AUBURN UNIVERSITY STUDENT LAUNCH PROJECT NOVA II. 211 Davis Hall AUBURN, AL CDR

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1 AUBURN UNIVERSITY STUDENT LAUNCH PROJECT NOVA II 211 Davis Hall AUBURN, AL CDR January 10, 2019

2 Contents List of Tables...7 List of Figures CDR Report Summary Payload Deployable Rover Changes Made Since PDR Changes to Vehicle and Recovery Changes to Payloads Project Plan Changes Vehicle Criteria Design and Verification of Launch Vehicle...15 Mission Statement...15 Vehicle Structure...16 Propulsion...20 Aerodynamics...22 Mass Statement Subscale Flight Results...29 Chosen Scaling factors...29 Updating the Simulation...32 Full Scale Modifications due to Subscale Recovery Subsystem...34 Recovery Subsystem Overview...34 AUBURN UNIVERSITY STUDENT LAUNCH 2

3 Structural Elements...35 Materials...40 Ejection System...44 Parachutes...47 Electronics Mission Performance Predictions...59 Flight profile predictions...61 Stability throughout Flight...64 Kinetic Energy, Descent Time and Drift Safety Launch Concerns and Operation Procedures Safety and Environment (Vehicle and Payload)...67 Personnel Hazard Analysis...67 Failure Modes and Effects Analysis...68 Environmental Concerns...68 Risk Mitigation and Impact Payload Criteria Soil Sample Rover...70 Overview of the Rover...70 Chosen Design Components...70 Mission Design...74 Electrical Design...77 Structural Design...78 AUBURN UNIVERSITY STUDENT LAUNCH 3

4 5.2 Altitude Control System...83 Altitude Control Overview...83 Team Derived Requirements...83 Design Overview...83 Control Logic and Electrical Design...90 Concept of Operations Project Plan Testing...96 Vehicle Body Testing...96 Recovery Testing...99 Rover Testing Altitude Control System Testing Requirements Verification Tables General Requirements Vehicle Requirements Vehicle Prohibitions Recovery Requirements Rover Requirements Safety Requirements Team Derived Requirements Line Item Budget Vehicle Body Budget Recovery Budget AUBURN UNIVERSITY STUDENT LAUNCH 4

5 Rover Budget Altitude Control Budget Educational Outreach Lab Supplies and Launch Fees Summary Table Funding Funding Sources Projected Total Expenses Budget Balance Project Timelines Rover Timeline Altitude Control System Timeline Educational Engagement Timeline Appendix A. Team General Information Team summary Adult Educators Team Mentor Student Team Leader Student Safety Officer Project Organization NAR/TRA Sections Appendix B. CDR Flysheet Appendix C. Launch Procedures and Checklists AUBURN UNIVERSITY STUDENT LAUNCH 5

6 Appendix D. Risk Assessment Tables D.1. Hazard Analysis Table D.2. Failure Modes and Effects Analysis Table D.3. Environmental Concerns Tables D.4. Risk Mitigation and Impact AUBURN UNIVERSITY STUDENT LAUNCH 6

7 List of Tables Table 1: Section Lengths Table 2: Motor Specifications Table 3: Fin Dimensions Table 4: Mass by section Table 5: Subscale Launch Data vs Simulation Table 6: BAE Options Pugh Chart Table 7: Parachute Materials Pugh Chart Table 8: Shroud Line Pugh Chart Table 9: Shock Cord Pugh Chart Table 10: Attachment Hardware Pugh Chart Table 11: 3D Printed Material Pugh Chart Table 12: Ejection Pugh Chart Table 13: Parachute Shape Pugh Chart Table 14: Parachute Sizing Details Table 15: Altimeter Pugh Chart Table 16: Microcontroller Pugh Chart Table 17: Servo Pugh Chart Table 18: Switch Pugh Chart - BAE Table 19: Switch Pugh Chart - Nose cone Table 20: Open Rocket Flight Simulation Data AUBURN UNIVERSITY STUDENT LAUNCH 7

8 Table 21: Component Masses Table 22: Descent Time Table 23: Kinetic Energy Upon Landing Table 24: Rocket Body Drift Table 25: Nose cone Drift Table 26: Fabrication Method Trade Study Table 27: Soil Recovery Choice Trade Study Table 28: Motor Choice Trade Study Table 29: Movement Choice Trade Study Table 30: Communication Method Trade Study Table 31: Microcontroller Trade Study Table 32: Drag Plate Design Trade Study Table 33: Funding Sources Table 34: Projected Future Expenses Table 35: Budget Balance Table 36: Project Deliverables AUBURN UNIVERSITY STUDENT LAUNCH 8

9 List of Figures Figure 1: Team Rocket Internals Design Figure 2: Simulated Model of Team Rocket Figure 3: Carbon Fiber Motor Mount Figure 4: Motor Thrust Curve Figure 5: Fin Rendering Figure 6: Booster Assembly Drawing Figure 7: Nose Cone Drawing Figure 8: Fin Drawing Figure 9: Full Rocket Drawing Figure 10: Acceleration Data from Subscale Launch Figure 11: Vertical Acceleration over Time in Simulation Figure 12: Subscale Flight Data Comparison of Velocity and Altitude Figure 13: Recovery Overview Diagram Figure 14: BAE Model Figure 15: Nose cone Full Model Figure 16: Nose cone Detail Model Figure 17: 6 Gore Template Figure 18: 8 Gore Template Figure 19: Hemispherical Parachute Figure 20: Nose cone Wiring Diagram Figure 21: PerfectFlite StratoLoggerCF AUBURN UNIVERSITY STUDENT LAUNCH 9

10 Figure 22: Adafruit Trinket Figure 23: Internal Components of Rocket Figure 24: Predicted Altitude over time for flight Figure 25: Predicted velocity over time for flight Figure 26: Predicted acceleration over time for flight Figure 27: Predicted drag coefficient over time for flight Figure 28: Stability Parameters over Time Figure 29: Rover Conops Figure 30: Rover Wiring Diagram Figure 31: Rover with Collection Unit Extended Figure 32: Rover with Collection Unit Extended Figure 33: Rover with Collection Unit Retracted Figure 34: Top of Rover Orientation System Figure 35: Side of Rover Orientation System (Note that although the sample collection arm is extended in this figure, it will be retracted in flight) Figure 36: Last year's three plate design Figure 37: Current Drag Plate System Design Figure 38: 4-pin male and female JST connectors Figure 39: Full Assembly Figure 40: Top View, Full Assembly Figure 41: Main support structure of the assembly Figure 42: Breadboard bracket Figure 43: Batteries bracket to assembly AUBURN UNIVERSITY STUDENT LAUNCH 10

11 Figure 44: Plate mounts, contains a notch to make sure rack gear is properly aligned Figure 45: Drag plate Figure 46: Rack gear, driven by spur gear which moves the plates Figure 47: Pinion Figure 48: Basic Wiring Diagram Figure 49: Simulink model, including the PID controller and plant. The EOM block and Projected Altitude block can be found in Figure 50 and Figure 51 respectively Figure 50: Equation of Motion block Figure 51: Projected Altitude block Figure 52: Altitude Control Concept of Operations Figure 53: Ejection Testing Figure 54: Shear Pin Testing Rig Figure 55: CFD Rocket Simulation Results Figure 56: Simulink results Figure 57: Overall Project Gantt Chart Figure 58: Rover Gantt Chart Figure 59: Altitude Control System Gantt chart Figure 60: STEM Outreach Gantt Chart Figure 61: Organizational Chart AUBURN UNIVERSITY STUDENT LAUNCH 11

12 1 CDR Report Summary Team Summary Project Name Mailing Address Mentor Name Nova II 211 Davis Hall Auburn, AL Dr. Eldon Triggs Certification Levels and Numbers Tripoli: Level 2, Certification #12139 NAR: Level 2, Certification #12139 Mentor Mentor Phone (344) Vehicle Summary Rocket Length/Diameter Final Motor Choice Mass (with Motor) Target Altitude Recovery Setup Rail Size in. (Length) /6.25 in. (Outer Diameter) Aerotech L1420R 47 lbm 4700 ft. One 30.8 in diameter drogue parachute at apogee, with an 8.28 ft. main deployed at 600 ft. and the nose cone descending under a separate 21.3 in. parachute from 700 ft rail /144 in length For additional information please see the CDR flysheet, attached in Appendix B. 1.1 Payload Deployable Rover This year, the Auburn University team has selected the deployable soil sample rover for their payload. The rover will move on a pair of treads, will collect a soil sample using a belt with buckets, and will deposit the soil into a sample container. A weight driven rover orientation system will be unlocked during descent so that after landing the rover will rotate and be oriented upright. AUBURN UNIVERSITY STUDENT LAUNCH 12

13 2 Changes Made Since PDR 2.1 Changes to Vehicle and Recovery Since the preliminary design review, only three major changes have been made to the vehicle and recovery systems. The length of the nose cone was changed from 20 to 24 in., with an attendant change to the length of the main parachute compartment from 30 to 26 in. to keep the overall length of the vehicle the same. The second and third change were to parachute size and the main parachute deployment altitude, reduced to 500 ft. from 600 ft. Slight increases in weight associated with the first change and the rover changes required larger parachutes and still keep the same descent kinetic energy. As a result, the rocket had a slower descent rate with all parachutes deployed. For drift values and total descent time to stay under the required limits, the team had to lower main parachute deployment height. 2.2 Changes to Payloads As both payload s designs matured from PDR, a few minor changes needed to be made. The rover s length, and the length of its supporting sled, have been increased by 2 ines to accommodate the batteries. In the initial PDR mock up the wires coming from the batteries were not accounted for. For the altitude control system, an MPU Axis IMU has been chosen instead of the ADXL-345 Accelerometer. The ADXL-345 could only measure the acceleration along three vectors, while the projected altitude equation used as the point of reference for the PID controller must know the acceleration along the gravity vector. This issue was noticed thanks to flying the ADXL-345 on the subscale flight test and noting that its ability to calculate accelerations was limited to along the vehicle axis and it could not determine its orientation. Thus, a change to a 9- DOF IMU was needed to read the acceleration along the gravity vector. 2.3 Project Plan Changes The primary project plan changes are mostly based around the fabrication of our full scale rocket and payloads. The team was able to confirm that it would have the resources and time to create both a conventional hand layup composite airframe and one out of an Open Architecture Composite Structure (O-ACS) in concert with Highland Composites. Throughout this report, all AUBURN UNIVERSITY STUDENT LAUNCH 13

14 masses and performance estimates are given with regards to the conventional airframe, as it is heavier and will possess worse performance and also will be made in house. Construction on the conventional full scale airframe began on schedule at the beginning of November, but due to delays in receiving parts from suppliers and the pressure of finals week was not completed before the end of November as predicted. Instead, full scale construction is predicted to be completed in the beginning of January. At the same time, the team began working with Highland on the O-ACS airframe, providing designs and molds. As a result, the team s predicted first full scale launch will be at the beginning of January with a conventional airframe, with February and March still open for backup launches or flights of the O-ACS airframe. Fabrication of the rover payload has been similarly delayed, as the CAD design took longer than expected, but the team has begun fabrication and still expects to fly the rover at the first full scale launch planned for the beginning of January. Next, the team was able to combine our Boy scouts STEM events and set a date, while also performing outreach at Loachapoka High School. For additional information on the team s STEM engagement activities, please see section 6.6.3: Educational Engagement Timeline. The team s budget has had additional costs but also additional funding; therefore, the overall outlook for the team still predicts a budget surplus, leaving funds for rebuilding the rocket in case of a catastrophic failure or reinvesting in the team s facilities at the end of the competition cycle. AUBURN UNIVERSITY STUDENT LAUNCH 14

15 3 Vehicle Criteria 3.1 Design and Verification of Launch Vehicle 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 to reach the goals set by NASA in this year s competition Mission Statement Figure 1: Team Rocket Internals Design The mission of the AUSL launch vehicle is to design and construct a lightweight, safe, and reliable vehicle. These motivators will ensure a vehicle that will allow for successful launches and the flexibility to adapt the design to various experimental payloads. Three driving requirements from the NASA Student Launch Handbook have been chosen to guide the success of this mission: 1. The vehicle must maintain stability of 2 or more calibers. 2. The vehicle must be recoverable and reusable. 3. The vehicle will not exceed Mach 1 during flight. These requirements serve as minimum standards set to achieve mission success. AUSL has also determined four team-specific requirements to drive the design of the launch vehicle: 1. The vehicle must have a minimum apogee of 4700 ft AGL. 2. The vehicle must have a factor of safety of at least Structural components must remain attached to launch vehicle. AUBURN UNIVERSITY STUDENT LAUNCH 15

16 4. The Vehicle must not bend too much at joints. These team requirements serve as additional constraints to assure the vehicle design is compatible with the mission. Vehicle Structure The structure of the launch vehicle has been designed to be able to withstand the forces the rocket will experience during operation. The launch vehicle body must be strong enough to maintain stable flights while accommodating all other subsystems and ensure they have adequate space and protection. The design of the structure requires heavy tradeoffs between strength, space, and weight. Figure 2: Simulated Model of Team Rocket The total length of the rocket is in. Component lengths are shown in Table 1. Section Length (in.) Nose Cone 24 Rover Section 25 Recovery Avionics Coupler External Wrap 2 Main Parachute Section 26 Drogue Parachute Section 12 Booster Section 34.5 Total Table 1: Section Lengths AUBURN UNIVERSITY STUDENT LAUNCH 16

17 The body tubes house all subsystems of the launch vehicle. These tubes comprise most of the vehicle body surface that is exposed to the airflow; therefore, the aerodynamic properties of the body tubes are directly related to the altitude gained by the vehicle. Additionally, as the largest structure in the rocket, the body tubes represent the largest collection of mass in the rocket, except for the motor. To ensure mission success, it is critical to select and design body tubes that can survive the stresses of high-powered flight while remaining light enough to achieve the mission altitude. 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 will be formed into an isogrid structure around a 6-in mandrel, which will already be prewrapped with a layer of carbon fiber to make bonding the internal systems to the body tubes easier. Isogrid structures are a lighter alternative to using a solid tube structure. For aerodynamic purposes, the openings of the isogrids will be filled with a lightweight, high strength polyethylene foam. This will give the body a lightweight, aerodynamic body. Using this wrapped isogrid method instead of using only filament wound carbon fiber, the mass of the body tubes will be decreased by approximately 20 to 30 percent while still maintaining the same compressive strength properties as a carbon fiber tube. The rover tube section, however, will be made out fiberglass for communication between the rover and the team during the mission Couplers The couplers serve as a joint between two body tube sections. The couplers must be able to withstand forces experienced during rocket ascent to keep the structure of the body attached. The upper body tube will be attached to the booster section with 4 aluminum bolts. The team has decided to use fiber glass to create the couplers. This choice of material reduces the risk of separation of the upper body from the booster section in mid-flight. With the trade-off of an increase in mass and more difficult construction for strength, fiber glass was considered a very safe and reliable option. Fiber glass also has the benefit of being non-conductive, so it will be ideal for making the coupler, which will be holding our electronics, the avionics bay. This will reduce the issues the team has had in the past when it comes to being able to communicate to the electronics inside the vehicle, such as GPS systems. AUBURN UNIVERSITY STUDENT LAUNCH 17

18 Bulkheads Bulkheads are typically flat plates used to increase the structural strength of a rocket. They are also used to create airtight spaces and to divide the body into separate compartments. In rockets, they are commonly used to separate payload bays and to mount equipment for avionics and payloads. For rockets similar in size to the Project Nova II 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 6 ines, and the bulkheads are designed to fit perfectly into this size. All bulkheads for this rocket will be between ines thick Centering Rings The purpose of the centering rings is to center a smaller cylindrical body or tube inside a tube of a larger diameter. In the case of high-powered model rocketry, centering rings can be used as an engine block in motor mounts. The Project Nova rocket will be using three centering rings. These centering rings are located in the engine tube, and they serve to attach to the fin set and to attach to the motor retention. The centering rings are made of carbon fiber, and they are manufactured using the Computer Numerical Control (CNC) machine at Auburn University Aerospace Design Lab due to the availability and the teams experience with using carbon fiber. The centering rings have an outer diameter of 6 ines with an inner diameter of 3 ines. The thickness of each ring is approximately 0.25 in. The centering rings have a mass of 3.65 oz., determined from sample pieces Motor Tube To contain the motor on the rocket, a carbon fiber motor tube is being used. The motor tube will be made by pre-impregnated solid carbon fiber wrapped around a mandrel that is the same diameter of the motor. The carbon fiber material was chosen for its strength relative to its weight when compared to a solid tube. AUBURN UNIVERSITY STUDENT LAUNCH 18

19 Figure 3: Carbon Fiber Motor Mount 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 purposes of the motor retention system are 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. 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. AUBURN UNIVERSITY STUDENT LAUNCH 19

20 Propulsion The motor selected for the competition is the Aerotech L1420R. This was the teams primary motor choice in PDR, and after minor modifications were made to the rocket, it still gave us the needed altitude for the rocket. The specifications are listed below in Table 2: Motor Specifications. Additionally, the thrust curve for this motor is shown in Figure: Motor Thrust Curve. Figure 4: Motor Thrust Curve This motor was chosen based on OpenRocket simulations, as it provides the roughly 7:1 thrust-toweight ratio desired for stable and predictable flight. In addition, as shown in the motor thrust curve above, the motor achieves a higher than average thrust after approximately one-quarter second, thus reaching the required 7-to-1 thrust ratio in about one-quarter second. Based on OpenRocket simulations, the motor provided an apogee of 5082 ft with a max acceleration of 230 ft/s 2, which delivers a max velocity of 612 ft/s. AUBURN UNIVERSITY STUDENT LAUNCH 20

21 Motor Specifications Manufacturer Motor Designation Diameter Length Total Impulse Total Motor Weight Propellant Weight Propellant Type Average Thrust Maximum Thrust Burn Time Aerotech L1420R 2.95 in 26.2 in 1038 lb sec 10.1 lbm 5.69 lbm Solid 326 lbf 374 lbf 3.18 sec Table 2: Motor Specifications AUBURN UNIVERSITY STUDENT LAUNCH 21

22 Aerodynamics The aerodynamics system requires the rocket to 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 because 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 0.3-in-thick carbon fiber flat plates. A rendering of the fin design is shown in Figure 5: Fin Rendering. Figure 5: Fin Rendering AUBURN UNIVERSITY STUDENT LAUNCH 22

