Student Launch. Enclosed: Flight Readiness Review. Submitted by: Rocket Team Project Lead: David Eilken. Submission Date: March 03, 2017

Transcription

1 University of Evansville Student Launch Enclosed: Flight Readiness Review Submitted by: Rocket Team Project Lead: David Eilken Submission Date: March 03, 2017 Payload: Fragile Material Protection Mentor: Dr. David Unger, NAR 89083SR Level 2 Submitted to: NASA Student Launch Initiative Program Officials Faculty of the UE Mechanical Engineering Program University of Evansville College of Engineering and Computer Science 1800 Lincoln Avenue; Evansville, Indiana P a g e

2 Table of Contents Table of Contents... ii List of Figures... v List of Tables... viii Nomenclature... xii FRR Summary... 1 Design Updates from Proposal... 2 Changes Made to Vehicle Criteria... 2 Changes Made to Payload Criteria... 2 Changes Made to Project Plan... 3 Vehicle Criteria... 4 Design and Construction of Vehicle... 4 Recovery Mission Performance Predictions Mission Performance Criteria Flight Simulations and Altitude Predictions Validity Assessment Actual Stability Margin Kinetic Energy Drift ii P a g e

3 Full Scale Flight Launch Day Conditions Flight Analysis Flight Results Payload Criteria Safety Personnel Hazard Analysis Failure Modes and Effects Analysis Environmental Considerations General Risk Assessment Launch Operations Procedures Parts Checklist Final Assembly Checklist Motor Preparation Recovery Preparation Setup on Launch Pad Ignitor Installation Launch Procedures Troubleshooting Post-Flight Inspection iii P a g e

4 Project Plan Testing Altimeter MTS (Bulkhead) Ejection Testing Parachute Deployment Force Testing Wind Tunnel Testing Scale Model Testing Payload Testing Full Scale Testing Requirements Compliance Budgeting and Timeline Budget Schedule References Appendix A Machine Prints Appendix B OpenRocket Simulation Appendix C Best Fit Curve Appendix D OpenRocket Simulation Appendix E Payload Part Specification iv P a g e

5 Appendix F Line Item Budget Appendix G Task Breakdown Appendix H Electrical Diagrams Appendix I Payload Accelerometer Graphs Appendix J Wind Tunnel Uncertainty Appendix K MTS Tensile Test Procedure List of Figures Figure 1 - Recovery electronics wiring diagram... 6 Figure 2 - Altimeter Wiring... 7 Figure 3 - Body Tube Figure 4 - Fin Stops Figure 5 - Painted Fins Figure 6 - Epoxy Nuts Figure 7 - Finished Nosecone Figure 8 - Epoxy Location for the Centering Rings Figure 9 - Complete coupling tube bulkhead assembly Figure 10 - Coupling tube with permanent bulkhead and all thread rods Figure 11 - Electronics sled assembly Figure 12 - Aft recovery mounting point Figure 13 - Cylinder 2 pin holes 3 inch spacing v P a g e

6 Figure 14 - Bulkhead Cylinder Figure 15 - Rough finish on bulkhead Figure 16 - Wire rope isolator pin and aluminum square assembly Figure 17- Final Assembly Figure 18 - Altimeter mounting assembly Figure 19 - Mounted O Ring Figure 20 - Battery Holder Figure 21 - Mounting Pins Figure 22 - Altimeter Mounting Assembly Figure 23 - Block diagram of recovery system Figure 24 - Full-Scale Simulation Figure 25 - Flight Simulation Input Data Figure 26 - Simulated Flight Configurations Figure 27 - Anticipated Motor Thrust Curve from OpenRocket Figure 28 - Flight 1 Actual vs OpenRocket Data Figure 29 - Flight 2 Actual vs OpenRocket Data Figure 30 - Flight 3 Actual vs OpenRocket Data Figure 31 - Actual Cp and Cg locations Figure 32 - Actual Altitude vs OpenRocket Altitude Figure 33 - Actual Data vs Regression Figure 34 - OpenRocket Data vs Regression... Error! Bookmark not defined. Figure 35 Predicted Coefficient of Drag During Flight Figure 36 - Final Design Assembly (new bolt and washer mounting) vi P a g e

