachute deploys at the Structural damage to rocket causing 5 4 Recovery systems will be thoroughly tested 2 53

155 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. 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 9.7: Risk Mitigation Tables - Wind Tunnel Testing Potential Risk Potential Effect Impact Risk Mitigation Risk2 Debris in the wind tunnel () 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 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 54

156 on the rocket in the future 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 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 60 ft per second. Only authorized personnel can operate the wind tunnel 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 Table 9.8: Risk Mitigation Table - Tensile Test Rig Potential Risk Potential Effect Impact Risk Mitigation Risk2 Object being tested is improperly aligned () 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 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 55

157 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 Unauthorized use (4) Damage to machine, personnel, and projects 5 2 Machine will be kept powered off in a locked lab when not in use 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 Table 9.9: Risk Mitigation Tables - Shear Pin Test Rig Potential Risk Potential Effect Impact Risk Mitigation Risk2 Shear pin being tested is improperly aligned () Results acquired from tests are incorrect and have a damaging effect on the rocket in the future 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 Fractured particles during test (2) Damage to eyes and body extremities when the item being tested fractures 4 4 All personnel must stay a safe distance away from tensile test rig while performing test. Safety eyewear must also be worn along with proper clothing covering body extremities 56

158 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 Unauthorized use (4) Damage to machine, personnel, and shear pin 5 2 Have machine locked up by an authorized worker and keep power off 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 to make sure they have the authorization to test a shear pin Recovery Risk Mitigation - Materials Table 9.20: Risk Mitigation Table - Kevlar Potential Risk Potential Effect Impact Risk Mitigation Risk2 Breathing in Fiber Dust () Severe respiratory problems 4 5 Respirators will be required when working with Kevlar 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 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 9 in 57

159 the case of any serious lacerations Leaving in direct sunlight (4) Discoloration of Kevlar Kevlar will be stored in closed containers in lab Table 9.2: Risk Mitigation Tables - Nylon Potential Risk Potential Effect Impact Risk Mitigation Risk2 Breathing in fiber dust () Respiratory Problems 4 4 Respirators will be required when working with Nylon 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 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 Table 9.22: Risk Mitigation Tables - Carbon Dioxide Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper Ventilation () CO2 gas can cause headaches, nausea, and loss of 5 4 Lab will be ventilated at all times when working with CO2. If 58

160 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 wellventilated and secure area, protected from the weather. Storage area temperatures will not exceed 00 F 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 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. 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 59

161 Table 9.23: Risk Mitigation Table - Black Powder Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper Ventilation () 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 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 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. 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 20 F in a cool dry place in a tightly sealed container. It will not be stored with any other flammables 60

162 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 Table 9.24: Risk Mitigation Table - Fiberglass Potential Risk Potential Effect Impact Risk Mitigation Risk2 Ventilation issues () 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 Potential Risk Potential Effect Impact Risk Mitigation Risk2 Injuries to hands and fingers from moving parts () Injury or loss of extremities 5 4 Using thick gloves to operate the sander. Turn the sander off when not in use. 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 6

163 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 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 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 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 Improper Work Attire (7) Damage to body 5 4 Proper clothing and PPE will be required to operate the orbital sander 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 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 Over-reaching (0) Severe cuts to body 5 2 Ensure proper footing and balance while operating 62

164 sander. Always turn off sander before performing another task Improper Tool Maintenance () 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 Over Exerting Tool (2) Causes damage to project due to excessive force applied to tool 3 3 Operators will be trained in proper operation of sander Improper Tool Replacement Parts (3) Tool becomes unusable 3 3 Only use replacement parts intended for the orbital sander Table 9.26: Risk Mitigation Table - Sewing Machine Potential Risk Potential Effect Impact Risk Mitigation Risk2 Sewing over fingers () 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 Pin misuse (2) Damage to body from the pins and damage to project 3 3 Proper training is required to use the sewing materials Improper machine use (3) Inexperienced personnel can damage material and damage self 5 3 Personnel must be trained before they can use sewing machine Cord can fray (4) Can cause a fire 5 3 Regular maintenance of machine will occur before and after use. Machine s 63

165 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 Table 9.27: Risk Mitigation Table - Hand Tools Potential Risk Potential Effect Impact Risk Mitigation Risk2 Improper use () 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 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 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. Flying Debris (4) Debris may cause eye and/or bodily damage 5 3 Proper clothing and eye protection are required to operate tools. 64

166 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. 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 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. Operations Failure Modes: Transportation to outreach site o Car accident Introduction/help students design their rockets o Children jam fingers o Children hurt by tools Multiple rocket launchings Rocket stands fall Rockets have mid-air collisions Rockets land in the woods Construction: Tools for rocket kits o Children incapable of using tools 65

167 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 Potential Risk Potential Effect Impact Risk Mitigation Risk2 Car Accident () 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. 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. 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. 66

168 Mid Air rocket collisions (4) Rockets would not reach highest altitude due to midair collision 2 Students rockets will be launched from significant distances from each other. Rockets will be launched one at a time 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. 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. Table 9.29: Risk Mitigation Table - Outreach Construction Potential Risk Potential Effect Impact Risk Mitigation Risk2 Children ignite motor at time other than directed () 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 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 67

169 USLI outreach team for the child Table 9.30: Risk Mitigation Table - Outreach Materials Potential Risk Potential Effect Impact Risk Mitigation Risk2 Children Break Rocket model () 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 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 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 68