23 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 with a density that is 1.03 oz/in 3 has been selected as the material due to its stiffness, strength, and light weight. The stiffness and strength of carbon fiber reduces change of fin flutter and increases the vehicles chance of success during flight. Each fin will have a surface area of in 2 (summing both sides), making the fin surface area total equal to in 2. The total component mass is 25.3 ounces. These dimensions provide the vehicle with a projected stability of 2.1 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: Fin Dimensions. Fin Dimensions Root Chord Tip Chord Height Sweep 6.0 in in 6 in in Sweep Angle 34.7 Thickness.3 in Table 3: Fin Dimensions Nose cone The coefficient of drag affects the overall performance of the rocket in flight. The goal for the team was to select a nose cone shape with a low drag coefficient to maximize performance. Utilizing the software OpenRocket, the four cone types were compared using the already chosen dimensions for the rocket. The team has decided to decrease the fineness ratio of the nose cone, decreasing to a near 4-to-1 ratio versus the initially designed near 5-to-1 ratio. This change has been made to reduce the length and overall weight of the rocket, and the team can easily acquire these nose cones by purchasing them from several vendors. The material for the nose cone has also changed to being fiber glass, as this is the most common material sold by these vendors. AUBURN UNIVERSITY STUDENT LAUNCH 23

24 Mass Statement The mass of the rocket and all its subsystems was calculated using optimal mass calculations from OpenRocket. In addition to using final masses from last year as a basis, a brick sample of carbon fiber was created to have an accurate density measurement since most of the parts 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 this model presents a sufficient estimate. As the program develops, the model will attain a higher and higher accuracy in its simulation. Section Mass (lb) Percentage Structure % Motor % Rover % Recovery % Airbrake % Total % Table 4: Mass by section AUBURN UNIVERSITY STUDENT LAUNCH 24

25 Figure 6: Booster Assembly Drawing AUBURN UNIVERSITY STUDENT LAUNCH 25

26 Figure 7: Nose Cone Drawing AUBURN UNIVERSITY STUDENT LAUNCH 26

27 Figure 8: Fin Drawing AUBURN UNIVERSITY STUDENT LAUNCH 27

28 Figure 9: Full Rocket Drawing AUBURN UNIVERSITY STUDENT LAUNCH 28

29 3.2 Subscale Flight Results Chosen Scaling factors A 2/3 scale was applied to the vehicle diameter and approximately applied to the length of the vehicle. The diameter of the rocket was the primary scaling factor as the diameter controls the area of the vehicle perpendicular to the flow and as such has the most significant impact on aerodynamic performance. The length was approximated to a 2/3 scale to ensure integrity of the subscale, but it was not exact due to the following safety concren. The team was unwilling to compromise the length of the recovery sections or the booster section due to the potential hazards that would be induced. For example, shrinking the recovery compartments further would have reduced the volume for parachutes too much, resulting in the team either having to use parachutes too small for safe descent or risk parachute deployment failure. Furthermore, shortening the booster section was simply not possible due to the size of the motor chosen for matching velocity while maintain the desired coupler overlap. Figure 10: Acceleration Data from Subscale Launch AUBURN UNIVERSITY STUDENT LAUNCH 29

30 Figure 11: Vertical Acceleration over Time in Simulation After flying the subscale, it was important to compare the team s OpenRocket simulation predictions with the flight data. As shown in the figure above, the subscale accelerometer data corresponds to the data from OpenRocket, showing similar accelerations and decelerations trends throughout flight. In particular, the two exhibit the same maximum acceleration during the boost phase and recovery during coast to altitude. During descent, it first appears that the two are very different, with the subscale flight recording large accelerations while the simulation predicts smooth descent. The team believes the oscillations on the real dataset are due to the rocket rocking in an irregular pendulum motion in the wind under parachute during descent, especially as all the irregular peaks seem to exhibit the same maximum and average to a much lower value. Knowing that the simulated behavior can be considered accurate is very important for the design of the altitude control system. Next, another figure shows a comparison between the subscale and simulation results that are comparing the vehicle s altitude and velocity. Once again, a good similarity between the two results can be seen. AUBURN UNIVERSITY STUDENT LAUNCH 30

31 Figure 12: Subscale Flight Data Comparison of Velocity and Altitude AUBURN UNIVERSITY STUDENT LAUNCH 31

32 Updating the Simulation Multiple simulations were run before and after the subscale launch. Initial simulations provided a predicted apogee of 5554 ft, and a static stability margin of 3.2 calibers. The simulated subscale weighed 22 pounds. Although the motor selected for the flight was simulating a higher apogee than desired, it was flown because it was readily available to the team and, of the alternatives, it most closely simulated the maximum velocity the full scale vehicle is expected to experience. The team s successful subscale launch took place on October 27, 2018 in Tuscaloosa, AL. Upon arrival, the rocket was assembled, and the altimeters and recovery system were tested to verify integrity. The rocket was successfully flown, launched, and recovered. The rocket turned out to be lighter by 2 pounds versus the predicted mass. OpenRocket Simulation Recorded Flight Data Updated Simulation Apogee (ft) Max Velocity (ft/s) Table 5: Subscale Launch Data vs Simulation The flight data did not match the simulated data, but the team believes this is primarily due to the change in weight of the rocket. When inputting accurate weight values into OpenRocket, values much closer to those achieved in the flight test were output. The results for the simulation with two pounds taken out at approximately the CG are shown in Table 5: Subscale Launch Data vs Simulation. The simulated altitude was then slightly below the flight data, but the max velocity was very similar. From the updated OpenRocket simulations, the coefficient of drag was determined to be 0.5. This was used to update the value used in the AUSL team s most recent full scale simulations discussed Mission Performance Predictions 3.4. AUBURN UNIVERSITY STUDENT LAUNCH 32

33 Full Scale Modifications due to Subscale The fins were made thicker to induce more drag on the rocket, this was due to how high the subscale flight flew. In order to reduce the maximum velocity and altitude, the most effective way without making major changes to the design was to make the fins slightly thicker. This increased the coefficient of drag and through that decreased the altitude. The subscale flight tests determined that the design is stable and will perform effectively. Because of the complex aerodynamic shape of the rocket, simulations were not exactly accurate at first, but with the subscale data the team is confident the current simulation iteration is far more accurate than before. More tests will be run to further improve and verify the accuracy of the simulation model. The design of the rocket will not change because of the latest subscale flight. The major insight from the subscale launch came from the vast overestimates of the weight of the different subsystem components. To ensure that this does not happen in the full-scale, each subsystem will be re-examined to ensure that the proper weight is simulated for each section. Of course, the team is happy to have come in under weight as opposed to over, and a lighter weight of the final rocket will be easily handled by the altitude control system. AUBURN UNIVERSITY STUDENT LAUNCH 33

34 3.3 Recovery Subsystem Recovery Subsystem Overview The Auburn Student Launch team will be employing a dual-stage, dual-system recovery approach this year. This approach was deemed necessary with team s integration of the rover payload. The rover payload requires a complete separation of the rocket at the nose cone without any dangling wires or shock cords. This would not be possible without 2 independent recovery systems. The lower recovery system consists of 2 parachutes, a drogue and a main, contained within their own separate housings. The housing for the drogue chute is placed just above the embedded systems portion of the rocket with the main parachute housing and Barometric Avionics Enclosure (BAE) respectively being stacked on top of that. The drogue chute will be deployed at apogee at a height of 4700 ft. followed by a main parachute deployment at 500 ft. Both parachutes will be deployed from their own individual compartments with redundant black powder charges and shear pins. These black powder charges will be ignited by electronic matches connected to redundant altimeters housed within the BAE. The upper recovery system will be housed in the nose cone and will be a servo controlled, locking arm release. This system consists of 2 altimeters, 2 microcontrollers, and 2 servos. These components would be assembled with the 2 different servos attached to the same location, above and below the locking arms in such a manner as to create a completely redundant mechanical release. At 700 ft, the altimeters will send a signal to their respective microcontrollers causing them to turn the servos, retracting the tabs holding the nose cone in place. This system will also include a single parachute and 2 bulk plates. The bulk plates will be placed on either side of the parachute and will ensure that the parachute or its shock cord do not become entangled in the electronics of the upper recovery system or the payload. The bulk plate between the upper recovery system electronics and the parachute will have a U-bolt mounted into it to anchor the parachute to the nose cone. The bulk plate between the payload section and parachute will have springs fixed to the payload side of the plate and will be compressed against the payload mount. The springs are in place to aid the nose cone in separating from the rocket body. This bulk plate will also be tethered to the shock cord of the parachute and will be pulled out of the rocket with the nose cone. Once the locking arms release at 700 ft., the springs, aided by the jerk of main parachute deployment at 500 ft, will push the nose cone and parachute out of the rocket and allow it to AUBURN UNIVERSITY STUDENT LAUNCH 34

35 descend safely to the ground. Figure 13: Recovery Overview Diagram below shows the different stages of the recovery system during descent. Figure 13: Recovery Overview Diagram The main parachute deployment height has decreased from 600 ft to 500 ft since PDR. This change was made because the estimated weight of the rocket increased as well as the size of the parachutes. The increased parachute size caused the time from apogee to touchdown to be greater than 90 seconds and thus the main deploy height was lowered to get total drift time back under the required time limit. Structural Elements The main component of Auburn's lower recovery system will be the BAE. The BAE will be formed by an 8 in. long cylindrical fiberglass coupler. The team has chosen this option versus a 12 in. BAE with carbon fiber board due to its many benefits. These include its shorter length, which will decrease the overall length of the rocket, to keep it within requirements. The weight of the 8 in. BAE is also an advantage because it is lighter than the 12 in. BAE due to the reduction in length and, therefore, material. Though its building process and installation are slightly more involved, the additively manufactured 8 in. BAE CAD model is much easier to modify than the carbon fiber board if changes need to be made moving forward. The 8 in. BAE with the 3D printed housing AUBURN UNIVERSITY STUDENT LAUNCH 35

36 provides a more optimal shield from back-pressure than the 12 in. BAE with the carbon fiber board because the design calls for a smaller cross-sectional hole for the charge wires, which reduces the chance of damage to the electronics or faulty readings by the altimeters. A Pugh chart detailing this comparison can be seen below in Table 6: BAE Options Pugh Chart. Categories Weight 8in BAE with 3D printed housing 12in BAE with carbon fiber board Mass Savings Length Ease of installation Shielding from back-pressure Total N/A Table 6: BAE Options Pugh Chart Within the 8 in. coupler, there will be a 3D printed housing for both the altimeters and their batteries to make efficient use of the space as seen in Figure 14: BAE Model. There will be three holes through the length of the housing to run charge wires and threaded rods. The hole for the wires will be in the center of the housing and half an in in diameter. It will continue down through the bottom bulk plate cap exclusively. The holes for the two threaded rods will have a 0.5 in. diameter and 2 in. away from either side of the centered charged wire hole. The rod holes will continue through both bulk plate caps. Rods will be fitted through these holes and secured with lock nuts on either side to secure the BAE together for the duration of the flight. The coupler forming the outside of the BAE will be sealed off with bulk plate caps on both ends. After the wires have been pulled through it, this hole will be sealed with epoxy to help eliminate the chance of back pressure causing damage to the altimeters or forces from parachute deployment pulling on the charge wires and damaging altimeter terminals. The BAE will serve as the coupler between the lower parachute housings and the payload section. Neither of these sections will separate once the rocket is assembled. AUBURN UNIVERSITY STUDENT LAUNCH 36

37 Figure 14: BAE Model Two additional holes will be drilled in the bottom bulk plate cap of the BAE for a U-bolt. This U- bolt will be the anchor point for the main parachute and will be attached with lock nuts on either side of the bulk plate to keep it from moving during parachute deployment. Centered on the outside of the BAE will be a 2 in. long ring of fiberglass that coincides with the diameter of the outside of the rocket. The switches and pressure holes for the altimeters will be located on this ring. The key switches located on the ring allow the team to externally arm the altimeters while the rocket is assembled so that the tube connections between the payload section, the BAE, and the main parachute housing are continuous and smooth. The stable tube connections minimize the impact on the aerodynamic performance of the rocket. In between the drogue and main parachute housings will be a 12 in. fiberglass coupler with a bulk plate fixed inside to separate the two compartments. The bulk plate will be fixed 6 in. away from either end of the coupler to separate the two sections evenly and give both the main and drogue tube a 6 in. shoulder. In this bulk plate there will be 5 holes, 4 of which will be used for U-bolts on either side of the plate secured in the same fashion as the one below the BAE. The 5 th hole will be used for the e-match wire for the drogue charge. Separation points on both sides of the coupler will each use 3 shear pins. Located in the nose cone will be a second altimeter bay to control the upper recovery system. A housing similar to the one seen in Figure 14: BAE Model will be used to secure all the upper recovery system electronics. The second servo will be mounted to the top of the bulk plate fixed to the bottom of the nose cone. The top of the housing will be tapered to reflect the taper of the nose cone, and an additional hole will be placed in the bottom of the housing to secure one of the servos. Both servos will be placed radially off center of the rocket by 2 in and attached to the actuation point of the locking arms via a 3:1 gear ratio. This will be done to utilize the gear ratio s AUBURN UNIVERSITY STUDENT LAUNCH 37

38 effect and create more torque to retract the arms. A detailed view of this assembly can be seen below in Figure 16: Nose cone Detail Model. There will be 3 slots cut in the nose cone and outer body of the rocket for the arms to slide into in order to lock the nose cone to the body. The arms, when locked into place, will be flush with the outer surface of the rocket in order to eliminate unnecessary drag during ascent. The arms will have a ring of support structured around them to reduce the moment on the arms themselves and strain on the servo. On the bottom of the nose cone shoulder there will be a bulk plate with U-bolt for the parachute to attach to. 3 threaded rods will pass through the bulk plate, arm support structure, bottom level and middle level of the altimeter housing with lock nuts on either side of each piece, totaling 6 lock nuts per rod. This will be done to ensure that each of these components stay in the same relative position to each other throughout the duration of the flight. The entire assembly will be fixed to the nose cone via 4 machine screws. These 4 screws will be secured in the middle level of the altimeter housing through the top of the shoulder of the nose cone. In order to keep the screws from causing part of the nose cone shoulder to be exposed, the end of the airframe tube will be notched. This will be done to allow the screws, switches, and pressure holes to slide down to where the nose cone is fully seated against the top of the rocket. This is also done to relieve any undue stress from boost on the locking arms by transferring those forces directly to the airframe. The CAD model of the upper recovery system can be seen below in Figure 15: Nose cone Full Model. AUBURN UNIVERSITY STUDENT LAUNCH 38

39 Figure 15: Nose cone Full Model AUBURN UNIVERSITY STUDENT LAUNCH 39

40 Figure 16: Nose cone Detail Model Materials The materials chosen to create the team's recovery subsystems are of the utmost importance. They must work properly for the team s payload, a rover that cannot be damaged when it returns to Earth. The materials of the parachutes, shock cords, and bulkhead attachments must be both strong and lightweight in order to be effective while not adding unnecessary weight to the rocket. The first material to consider is that of the parachute. The parachutes must be both strong and lightweight to work to add minimal mass to the rocket. Ideally, the parachute material should also hold up to multiple tests and launches without being badly damaged. The material the team has chosen is ripstop nylon because of its superior strength and low weight. Ripstop nylon has a tensile strength of 1500 psi and can withstand stretching up to 40 percent of its length without ripping (compared to cotton s 400 psi and 10 percent stretch), making it the superior option for the parachute. A full comparison is detailed in a Pugh chart below in Table 7: Parachute Materials Pugh Chart. AUBURN UNIVERSITY STUDENT LAUNCH 40

41 Weight Ripstop Nylon* Cotton Strength Weight Cost Total N/A Table 7: Parachute Materials Pugh Chart The next two materials to be considered are the shroud lines and shock cord for the parachutes. The shroud lines connect the parachute to the shock cord, which in turn connects to the rocket body, and for this reason it is crucial for the team to choose a good material for each. The team chose to use paracord for the shroud lines of both parachutes, and 0.5-in tubular nylon for all shock cords. Paracord has the least mass and volume but comparably inferior strength to tubular nylon, making it ideal for shroud lines but too weak to be used on shock cords. Because of this, the team has chosen 0.5 in. tubular nylon for the shock cords. It is very light and takes up the least amount of space while still being able to withstand the shock of deployment. It is easy to handle, cost effective, and includes a wrap-around webbing that increases the overall strength per in. Its high flexibility and pliability allow it to easily glide over rough surfaces and pack into the vehicle body compactly. Pugh charts detailing the decision process for both shroud lines and shock cord can be seen below in Table 8 and Table 9. Weight Paracord*.5 Tubular Nylon 1 Tubular Nylon Kevlar Size Weight Strength Cost Total N/A Table 8: Shroud Line Pugh Chart AUBURN UNIVERSITY STUDENT LAUNCH 41

42 Weight Paracord.5 Tubular Nylon* 1 Tubular Nylon Kevlar Strength Size Weight Cost Total N/A Table 9: Shock Cord Pugh Chart Though the material of the shock cord is important, the team must also consider the length of shock cord used. The shock cord must be long enough to allow adequate separation between the sections of the rocket and to minimize rebound when the shock cord is fully extended. This prevents any section from causing damage to adjacent ones. However, unnecessarily long shock cord adds extra mass and volume, which can make parachute ejection more difficult. Generally, the length of shock cord needed is proportional to rocket length, and can be estimated by the following equations: LL dddddddddddd = 3.5 (bbbbbbbbbbbbbb ssssssssssssss llllllllllh + dddddddddddd ssssssssssssss llllllllllh) LL mmmmmmmm = 4 (uuuuuuuuuu tttttttt llllllllllh + nnnnnnnn cccccccc llllllllllh) Using a booster section length of 35 ines, drogue section length of 12 ines, nose cone length of 24 ines, and upper tube length of 51 ines (which includes the rover section, recovery bay, and main parachute compartment), this becomes: LL dddddddddddd = 3.5 (35 iiii + 12 iiii) = iiii LL mmmmmmmm = 4 (53 iiii + 24 iiii) = 308 iiii For the drogue, this length was increased to 15 ft (180 ines) to further prevent rebound damage. The main shock cord length was decreased to 23 ft (276 ines) to reduce mass and volume within the main parachute section, allow for easier parachute deployment, and aid in separating the nose cone from the rocket body due to the slightly increased jerk when the main parachute inflates. A shorter shock cord can cause some rebound damage if the ejection charge is too great; however, AUBURN UNIVERSITY STUDENT LAUNCH 42