7 Figure 37 - Exploded view of payload assembly (annotation following) Figure 38 - Exploded view base spring attachment Figure 39 - CR1-400 wire rope isolator pin and plate assembly Figure 40 - Fracture mechanics clevis grip attached to U bolts Bulkhead assembly for MTS testing Figure 41 - The Assembly Mounted into the MTS Machine Figure 42 Variable Frequency Drive Figure 43- Strain Gage (From Vishay website) Figure 44 - Strain Indicator Figure 45 - Air Fan Figure 46 - Wind Tunnel Figure 47 - Example of wiring strain gage to strain indicator Figure 48 - Wiring Diagram (strain gage to strain indicator) Figure 49 - Pareto Chart Figure 50 - Accelerometer Data Full-scale Flight Figure 51 - Sectional Budget Amounts Figure 52 - Gantt Chart Figure 53 Aft Body Tube Drawing Figure 54 - Bow Body Tube Drawing Figure 55 - Fin Drawing Figure 56 - Motor Drawing Figure 57 - Nosecone Drawing Figure 58 - Launch Vehicle Drawing vii P a g e

8 Figure 59 Recovery bulkhead drawing Figure 60 - Payload Main bulkhead residing in Cylinder Figure 61 - Payload assembly general dimensions Figure 62 - Recovery attachment bulkhead and hardware Figure 63 - Altimeter Mounting Plate Piece Figure 64 - Altimeter Mounting Plate Vertical Figure 65 - Metal O-Ring Figure 66 Propulsion Section Figure 67 Inner Tube Figure 68 - Centering Ring Figure 69 - Thrust Plate Figure 70 - Inner Cylinder Figure 71 - Payload Coupler Figure Degree Cotton Fill Large Bulb Figure Degree Paper Fill Large Bulb Figure Degrees Packing Peanuts Large Bulb Figure Degrees Large Bulb Only Figure Degrees DogBrag Fill Large Bulb Figure Degrees Base Value List of Tables Table 1 - Vehicle Specifications... 4 Table 2 - System Level Functional Requirements... 9 viii P a g e

9 Table 3 - Simulation Summary Different Launch Configurations Table 4 - Rail Exit Velocity on Different Flights Table 5 - Mach Number on Different Flights Table 6 Impact of Wind Speed on Altitude Table 7 - Actual Stabilities Table 8 - Predicted kinetic energy of launch vehicle sections Table 9 - Predicted drift distance for selected wind speeds Table 10 - Actual Flight vs Predicted Flights Summary Table 11 - Definitions for Hazard and Failure Mode Analyses Table 12 - Personnel Hazard Analysis - Epoxy Table 13 - Personnel Hazard Analysis - Launch Operations/Post-Launch Inspection Table 14 - Personnel Hazard Analysis - Testing Table 15 - Personnel Hazard Analysis - Fabrication Table 16 - Personnel Hazard Analysis - Education Engagement Outreach Events Table 17 - Failure Modes and Effects Analysis - Design/Fabrication Table 18 - Failure Modes and Effects Analysis - Payload Table 19 - Failure Modes and Effects Analysis - Payload Integration Table 20 - Failure Modes and Effects Analysis - Recovery System Table 21 - Failure Modes and Effects Analysis - Testing Table 22 - Failure Modes and Effects Analysis - Launch Support Equipment Table 23 - Failure Modes and Effects Analysis - Launch Operations Table 24 - Environmental Consideration Hazard Analysis Table 25 - General Risk Assessment ix P a g e