170 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. 69

171 Section 9: Project Plan Budget Plan The budgets displayed in Table 0. 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 $2274.0, putting us well below the maximum budget of $7,500 outlined in requirement.4 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 $2,274.0 for the rocket on the pad and $4,000 for travel, this leaves $6, 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 0.. Table 0.: Initial Budget Estimates Item Price Source Rocket on the Pad Carbon Fiber $ US Composites Resin $75.50 US Composites Paint $70.00 Eastwood Aerotech L520T Motor $69.99 Sirius Rocketry RMS 75/3840 Motor Case and Associated Hardware $ Apogee Rockets Rail buttons $6.00 Chris's Rocket Supplies 70

172 Ripstop Nylon $ Rockywoods Tubular Nylon $90.00 Chris's Rocket Supplies Altus Metrum (x2) $ Chris's Rocket Supplies CO2 System $80.00 Tinder Rocketry ABS Material $ HDPE $70.00 McMaster-Carr Tender Descender $85.00 Apogee Components Total $2,274.0 Travel and Lodging Team Travel to Competition in Huntsville $4, *assuming 30 members will be traveling Funding Plan The team has secured funding from the sources presented in Table 0.2. 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 0.2: Funding Sources Source Amount Alabama Space Consortium $3,000 7

173 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 0.2 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. 72

174 Table 0.3: Launches and Vehicle Timeline Table 0.2 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 0.4: Subsystem Timeline Table 0.3 shows the competition milestones set forth by NASA in the 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. 73

175 Table 0.5: Competition Timeline 74

176 Section 0: 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 7 th 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 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, 75

177 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.: Picture from Rocket Week 204 Rocket Week Plan of Action Day : 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 wellprepared 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 76

178 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 7 th 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. 77

179 Figure.2: A photo taken from DMS 7th Grade Rocket Week in April 204 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. 78

180 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, 206 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 204 and 205, 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 79

181 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:. 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.. 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:. The law of action-reaction 2. How rocket engines work 3. How satellites stay in orbit 80

182 4. How satellite pictures of Earth and pictures of other planets are made and transmitted. Do TWO of the following:. 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:. 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:. 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. 8

183 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 206, 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.3. Figure.3: Space Exploration Merit Badge 82

184 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 206 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. 83

185 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.. 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 206, just like any rocket launch held by NAR and TRA. 0. 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:. 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. 84

186 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 9, 205 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,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.4: A Photo taken from Auburn Junior High School Engineering Day 85

187 Section : 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. 86

188 Appenix A: References Knowles, Vern. Ejection Charge Sizing. Vern s Rocketry Web. 02 October 205. Niskanen, Sampo. "OpenRocket Technical Documentation." OpenRocket Simulator. 0 May 203. Web. 4 Jan < Howard, Zachary. "How to Calculate Fin Flutter Speed." 9 July 20. Web. 4 Jan < 87

189 Appenix B: Risk Mitigation Matrices Airframe Section Airframe Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High Medium,2 3,4,5,7 Low 6 Very Low 88

190 Autoclave Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High 8 High Medium 3,4 2,6,9 7 Low 5 Very Low Filament Winder Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High 89

191 High Medium 2,3,4 Low Very Low Carbon Fiber Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High Medium 2,3 Low 4 Very Low 90

192 Epoxy Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High 2 High 4 Medium 7,8,9 3 5 Low 6 Very Low Airframe Environment Effects Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High 9

193 High Medium 2 Low 3 Very Low Testing Wind Tunnel Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High Medium 3,4,5 Low 2 92

194 Very Low Tensile Test Rig Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High,2 Medium 5 Low 3,4 Very Low Scientific Payloads Payload Fairing Operations Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe 93

195 Very High High Medium 3 Low 5 4 8,2 Very Low 7 6 Testing Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High 2,3 Medium 0 Low

196 Very Low,2,6,7, 3,4,5 Construction Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High 5 3,4 Medium Low,2 Very Low Scientific Payload - WAFLE Operations Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe 95

197 Very High High Medium Low,2 3 4,5 7 Very Low 0 6,8,9 2 Testing Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High Medium Low 3,2 6 96

198 Very Low 5 4 Construction Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High 4 3 Medium 2 Low Very Low 6 5 Materials Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High 97

199 High Medium Low 5,7,8 Very Low 2,3,9 4,6 Recovery Section Flight Recovery Operations Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High,3,7,9 Medium 5 2,4,6,8 Low 98

200 Very Low Wind Tunnel Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High Medium,3,4,5 Low 2 Very Low Tensile Test Rig Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High 99

201 High,2 Medium 5 Low 3,4 Very Low Shear Pin Test Rig Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High,2 Medium 5 Low 3,4 Very Low 200

202 Kevlar Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High 2 Medium 3 Low Very Low 4 Nylon Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High 2 20

203 Medium 3 Low Very Low Carbon Dioxide Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High 3 5 Medium 4 2 Low Very Low Black Powder Risk Matrix 202

204 Impact Probability Negligible Minor Moderate Significant Severe Very High High,3,4,5 Medium 2 Low Very Low Fiber Glass Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High 2 High 203

205 Medium Low Very Low Orbital Sander Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High 8 High,2,4,6,7 Medium 2,3 9 3,5, Low 0 Very Low Sewing Machine Risk Matrix 204

206 Impact Probability Negligible Minor Moderate Significant Severe Very High High Medium 2,5 3,4 Low Very Low Hand Tools Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High,2,3,5,6 205

207 Medium 4 Low Very Low Outreach Section Outreach Operations Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High Medium Low 4 2,5,6 3 Very Low 206

208 Construction Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High Medium Low 2 Very Low Materials Risk Matrix Impact Probability Negligible Minor Moderate Significant Severe Very High High 207

209 Medium Low,2 Very Low 208

210 Appenix C: Competition Calendar 209

211 20

212 2

213 22

214 23

215 24

216 25