43 with the team s calculated ejection charge sizes, 23 ft of shock cord is adequate to avoid damaging the rocket. How the parachute is anchored to the rocket body is another area of importance. The team will use U-bolts to secure our parachutes to our rocket. The eye bolt is more susceptible to causing the shroud lines to tangle up and only provides one point of attachment. The U-Bolt provides two attachment points and is less susceptible than the eye bolt to cause entanglement of the parachutes. A Pugh chart comparing these 2 items is shown below in Table 10: Attachment Hardware Pugh Chart. Weight U-Bolt* Eye-Bolt Chance of Entanglement Attachment points Weight Ease of Installation Total N/A Table 10: Attachment Hardware Pugh Chart Several pieces within this year s rocket will be 3D printed, a few of which will bear some load. Because of this, the team has analyzed traditional ABS and a new product called Onyx produced by Markforge. Onyx is a product consisting of nylon and chopped carbon fiber that is 1.4 times stronger than traditional ABS. A Pugh chart comparing the 2 materials can be seen below in Table 11: 3D Printed Material Pugh Chart Weight ABS Markforge Onyx Strength Weight Cost Total N/A Table 11: 3D Printed Material Pugh Chart AUBURN UNIVERSITY STUDENT LAUNCH 43

44 Ejection System A mechanical release and 4F black powder have been chosen for the ejection of the team's parachutes. A comparison of these two systems and a CO2 pressure event can be found in Table 12. The team will use a mechanical release for the nose cone separation, which was chosen because of its ease of launch-day installation and its lack of generated heat and pressure. The team has determined that this lack of heat and pressure will give the rover payload a much lower chance of being damaged than a standard black powder charge separation would. This makes the mechanical separation a safer and less failure-prone choice. The main setback of this system is design complexity, but this additional preparatory work is countered by a desirable decrease in work the day of the launch and increase in the safety of the payload. Black powder is an effective, reliable means of pressurization that the team has had success with in the past. The team will continue to use this method for the main body separations since there is not a concern of damage to a payload in this section. Compared to CO2 ejection, black powder can produce greater pressures per cubic in required to house the system. For the black powder events, two charges are placed within 3D printed charge cups and armed with electronic matches: the first charge is the primary means of ejection, and the other will be a backup charge for redundancy and to decrease the chances for failure of the recovery system. The redundant apogee charge is set to fire 1 second after the first charge. The charges are not ignited at the same time as that has the potential to cause damage to the airframe and failure of the system. AUBURN UNIVERSITY STUDENT LAUNCH 44

45 Weight Mechanical Release Black Powder CO2 pressure event Simplicity of design Heat generated Pressure generated Potential damage to Payload Ease of installation 3 2* Volume required Total N/A Table 12: Ejection Pugh Chart * No pressure generated; ease of separation is a comparable measure used in this case. Black powder, when ignited, can be approximated as an ideal gas. As such, the Ideal Gas Law can be used to calculate an estimate amount of black powder needed for ejection. The Ideal Gas Law is as follows: Equation 1 PP VV = nn RR TT iiii llllll RR = llllll RR TT = 1260 RR Ignoring the volume of the parachutes and other recovery systems contained within the two sections, a conservative volume for the drogue and main parachute sections can be calculated with an inner diameter of 6 in. and respective internal usable lengths of 8 in. and 25 in. as: AUBURN UNIVERSITY STUDENT LAUNCH 45

46 VV DDDDDDDDDDDD = AA LL = ππ 4 dd2 LL = iinn 3 VV MMMMMMMM = AA LL = ππ 4 dd2 LL = iinn 3 Using Equation 1, the team calculated the amount of black powder needed to produce pressures sufficient enough to shear the nylon screws. With 3 nylons screwed, each rated for 50 psi of shear, pressure required to shear the screws is: Equation 2 PP = 3 FF ππ = 5.3 pppppp dd2 4 Which, in conjunction with Equation 2, yields charge sizes for 4F black powder of: nn DDDDDDDDDDDD = nn MMMMMMMM = PP VV RR TT gg = 1.62 gg llllll PP VV RR TT gg = 5.07 gg llllll The estimates provided above were determined to be the minimum amount of black powder required for a successful separation and were verified via ground testing. Both charge sections are filled with fireproof cellulose insulation (colloquially known as barf ) to protect the parachutes from the heat of ejection. The team then applied a factor of safety of 1.5 to both charge values to guarantee separation on launch day. A second round of ground ejection tests were performed with 3.25 g drogue and a 7.61 g main charge to ensure that the increased charge size would not compromise the structural integrity of the airframe or singe the parachutes. The airframe and parachutes were examined after the completion of the test and was proved successful when neither the airframe nor parachutes showed signs of damage. AUBURN UNIVERSITY STUDENT LAUNCH 46

47 Parachutes Auburn s double system recovery approach makes use of three separate parachutes designed and constructed in house by the Auburn University Student Launch team. Following the first event at apogee, the drogue will be deployed from the middle of the rocket with both halves still connected. This connection will be maintained by equal lengths of shock cord anchored to both halves of the rocket and connected at the base of the drogue parachute shroud lines. This will stabilize the descent until main parachute deployment. At 700 ft, 200 ft prior to main deployment, the nose cone recovery system will activate, the nose cone will completely separate from the rocket body, and the nose cone parachute will deploy. When the rocket reaches 500 ft in altitude, a second set of charges from the BAE will push off the parachute coupler to release the main parachute. A spill hole 20% of the total base diameter of the chutes will be added to the main parachute. The 20% diameter of the spill hole is used because it only reduces the area of the parachutes by 4%, allowing enough air to go through the spill hole to stabilize the rocket without drastically altering the descent rate. The team will be using a circular design for our drogue parachute. A circular design is best for the drogue parachute in this case because it stabilizes the rockets descent while keeping the descent rate high enough to fulfill the descent time requirement. It also has the added benefit of being easy to manufacture and pack on launch day. A circular drogue parachute size can be found by calculation based on the length and diameter of the rocket body. Equation 3 DD DDDDDDDDDDDD = 4 LL TTTTTTTT DD TTTTTTTT ππ The team s rocket will have a length of 124 in. and a diameter of 6 in, yielding a drogue chute diameter of in. The main parachute and nose cone parachute will be hemispherical in shape. The hemispherical design will produce the most drag, allowing the rocket to have maximum drag with minimum weight. A Pugh chart can be seen below in Table 13 comparing different parachute shapes. The AUBURN UNIVERSITY STUDENT LAUNCH 47

48 shape of the main parachutes and their gores for both 8 and 6 gore configurations can be seen in Figure 17: 6 Gore Template and Figure 18: 8 Gore Template. Baseline Square Circular Hemispherical Drag Produced Ease of Manufacturing Stability Total N/A Table 13: Parachute Shape Pugh Chart Figure 17: 6 Gore Template AUBURN UNIVERSITY STUDENT LAUNCH 48

49 Figure 18: 8 Gore Template Figure 19: Hemispherical Parachute When determining the size of the parachutes, the team must first consider kinetic energy since it is directly related to descent velocity. A low kinetic energy is important to both the potential safety of bystanders and reusability of the rocket. The kinetic energy of the rocket upon landing can be calculated using the following formula: Equation 4 KKKK = 1 mm VV2 2 AUBURN UNIVERSITY STUDENT LAUNCH 49

50 Where m is mass and V is descent velocity. Mass in this case is not the mass of the entire rocket. Since the rocket is drifting down in multiple pieces that are simply tethered and not rigidly connected, each piece will have a unique kinetic energy value. Therefore, for the equation above, m will be the mass of the heaviest piece under each parachute. The team will use this equation to solve for the descent velocity to determine the size of the parachute. Solving the above equation for V yields: Equation 5 22 KKKK VV = mm Auburn Student Launch will apply a factor of safety of 1.5 to the competition kinetic energy requirement of 75 ft-lbs of force upon impact for a target maximum kinetic energy of 50 ft-lbs upon landing for a single section. Using this value and an initial mass estimate of 18 lbm for the upper section, the equation above yields: VV = ffff 50ffff llll ss 2 = ffff 18llbb mm ss Applying the same technique to the nose cone with an initial mass estimate of 7 lbm, the descent velocity equation shows: VV = ffff 50ffff llll ss 2 = ffff 7llbb mm ss Parachute areas for hemispherical shaped chutes are determined using the following equation: Equation 6 AA = 22 FF ρρ CC DD VV 22 Where F is force, ρ is density of the air, CD is the drag coefficient and V is the descent velocity calculated in the previous equations. This F force is the force that the parachute experiences from the mass of all rocket body pieces multiplied by the acceleration due to gravity. The team will use AUBURN UNIVERSITY STUDENT LAUNCH 50

51 this equation to calculate an appropriate area for the main parachute so that the kinetic energy of the rocket does not exceed 75 ft-lbs during recovery and remains within safe limits. Testing on hemispherical parachutes done in prior years by Auburn University Student Launch has concluded that hemispherical parachutes have a drag coefficient of Using this data, a mass of 41.9 lbs after a burnout with 7 lbs of that being in the nose cone, and descent velocities of ft/s and ft/s calculated above, the team has used Equation 6 to determine the size of the main and nose cone parachutes. AA MMMMMMMM = AA NNNNNNNNNNNNNNNN = llbbmm 32.2 ffff ss llbb 2 mm ffff = fftt2 fftt ss 2 (7.0 llbbmm 32.2 ffff ss 2) llbb 2 mm fftt = 9.79fftt2 fftt ss The descent velocity needs to be under the drogue later to help determine drift, so that is computed here as well. Since the drogue parachute is circular, we will need a different value for the coefficient of drag than what we have been using previously for the hemispherical parachutes. The team has analyzed several technical reports and determined that the coefficient of drag of circular parachutes is 0.8. Using that value and burnout weight of 41.9 lbs, we can calculate the descent rate of the rocket under drogue by solving Equation 6 for V. Equation 7 VV = 2 FF ρρ CC DD AA VV = llllll 32.2 ffff2 ss llll mm ffff ffff2 = ffff ss AUBURN UNIVERSITY STUDENT LAUNCH 51

52 Main Nose Cone Area ft^ ft^2 Diameter 8.94 ft in Diameter of Spill hole in 5.99 in Number of Gore 8 6 Width of Each Gore at Base in in Height of Each Gore in in Circumference at Base in in Table 14: Parachute Sizing Details Electronics There will be a total 4 altimeters, 2 microcontrollers and 2 servos within recovery system for the AUSL team s rocket. Two of the altimeters will be mounted inside the BAE and be used to provide doubly redundant deployment for drogue and main parachute deployment at predetermined altitudes. The other two altimeters, both microcontrollers and both servos will be contained within the nose cone recovery system. The electronics in the nose cone will operate as a sub-system in the following fashion. Once an altimeter determines that the rocket is at the correct altitude, it will send an electrical impulse to its microcontroller the same way it would if an electronic match were attached to it. The microcontroller will be programed in such a manner that it is waiting for this signal to arrive from the altimeter. Once the signal is received, the microcontroller will begin turning its attached servo motor 170 degrees to completely retract the locking arms that are holding the nose cone to the rocket. There will be two of these identical and isolated sub-systems contained within the nose cone recovery system to provide complete double redundancy. A single servo will be able to fully retract the locking arms. A wiring diagram can be seen below in Figure 20: Nose cone Wiring Diagram. AUBURN UNIVERSITY STUDENT LAUNCH 52

53 Figure 20: Nose cone Wiring Diagram There will a total of 4 pull pin switches for the nose cone electronics; one for each altimeter and microcontroller. The arming of the nose cone electronics will be done in two separate stages. The day prior to launch day all the nose cone electronics will be tested and sealed into the nose cone with all 4 of the pins being inserted into the switches to turn everything off and to retract the locking arms. On launch day once the nose cone parachute is packed, the nose cone will be inserted into the top of the rocket and the switches for the microcontrollers will have their pins removed. This will turn the microcontrollers on and actuate the locking arms locking the nose cone in place. Once the rocket is on the launch pad, the switches attached to the altimeters will have their pins pulled to turn on the altimeters in flight mode. The team will bring a 6-ft. step ladder to the pad on launch day to access the altimeter pins since these pins will be at the top of the rocket and will be roughly ft off the ground. The team evaluated 5 different altimeters to determine which would be best suited for this project. The team has decided to use 4 Perfectflite StratologgerCFs for all the altimeters over the AltusMetrum Telemetrum, AltusMetrum TeleMega, Sparkfun s XA1110 GPS Breakout, and Sparkfun s MPL3115A2 Pressure Sensor Breakout. The most important feature the team considered when selecting the altimeters was how difficult it would be to interface with. The StratologgerCF has a serial stream of altitudes available via the data port. We will be using it because a serial data stream AUBURN UNIVERSITY STUDENT LAUNCH 53

54 is easy to parse and easy to connect to a microcontroller as opposed to the I 2 C protocol of the XA1110 and MPL3115A2. While we have since chosen to simply receive a pulse from the electronic match trigger, the team wanted to keep multiple connection options while deciding on an altimeter. The TeleMega and TeleMetrum documentation provided no information on features that would have made them useful in our application, such as reading data with a microcontroller or controlling servos with the altimeter itself. Size was also a consideration, as we are installing the system into the nose cone around a rover. The Stratologger CF is the compact footprint version of the original Stratologger, at just 2 long and.84 wide. The weight was considered, with the tiny MPL311A2 being the lightest and the StratologgerCF also very light, just 0.38 oz. The weights of the different altimeters were not significant enough to sway our decision based on difficulty of use. The cost was also considered because of the massive disparity between the features we needed and the cost of each product. The TeleMetrum and TeleMega were essentially unusable for this purpose despite being the most expensive products compared. A full comparison of the altimeters as well as a picture of the Perfectflite StratologgerCF can be seen below in Table 15: Altimeter Pugh Chart and Figure 21: PerfectFlite StratoLoggerCF. Criteria Weights StratologgerCF TeleMetrum TeleMega XA1110 Breakout MPL3115A2 Breakout Ease of Interface Weight Cost TOTAL N/A Table 15: Altimeter Pugh Chart AUBURN UNIVERSITY STUDENT LAUNCH 54

55 Figure 21: PerfectFlite StratoLoggerCF The microcontroller that the team has chosen to use for the nose cone electronics is the Adafruit Trinket pictured in Figure 22: Adafruit Trinket. Like the other microcontrollers, the Trinket is programmed with the Arduino IDE and is available in 3.3V or 5V logic levels. Because of the interface between the servo and altimeter, we chose the 5V model. In terms of processing speed and memory, the Trinket is one of the slowest and has the least amount of memory available, but these categories were weighted lowest because the microcontroller will not be performing intensive tasks requiring complex code or very fast response time. The two most important qualifiers were ease of integration, how much extra work would be needed to use the microcontroller with the other components, and the size, as we would need to fit the microcontroller in the nose cone with a battery, servo, and altitude sensor. Size was chosen as the most important factor because a smaller microcontroller would allow much more design flexibility. The Trinket was the smallest at 1.2 x 0.6. A full comparison of the microcontrollers as well as a picture of the Adafruit Trinket can be seen below in Table 16: Microcontroller Pugh Chart and Figure 22: Adafruit Trinket. AUBURN UNIVERSITY STUDENT LAUNCH 55

56 Criteria Weights Adafruit Trinket Arduino Uno Arduino Nano Size Ease of Integration Programming Interface Processing Speed Memory TOTAL N/A Table 16: Microcontroller Pugh Chart Figure 22: Adafruit Trinket The servo that the team has selected is the HS-485HB. Price and weight were the lowest weighted criteria for the servo because price is a much lower priority than finding a good servo that will work in our application. Weight is not as important in the nose cone area and, again, getting a light servo is a much lower priority than ensuring that our servo will work. Size was the next criterion as a smaller servo will give us more flexibility in mounting solutions and orientation. The HS- AUBURN UNIVERSITY STUDENT LAUNCH 56

57 485HB is 1.57 x 0.78 x 1.5 and 1.6 oz. Torque was the second most important factor as more torque will almost certainly aid performance and repeatability while having practically no detriments. The HS-485HB has a stall torque of 83.3 oz/in. at 6V and is essentially a faster and higher torque version of another Hitec model. Ease of integration was deemed the most important factor as a continuously rotating servo like the FS5103R would make a release mechanism more complex and a servo with a nonstandard protocol would make the electronic interface more difficult. A full comparison of the two servos can be seen below in Table 17: Servo Pugh Chart. Criteria Weights HS-485HB FS5103R Ease of Integration Torque Size Weight Cost TOTAL N/A Table 17: Servo Pugh Chart The final component to decide were the switches for the various electronic components. We decided to use pull-pin switches for the nose cone recovery system and key switches for the lower recovery system. The pull pin switches were chosen for the nose cone primarily because of their vibration resistance and minimal aerodynamic impact. Pull pin switches require only a small hole on the outside of the rocket while all the other options require some protrusion, which could affect trajectory when mounted just below the nose cone. Pull pin switches also are not significantly affected by vibration or impact, unlike toggle switches. The team selected key switches for the BAE because of the ease of installation and resistance to vibration and accidental actuation. Key switches do not require any mount inside the body tube, just a hole drilled in the tube. Key switches are also the most secure switch against accidental activation, as a key is required to flip the switch whereas a pull pin switch or toggle switch could be more easily triggered by accidentally bumping the switch or pulling the pin. Aerodynamic impact is minimal due to the sloped bezel and is not particularly significant at the location of the BAE. The team has flown multiple rockets with key AUBURN UNIVERSITY STUDENT LAUNCH 57

58 switches in the past, including the subscale model, with no ill effects. Pugh Charts can be seen below in Table 18: Switch Pugh Chart - BAE and Table 19: Switch Pugh Chart - Nose cone detailing these decisions. Criteria Weights Key Switch Pull-pin Switch Flip Switch Vibration Resistance Aerodynamic Profile Ease of Installation TOTAL N/A Table 18: Switch Pugh Chart - BAE Criteria Weights Key Switch Pull-pin Switch Flip Switch Vibration Resistance Aerodynamic Profile Ease of Installation TOTAL N/A Table 19: Switch Pugh Chart - Nose cone AUBURN UNIVERSITY STUDENT LAUNCH 58

59 3.4 Mission Performance Predictions Figure 23: Internal Components of Rocket The chosen target altitude for the team is 4700 ft. This number was chosen because the team feels that it provides sufficient margin less than the predicted final altitude from the most up to date simulation, which was 5082 ft. This leaves room for the mass of the vehicle to increase or decrease and still have a final altitude above the 4700 ft target, giving the altitude control system room to respond. In order to get an initial, fairly accurate projection of the altitude, the team used OpenRocket, a freeware program designed to calculate various parameters in rocket flight. Given the team s experience with this software in the previous years, the team is confident in the ability of OpenRocket to produce accurate estimates of the altitude. With the current motor selection, the Aerotech L14210R, the current projected apogee for the vehicle is 5082 ft. While this is currently above the chosen target altitude, the team requires the projected altitude to be above the target altitude for the altitude control module to control the rocket s final altitude. Additionally, since mass increases of up to 25% are very common in the manufacturing process, the team considers this to be an acceptable initial altitude estimate that will decrease throughout the project. Below is the average data from the standard OpenRocket simulation out of many different runs. AUBURN UNIVERSITY STUDENT LAUNCH 59