10 Table 26 - Parts Checklist - Propulsion Table 27 - Parts Checklist - Aerodynamics Table 28 - Parts Checklist - Main Payload Table 29 - Parts Checklist Electronics Payload/Avionics Bay Table 30 - Parts Checklist - Recovery Table 31 - Parts Checklist - Safety and Education Table 32 - Parts Checklist - Miscellaneous Table 33 - Final Assembly Checklist - General Set Up Table 34 - Final Assembly Checklist - Comprehensive Structural Inspection Table 35 - Final Assembly Checklist - Electronics Table 36 - Final Assembly Checklist - Payload Table 37 - Final Assembly Checklist - Recovery System Table 38 - Final Assembly Checklist - Motor/Ejection System Preparation Table 39 - Final Assembly Checklist - Secure Attachment Inspection Table 40 - Final Assembly Checklist - Launch Pad/Pre-Launch Inspection Table 41 - Motor Preparation Checklist Table 42 - Recovery Preparation Checklist Table 43 - Launch Pad Configuration Checklist Table 44 - Ignitor Installation Checklist Table 45 - Launch Procedures Checklist Table 46 - Troubleshooting - Cracking in Main Body Tube or Subsection Table 47 - Troubleshooting - Insecure Fit Between Adjoining Subsections Table 48 - Troubleshooting - Unresponsive or Malfunctioning Electronics x P a g e

11 Table 49 - Troubleshooting - Insecure Connection Between Launch Rail and Launch Pad 126 Table 50 - Post-Flight Inspection Checklist Table 51 - Test Results Table 52 - MTS Test Results Table 53 - Results of ejection testing Table 54 - Maximum force on launch vehicle during descent Table 55 - Testing Apparatus Components Table 56 - Inputs for Uncertainty analysis Table 57 - Spring Constant Test Values Table 58 - Charpy Impact Acceleration Test Data Table 59 - Fragile Material Sample Testing Table 60 - Full Scale Flight Results Table 61 - NASA Requirement Compliance Table 62 - Team Requirement Compliance Table 63 - Sources of Funding Table 64 - Sectional Budget Breakdown Table 65 - Critical Dates xi P a g e

12 Nomenclature AP Ammonium Perchlorate Composite BH Bulkhead Cp Specific Heat with constant pressure [kj/kmol-k] CR Centering Rings E modulus of elasticity [psi] F Force [lbf] f operational frequency [Hz] f n natural frequency [Hz] F. O. S Factor of Safety h enthalpy [kj/kmol] Combustion Analysis Section h thickness [in] h f o Enthalpy of Formation [kj/kmol] I second moment of inertia [in 4 ] k stiffness [psi] KE Kinetic Energy L length [in] xii P a g e

13 M moment [lbf-in] m mass [lbm, kg] kg will be specified in the equation otherwise it is lbm n moles [kmol] P pressure [kpa] Q heat [kj] R radius [in] R Gas Constant [=8.314 kj/kmol-k] Combustion Analysis Section r frequency ratio SA surface area [in 2 ] T temperature [K] t time [s] V volume [in 3 ] v velocity of the rocket at burnout [m/s] v f ground impact velocity [ft/s] v d descent rate [ft/s] ω circular natural frequency [rad/sec] w work [kj] xiii P a g e

14 FRR Summary Project ACE will field a long, 32.6-pound carbon fiber and aluminum based rocket. The leading tip of the rocket begins with a G-10 Fiberglass, 22, ogive nosecone. Contained in a pressure-equalizing compartment in the nosecone sits the official altimeter as well as a GPS tracking system. Just aft of this compartment are four threaded rods for fastening ballast. A fragile material protection system resides below the nosecone. The payload contains concentric cylinders, connected by an array of springs and wire-rope isolators selected through extensive mathematical modeling. The innermost cylinder, where the fragile material is to be contained, features a variable position cap and fill material to ensure that the fragile material will be contained under sufficient pressure regardless of volume. It is the team s objective to produce a successful payload that provides meaningful vibration and impulse reduction information. Moving aft from the payload is the recovery system. This system features completely redundant separation circuits. At apogee, a 24 drogue chute ejects, followed by a 96 main chute at 750. At the aft end of the rocket is the propulsion section. A 75-mm L-850W Aerotech motor propels the rocket for just over four seconds to an altitude of one mile. A extruded aluminum launch rail has been selected to achieve an acceptable rail-exit velocity. The motor is held in place via 6061-T6 Aluminum centering rings and thrust plate. All components are housed in two carbon fiber body tubes. The fins, which adhere to the centering rings and body tubes, are made out of G-10 Fiberglass and have a clipped delta design. Each system is covered in much more depth in the Vehicle Criteria section of this report. For specific team information, such as the mentor and mailing address, please see the cover page of this report. For more quick facts on the rocket please reference the associated milestone review flysheet. 174 P a g e