60 Flight Simulation Data (Wind = 0 mph) Maximum Velocity 708 ft/s Maximum Acceleration 269 ft/s 2 Launch Weight Burnout Weight Length Maximum Diameter Launch Stability Velocity off Rod 47 lbm 41.1 lbm 124 in 6.25 in 2.2 calibers 70.7 ft/s Table 20: Open Rocket Flight Simulation Data AUBURN UNIVERSITY STUDENT LAUNCH 60

61 Flight profile predictions The expected components masses were predicted using a combination of OpenRocket predictions and extrapolations from the weights of completed components from previous years. Component Mass (lbm) Nose Cone 3.6 Rover Body Tube 4.6 Main Parachute Body Tube 3.6 Drogue Chute Body Tube 1 Booster Section Body Tube 2.7 Fins (4) 1.6 Centering Rings (3).6 Motor 10.1 Motor Retention.3 Motor Tube 1 Rover and Bay 8 Airbrake System 2 Bulkheads.9 Electronics 2 Recovery System 5 Total 47 Table 21: Component Masses The simulated motor thrust curve for the selected motor can be seen in Figure 4: Motor Thrust Curve. The team is confident that the Aerotech L1420R motor will provide adequate thrust to AUBURN UNIVERSITY STUDENT LAUNCH 61

62 propel the launch vehicle to above the target altitude, enabling the altitude control module to adjust the final altitude to 4700 ft. Using the given component masses, the updated coefficient of drag from subscale, and the L1420R motor, the following trajectory data is predicted for a full scale flight without an active altitude control system. Figure 24: Predicted Altitude over time for flight Figure 25: Predicted velocity over time for flight AUBURN UNIVERSITY STUDENT LAUNCH 62

63 Figure 26: Predicted acceleration over time for flight Figure 27: Predicted drag coefficient over time for flight AUBURN UNIVERSITY STUDENT LAUNCH 63

64 Stability throughout Flight Figure 28: Stability Parameters over Time For the rocket to be stable during flight, the center of pressure must be located aft of the center of gravity. NASA requirement 2.17 (See Section Vehicle Requirements) states that the vehicle must have a stability margin above 2 calibers at rail exit. As can be seen from Figure 28: Stability Parameters over Time, the team s rocket will have a stability margin of 2.2 at rail exit. As the rocket continues to accelerate the stability will continue to rise, with an average value for the entire ascent phase of flight predicted to be The stability decreases after burnout, but the rocket retains a stability margin of above 2 until right at apogee, when the velocity becomes too low for aerodynamic forces to stabilize the rocket. Of course, this is not a problem because the drogue deploys to stabilize the rocket during descent. The location of CG and CP during ascent can be seen. Kinetic Energy, Descent Time and Drift The amount of time it will take the rocket to drift down under a specific parachute can calculated by the following equation using the descent velocity computed in Section 3.3.5: Equation 5. AUBURN UNIVERSITY STUDENT LAUNCH 64

65 Equation 8 tt dddddddddddddd = Δh VV DDDDDDDDDDDDDD Using Equation 8, we produced the following table regarding the descent time for the different sections of the rocket: Drogue Parachute (4700 ft) Main Parachute (500 ft) Nose cone Parachute (700 ft) Total Descent Time Rocket Body sec sec sec Nose cone sec sec sec Table 22: Descent Time Using from Equation 4 section 3.3.5, the team can validate that the rocket will not exceed 75 ft-lbs of kinetic energy (NASA requirement 3.3. That verification can be seen below in Table 23: Kinetic Energy Upon Landing. Booster Section (15 lbs) Drogue Tube and Coupler (2 lbs) Avionics Section (18 lbs) Nose cone (7 lbs) ft-lbs 5.55 ft-lbs ft-lbs ft-lbs Table 23: Kinetic Energy Upon Landing The distance the rocket will drift during descent can be estimated with the following equation. Equation 9 DDDDDDDDDD = WWWWWWWW SSSSSSSSSS tt dddddddddddddd 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 three stages of descent. First, the rocket descends under the drogue parachute from an altitude of 4700 ft. to 700 ft. At 700 ft. the second event occurs, and the nose cone then separates from the rocket body, continuing the rest of the way to the ground under a separate main parachute. A third event occurs at 500 ft when a black powder charge separates the parachute coupler and pushes the main parachute out. Using this information and the weight of the rocket, the team was able to calculate an estimated descent velocity. This descent velocity is then used to ensure drift is kept under a maximum of 2,500 ft away from the launch pad. Separate tables below predict drift for the rocket body and the nose cone. AUBURN UNIVERSITY STUDENT LAUNCH 65

66 Wind Speed (mph) Wind Speed (ft/s) Drift Under Drogue(ft) Drift Under Main Parachute (ft) Total Drift of Rocket Body (ft) Table 24: Rocket Body Drift Wind Speed (mph) Wind Speed (ft/s) Drift Under Drogue(ft) Drift Under Nose cone Parachute (ft) Total Drift of Nose cone (ft) Table 25: Nose cone Drift AUBURN UNIVERSITY STUDENT LAUNCH 66

67 4 Safety 4.1 Launch Concerns and Operation Procedures The team has created a list of procedures for each sub team to follow at launch. These procedures have step-by-step instructions on how to prepare the rocket for launch in the safest way possible. Included in the steps are warnings of hazards, when to use PPE, and which steps are critical to project success. The procedures are based off those used for subscale launches with some changes made to accommodate the transition to full-scale. Each team lead will have a copy of their sub team s procedures at launch and Jackson will be checking in on each team to ensure all safety precautions are being followed. Once a sub team is ready for launch, Jackson and that team lead will check that each step was completed. 4.2 Safety and Environment (Vehicle and Payload) Upon review, all safety tables have been updated to give a better understanding of the effects that each hazard is capable of. Previously, some mitigations were less of a tangible fix and more of a do it right instruction. These have been changed to be focused on mitigating the cause instead of relying on getting the design right. Some of the mitigations to hazards have been verified and their tables have been updated accordingly, while others have planned tests for verification during coming months. Personnel Hazard Analysis Workplace Hazard Analysis: All construction of the rocket is done in on-campus labs. All personnel are expected to adhere to lab rules while using the workspaces. The hazards associated with these labs have been assessed and the results are in Appendix C.1 STEM Outreach Hazard Analysis: STEM Outreach events are conducted at numerous locations over the course of the season. Personnel will work closely with participants and their guardians to AUBURN UNIVERSITY STUDENT LAUNCH 67

68 ensure all safety needs are met. The potential hazards that could occur at these events have been assessed and the results are in Appendix C.1. Failure Modes and Effects Analysis Stability and Propulsion FMEA: The hazards associated with the performance of the rocket during flight and the motor are in Appendix C.2. Dr. Eldon Triggs is the only team member who has the required certification to handle the motor and thus will be responsible for the transportation, storage, and handling of the motor. Vehicle Assembly and Launch Procedures FMEA: Launch day is an enjoyable time for the entire team, however improper assembly of the rocket could easily lead to a catastrophic failure resulting in mission failure. Therefore, it is important to be aware of all hazards associated with the assembly, launch of the rocket, and surroundings to ensure all personnel stay aware of the safety concerns. The potential hazards are located in Appendix C.2. Recovery FMEA: The rocket design this year includes 3 separate recovery events with redundant systems. The hazards for these redundant systems are identical, therefore they are only listed once in the table. The hazards associated with the recovery of the rocket are located in Appendix C.2. Payload and Embedded Systems FMEA: This year we are implementing an altitude control system involving drag plates. Our chosen payload this year is the rover. All hazards associated with these systems are located in Appendix C.2. Environmental Concerns Vehicle Effects on Environment: The goal of NASA Student Launch is to provide viable designs for future NASA missions. If the designs presented are harmful to the environment, then they are not a viable design and fail to meet the goal of the competition. For this reason, it is extremely important that our rocket has a minimal impact on the environment, if any at all. The hazards our rocket poses to the environment and mitigation techniques are located in Appendix C.3. Environment Effects on Vehicle: The majority of environmental effects that we have assessed can be avoided by following NAR Safety Code. The effects assessed are located in Appendix C.3. AUBURN UNIVERSITY STUDENT LAUNCH 68

69 Risk Mitigation and Impact Risk Mitigation and Impact Table: Currently, these tables include some of the higher impact risks that the team has considered, as well as a few risks that we have experienced this season. More focus will be put into this table between now and CDR in order to avoid these risks when the team begins work on the full-scale rocket. This table is located in Appendix C.5. Quantification of Risk Impacts Table: This table quantifies the impacts of the risks assessed in the Risk Mitigation and Impact Table. The impacts are separated into 3 categories: time, cost, and design. All assessments of impact to time are estimates based on the Project Timeline seen in Figure 4. All assessments of impact to cost are based on the sum of the parts lost or damaged in the risk which can be found in their respective subsections budget table. Also included in the cost is any item that is purchased at a launch site as an immediate replacement and is labeled as such. The assessments of impact to design are based on whether or not the risk affects the current project design, and only states what aspects of the design would need to be changed. This table is located in Appendix C.5. AUBURN UNIVERSITY STUDENT LAUNCH 69

70 5 Payload Criteria 5.1 Soil Sample Rover Overview of the Rover The objective of the rover is to satisfy NASA Student Launch Requirements 4.3-Deployable Rover / Soil Sample Recovery Requirements. For the rover to completely succeed in its mission it must satisfy each requirement listed. Overall, the rover begins its mission retained within the launch vehicle by a fail-safe active retention system facilitated by a servo. Once landed after launch, the ground crew awaits the go ahead from the Remote Deployment Officer to remotely activate the rover to deploy from the internal structure of the launch vehicle. After deployment, the rover autonomously moves at least 10 ft away from its deployment site and collects a soil sample. The soil sample is at least 10 ml in volume and the sample compartment subsequently closes to protect the soil after collection. The batteries of the rover are also brightly colored, clearly marked as a fire hazard, easily distinguishable from other parts, and protected from impact with the ground. Chosen Design Components The following sections are repeated and abbreviated from the PDR report to reiterate which design alternative was chosen and why Fabrication Method The mechanical design of the rover is dependent upon the manufacturing methods and materials which are used for fabrication. Considering the limitations associated with each manufacturing process and material, three methods were taken into consideration: 3D plastic printing, traditional metal machining (mills, lathes), and a stored supply of prefabricated parts. Item 3D Printing Traditional Prefabrication Machining Parts/Assembly Cost 5 (Low) 1 (High) 2 (Med-High) Complexity 4 (Med-Low) 1 (High) 5 (Low) Customization 5 (High) 2 (Med-Low) 1 (Low) Totals Table 26: Fabrication Method Trade Study AUBURN UNIVERSITY STUDENT LAUNCH 70

71 The driving factors in choosing to 3D print all components of the rover were cost and simplicity for a non-complex fabrication. Therefore, with the exception of the rover s electronics, all structural components of the rover will be fabricated using the Auburn University Aerospace Engineering department s MarkForged composite 3D printer. Structural components will be printed using onyx, and moving components (such as tank treads) will be printed using nylon Soil Recovery Method The system for sample recovery and its corresponding method were considered first, as the rest of the rover must operate around this system. We considered methods already in existence with modifications to meet our purpose. The main objectives were a conveyor belt type collection unit, an auger type collection unit, a drag type collection unit, and a dig and scoop collection unit. Item Conveyor Auger Drag Dig/ Scoop Complexity 4 (Med-Low) 1 (High) 2 (Med-High) 1 (High) Reliability 5 (High) 1 (Low) 5 (High) 4 (Med-High) Total Table 27: Soil Recovery Choice Trade Study The conveyor belt type collection unit was chosen for its reliability and simplicity. The collection system is the main purpose of the rover, so the rest of the rover must be configured around this. All above options were considered and elaborated on within the team. This allowed for a full system sketch. The belt collection unit allowed for the greatest level of customization and modification to complete the task Motor Choice When selecting motors to drive the rover, only torque, weight, and cost were constraints were taken into account. The rover sub-team quickly decided that speed of the vehicle was of little importance and therefore not considered when deciding on motors, leaving torque as the only mechanical constraint. With strict size limitations in place due to the small rover housing bay and rocket body, small size was of great importance. The motors will be connected to the Arduino Uno via an L298N motor driver that will regulate current to the motors. AUBURN UNIVERSITY STUDENT LAUNCH 71

72 Item Pololu Micro Digi-Key SparkFun Pololu Metal Metal HPCB Stepper Motor Stepper Motor Gearmotor 12V 12V 12V 12V Torque (oz-in) 5 (125) 2 (28) 2 (32) 5 (160) Weight (g) 5 (10.5) 1 (200) 1 (200) 2 (46) Cost ($) 4 (25) 5 (14) 5 (15) 4 (22) Totals Table 28: Motor Choice Trade Study The team selected the Pololu Micro Metal HPCB motor due to the large torque delivered in a micro volume at a reasonable price. Additionally, the testing sub-team determined that powering the motors for the amount of time required competition will not be problematic. At full speed and in the final power configuration, the motors ran for more than 3 hours, which is plenty of time to accomplish the distance requirement; however, more testing will be conducted as the design progresses Movement Choice To facilitate rover movement, 3 options were contemplated. These included 4 circular wheels each with a drive motor, a pair of treads with a single drive motor on each side, and a spider leg concept. The surface we expect to encounter at Bragg Farms in Huntsville, Alabama was the driving factor in our consideration. All methods of movement can be developed in house, but all had different levels of complexity. The four wheeled design would add complexity as two additional electric motors are required and therefore localizing the weight at 4 points along the rover. However, it would provide the most power to the wheels. The treaded design provides the most contact area with the surface. The spider leg concept added more complexity to the design but allowed for greater grip in the potentially muddy terrain. Item Wheels Treads Spider Terrain Grip 1 (Low) 3 (Med) 5 (High) Movement (Slope) 1 (Low) 2 (Med-Low) 1 (Low) Movement (Flat) 5 (High) 4 (Med-High) 1 (Low) Soil Recovery Deployment Ease 5 (High) 5 (High) 1 (Low) Total Table 29: Movement Choice Trade Study AUBURN UNIVERSITY STUDENT LAUNCH 72

73 Due to the light weight of the proposed rover, a large amount of power to facilitate movement is not required. What is required traverse a recently tilled Alabama red clay field in Alabama is a large surface area to be able to maintain grip on the soil. Therefore, after taking into consideration of all variables, the team decided to use the treaded concept Communication Method Selecting the method which we would communicate with the rover to activate it was one of the bigger challenges the rover team faced. Many traditional options of communication were not possible based on competition constraints (range, launch location, etc.). DX6 6-Channel DSMX Item XBee Pro S5B ESP8266 Wi-Fi Transmitter Gen 3 with AR6600T Receiver (SPM6755) Range (m) 5 (14k) 1 (92) 3 (1600) Power 3 (3.3V ma) 3 ( V mA) 3 (3.9-9V) Size 5 (Small) 5 (Small) 5 (Small) Price ($) 3 (100) 5 (7) 1 (230) Total Table 30: Communication Method Trade Study The XBee Pro S3B was selected based on its small size and range. Again, size was a design constraint that was considered in almost every step of the design process. In addition to these properties the XBee is a versatile communication platform with different methods of communicating to our board either via the serial port or through the digital input output pins. This will simplify development of both transmitting and receiving the launch signal by utilizing a signal pair of modules without significant add-ons to facilitate communications. The price is also a gain against some of the expensive alternatives that fit our transmission range. The XBee will attach to the Arduino Uno via an SD Wireless Shield and will receive a signal the rover team sends to it over a laptop or separate controller (internet connection not necessary). AUBURN UNIVERSITY STUDENT LAUNCH 73

74 Microcontroller Choice Since the rover will be inside a small compartment within the rocket, it needed something that is small and can be powered by simple batteries at the lowest possible cost. The commands we are giving the rover are simple, so computational power is not a serious problem. The microcontroller boards the team considered are compared in Table 31: Microcontroller Trade Study. Item Arduino Uno R3 Raspberry Pi Arduino Mega Voltage Required (V) 4 (5) 4 (5) 4 (5) Weight (oz) 4 (1.6) 2 (4.8) 4 (1.3) Cost ($) 3 (22) 2 (35) 2 (38.50) Total Table 31: Microcontroller Trade Study The team chose the Arduino Uno because of its low power requirement and lightweight design, which we can obtain at a reasonable price. Mission Design Concept of Operations The rover concept of operations begins at the launch field. Batteries will be plugged in, and the rover will turn on. The initialization code will tell the rover Active Retention System (ARS) to open. The rover will be placed on the sled mounted in the launch vehicle. A wireless signal will be sent to tell the rover to activate the ARS and the servo will close, attaching the rover to the launch vehicle. From there, the nose cone will be attached to the vehicle, and the vehicle will be taken to the launch location. Following ascent, descent, separation, landing, and direction from the RSO a second wireless signal will be sent to the rover via an XBee Pro S3B which has an unobstructed range of greater than a mile. The Arduino Uno microcontroller will take this message and instruct the ARS to release the rover from the vehicle. There will then be a 5 second pause to allow complete release; then the two drive DC motors will propel the rover forward. A 10-DOF accelerometer will monitor the movement of the rover. Should the rover encounter an object that AUBURN UNIVERSITY STUDENT LAUNCH 74

75 stops forward progression, the accelerometer will sense this and tell the rover to stop, back up, turn, and continue once again. The accelerometer will also tell when the rover reaches at least 10 ft. When it does reach 10 ft, the soil recovery routine will begin. A servo will extend the recovery arm until it reaches the ground, when the belt drive DC motor will begin turning. The belt will turn for a predetermined amount of time to collect 10 ml of soil. Then, the belt drive DC motor will stop, and the sample door DC motor will close the sample door. The rover will then enter standby mode and wait for a signal from the XBee to open the sample door. This is graphically displayed in Figure 29: Rover Conops Completing NASA Requirements Figure 29: Rover Conops The payload was designed with the sole purpose of completing the requirements set forth by NASA. We chose to complete the rover/ soil collection requirements, because we could accomplish those requirements the best with the experience and equipment available at Auburn University. Since the most important aspect of the requirements is the soil collection, the rover was designed around our chosen collection method. A conveyor belt style collection unit will collect 10 ml of soil as defined by NASA requirement The next requirement we considered during the design process was NASA requirement 4.3.4: After deployment, the rover will AUBURN UNIVERSITY STUDENT LAUNCH 75