15 Design Updates from Proposal Changes Made to Vehicle Criteria The nosecone shoulder was shortened from 5.25 inches to 3 inches. The change accounts for spring oscillation from the main payload, which is located below the nosecone in the bow body tube. A 3 8 inch hole was drilled in the furthest aft centering ring. The hole allows the furthest aft rail button to be tightened and loosened as needed by giving access to the interior threads of the rail button. An aluminum nut holds the rail button to the airframe and the rail button assembly is now removable if necessary through the aft centering ring. It was observed that after sub-scale test flights, the quarter-inch quick links used to secure the recovery harnesses to the launch vehicle body tubes had become mildly deformed, making it irksome to tighten or loosen them. Consequently, larger quick links were initially selected for the full-scale launch vehicle. The shift in mass associated with these larger quick links created a low stability off the launch rod for the second full-scale test flight. In order to return stability to an acceptable value, the original quarter-inch quick links were reemployed and will be used on all launch configurations moving forward. Changes Made to Payload Criteria The main change the payload saw was the mounting of the base springs. Due to epoxy failing during impact and welds weakening the integrity of the spring, a new design was developed and used in testing. This design is spoken about in detail in the Payload section of this report. One other decision that was made through the testing of the payload was the final 2 P a g e

16 choice of fill material for Cylinder 1 (the innermost cylinder). The selection is a combination of shredded paper and cotton filling and the rationale behind the decision is described in detail in the payload testing section of this report. During testing, epoxy continued to fail on mating surfaces especially where the wire rope isolators were adhered to Cylinder s 1and 2. To combat this failure, pins were used and inserted into holes in both cylinders with epoxy to add further strength. Thus, for the payload as a whole, conceptual changes were not made but small changes to the mounting design were made. Changes Made to Project Plan A few changes were made to both the schedule and budget. Project ACE s build phase extended about one week longer than anticipated mainly due to the redesign of payload mounting. The redesign resulted in a multi-week delay in testing. Project ACE was also forced to launch one week late due to weather. The team launched with BluesRocks Rocketry in Elizabethtown, KY instead of Laünch Crüe. These scheduling changes are reflected in the Schedule section of this report, but Project ACE is on schedule once again. Slight alterations were made to the budget to accommodate sections with unforeseen costs. In essence, funds for sections of the rocket that were under budget were allocated to administrative and payload sections to cover overrun costs. Project ACE remains under budget, but more detail on the allocation of funds can be found in the Budget section of the report. 3 P a g e

17 Vehicle Criteria Design and Construction of Vehicle Design Features Structural Elements Vehicle Overview The vehicle specifications can be seen in Table 1. The overall length of the rocket is inches with a diameter of 5.5 inches. Table 1 - Vehicle Specifications Component Dimension Material Bow Body Tube 48 inches Carbon Fiber Aft Body Tube 41 inches Carbon Fiber Nosecone inches G10 Fiberglass Bulkhead/Centering Ring ¼ inch Aluminum Coupler 12 inches G10 Fiberglass Body Tubes The body tubes provide the structural rigidity necessary for housing the internal components as well as undergoing flight/recovery stresses. These tubes also account for the bulk of the mass of the airframe and provide a large surface area for airflow while in flight. To guarantee a successful flight, all these factors must be accounted for in the material selection of the body 4 P a g e