76 autonomously move at least 10 ft. (in any direction) from the launch vehicle. Once the rover has reached its final destination, it will recover a soil sample. To facilitate this, a skid steer, treaded design was chosen for locomotion, and a 10-DOF IMU will be used to determine the rover s spatial position. Next, we completed NASA requirement 4.3.2, The rover will be retained within the vehicle utilizing a fail-safe active retention system, by using a servo driven lock that will be active during ascent and decent. The rover will not activate upon landing until it receives a message via an XBee Pro S3B operated by the micro-controller on-board; this satisfies NASA requirement requiring a Remote Deployment Officer so be present when activating the rover. A door will slide closed after the conveyor belt style collection unit stops turning; sealing the sample and satisfying NASA requirement The batteries will be wrapped in brightly colored red tape and in an easily accessible compartment so satisfy NASA requirements and All of these will be completed while keeping the rover small enough to be deployed from the inside of the launch vehicle as defined by NASA requirement Completing Team Set Requirements During the design process, the team developed Team Derived requirements that had to be satisfied alongside the NASA requirements. The payload team derived 3 requirements for the rover: The rover and orientation system must fit inside the 6 in diameter launch vehicle , The rover must exit the launch vehicle in the correct orientation before completing the experiment , and The rover must be able to have enough battery power to be able to operate after sitting on the launch pad for up to 2 hours Requirements and were taken into consideration at the same time. Since the rover had to exit the launch vehicle in the proper orientation, a Rover Orientation System (ROS) was required. The ROS consists of 3 parts: a sled that the rover would sit on, a bearing at the aft of the sled to fasten the ROS to a launch vehicle bulk head, and a bearing at the most forward section of the ROS to support it. It was determined that a typical ball bearing would not be suitable for this support, so we designed our own. As seen in Figure 35: Side of Rover Orientation System, this bearing consists of a 3D printed ring with 3 press fit type transfer units integrated into it. The inner diameter where the transfer units sit are not less than that of a ball bearing, but where the transfer units are not provides up to an in bigger diameter than that of a ball bearing. This extra space allows for the rover to be wider and to satisfy Team Derived requirements and To satisfy requirement 6.1.1, the team is using lithium AUBURN UNIVERSITY STUDENT LAUNCH 76

77 polymer batteries that provide more than 5 times as many milliamp hours as the 9 Volt batteries used on last year s rover. These batteries will be tested. If they do not meet the requirements, we can purchase batteries with a greater mah rating of the same size, and they will operate the same. Electrical Design There are 11 electrical components and 2 batteries that make up the rover: Arduino Uno (base layer) Microcontroller that operates the Rover 10 DOF Pololu AltIMU v5 (Gyro, Accelerometer, Compass, and Altimeter) Accelerometer that allows the rover to know its special orientation XBee shield Interface between microcontroller and communication device XBee Pro S3B Device that facilitates communication between remote and rover Adafruit Motor Shield Interface between microcontroller and motors/ servos DC Motor: Left drive Motor used to propel rover DC Motor: Right drive Motor used to propel rover DC Motor: Belt drive Motor used to turn conveyor belt on soil recovery system DC Motor: Sample door Motor used to open and close soil containment compartment Servo: Belt arm 270 servo used to extend and retract soil collection arm Servo: Active Retention System 180 servo used to retain rover in launch vehicle during ascent and descent These components will be wired as shown in Figure 30: Rover Wiring Diagram. An Arduino Uno will be used as the microcontroller because of its low power requirement and lightweight design, which we can obtain at a reasonable price. A 10 DOF Pololu AltIMU v5 breakout will provide spatial awareness so the rover can know when it hass moved 10 ft from the launch vehicle. It will be wired to the microcontroller. Also connected to the microcontroller will be the XBee Shield which will act as the interface for the XBee Pro S3B. The XBee Pro S3B will provide communication for the rover. It was chosen, because a device that uses radio signals was required for the long range and possibility of poor weather on the day of the launch. The final shield on the top of the microcontroller stack will be an Adafruit motor shield. It will control the 4 Pololu Micro Metal HPCB motors that were chosen due to the large torque delivered in a micro volume at a AUBURN UNIVERSITY STUDENT LAUNCH 77

78 reasonable price. It will also control the servos used to direct the ARS and belt style collection unit. Figure 30: Rover Wiring Diagram Structural Design The rover will be made entirely using additive manufacturing, and CAD models for all parts can be found in Figure 31 to Figure 35 at the end of this section. The chassis of the rover will be one piece and 3D printed out of Onyx. This will provide structural integrity to hold up against the forces experienced during ascent and landing. The rollers and sprockets for the tracks will also be 3D printed out of Onyx, but the tracks and conveyor belt will be printed out of vinyl to allow them to bend without breaking. The tracks rollers and sprockets will be connected to the chassis via AUBURN UNIVERSITY STUDENT LAUNCH 78

79 metal bolts and nuts and tightened so that the rollers can still freely rotate. The sprocket will be attached to the DC motor to turn the tracks. The arm for the soil collection unit will also be 3D printed in house. The structural component of the arm will be printed out of Onyx and the belt will be printed out of vinyl. A sprocket will turn the belt but there will be no rollers. Due to the limited time the belt will spin, the Vinyl is expected to hold up against wear. A DC motor will turn the sprocket connected to the belt. A servo mounted to the rover will connect the arm to the rover and allow for it to be deployed. To safely contain the soil after collection, a separate DC motor will roll a cover into place to enclose the soil onboard. The Active Retention System will consist of a servo, a servo housing, a locking arm, and a U-bolt. A servo will provide the force required to hold the rover in place, and the servo housing will hold the servo to the rover. The arm will be 3D printed and in the shape of an L that will hook into the U-bolt mounted to the Rover Orientation System. The rover needs to exit the launch vehicle in the proper orientation. This requires a system separate from the rover. To complete this, we have developed the Rover Orientation System. The ROS consists of 3 parts: a sled that the rover would sit on, a wall and bearing at the aft of the sled to fasten the ROS to a launch vehicle bulk head, and a bearing at the most forward section of the ROS to support it. It was determined that a typical ball bearing would not be suitable for this support, so we designed our own. As seen in Figure 35: Side of Rover Orientation System this bearing consists of a 3D printed ring with 3 press fit type transfer units integrated into it. The inner diameter where the transfer units sit are not less than that of a ball bearing, but where the transfer units are not provides up to an in bigger diameter than that of a ball bearing. The U-bolt mentioned in the previous section will be mounted on the aft wall which rotates with the sled. When the launch vehicle is preparing to launch, the rover will be placed in the ROS which is already mounted inside the launch vehicle. AUBURN UNIVERSITY STUDENT LAUNCH 79

80 Figure 31: Rover with Collection Unit Extended Figure 32: Rover with Collection Unit Extended AUBURN UNIVERSITY STUDENT LAUNCH 80

81 Figure 33: Rover with Collection Unit Retracted Figure 34: Top of Rover Orientation System AUBURN UNIVERSITY STUDENT LAUNCH 81

82 Figure 35: Side of Rover Orientation System (Note that although the sample collection arm is extended in this figure, it will be retracted in flight) AUBURN UNIVERSITY STUDENT LAUNCH 82

83 5.2 Altitude Control System Altitude Control Overview The team will be employing an altitude control mechanism as a secondary payload. Its main objective is to ensure the vehicle hits the target altitude as accurately as possible, as well as provide flexibility with the mass of the vehicle. The system will be armed on the launchpad by turning it on via key switch. It will wait until it has sensed that the vehicle has launched and the motor has burned out before it will deploy its drag mechanism. It will continue controlling the apogee of the vehicle until it reaches apogee, where it will retract all drag components and enter standby mode, having completed its mission. The altitude control system is inactive during descent and will have no protruding surfaces upon landing, and also has no transmitting electronics to interfere with other aspects of the design. Team Derived Requirements The team has devised its own set of requirements for altitude control. It currently has four requirements, Requirement states the drag plates must be able to withstand 80 pounds of distributed normal force. CFD analysis showed that each plate would be subjected to a maximum of pounds during flight. The plates were designed with a safety factor of 2 to ensure the plates do not fail in flight. Requirement was derived to exceed Vehicle Requirement 2.11 to ensure the system can operate after an extended pad stay. Requirement requires the system to be completely embedded in the vehicle, which gives the vehicle a clean profile, and allows much of the system to be assembled separately from the vehicle, reducing the complexity of the vehicle s assembly as a whole. Lastly, requires the system to be fully autonomous. This requirement reduces the risk of human error during flight (e.g. forgetting to send the command to begin deploying the plates in flight, or sending the command too early, etc.). In addition, it reduces the complexity of the electrical components, since it does not require an extra component to receive commands from an external source. Design Overview This year s design is similar to last year s, in that the system will use plates embedded within the vehicle to provide drag. Grid fins were also considered since the team has previous experience with manufacturing and using them, and skin flaps were considered as well. AUBURN UNIVERSITY STUDENT LAUNCH 83

84 Properties 2 Plates 3 Plates Grid Fins Skin Flaps Simplicity Strength Max drag created Drag when retracted Cost Weight Total Score Table 32: Drag Plate Design Trade Study Two plates provide the best option of the four. Three plates were chosen since they provided the most leverage for the center piece to move them as well as provide sufficient drag. Figure 36: Last year's three plate design However, the size of the plates were still not large enough to provide a large amount of drag, and consequently there was a small window where the system can control the apogee of the vehicle. Instead, two plates will be attached to mounts that are moved along aluminum rods by a rack and pinion mechanism. This top mounted design allows the plates to be retracted fully against one another, maximizing the surface area of the plates. AUBURN UNIVERSITY STUDENT LAUNCH 84

85 Figure 37: Current Drag Plate System Design All components will be attached to a central structure, including the aluminum rods. This structure will be bolted to the coupler with four 0.25 in bolts, through the holes at the end of each arm. The full assembly is shown in Figure 39. Aluminum rods were chosen as they provide plenty of strength to resist the bending moments caused by the plates, as well as providing a smooth surface for the mounts to slide along. The assembly s design is mostly unchanged from PDR, however there are a few modifications. The electronics will be mounted to brackets which are attached to the central main structure. This allows the electronics to be soldered to breadboards, of which the team has an ample supply, as well as take advantage of some of the empty space above the assembly. Custom brackets will also be printed to hold the 9V batteries, since existing battery holders did not conform well to the layout of the assembly. Every custom component will be printed with Onyx. In addition to the electronics, JST connectors will be soldered to each breadboard, which simplifies the connections between each breadboard and emphasizes the modular design of the system. Figure 38: 4-pin male and female JST connectors AUBURN UNIVERSITY STUDENT LAUNCH 85

86 Figure 39: Full Assembly Figure 40: Top View, Full Assembly AUBURN UNIVERSITY STUDENT LAUNCH 86

87 Figure 41: Main support structure of the assembly Figure 42: Breadboard bracket AUBURN UNIVERSITY STUDENT LAUNCH 87

88 Figure 43: Batteries bracket to assembly Figure 44: Plate mounts, contains a notch to make sure rack gear is properly aligned AUBURN UNIVERSITY STUDENT LAUNCH 88

89 Figure 45: Drag plate Figure 46: Rack gear, driven by spur gear which moves the plates AUBURN UNIVERSITY STUDENT LAUNCH 89

90 Figure 47: Pinion Control Logic and Electrical Design A PID controller will be used to control the altitude of the vehicle, using the projected altitude of the vehicle as its point of reference. The projected altitude equation (which will be derived later in this section) only requires the altitude, velocity, and acceleration of the vehicle. These values can be calculated using an altimeter and an accelerometer. The accelerometer of choice is the MPU- 9250, a 9-DOF IMU which can read both the acceleration and orientation of the vehicle, which is necessary to compute the acceleration along the gravity vector. The altimeter is a MPL3115A2, which is an altitude, pressure, and temperature sensor in one package. It can provide the altitude within 1 ft of the actual altitude. In addition, data will be written to an SD card. The microcontroller of choice is a STM32f103C8T6, known colloquially as the Blue Pill. The Blue Pill contains a 32-bit processor clocked at 72 MHz, with 128 KB of memory and 20 KB of SRAM. Previously the team used an Arduino Uno due to familiarity, however last year there were issues with the lack of memory on the Uno as the software grew more complex. The Blue Pill offers an increase in memory in a smaller package, as well as increased computational power. AUBURN UNIVERSITY STUDENT LAUNCH 90

91 The motor used to drive the plates will be a 280 RPM Planetary Gear Motor with an encoder from ServoCity. The encoder allows the microcontroller to track the position of the plates and allows it to fine tune the position of them according to the PID controller. It is similar to the motor that was used last year, but it is smaller and lighter. A comprehensive list of components is provided below. STM32f103C8T6 Microcontroller MPU Axis IMU for acceleration and pitch MPL3115A2 Precision Altimeter Adafruit MicroSD Breakout Board allows microcontroller to write to SD card 280 RPM Planetary Gear Motor with Encoder (ServoCity) 2 9V Batteries The basic wiring diagram can be seen in Figure 48. Figure 48: Basic Wiring Diagram As stated previously, the equation for the projected altitude can be found using the acceleration, velocity, and altitude of the vehicle. Starting with the following equation, where, yy mmmmmm = vv tt 2 2gg llll(vv 0 2 +vv 2 tt vv 2 ), tt AUBURN UNIVERSITY STUDENT LAUNCH 91

92 vv tt 2 = 2mmmm CC dd AAAA However, CC dd is not known accurately at all times during flight. Instead, using the following equation, and solving for CC dd results in, aa = gg CC ddρρvv 2 2mm CC dd = 2mm (aa + gg) ρρρρvv2 Plugging this back into the equation for vv tt and simplifying gives us the following equation, vv tt 2 = vv2 gg (aa + gg) This eliminates the need for CC dd, as well as ρρ. Now there is an equation for vv tt as a function of acceleration and velocity. Plugging this new equation for vv tt into the equation for yy mmmmmm and simplifying results in the following equation, vv 2 yy mmmmmm = 2(aa + gg) ln aa gg + yy 0 From here the acceleration and velocity can then be calculated using an accelerometer and altimeter. Using this equation, the PID controller was designed using Simulink, and the gain values were computed accordingly. The model is outlined below. AUBURN UNIVERSITY STUDENT LAUNCH 92

93 Figure 49: Simulink model, including the PID controller and plant. The EOM block and Projected Altitude block can be found in Figure 50 and Figure 51 respectively The Equation of Motion block and Projected Altitude block are expanded below. Figure 50: Equation of Motion block AUBURN UNIVERSITY STUDENT LAUNCH 93

94 Figure 51: Projected Altitude block Concept of Operations The system is completely autonomous and does not receive any external commands during operation. The system is turned on via key switch on the launch pad, at which point it initializes and enters standby mode. Internally, there are four modes of operation: standby, launch, coast, and shutdown. On startup, the microcontroller will read the current altitude and set the target altitude accordingly. Afterwards, it will continue to read the altitude until it senses an increase in altitude of 250 ft. It will then switch to launch mode, where it will begin logging data and reading the accelerometer data. Once it senses that the motor has burned out (acceleration goes from positive to negative), it will switch to coast mode. The current mode must be launch or else it cannot be switched to coast mode, as an extra point of safety. During coast, the system will begin slowing the ascent of the vehicle by extending the plates, which is controlled by the PID controller. After it has reached apogee and the system senses that the vehicle s vertical velocity is negative, it will enter shutdown mode, where it will retract the drag plates and close the data file. After all the necessary actions are completed, the microcontroller will enter sleep mode. AUBURN UNIVERSITY STUDENT LAUNCH 94

95 Figure 52: Altitude Control Concept of Operations AUBURN UNIVERSITY STUDENT LAUNCH 95

96 6 Project Plan 6.1 Testing Vehicle Body Testing Tensile Testing of Composite and 3D Printed Material Test Objective This test aims to verify that the characteristics of all materials used to construct any portion of the rocket are consistent with expected values and can withstand expected launch stresses and loads. This test will fulfill Auburn requirement Success Criteria This test will be utilized to gather material data on the various materials being used to construct the launch vehicle in order to determine calculate the necessary component dimensions required for successful launch. As such, these experiments will have no success criteria. Test components - Onyx 3D-printed Carbon Fiber - Epoxy Carbon Fiber - Epoxy Fiberglass - Instron Multipurpose Testing Machine Procedure 1. The dimensions of numerous samples of Onyx 3D-printed Carbon Fiber constructed up to ASTM standards will be measured and cataloged. This data will be input into the Instron machine. 2. The material will be placed between the grips of the apparatus and secured into place by means of a hydraulic system. 3. The load and strain will be zeroed out on the computer interface, after which the tensile test will be initiated. AUBURN UNIVERSITY STUDENT LAUNCH 96

97 4. The data will be analyzed and plotted until the max stress and load of each sample are reached and the structural integrity compromised, at which point the machine will automatically end the test. 5. Once this process is repeated for all samples of the Onyx material, the previous steps will be repeated for Epoxy Carbon Fiber and Epoxy Fiberglass. Safety Notices - There is a risk of flying debris due to the breaking of the material samples. This will be counteracted via use of protective eye ware Compression Testing of Composite and 3D Printed Material Test Objective This test aims to verify that the characteristics of all materials used to construct any portion of the rocket are consistent with expected values and can withstand expected launch stresses and loads. This test will fulfill Auburn requirement Success Criteria This test will be utilized to gather material data on the various materials being used to construct the launch vehicle in order to determine calculate the necessary component dimensions required for successful launch. As such, these experiments will have no success criteria. Test Components - Onyx 3D-printed Carbon Fiber - Epoxy Carbon Fiber - Epoxy Fiberglass - Instron Multipurpose Testing Machine Procedure AUBURN UNIVERSITY STUDENT LAUNCH 97

98 1. The dimensions of numerous samples of Onyx 3D-printed Carbon Fiber constructed up to ASTM standards will be measured and cataloged. This data will be input into the Instron machine. 2. The material sample will be placed within the apparatus and lined up by eye to center it between two solid contact points. The machine will then be more accurately positioned using various fine adjustment dials to ensure the compression plates are slightly in contact with the sample. 3. The load and strain will be zeroed out on the computer interface, after which the compression test will be initiated. 4. The data will be analyzed and plotted until each specimen reaches -30% elongation at which point the machine will automatically end the test. 5. Once this process is repeated for all samples of the Onyx material, the previous steps will be repeated for Epoxy Carbon Fiber and Epoxy Fiberglass. Safety Notices - There is a risk of flying debris due to the breaking of the material samples. This will be counteracted via use of protective eyeware Flexure Testing of Composite and 3D Printed Material Test Objective This test aims to verify that the characteristics of all materials used to construct any portion of the rocket are consistent with expected values and can withstand expected launch stresses and loads. This test will fulfill Auburn requirement Success Criteria This test will be utilized to gather material data on the various materials being used to construct the launch vehicle in order to determine calculate the necessary component dimensions required for successful launch. As such, these experiments will have no success criteria. Test Components AUBURN UNIVERSITY STUDENT LAUNCH 98

99 - Onyx 3D-printed Carbon Fiber - Epoxy Carbon Fiber - Epoxy Fiberglass - Instron Multipurpose Testing Machine Procedure 1. The dimensions of numerous samples of Onyx 3D-printed Carbon Fiber constructed up to ASTM standards will be measured and cataloged. This data will be input into the Instron machine. 2. The material sample will be placed within the apparatus and lined up by eye to center it between two solid contact points. The machine will then be more accurately positioned using various fine adjustment dials to ensure the contact arm is slightly in contact with the sample. 3. The load and strain will be zeroed out on the computer interface, after which the Flexure test will be initiated. 4. The data will be analyzed and plotted until the max flexure load of each sample is reached, at which point the machine will automatically end the test. 5. Once this process is repeated for all samples of the Onyx material, the previous steps will be repeated for Epoxy Carbon Fiber and Epoxy Fiberglass. Safety Notices - There is a risk of flying debris due to the breaking of the material samples. This will be counteracted via use of protective eyeware. Recovery Testing Separation Testing Test Objective This test aims to ensure the recovery systems of the launch vehicle can properly deploy. Testing criteria include ensuring the proper amount of black powder is used and ensuring AUBURN UNIVERSITY STUDENT LAUNCH 99

100 the body section provides adequate compression. This test will fulfill NASA requirement 3.2. Success Criteria This test will be considered successful if the recovery sections can both fully separate without any damage to the parachute systems. Test components Recovery Barometric Avionics Enclosure (BAE) structure - Launch configuration upper or lower recovery section, secured with shear pins, including - Main parachute (with drogue in upper section) - Shock cord - Recovery wadding - Assembled black powder charges - Nose cone or first lower body section coupler (depending on upper or lower section) - Launch configuration upper or lower recovery section, secured with shear pins, including - Electronic matches - Ignition system - Fire extinguisher (have not needed to use it, but always will be located beforehand) Procedure 1. Fill charge cups according to equation established in the Recovery Section, or to increased amount based on the result of the previous test/launch (the equation from the recovery section is an ideal case and has been found from previous years to undershoot the required amount). 2. Assemble structural components of the BAE with the electronic matches threaded through to reach the charge cups. 3. Attach recovery section tube to the BAE, packing charges, wadding, parachutes, etc. as they would be for a launch. AUBURN UNIVERSITY STUDENT LAUNCH 100

101 4. Seal the recovery section with the coupler or nose cone that completes the enclosure, securing with shear pins. 5. Place the assembled system on the ground or a test stand away from all personnel. 6. Attach the electronic matches to the power supply, and after verifying everyone is at a safe distance, fire the charges. 7. If the components separate, record the amount of black powder used. If the charges do not separate the components, disassemble the rocket and increase the charge. Safety Notices - As this test involves the use of combustibles, there is a risk of fire. As such, there will always be a fire extinguisher on hand during testing. - This test involves launching segments of the rocket multiple ft, which could result in the testers being struck. This can be counteracted by establishing a 5 ft clear zone around the testing area, and by ensuring no test personnel are standing directly in the path of the rocket segments. Results Separation testing for the subscale launch vehicle was a complete success. Both segments of the rocket separated cleanly. The first iteration of full-scale separation testing was a failure. The top segment of the recovery segment separated cleanly, but the bottom section failed to separate over the course of several tests. We believe this is because the recovery section was not fully wired for this test, thus pressure was able to escape through holes normally plugged by wires. This oversight will be rectified in the planned re-test. Figure 53: Ejection Testing AUBURN UNIVERSITY STUDENT LAUNCH 101

102 Shear Pin Testing Test Objective The shear pin test will ensure the shear pins used by the recovery systems are strong enough to withstand the initial impulse applied by the deployment of the drogue parachute while still allowing for main separation. This test will fulfill NASA requirement 2.11 and Auburn requirement Success Criteria This test will be considered successful if the experimental data collected matches the manufacturer strength specifications. Figure 54: Shear Pin Testing Rig Test components - Instron Multipurpose Testing Machine Nylon Shear Pins - ABS Shear pin testing rig -.5 in diameter rod Procedure 1. The test rig and rod will be affixed in opposing positions in the Instron machine. 2. Once they are in place, a shear pin will be inserted into the appropriately sized hole in the side of the test rig. 3. The Instron will be set to perform a compression test, which will drive the rod through the opening in the test rig. Because the shear pin is in a perpendicular orientation to the movement of the rod, the test will output the strain that the pin undergoes. 4. Gather the collected data and compare to the expected forces calculated by the recovery team. AUBURN UNIVERSITY STUDENT LAUNCH 102

103 Safety Notices - There is a risk of flying debris due to the breaking of the material samples. This will be counteracted via use of protective eyewear. Results Testing proved to be a complete success. The shear pins failed at the manufacturer rating of 45 lbf, as can be seen below, thus confirming our required force calculated for separation. Rover Testing Rover Battery Testing Test Objective The rover and its supporting electronic equipment must remain operable after an extended wait period on the launch pad. This test will fulfill NASA requirement 2.11 and Auburn requirement Success Criteria This test will be considered a success if both battery systems are able to supply power to the rover systems after four hours of idle power usage. Test Components AUBURN UNIVERSITY STUDENT LAUNCH 103

104 - Rover servos and motors - Rover batteries (both large and small) - Low voltage alarm Procedure 1. The batteries will be securely connected to their respective systems. 2. The systems will be remotely activated and left to run for 4 hours. This means activating all motors in the system and allowing them to spin at full power for 4 consecutive hours after reception of the remote system. 3. During these four hours, the motors will be checked for differences in generated pitch, as this may be an indication of a change in the amount of power being supplied to the motors. This should be performed every 20 minutes. 4. The test will be considered successful if all batteries operate continuously without any low voltage alarms. 4 hours will be double the amount of time any one system should have to operate at launch. (NOTE: A low voltage alarm must be connected to both batteries. If a low voltage alarm sounds, testing should stop immediately, and batteries should be removed from the systems. Please put batteries into a battery case when done. An observer must be present in the testing area for the duration of the test. Only one observer may note the changes in motor pitch.) Safety Notices - Batteries are prone to ignite when improperly charged, this is counteracted by having two people present during charging to ensure proper charging procedure is followed. Additionally, all charging batteries will be stored in a fire-retardant bag. - There is a risk of electrical shock when doing testing with batteries. This will be mitigated with shock resistant gloves and peer observation. Altitude Control System Testing CFD Simulation Test objective AUBURN UNIVERSITY STUDENT LAUNCH 104

105 This test will provide simulated data on the drag characteristic of the drag plates at various states of extension. This data will be used to ensure the PID controller is properly calibrated. This test will fulfill and Auburn requirement Success Criteria The data collected by this experiment will be used to calibrate the altitude control system. As such, there is no success criteria. Test components - Ansys Fluent CFD program - CAD model of launch vehicle with drag plates at various extension Procedure 1. A CAD model of the launch vehicle with retracted drag plates will be imported into the Ansys Fluent CDF software. 2. Once the model of the vehicle is imported into Fluent s 3D environment, a mesh of the test control volume will be generated to simulate a compressible atmosphere. 3. Once the mesh is completed, the flow will be simulated to a Mach number of 0.65 and drag data will be gathered from the flows interaction with the vehicle model. 4. Once all data has been collected, the process will be completed with a CAD model with extended drag plates. 5. The data from the two simulations will be used to calibrate the altitude control system. Safety Notices - N/A Results The CFD simulation showed the drag plates, when fully extended, provide a 90% increase in total drag on the launch vehicle. This data has been used to calibrate the altitude control system, which now awaits a full-scale test flight to confirm the data gathered from this test. AUBURN UNIVERSITY STUDENT LAUNCH 105

106 Figure 55: CFD Rocket Simulation Results Altitude Control System Battery Testing Test Objective It is necessary to ensure the Altitude control system can remain powered and functional after extended idle time on the pad. To ensure this, the system will be evaluated over the course of a 4 hr test to determine the drain rate and longevity of the system s battery. This test will fulfill NASA requirement 2.11 and Auburn requirement Success Criteria This test will be considered a success if both battery systems are able to supply power to the rover systems after four hours of idle power usage. Test Components - Altitude control system - 9-volt battery - Volt meter Procedure AUBURN UNIVERSITY STUDENT LAUNCH 106

107 1. The battery will be connected to the altitude control system, and the system will be turned on. 2. The volt meter will be used to measure the voltage drop across the battery at 20 min intervals for 4 hr (should the voltage drop to zero, the test will be terminated immediately). 3. Once the test is completed, the voltage results will be examined to determine whether the power supply is sufficient. Safety Notices - There is a risk of electrical shock when doing testing with batteries. This will be mitigated with shock resistant gloves and peer observation Altitude Control System Virtual Launch Test Objective This test will determine whether the PID controller in the control system can properly interpret flight data. Additionally, the test will demonstrate whether the systems hardware will function properly. This test will fulfill Auburn requirement 6.4.4, 6.4.5, and Success Criteria This test will be considered a success if the final altitude is within ten ft of the target altitude. Test Components Altitude Control System Procedure 1. A simulated set of flight input data will be written. 2. The simulated data will be fed into the PID controller in the altitude control system, which will cause the systems drag plates to actuate. AUBURN UNIVERSITY STUDENT LAUNCH 107

108 3. Once the simulation is completed the output data will be examined to determine whether the controller properly interpreted the data it was given. Additionally, the movements of the drag plates will be reviewed to ensure they functioned properly. 4. There is a risk of electrical shock when doing testing with batteries. This will be mitigated with shock resistant gloves and peer observation. Results The Simulink model was run with the PID controller deactivated to verify the accuracy of the simulation. The results were within 10 ft of the OpenRocket results, providing a sufficiently accurate simulation to tune the PID controller. The controller was able to slow the vehicle down to hit an altitude of 4700 ft, exactly at the target altitude. The results are outlined in the figure below. Safety Notices - N/A Figure 56: Simulink results AUBURN UNIVERSITY STUDENT LAUNCH 108

109 6.2 Requirements Verification Tables General Requirements Requirement Number Description Verification Method Students on the team will do 100% of the project, including design, construction, written reports, Demonstration: During the course of the presentations, and flight preparation competition, the team s mentor only 1.1 with the exception of assembling the assembles the motor and handles the motors and handling black powder or ejection charges and electronic matches, any variant of ejection charges, or with all other tasks performed by students preparing and installing electric on the team. matches (to be done by the team s mentor). The team will provide and maintain a project plan to include, but not limited Inspection: The team includes a detailed to the following items: project project timeline, educational engagement 1.2 milestones, budget and community plan, and safety risk and mitigations with support, checklists, personnel proposal, and release additional assignments, STEM engagement documentation as early as possible. events, and risks and mitigations. Foreign National (FN) team members 1.3 must be identified by the Preliminary Demonstration: Foreign National team Design Review (PDR) and may or may members have been be identified by PDR not have access to certain activities and understand that they may be during launch week due to security separated from the team during certain restrictions. In addition, FN s may be launch week activities. separated from their team during certain activities. AUBURN UNIVERSITY STUDENT LAUNCH 109

110 The team must identify all team members attending launch week activities by the Critical Design Review (CDR). Team members will include: Students actively engaged in the project throughout the entire year One mentor (see requirement 1.13) No more than two adult educators. The team will engage a minimum of 200 participants in educational, handson science, technology, engineering, and mathematics (STEM) activities, as defined in the STEM Engagement Activity Report, by FRR. To satisfy this requirement, all events must occur between project acceptance and the FRR due date and the STEM Engagement Activity Report must be submitted via within two weeks of the completion of the event. A sample of the STEM Engagement Activity Report can be found on page 33 of the handbook. The team will establish a social media presence to inform the public about team activities. Teams will all deliverables to the NASA project management team by the deadline specified in the handbook for each milestone. In the event that a deliverable is too large to attach to an , inclusion of a link to download the file will be sufficient. Demonstration: The team has identified all team members attending launch week activities by this CDR. Members attending launch week will be students, one mentor, and two adult educators Demonstration: The team continues to engage far more than 200 participants in educational STEM activities prior to FRR. All events will have their Educational Activity Report submitted via within two weeks of the event. Demonstration/ Inspection: The team has an Instagram and Facebook account and updates both regularly. Demonstration: The team s all deliverables to the NASA project team prior to the specified deadlines, utilizing a link to download if the deliverable is too large. AUBURN UNIVERSITY STUDENT LAUNCH 110

111 All deliverables must be in PDF format. In every report, teams will provide a table of contents including major sections and their respective subsections. In every report, the team will include the page number at the bottom of the page. The team will provide any computer equipment necessary to perform a video teleconference with the review panel. This includes, but is not limited to, a computer system, video camera, speaker telephone, and a sufficient Internet connection. Cellular phones should be used for speakerphone capability only as a last resort. All teams will be required to use the launch pads provided by Student Launch s launch services provider. No custom pads will be permitted on the launch field. Eight foot 1010 rails and 12 foot 1515 rails will be provided. The launch rails will be canted 5 to 10 degrees away from the crowd on launch day. The exact cant will depend on launch day wind conditions Each team must identify a mentor. A mentor is defined as an adult who is included as a team member, who will be supporting the team (or multiple teams) throughout the project year, and may or may not be affiliated with the school, institution, or organization. The mentor must maintain a current Demonstration: All deliverables are in PDF format. Demonstration: All reports include a table of contents. Demonstration: All reports include a page number at the bottom of the page. Demonstration: The team utilizes a conference room within the Aerospace Engineering building that provides a video camera, speaker telephone and internet connection, while a team member provides the computer. Test: The rocket is designed to fit on 1010 launch rails. During construction it will be verified to fit on the planned type of launch rails using a test segment of the same type of launch rail. Demonstration: The team has identified its mentor as Dr. Eldon Triggs. Dr. Triggs has a current level 2 certification from the Tripoli Rocketry Association and is in good standing. AUBURN UNIVERSITY STUDENT LAUNCH 111

112 certification, and be in good standing, through the National Association of Rocketry (NAR) or Tripoli Rocketry Association (TRA) for the motor impulse of the launch vehicle and must have flown and successfully recovered (using electronic, staged recovery) a minimum of 2 flights in this or a higher impulse class, prior to PDR. The mentor is designated as the individual owner of the rocket for liability purposes and must travel with the team to launch week. One travel stipend will be provided per mentor regardless of the number of teams he or she supports. The stipend will only be provided if the team passes FRR and the team and mentor attend launch week in April. 2.1 Vehicle Requirements Requirement Number Requirement Statement The vehicle will deliver the payload to an apogee altitude between 4,000 and 5,500 ft above ground level (AGL). Teams flying below 3,500 ft or above 6,000 ft on Launch Day will be disqualified and receive zero altitude Verification Method and Execution of Method Analysis, Demonstration, Testing: The team will design the rocket to carry the payload to the chosen altitude. They will launch the flight vehicle and check the altimeters after flight. AUBURN UNIVERSITY STUDENT LAUNCH 112

113 points towards their overall project score. Teams shall identify their target altitude goal at the PDR milestone. The declared target altitude will be used to determine the team s altitude scoring during Launch Week. The vehicle shall carry one commercially available, barometric altimeter for recording the official altitude used in determining the altitude award winner. The Altitude Award will be given to the team with the smallest difference between their measured apogee and their official target altitude on launch day. Each altimeter will be armed by a dedicated mechanical arming switch that is accessible from the exterior of the rocket airframe when the rocket is in the launch configuration on the launch pad. Demonstration: The team has determined its altitude goal to be 4700 ft by PDR. Inspection, Demonstration: The team will purchase and calibrate one commercially available altimeter and fly it in the launch vehicle. Demonstration: The rocket has been designed with a arming switch incorporated into the rocket airframe in a coupler. This location is available to the team to activate the altimeters from the launch configuration. 2.5 Each altimeter will have a dedicated power supply. Inspection, Demonstration: The team has designed altimeters with separate power supplies Each arming switch will be capable of being locked in the ON position for launch (ie. Cannot be disarmed due to flight forces) The launch vehicle shall be designed to be recoverable and reusable. Reusable defined as being able to launch again on the same day without repairs or modifications. Inspection, Demonstration: The team has designed for and tested that the arming switch is capable of being locked in the ON position. Testing, Analysis, Demonstration, Inspection: Trajectory simulations and testing will ensure the launch vehicle is recoverable and reusable. The subscale design has shown that the AUBURN UNIVERSITY STUDENT LAUNCH 113

114 rocket can be recoverable and reusable The launch vehicle shall have a maximum of four (4) independent sections. An independent section is defined as a section that is either tethered to the main vehicle or is recovered separately from the main vehicle using its own parachute. Coupler/airframe shoulders which are located at in-flight separation points will be at least one (1) body diameter in length. Nose cone shoulders which are located at in-flight separation points will be at least ½ body diameter in length. The launch vehicle shall be limited to a single stage. The launch vehicle will be capable of being prepared for flight at the launch site within 4 hours, from the time the Federal Aviation Administration flight waiver opens. The launch vehicle will be capable of remaining in launch-ready configuration at the pad for a minimum of 2 hours without losing the functionality of any critical on-board component. Demonstration: Team has designed and will build a launch vehicle that can have, but does not require, four independent sections. Inspection, Analysis, Testing: The team has designed the rocket with couplers/shoulders at separation points to be at least one body diameter in length each. Demonstration: The team has designed the flight vehicle with a nose cone shoulder that is going to be separated mid-flight with at least ½ body diameter in length. Demonstration: Team has designed and build a single-stage launch vehicle. Demonstration: The team will be timely and organized to ensure vehicle is prepared on time. The team practices this timely process each time they launch at a field. Demonstration: The team has designed the vehicle with ability to remain launch-ready for at least two hours. The team will continue testing the batteries to find the maximum time the electronics can sit active. AUBURN UNIVERSITY STUDENT LAUNCH 114

115 The launch vehicle shall be capable of being launched by a standard 12 volt direct current firing system. The firing system will be provided by the NASAdesignated launch services provider. The launch vehicle shall require no external circuitry or special ground support equipment to initiate launch (other than what is provided by launch services provider). The launch vehicle will 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). Testing: The team has designed the rocket to be launched by a 12 volt direct current. The team will test the launch configuration s ability to launch with prior launch. Demonstration: The team has designed a vehicle requiring no external circuitry or special ground support equipment. Demonstration: The team has designed the vehicle to implement commercially available, certified motors Final motor choices must be made by the Critical Design Review (CDR) milestone. Demonstration: The team has finalized the design enough by CDR to determine which motor the team will use for competition Any motor changes after CDR must be approved by the NASA Range Safety Officer (RSO) and will only be approved if the change is for the sole purpose of increasing the safety margin. A penalty against the team s overall score will be incurred when a Demonstration: The team has tested their rocket and the safety of their design before CDR with the sub scale to prevent this from occurring. If it does, the team will seek approval from the NASA RSO and take on the penalty. AUBURN UNIVERSITY STUDENT LAUNCH 115