18 tubes. Carbon fiber was selected to provide a lightweight frame for the launch vehicle (0.658 oz/in 3 ) while also providing a higher tensile strength than that of fiberglass or BlueTube. Nosecone The nosecone must withstand the forces of in-flight airflow and vehicle recovery, however, both of these forces are minimal and do not require the increased strength provided by carbon fiber. The nosecone is also smaller than the body tubes and provides less of a weight reduction from carbon fiber to fiberglass. Lastly, the official scoring altimeter of Project ACE is housed in the nosecone. The altimeter is specified to not be housed in carbon fiber for transmission purposes. For these reasons, it was chosen to use fiberglass instead of carbon fiber. Coupling Tube The coupling tube serves as the joint between body tubes and the housing for the recovery electronics. This coupler separates from the remainder of the launch vehicle during the recovery process. Aluminum caps seal the space on either side of the coupler, and threaded aluminum rods connect the aluminum caps. The caps and rods bear the stresses of the recovery process. For this reason, an inexpensive material was able to be chosen for the coupler. BlueTube was the material chosen because it is readily available from many manufacturers at a low cost. Bulkheads/Centering Rings The bulkheads and centering rings provide additional structural integrity for the launch vehicle. They also serve as possible mounting points for vehicle components such as the motor retention system or shock cords. Lastly, they are used to separate the recovery section from the payload and the propulsion section. Aluminum was chosen for the bulkheads for its high tensile strength (300 MPa) to ensure the success of the crucial functions these components perform. 5 P a g e

19 Electrical Elements Recovery electronics are connected using 20-gauge wire. Screw terminals are used to make electrical connections with the terminal blocks and recovery altimeters. Connections to the rotary arming switches are soldered. 4-pin Molex connectors are soldered in series with the signal wires from each altimeter to allow the electronics sled to be removed from the coupling tube. To allow for easier management, all wire pairs were twisted neatly and some wiring was secured to the sled using metal retainers. A complete wiring diagram of the recovery electronics is shown in Figure 1. Figure 1 - Recovery electronics wiring diagram The scoring altimeter has two connections. The first, which runs to the battery, connects into the socket at the bottom of the altimeter. On the other end of the wire, it is soldered into the battery. The second connection is the switch to turn the altimeter on or off. The toggle switch is soldered to two leads which are locked into place on the altimeter by a terminal block. Both the 6 P a g e

20 attached toggle switch and the battery connection can be seen on Figure 2 in its respective socket. Drawing and Schematics Figure 2 - Altimeter Wiring A full listing of dimensioned drawings can be seen in Appendix A. Flight Reliability Mission Success Criteria Listed below are the mission success criteria determined by the Project ACE team. 1. Aerodynamics a. The airframe, nose cone, and fins remain intact for the duration of the flight. b. The airframe, nose cone, and fins are reusable for any following flights. 7 P a g e

21 c. The airframe and nose cone will protect all internal components from damage from external sources. 2. Propulsion a. The vehicle will attain an apogee between 5,125 feet and 5,375 feet. b. The vehicle will remain below Mach 1. c. The motor mount will withstand propulsion forces and remain reusable for any following flights. 3. Recovery a. The drogue parachute and main parachute are ejected at apogee and 750 feet, respectively. b. The drogue parachute and main parachute inflate successfully following ejection. c. The maximum kinetic energy of any independent section of the rocket is less than 75 ft-lbf at landing. 4. Electronic Payload a. The data sent from the electronic payload is received remotely during and after the vehicle s flight. b. The electronic payload withstands flight forces and remains reusable for any following flights. c. The electronic payload accurately determines the apogee of the rocket. 5. Main Payload a. The fragile object(s) remain undamaged. 8 P a g e