116 motor change is made after the CDR milestone, regardless of the reason. Pressure vessels on the vehicle shall be approved by the RSO and shall meet the following criteria: The minimum factor of safety (Burst or Ultimate pressure versus Max Expected Operating Pressure) shall be 4:1 with supporting design documentation included in all milestone reviews. Each pressure vessel will include a pressure relief valve that sees the full pressure of the tank and is capable of withstanding the maximum pressure and flow rate of the tank. Full pedigree of the tank shall be described, including the application for which the tank was designed, and the history of the tank, including the number of pressure cycles put on the tank, by whom, and when. Inspection: The team has designed the rocket this year not to utilize any pressure vessels. Inspection: The team has designed the rocket this year not to utilize any pressure vessels. Inspection: The team has designed the rocket this year not to utilize any pressure vessels. Inspection: The team has designed the rocket this year not to utilize any pressure vessels The total impulse provided by a College and/or University launch vehicle shall not exceed 5,120 Newton-seconds (L-class). Demonstration, Analysis: The team has chosen a motor with a total impulse that does not exceed 5,120 Newton-seconds (L-class) The launch vehicle shall have a minimum static stability margin of 2.0 at the point of rail exit. The launch vehicle shall accelerate to a minimum velocity of 52 fps at rail exit. Testing, Demonstration, Analysis: The team has designed and will test the vehicle to ensure that it has a stability margin of 2.0 at the point of rail exit Demonstration, Analysis, Testing: The team has designed and tested the vehicle to ensure that it s minimum velocity at rail exit is at least 52 fps. AUBURN UNIVERSITY STUDENT LAUNCH 116

117 All teams shall successfully launch and recover a subscale model of their rocket prior to CDR. The subscale model should resemble and perform as similarly as possible to the full-scale model, however, the fullscale shall not be used at the subscale model. The subscale model will carry an altimeter capable of reporting the model s apogee altitude. The subscale rocket must be a newly constructed rocket, designed and built specifically for this year s project. Proof of a successful flight shall be supplied in the CDR report. Altimeter data output may be used to meet this requirement. Demonstration, Testing: The team flew their subscale to ensure there was a successful launch prior to CDR. Demonstration: The subscale model was designed to resemble and perform similarly to the full-scale model. Demonstration: An altimeter capable of reporting the model s apogee altitude was implemented on the subscale model. Demonstration: The team constructed a subscale specifically for this year s vehicle. Testing, Demonstration: The team used altimeter data and photographic evidence as proof for a successful flight All teams will complete demonstration flights as outlined below All teams shall successfully launch and recover their full-scale rocket prior to FRR in its final flight configuration. The rocket flown at FRR must be the same rocket flown of launch day. The following criteria must be met during the fullscale demonstration flight: The vehicle and recovery system shall have functioned as designed Testing, Demonstration, Analysis: A test of the rocket will be exhibited, demonstrating all hardware functions properly. Testing: The team has defined the design and function of the vehicle and AUBURN UNIVERSITY STUDENT LAUNCH 117

118 The full-scale rocket must be a newly constructed rocket, designed and built specifically for this year s project. recovery system and has shown it to work for subscale and will show it working for full-scale. Demonstration: The team will build a vehicle specifically for this year s competition The payload does not have to be flown during the full-scale test flight. The following requirements still apply: If the payload is not flown, mass simulators shall be used to simulate the payload mass. The mass simulators shall be located in the same approximate location on the rocket as the missing payload mass. Testing, Demonstration, Analysis: The team will either prepare the payload to go in this flight or a payload mass for a correct simulation of the flight. Inspection: An inspection of the rocket payload will be done by the team to ensure it is properly placed in the expected location If the payload changes the external surfaces of the rocket (such as with camera housings or external probes) or manages the total energy of the vehicle, those systems shall be activated during the full-scale Vehicle demonstration flight Demonstration, Testing: The team only has switches to turn on electronics for the payload in the rocket and will always be with the rocket Teams shall fly the launch day motor for the Vehicle Demonstration Flight. The RSO may approve use of an alternative motor if the home launch field cannot support the full impulse of the launch day motor or in other extenuating circumstances. The vehicle shall be flown in its fully ballasted configuration during the fullscale test flight. Inspection: The team plans on using the same rocket for the final competition launch day that they can legally fly on the closest homefield. Testing, Demonstration: The team plans to test and prepare the amount of ballast that will be used on all flights leading up to competition. AUBURN UNIVERSITY STUDENT LAUNCH 118

119 After successfully completing the fullscale demonstration flight, the launch vehicle or any of its components will not be modified without the concurrence of the NASA Range Safety Officer (RSO). Proof of a successful flight shall be supplied in the FRR report. Vehicle Demonstration flights must be completed by the FRR submission deadline (March 4, 2019). Demonstration: The team will not alter the rocket after the full-scale demonstration launch Demonstration, Testing: The team will use altimeters to verify a successful flight Demonstration: The full-scale flight will be completed by March 4th, Payload Demonstration Flight - All teams will successfully launch and recover their full-scale rocket containing the completed payload prior to the Payload Demonstration Flight deadline (March 25, 2019). Demonstration: The payload demonstration flight will be completed by the FRR submission (March 4 th, 2019) The payload must be fully retained throughout the entirety of the flight, all retention mechanisms must function as designed, and the retention mechanism must not sustain damage requiring repair. The payload flown must be the final, active version. If the above criteria is met during the original Vehicle Demonstration Flight, occurring prior to the FRR deadline and the information is included in the Demonstration, Testing: The team will demonstrate the payload being fully secured during flight. Demonstration, Testing: The team will use the final active payload during flight. Demonstration: The additional flight will not be flown, if the above criteria is met. AUBURN UNIVERSITY STUDENT LAUNCH 119

120 FRR package, the additional flight and FRR Addendum are not required. Payload Demonstration Flights must be completed by the FRR Addendum deadline. An FRR Addendum will be required for any team completing a Payload Demonstration Flight or NASArequired Vehicle Demonstration Reflight after the submission of the FRR Report. Teams required to complete a Vehicle Demonstration Re-Flight and failing to submit the FRR Addendum by the deadline will not be permitted to fly the vehicle at launch week. Teams who successfully complete a Vehicle Demonstration Flight but fail to qualify the payload by satisfactorily completing the Payload Demonstration Flight requirement will not be permitted to fly the payload at launch week. Teams who complete a Payload Demonstration Flight which is not fully successful may petition the NASA RSO for permission to fly the payload at launch week. Demonstration: The payload demonstration flight will be completed by March 25th, Analysis: The FRR Addendum will be finished if a re-flight is required. Analysis: The team plans on having the demonstration flight completed and the FRR submitted on time. If not, the team will accept the punishment for not doing so. Analysis: The team will abide by the rules and not fly with the payload if that is the case. Analysis: The team will listen to the NASA RSO on permission for flying with the payload Any structural protuberance on the rocket will be located aft of the burnout center of gravity. The team s name and launch day contact information shall be in or on the rocket airframe as well as in or on Demonstration: Team will design all structural protuberances on the vehicle to be aft of the burnout center of gravity. Demonstration: The team will design the vehicle as to have contact information and the team s name on AUBURN UNIVERSITY STUDENT LAUNCH 120

121 any section of the vehicle that separates during flight and is not tethered to the main airframe. the airframe of each piece of the rockets. AUBURN UNIVERSITY STUDENT LAUNCH 121

122 Vehicle Prohibitions Requirement Number Requirement Statement The launch vehicles shall not utilize forward canards. The launch vehicle shall not utilize forward firing motors. The launch vehicle shall not utilize motors that expel titanium sponges (Sparky, Skidmark, MetalStorm, etc.) Verification Method and Execution of Method Demonstration - The team demonstrated how the launch vehicle does not utilize canards with subscale. Demonstration - The team has designed the vehicle so that it does not utilize forward firing motors. Demonstration- The team does not utilize a motor that expels titanium sponges The launch vehicle shall not utilize hybrid motors. Demonstration - The team will not utilize a hybrid motor The launch vehicles shall not utilize a cluster of motors. The launch vehicle shall not utilize friction fitting for motors. The launch vehicle shall not exceed Mach 1 at any point during flight. Vehicle Ballast shall not exceed 10% of the total weight of the rocket. Demonstration - A demonstration and inspection of the launch vehicle shall be carried out to validate it does not use a cluster of motors. Demonstration - The team has designed the vehicle so that it does not utilize friction fitting for the motor. Demonstration,Testing, Analysis - The team will test and demonstrate to ensure that the vehicle does not exceed Mach 1 at any point during the flight. Demonstration, Testing,Analysis- AUBURN UNIVERSITY STUDENT LAUNCH 122

123 Transmissions from onboard transmitters will not exceed 250 mw of power. Excessive and/or dense metal will not be utilized in the construction of the vehicle. Use of lightweight metal will be permitted but limited to the amount necessary to ensure structural integrity of the airframe under the expected operating stresses. Recovery Requirements The team has designed ballast so that it does not exceed 10% of the total weight of the rocket. Demonstration, Testing- The team has designed transmitters that will not exceed 250 mw of power. Demonstration- The team has designed the vehicle to ensure it does not use dense metals. Requirement Number Description Verification Method The launch vehicle will stage the deployment of its recovery devices, where a drogue parachute is deployed at apogee and a main parachute is deployed at a lower altitude. Tumble or streamer recovery from apogee to main parachute deployment is also permissible, provided that kinetic energy during drogue-stage descent is reasonable, as deemed by the RSO. The main parachute shall be deployed no lower than 500 ft. The apogee event may contain a delay of no more than 2 seconds. Each team must perform a successful ground ejection test for both the drogue and main parachutes. This The team has designed to stage the deployment of our recovery devices with a drogue parachute deployed at apogee, currently specified at 5111 ft. before altitude control. The team will stage the deployment of our recovery devices with the main parachute at 500 ft. A redundant altimeter will be in place with an apogee delay of 1 second. Prior to the initial subscale and full scale launches the team have performed a AUBURN UNIVERSITY STUDENT LAUNCH 123

124 must be done prior to the initial subscale and full-scale launches. At landing, each independent section of the launch vehicle will have a maximum kinetic energy of 75 ft-lbf. The recovery system electrical circuits will be completely independent of any payload electrical circuits. All recovery electronics will be powered by commercially available batteries. The recovery system will contain redundant, commercially available altimeters. The term altimeters includes both simple and more sophisticated flight computers. Motor ejection is not a permissible form of primary or secondary deployment. Removable shear pins will be used for both the main parachute compartment and the drogue parachute compartment. Recovery area will be limited to a 2,500 ft. radius from the launch pads. Descent time will be limited to 90 seconds (apogee to touch down). An electronic tracking device will be installed in the launch vehicle and will transmit the position of the tethered ground ejection test for both the drogue and main parachute. The team will calculate and test both sections of our launch vehicle to ensure that a maximum kinetic energy of 75 ftlbf at landing is not exceeded. The team have created independent circuits for our recovery system so that they are independent of any payload electrical circuits. All recovery electronics included in the recovery systems are powered by commercially available batteries. The recovery systems contain TeleMetrum and TeleMega altimeters for the BAE, and two Stratologger altimeters for the nose cone. The recovery system do not use motor ejection as a form of primary or secondary deployment. Removable shear pins are used for both the main parachute and drogue parachute compartments. The team will plan to ensure recovery area will be less than a 2,500 ft. radius from the launch pads. The team has designed to limit the descent time to under 90 seconds. The team has installed an electronic tracking device in the launch vehicle that will transmit the position of both AUBURN UNIVERSITY STUDENT LAUNCH 124

125 vehicle or any independent section to a ground receiver. Any rocket section or payload component which lands untethered to the launch vehicle will contain an active electronic tracking device. The electronic tracking device(s) will be fully functional during the official flight on launch day. The recovery system electronic will not be adversely affected by any other on-board electronic devices during flight (from launch until landing). The recovery system altimeters will be physically located in a separate compartment within the vehicle from any other radio frequency transmitting device and/or magnetic wave producing device. The recovery system electronics will be shielded from all onboard transmitting devices to avoid inadvertent excitation of the recovery system electronics. The recovery system electronics will be shielded from all onboard devices which may generate magnetic waves (such as generators, solenoid valves, and Tesla coils) to avoid inadvertent excitation of the recovery system. The recovery system electronics will be shielded from any other onboard devices which may adversely affect independent sections to a ground receiver. Both independent sections of the launch vehicle will contain and active electronic tracking device. The team will test the electronic tracking devices to ensure their functionality for the official flight on launch day. The team have tested the recovery system electronic to ensure it is not affected by any other on-board electronic devices. The recovery system altimeters are in separate compartments within the vehicle from any other radio frequency transmitting device and/or magnetic wave producing device. The team ensures that the recovery system electronics will be shielded from all transmitting devices to avoid any excitation of said electronics. The recovery electronics are sealed in their own separate compartment separate from all other magnetic wave including devices in the rocket. The recovery electronics are sealed in their own separate compartment from all other magnetic wave inducing devices in the rocket. AUBURN UNIVERSITY STUDENT LAUNCH 125

126 the proper operation of the recovery system electronics. AUBURN UNIVERSITY STUDENT LAUNCH 126

127 Rover Requirements Requirement Number Description Verification Method 4.2 Design: The team has chosen the Each team will choose one experiment Deployable Rover/Soil Sample option from the following list. Recovery option for the experiment An additional experiment (limit of 1) is allowed, and may be flown, but will not contribute to scoring. Design: The team is doing one additional experiment of an altitude control system with drag plates If the team chooses to fly an additional Demonstration: The team includes all experiment, they will provide appropriate documentation for the appropriate documentation in all additional experiment in all the design design reports, so the experiment may reports. be reviewed for flight safety Design: Tracked rover and orientation Teams will design a custom rover that system has been designed to be will deploy from the internal structure completely housed inside the body of of the launch vehicle. the rocket. The rover will be retained within the vehicle utilizing a fail-safe active Design: The rover will be using a retention system. The retention system servo driven lock that will be active will be robust enough to retain the during ascent and decent. rover if atypical flight forces are experienced At landing, and under the supervision of the Remote Deployment Officer, the team will remotely activate a trigger to deploy the rover from the rocket. Design: The rover will not active upon landing until it receives a message via an XBee Pro S3B operated by the micro-controller on-board Design: To facilitate this, a skid steer, After deployment, the rover will treaded design was chosen for autonomously move at least 10 ft. (in locomotion, and a 10-DOF IMU will any direction) from the launch vehicle. be used to determine the rover s Once the rover has reached its final spatial position. At the final destination, it will recover a soil destination soil recovery device will sample. activate to recover the sample. AUBURN UNIVERSITY STUDENT LAUNCH 127

128 The soil sample will be a minimum of 10 milliliters (ml). The soil sample will be contained in an onboard container or compartment. The container or compartment will be closed or sealed to protect the sample after collection. Teams will ensure the rover s batteries are sufficiently protected from impact with the ground. The batteries powering the rover will be brightly colored, clearly marked as a fire hazard, and easily distinguishable from other rover parts Design: A conveyor belt style collection unit will collect 10 ml of soil as defined by NASA requirement. Testing: This device will be tested to ensure it collects a minimum of this amount. Design: A door will slide closed after the conveyor belt style collection unit stops turning, sealing the sample. Testing: A slot for the batteries will be a priority when designing the rover so they will be integrated seamlessly in a safe, padded location. Design: All hazardous elements of the rover will be marked with a bright red color and be visible from the outside without moving parts to access them. AUBURN UNIVERSITY STUDENT LAUNCH 128

129 Safety Requirements Requirement Number Description Each team will use a launch and safety checklist. The final checklists will be included in the FRR report and used during the Launch Readiness Review (LRR) and any launch day operations. Each team must identify a student safety officer who will be responsible for all items in section 5.3. Monitor team activities with an emphasis on Safety during: 1. Design of the vehicle and payload 2. Construction of the vehicle and payload 3. Assembly of the vehicle and payload 4. Ground testing of the vehicle and payload 5. Sub-scale launch test(s) 6. Full-scale launch test(s) 7. Launch Day 8. Recovery activities 9. STEM Engagement Activities Verification Demonstration - The team develops and utilizes their launch and safety checklists at all flights (testing and at competition) and at LRR The team has identified Jackson Treese as the student safety officer Demonstration- The Safety team is present during all activities by being embedded within every team, and monitors all activities to ensure that they are performed safely Implement procedures developed by the team for construction, assembly, launch and recovery activities Manage and maintain current revisions of the team s hazard analysis, failure analysis, and procedures and MSDS/inventory data Assist in the writing and development of the team s hazard analyses, failure modes analyses, and procedures. During test flights, teams will abide by the rules and guidance of the local rocketry club s RSO. The allowance of certain vehicle configurations and/or payloads at the NASA Student Launch does not give explicit or implicit authority for teams Demonstration During their duties while embedded in the other teams, safety liaisons record the procedures that their teams follow, recording best practices and ensuring that safety checks are implemented. Demonstration The safety team maintains current revisions of all required documents Demonstration The safety team assists in writing all required documents Analysis, Demonstration During test flights and design, the team ensures that all aspects of the rocket are compliant with launch site requirements. AUBURN UNIVERSITY STUDENT LAUNCH 129

130 to fly those vehicle configurations and/or payloads at other club launches. Teams should communicate their intentions to the local club s President or Prefect and RSO before attending any NAR or TRA launch. 5.5 Teams will abide by all rules set forth by the FAA. Demonstration The team abides by all rules set forth by the FAA AUBURN UNIVERSITY STUDENT LAUNCH 130

131 6.3 Team Derived Requirements Requirement Number 6.1 Description Justification & Verification In addition to the requirements set forth by NASA, we have self-imposed restrictions that were considered when planning the rover. Extra information can be seen under the Rover Completing Team Set Requirements section Justification: The team is limiting themselves to the The most important restriction required the dimensions of a 6-in diameter body tube. rover and orientation system to fit inside the 6- Verification: The team has designed and will test the in diameter launch vehicle. system to ensure it will fit in the allotted space. Justification: The rocket can land in any orientation and the rover works in one orientation, so the rover The rover must exit the launch vehicle in the must be oriented correctly. correct orientation before completing the Verification: The team has planned to use a rover experiment. orientation system to ensure the rover is oriented correctly. A ball bearing system was designed by the team to meet the requirements of the system. Justification: Launch days can be unpredictable, and the rocket may sit on the pad for an indeterminate amount of time. Verification: The team has designed the rover and its The rover must be able to have enough battery batteries to remain active for at least 2 hours on the power to be able to operate after sitting on the launch pad. The team will be using lithium polymer launch pad for up to 2 hours. batteries that provide more than 5 times as many milliamp hours as the 9 Volt batteries used on last year s rover. If they do not meet the requirements, we can purchase batteries with a greater mah rating of the same size and they will operate the same. Due to the specific design of the vehicle body, there are specific team requirements based on the design The Rocket must have a structural factor of safety of at least 2.0. Justification: Composites are brittle, showing little sign of failure before breaking, and when shattered leaves jagged, dangerous edges. Plus, if a structural component has shattered the mission will likely fail AUBURN UNIVERSITY STUDENT LAUNCH 131