22 Propulsion Aerodynamics b. The force acting on the payload is reduced by 50% for each of the areas of interest: (thrust curve, parachute deployment, and landing.) c. The acceleration acting on the payload is reduced by 50% for each of the areas of interest: (thrust curve, parachute deployment, and landing.) Flight Reliability Confidence The system level functional requirements can be seen in Table 2 where the severity and likelihood of failure in each mission success criteria and the action performed to mitigate these failures are described. Table 2 - System Level Functional Requirements Section Success Criteria Explanation The airframe, nose cone, and fins should remain intact for the duration of the flight. The airframe, nose cone, and fins should be reusable for any following flights. The airframe and nose cone should protect all internal components from damage from external sources. The vehicle should attain an apogee between 5,125 feet and 5,375 feet. A failure of the airframe during flight could cause a complete failure in the launch vehicle s flight ability. However, the use of carbon fiber mitigates this risk to a very low likelihood. Reusability of parts is not detrimental to the project; new parts can be purchased. The most likely part to fail is a fin upon recovery, thus warranting a medium likelihood of failure. Damage to internal components can be detrimental to the launch vehicle s ability to deploy the recovery system. However, a carbon fiber airframe mitigates this risk to a very low likelihood. Motor variations and launch day conditions both contribute to apogee variations from the fullscale test. However, the team s allocated window should account for these. Severity of Failure Significant Minor Major Medium Likelihood of Failure Low Medium Low Low 9 P a g e

23 Electronic Payload Recovery Section Success Criteria Explanation The vehicle should remain below Mach 1. The motor mount should withstand propulsion forces and remain reusable for any following flights. The drogue parachute is successfully deployed at apogee. The main parachute is successfully deployed at 750 feet. The drogue parachute and main parachute inflate successfully following ejection. The maximum kinetic energy of any independent section of the rocket is less than 75 ft-lbf at landing. This is a requirement from NASA. The vehicle has not been designed to withstand transonic forces. The anticipated Mach number is Motor mount/retention failure could cause a poor flight, no flight, or safety hazard. This is mitigated by using aluminum and high strength epoxy. If the drogue parachute does not deploy at apogee, the main parachute will deploy at high velocity. This could result in damage to the parachute or airframe. If the main parachute does not deploy, the launch vehicle will descend under only the drogue parachute. This would result in excessive ground impact speed. A partially-inflated parachute is much less effective at slowing the launch vehicle during its descent. This could result in excessive ground impact speed. Excessive kinetic energy on landing could result in damage to the fragile payload or airframe. Severity of Failure Significant Major Major Major Major Major Likelihood of Failure Low Low Low Low Low Low The data sent from the electronic payload should be able to be received remotely during and after the vehicle s flight. If the data is not received remotely after the flight, the team will not be scored. Low Low 10 P a g e

24 Main Payload Section Success Criteria Explanation The electronic payload should withstand flight forces and remain reusable for any following flights. The altimeter notwithstanding the forces will prevent the altimeter from being reused, and the team cannot be scored. Severity of Failure Low Likelihood of Failure Low The electronic payload should accurately determine the apogee of the rocket. If apogee is not detected accurately, it will affect the score of the team. Low Low The fragile object(s) should remain undamaged. The force felt by the payload should be reduced by 50% for each of the areas of interest: takeoff (thrust curve, parachute deployment, and landing.) To properly reduce the risk of damage to any and all unknown fragile material, the desired reduction of force felt by the payload should be reduced by a minimum of 50 percent for the most extreme forces exerted throughout flight. Major Medium The Acceleration felt by the payload should be reduced by 35% for each of the areas of interest: (thrust curve, parachute deployment, and landing.) To reduce the force at the maximum and minimum points of spring displacement, total acceleration of the payload should be reduced by a minimum of 35 percent. Significant Low Construction Process Body Tubes / Nosecone One of the first steps that was taken in the construction of the body tubes was to cut down the bow body tube to forty-one inches in length, using a horizontal band saw. This can be seen in 11 P a g e

25 Figure 3. The body tube was then filed smooth using a file. The next step was drilling three 1 8 inch diameter holes in the bow body tube using a CNC. These holes were used to bolt the nosecone into the bow body tube. Three holes were drilled for the rail buttons on the side of the aft body tube using a CNC. Figure 3 - Body Tube The locations of the rail buttons were determined using the center of gravity and center of pressure. The rail buttons are attached to the body tube using a nut and bolt. Two of the three rail buttons are accessible, and can be removed from the rocket. The first rail button was fastened onto the aft body tube in the recovery section, the second rail button is fixed on the exterior of the aft body tube between two bulk heads, and the third is accessible through a hole that was drilled in the lower bulk head. The second rail button was not only bolted together, but also epoxied in place to ensure that it would not move or break free. The aft body tube was slotted using a CNC. These slots allowed the fin tabs to be inserted into the body tube. 12 P a g e

26 The fins were made from two, 2 x1 x1/8 sheets of G10 fiberglass. These were hand cut to the desired fin dimensions and were beveled using a freestanding horizontal belt sander. The dimensions of the fins are: 5.5 inches tall, the tip chord is 5.8 inches long, the root chord is 7.5 inches long, the fin tab has a height of 1.2 inches tall, and the fin tab length is five inches. 3 inch 8 thick ABS Plastic fin stops, pictured in Figure 4, were manufactured using a 3-D printer. The fin stops are inserted into the body tube flush with the centering ring, and epoxied on all contact surfaces to ensure a solid fit. These fin stops fit between each of the already epoxied fins, and provide extra internal support for the fins. Then the fins were painted for aesthetics, seen in Figure 5. Figure 4 - Fin Stops 13 P a g e

27 Figure 5 - Painted Fins The shoulder of the nosecone was reduced to three inches to allow the main payload to oscillate. The nosecone was attached to the body tube using three epoxy nuts, and bolts. An example of an epoxy nut can be seen in Figure 6. These epoxy nuts were epoxied using RocketPoxy to the inner diameter of the body tube, concentric with the bolt hole. The holes in the nosecone were lined up with the holes in the body tube before the bolts were tightened to form a snug fit. The nosecone was also painted for aesthetics, as seen in Figure 7. Figure 6 - Epoxy Nuts 14 P a g e

28 Figure 7 - Finished Nosecone Propulsion The propulsion section was constructed over a period of five days to ensure the epoxy was completely set before moving to the next section of the construction process. The first part of the process was milling out the center of the engine block (bulkhead) to be certain that the inner tube was in the middle of the plate. After the milling was completed, the blue tube cut to a length of inches was epoxied to the milled portion of the engine block and allowed to cure overnight. The next day, the engine block and inner tube assembly were epoxied into the body of the rocket 21 in from aft end of the rocket. The bulkhead was epoxied on the bow and aft side for a secure bond. Once the epoxy was applied, a loose centering ring was placed at the aft end of the body tube to make sure the bulkhead and epoxy were set completely in-line with the body tube. Once the bulkhead epoxy dried, the first centering ring was epoxied to the body tube 19 in from aft end of the rocket. The centering ring was epoxied on both the outer and inner edge around the body tube and the inner tube for a secure bond. Again, a loose centering ring was applied to ensure the centering right would not set-up at an angle. Figure 8 shows the location where the epoxy is applied to the centering rings. 15 P a g e

29 Outer Area Epoxy Location Inner Area Epoxy Location Figure 8 - Epoxy Location for the Centering Rings Once the first centering ring was set, the centering ring that sits in front of the fin tables was inserted. A fishing wire was tied around the centering ring so it could be pulled back to the front of the tabs for adjustability. When the fins were placed into the body tube, the centering ring was pulled aft to sit against the front of the fin tabs. With the centering ring in the correct position, the fins were removed to epoxy the circumference of the outer and inner edges of the centering ring. Once the epoxy was applied, a loose centering ring was inserted over the inner tube to ensure the entire assembly would not fall at an angle with the epoxy setting. With the centering ring dry, epoxy was applied to the area where the fins would sit against the centering ring. A more detailed description on how the fins and fin stops were assembled is included in the aerodynamic construction section. When all fins were epoxied in place, the last centering ring was inserted onto the inner tube resting against the back of the fin 16 P a g e