132 6.3 Verification: The testing team will continue to perform materials testing on all materials that the team receives, as materials properties can vary for composites with age. Due to the specific design of the recovery system, there are self-imposed requirements the team placed on itself Once the BP charge cups have been wired to the altimeter BAE, the recovery system will not be armed until the rocket is on the launchpad. Wires need to come out of the drogue compartment post main separation The nose cone of the rocket needs to be jettisoned before landing without using an explosive charge The three shear pins holding in the main parachute must be able (in concert) withstand the twice the expected mass of the main parachute section, recovery avionics, rover section and nose cone. Justification: There is a possibility of premature detonation of the BP charge cups, and there is a requirement that the recovery electronics must have an external arming switch (HB 2.4). Verification: The team determined it is unsafe to arm the recovery system until the rocket is on the launchpad and designed the system to reflect that. Justification: If the wires do not come out of the drogue compartment post separation, the team is concerned that they will hold the bulk plate in front of the main parachute compartment in place, preventing separation. Verification: Ground testing with separating the compartments manually (i.e. just sliding the rocket apart at the required joint with wires passed through) shows that if packed correctly the wires will pull out easily. Justification: To facilitate rover deployment the nose cone needs to be out of the way of the rover prior to landing, but using a pressurized method may damage the rover. Verification: When constructed, testing will ensure that the nose cone will be able to fall out of the upper section under the influence of gravity alone. Justification: Due to the unique configuration of the recovery system this year, the main parachute will be held in place by three shear pins during descent under drogue. If the shear pins were to break, the parachute would be released early and the team would violate AUBURN UNIVERSITY STUDENT LAUNCH 132

133 6.4 NASA requirement 3.9 and 3.10 on descent time and drift. Verification: Extensive ground testing (see Section ) will ensure that the shear pins will be capable of withstanding far in excess of the required loads, and if not additional shear pins will be used. Since the team is partaking in an additional experiment with the altitude control system with the drag plates, the team will derive requirements for the system. Justification: At maximum velocity and plates fully extended the plates will experience no more than pounds of force. The drag plates must withstand 80 pounds of Verification: CFD analysis showed that each plate normal distributed force. would be subjected to a maximum of pounds during flight. The plates were designed with a safety factor of 2 to ensure the plates do not fail in flight Batteries must last longer than three hours. Justification: Exceed Vehicle Requirement Ensure system can function after an extended pad stay hours. Justification: The vehicle will have a clean profile System must be completely embedded within when plates are fully retracted. In addition, it allows vehicle when plates are fully retracted. the assembly to be separate from the rest of the vehicle System must be fully autonomous. Justification: Reduces complexity of electrical components and reduces human error risk. Verification: Software will be written to simulate the vehicle in flight when system is fully assembled. 6.5 The team will follow and maintain these STEM Outreach requirements set by the team: Team members will be trained in the going on of different STEM outreach events to ensure they are prepared for them. Justification: Untrained team members might cause errors in the event flow and the building of rockets at events. Verification: The outreach team plans on holding training and information sessions for the team overall, so everyone is ready for events. AUBURN UNIVERSITY STUDENT LAUNCH 133

134 Justification: Maintained relationships creates familiarity with the organization in the community The team must maintain ongoing relationship with events to ensure stable reliability for events, while always trying for more events. and makes other more open to work with the team. Verification: The team continues to plan and work at events if they run year to year. The outreach team as already started communications and planning the yearly events. 6.6 The team will follow and maintain these Testing requirements established by the team. All team created components integrated into the Justification: It is critical to ensure all components launch vehicle will be thoroughly tested and launched will perform as needed before, during, certified on the ground before being certified for and after flight. As such, these components must flight be pre-certified to ensure launch success AUBURN UNIVERSITY STUDENT LAUNCH 134

135 6.4 Line Item Budget The following tables represent the current expenses as of December 31, 2018 for each team. Onyx 3D printer filament is used by all the teams, but the three rolls purchased to date have been divided into one roll each for Vehicle, Rover and Altitude Control. Vehicle Body Budget Vehicle Expenses To Date Item Cost Per Unit Unit Quantity Total $0 (Donated Per Prepreg Carbon Fiber by GKN square 9 $0 Aerospace) yard Aerotech K-1275R $ Per unit 2 $ RMS 54/2560 Motor Case and Associated Hardware $220 Per unit 1 $220 4" Fiberglass Coupler $26.78 Per unit 3 $80 Rail Button Material $30.69 Per bar stock 1 $31 4:1 Nose cone $ per unit 1 $166 Onyx Printer Filament $ Per Roll 1 $198 Aerotech L1420R $ Per unit 1 $ RMS 75/5120 Motor Case and Associated Hardware $390 Per unit 1 $390 6" Fiberglass Coupler $69 Per unit 3 $207 Total $1,868 AUBURN UNIVERSITY STUDENT LAUNCH 135

136 Recovery Budget Recovery Expenses To Date Item Cost Quantity Total Ripstop Nylon (yards) $ $ Shear Pins $ $17.35 Parachute Swivels $ $37.98 Stratologgers $55 4 $220 Adafruit Trinket $14 1 $14 Nylon Thread $8 3 $24 Tubular Nylon $1 50 $50 Paracord $5 1 $5 Total $473 Rover Budget Rover To Date Item Price Quantity Total Transfer Units BRUPS-15-S-NBK Press Fit Type Ball Transfer Unit $15 3 $ Servo LewanSoul LD-3015MG $ $16.49 Turn table ball bearing $10 1 $ Servo Adafruit TowerPro SG-501 $ $16.99 Plastic Shelving Tower $ $24.08 Bread Boards $ $14.98 LiPo 800mAh 2s 7.4V $ $14.99 LiPo 1400mAh 2s 7.4V $ $18.99 Roll of ABS printer filament $ $21.99 Roll of Onyx printer filament $198 1 $198 Sum $ AUBURN UNIVERSITY STUDENT LAUNCH 136

137 Altitude Control Budget Altitude Control Expenses to Date Items Name Price Accelerometer ADXL345 $17.50 SD Card Datalogger MicroSD card breakout board+ $7.50 Altimeter/Pressure Sensor MPL3115A2 $9.95 RTC DS3231 $ V Batteries Duracell $10.00 Onyx Structure (one roll of material) MarkForged $ Microcontroller STM32F103C8T6 Blue Pill $2.50 Motor Driver Adafruit DRV8833 $4.95 Total $ Educational Outreach Education Outreach Current Item Cost 1.5" Carboard Motor Mounts $41.59 Plywood and Tools $ " Carboard Airframes $54.39 Estes Educator Rockets Straw Rockets $20.00 Estes Educator Rockets $2, Estes A8-3 Motors $1, Total Cost $3,578 AUBURN UNIVERSITY STUDENT LAUNCH 137

138 Lab Supplies and Launch Fees Lab Supplies And Launch Fees (actual) Item Cost Per Unit Unit Quantity Total Calipers $9.99 Per unit 4 $39.96 Cutting Mats $29.96 Per unit 2 $59.92 Silicone Heat Hose $71.20 Per unit 1 $71.20 Vacuum Pump $79.99 Per unit 1 $79.99 Pump Oil $16.08 Per unit 1 $16.08 Vacuum Valves $75.00 Per unit 1 $75.00 Hose Clamps $14.89 Per unit 1 $14.89 Hammers $8.53 Per unit 2 $17.06 Bama Blast Off Registration $5.00 Per team member 7 $35.00 Screwdriver Set $27.99 Per Set 1 $27.99 Restock Allen Wrench sets $6.14 Per Set 5 $30.69 Nitrile Gloves $86.38 Per Box 1 $86.38 Total $ Summary Table Budget Summary Vehicle $1,868 Recovery $473 Rover $ Altitude Control $ Education/ Outreach $3,578 Lab Supplies and Launch Fees $ Total $7, AUBURN UNIVERSITY STUDENT LAUNCH 138

139 6.5 Funding Funding Sources The team s current funding level is in Table 33: Funding Sources. This table represents the current funds that have been donated to AUSL. The team expects to receive further donations over the course of the year, but for now it will operate under the assumption that these funds are all that are available. Compared to PDR, the team has received more money in donations from current Auburn Alumni. The team continues to search for more funding sources to allow for major improvements to facilities if the team can maintain the current level of budget surplus. In addition to the generous donations below, the team must also thank GKN Aerospace in Tallassee, Alabama for their support. Instead of providing monetary support, GKN donates surplus composites to Auburn University, which the team is in turn able to use to manufacture our airframe. Source Amount Alabama Space Grant Consortium $14,000 Dynetics $2,500 Current Auburn Alumni Donations $2,000 Total Funding $18,500 Table 33: Funding Sources AUBURN UNIVERSITY STUDENT LAUNCH 139

140 Projected Total Expenses The following represent the few additional materials and tools that the team expects to purchases during the remainder of the competition, as well as the projected travel expenses. Note that no projected future expenses are currently listed for the Recovery team, and most of the components have been purchased for the other teams by this stage in the competition as well. A stock of all the tools, glue, batteries, etc. required for the educational team is maintained from previous years, but based off of previous years the team expects to spend around $100 restocking these supplies. Team Vehicle Item Carbon fiber and resin for open weave structure. Fiberglass Coupler for O-ACS full scale Aerotech L1420R for O-ACS and competition Cost Per Unit Quantity Total $284 2 $568 $69 3 $207 $ $ Rover Spool of Vinyl Print Filament $50 3 $150 Altitude Control 280 RPM Premium $ $49.99 Glue for Rockets $50.00 estimated $50.00 STEM Outreach Launch Controller Batteries $30 estimated $30 Rocket Building Supplies (X-acto Knives, Sandpaper) $20 estimated $20 Hotel and Travel $2,500 Promotional Materials $250 Total Future Expenses $4, Table 34: Projected Future Expenses AUBURN UNIVERSITY STUDENT LAUNCH 140

141 Budget Balance A summary of the results of the previous sections can be seen below in Table 35: Budget Balance. Project Nova II has seen a cost rise since PDR of around $1200, which can mostly be attributed to an error in the PDR report that neglected purchasing full scale motors as either a current or future expense. Nevertheless, the team still predicts that it will be able to come in well under budget based off of to date expenses plus further predictions for the rest of the year. This means the team will have the funding required into respond to any unexpected contingencies such as a total loss of the rocket due to a motor catastrophic failure. Hopefully, no such unexpected additional massive losses of funds will occur, and the team will have a large surplus at the end of year to invest into improved facilities. Category Numerical Value Funding Sources $18500 Total costs to date $ Total projected costs + $1500 more unexpected expenses $5, Budget Balance $5, Table 35: Budget Balance AUBURN UNIVERSITY STUDENT LAUNCH 141

142 6.6 Project Timelines A current Gantt chart for the overall project can be seen in Figure 57. Figure 57: Overall Project Gantt Chart Deliverables Description Conceptual Design Initial configuration brainstorming, payload selection, vehicle sizing and trade studies Proposal Time devoted to writing the proposal for competition CFD Testing Testing on rough body shape to develop drag plate system, see section for more detail Materials Testing See section AUBURN UNIVERSITY STUDENT LAUNCH 142

143 Subscale Fabrication Construction of first subscale Trade Studies Detail trade studies for full scale rocket, informing and being informed by Subscale construction Configuration Testing Using subscale hardware to test large aspects of design, such as unique recovery setup (See 3.2 and ) PDR Time allotted for writing the PDR report First Subscale Flight 10/06/2018, successful except for launch instability Subscale Redesign Analysis of launch data to determine cause of failure Subscale Modification Changes to compartmentalization and fin sizing Payload fabrication Start of design/build of payloads (See 0 and 6.6.2) Subscale Reflight 10/27/2018, fully successful flight see Full Scale Fabrication (Solid Carbon) Construction of primary hand layup airframe Full Scale Fabrication (O-ACS) Construction of experimental, lighter airframe CDR Time allotted for writing the CDR report Payload integration into vehicle Work on payload interface with vehicle and flight Full scale w/ Payload flight window Time period of potential scored full scale flights First Full Scale Flight 1/12/2019, Planned first full scale flight FRR Time allotted for writing the FRR report Competition Preparation Rocket fair display, shirts arrive, additional prep Huntsville Competition USLI competition in Huntsville PLAR Time allotted for writing the PLAR report Table 36: Project Deliverables. This schedule has seen few changes since the PDR report, and most have been with regards to the fabrication of the full scale rocket. As mentioned in section 2.3: Project Plan Changes, this revolves around the team being able to put a dates on construction of both our conventional hand layup composite airframe and our Open Architecture Composite Structure (O-ACS) airframe. Since the terms on the Gantt chart for the overall project are broad, Table 36: Project Deliverables describes each deliverable in more detail. AUBURN UNIVERSITY STUDENT LAUNCH 143

144 Deliverables Description Conceptual Design Initial configuration brainstorming, payload selection, vehicle sizing and trade studies Proposal Time devoted to writing the proposal for competition CFD Testing Testing on rough body shape to develop drag plate system, see section for more detail Materials Testing See section Subscale Fabrication Construction of first subscale Trade Studies Detail trade studies for full scale rocket, informing and being informed by Subscale construction Configuration Testing Using subscale hardware to test large aspects of design, such as unique recovery setup (See 3.2 and ) PDR Time allotted for writing the PDR report First Subscale Flight 10/06/2018, successful except for launch instability Subscale Redesign Analysis of launch data to determine cause of failure Subscale Modification Changes to compartmentalization and fin sizing Payload fabrication Start of design/build of payloads (See 0 and 6.6.2) Subscale Reflight 10/27/2018, fully successful flight see Full Scale Fabrication (Solid Carbon) Construction of primary hand layup airframe Full Scale Fabrication (O-ACS) Construction of experimental, lighter airframe CDR Time allotted for writing the CDR report Payload integration into vehicle Work on payload interface with vehicle and flight Full scale w/ Payload flight window Time period of potential scored full scale flights First Full Scale Flight 1/12/2019, Planned first full scale flight FRR Time allotted for writing the FRR report Competition Preparation Rocket fair display, shirts arrive, additional prep Huntsville Competition USLI competition in Huntsville PLAR Time allotted for writing the PLAR report Table 36: Project Deliverables AUBURN UNIVERSITY STUDENT LAUNCH 144

145 Rover Timeline Figure 58: Rover Gantt Chart As can be seen from the figure above, the rover has experienced some delays compared to the PDR report. The CAD design for the Rover took longer than expected, as difficulty was encountered with fitting all the required electronics within the frame and the design passed through multiple iterations. This in turn delayed the Sled CAD as well, as the two are necessarily closely linked. However, after a short holiday break the team has begun fabricating in order to fly on the first full scale launch planned for the beginning of January. AUBURN UNIVERSITY STUDENT LAUNCH 145

146 Altitude Control System Timeline Figure 59: Altitude Control System Gantt chart A quick glance at Figure 59 shows that the altitude control system timeline remains the same since PDR, and that the team is currently in a cycle of assembling and iterating on the design in order to integrate the altitude control system into the full scale rocket for launch. AUBURN UNIVERSITY STUDENT LAUNCH 146

147 Educational Engagement Timeline Figure 60: STEM Outreach Gantt Chart The AUSL team has continued to perform STEM outreach events, and has acquired the large stock of rocket building kits necessary for all of this year s remaining currently planned outreach events with a significant surplus. This allows the team to continue to flexibly respond to new outreach opportunities beyond those currently planned. Compared to PDR, the team was able to set a date in November and complete our Auburn Junior High School E-Day event, complete our Sanford Middle School STEM day, and add and participate in a new event at Locapoka High School. Additionally, the team was able to set an exact date with the Boy Scouts, and have continued to talk to the local Girl Scouts troops, although a final date has not yet been set for our event. For additional information on our completed events, please see our STEM engagement forms submitted separately after each event. AUBURN UNIVERSITY STUDENT LAUNCH 147

148 Appendix A. Team General Information For reference, all general information from the proposal is repeated below. 6.7 Team summary General Team Information University Affiliation Mailing Address Title of Project Auburn University 211 Davis Hall Auburn, AL Nova II Date of Proposal September 12, 2018 Experiment 6.8 Adult Educators Option 2: Deployable Rover Adult Educator Name Title Dr. Brian Thurow Aerospace Engineering Department Chair Phone Davis Hall Address Auburn, AL Adult Educator AUBURN UNIVERSITY STUDENT LAUNCH 148

149 Name Title Robert Kulick Aerospace Engineering Faculty Advisor Phone Davis Hall Address Auburn, AL Team Mentor Mentor Name Title Dr. Eldon Triggs Lecturer, Aerospace Engineering, Mentor Certification Tripoli Rocketry Association Level 2 trigged@auburn.edu Phone Davis Hall Address Auburn, AL Student Team Leader Student Project Lead Name Title Bryce Gardner Senior in Aerospace Engineering blg0010@auburn.edu Phone South Gay St Apt. 106 Address Auburn, AL AUBURN UNIVERSITY STUDENT LAUNCH 149

150 6.11 Student Safety Officer Student Safety Officer Name Title Jackson Treese Junior in Aerospace Engineering Phone E University Dr. Apt 8D Address Auburn, AL Project Organization This year, the team has four technical development teams and four operations support teams. All students that are not in a leadership position are a member of a technical team. The technical teams are vehicle body, recovery, payload, and altitude control. They are responsible for the design and fabrication of major sections of the rocket. The support teams are systems engineering, educational engagement, safety, and paper editing. Members of the support teams are embedded in the other teams. This is essential for the systems engineering and safety teams to perform their roles by engaging them at the lowest level of organization. This arrangement also provides additional experience for members of support teams, especially educational engagement and paper editing, which may otherwise not have the opportunity to work in a technical role. During educational events, members of the outreach team organize and coordinate while members of all teams volunteer. An organizational chart of the team can be seen in Figure 61: Organizational Chart. AUBURN UNIVERSITY STUDENT LAUNCH 150

151 6.13 NAR/TRA Sections The Auburn Student Launch team is planning to attend launches hosted by Phoenix Missile Works (PMW) and the South Eastern Alabama Rocket Society (SEARS). PMW (Tripoli Section #81) will be hosting a new launch event called Bama Blastoff in Aliceville, Alabama on October SEARS (NAR Section #572) will host a launch on the first Saturday of every month during Figure 61: Organizational Chart AUBURN UNIVERSITY STUDENT LAUNCH 151