30 tabs. The centering ring was epoxied in place. Finally, epoxy was applied to the inner surface of the retention system and then placed onto the inner tube. Then epoxy was applied to the outside of the retention system making a fillet between the blue tube and the retention system. Recovery The construction of the full-scale recovery system began with the assembly of the coupling tube. A stock 12-inch blue tube coupler of 5.48-inch OD was selected to house the electronics. First, a 1-inch ring of 5.5-inch OD blue tube was epoxied around the middle of the coupling tube. This ring provides a smooth, continuous surface as air passes from one body tube to the next. It also locates the coupling tube vertically within the body tubes and allows access to the recovery electronics through pressure sampling holes. Four pressure sampling holes of inch diameter (as specified by the PerfectFlite Stratologger CF manual) were drilled through the ring and coupling tube, spaced equally around the circumference of the tube. Finally, two smaller rings of blue tube were epoxied to the inside of each end of the coupling tube, leaving a inch shoulder to locate the bulkheads. Two 0.25-inch thick aluminum bulkheads of inch OD were machined using a CNC mill. Holes of 0.25-inch and inch diameter were drilled on perpendicular axes to accommodate all thread rods and U-bolts, respectively. Each bulkhead received a steel U-bolt secured with hex nuts, flat washers, a steel backing plate, and lock washers. Epoxy was applied around the washers and nuts after assembly to ensure the bulkhead was airtight. One bulkhead, hereafter referred to as the permanent bulkhead, received two 14-inch steel all thread rods secured with hex nuts, flat washers, and lock washers. The all thread rods were first located such that they spanned the entire length of the coupling tube with equal lengths protruding from each 17 P a g e

31 bulkhead when fully assembled. They were then epoxied to the permanent bulkhead with all mounting hardware previously described. Two ejection charge wells made from 1-inch OD aluminum tubing were then epoxied to the outside of each bulkhead. A completed bulkhead is shown in Figure 9. EJECTION WELL U-BOLT MOLEX CONNECTOR TERMINAL BLOCK Figure 9 - Complete coupling tube bulkhead assembly After the all thread rods had been permanently fixed, the construction of the electronics sled could begin. First, a brass tube of 0.25-inch ID was slid over each all thread rod. The brass tubes each received a thin bead of epoxy along their lengths before being pressed against a sheet of balsa wood. Once dry, the tubes were correctly located to match the all thread rods, and additional epoxy was applied to secure the balsa wood sled to the brass tubes. Mounting holes were then drilled to accommodate the altimeters and batteries. The altimeters were secured using #4 bolts and nuts, while the batteries were secured using plastic zip-ties. Next, each altimeter 18 P a g e

32 arming switch was mounted through a hole in a small piece of balsa wood. This assembly was epoxied to the electronics sled such that the switched faced out radially, in line with opposite pressure sampling holes. Next, the permanent bulkhead was epoxied into the coupling tube at an orientation that aligned the arming switches with the pressure sampling holes, being careful to create an airtight connection around the circumference of the bulkhead. A bead of silicone rubber was applied around the shoulder where the other bulkhead, hereafter referred to as the removable bulkhead, was to rest. This ensured an airtight seal around the removable component. The coupling tube assembly with permanent bulkhead and all thread rods is shown in Figure 10. COUPLING TUBE ALL THREAD ROD MOLEX CONNECTOR Figure 10 - Coupling tube with permanent bulkhead and all thread rods With the electronics sled removed from the coupling tube, wires were soldered to each terminal of the arming switches and connected to the dedicated switch leads of each altimeter. 19 P a g e

33 Connections were made between each battery and the dedicated power leads of the corresponding altimeter. To allow for easy replacement of spent igniters, terminal blocks were epoxied to the outside of each coupling tube bulkhead. A inch hole was then drilled in each bulkhead to allow the passage of wires from the interior of the tube to the terminal blocks. Two pairs of wires (one for each recovery event s redundant igniters) were connected to each terminal block and fed through the bulkheads before being epoxied to create an airtight seal, as shown previously in Figure 9. The four wires concerned with main parachute deployment (from the aft-most coupling tube bulkhead) were connected in pairs to the primary and backup altimeter s MAIN leads, and the four wires concerned with drogue parachute deployment (from the bow-most coupling tube bulkhead) were connected in pairs to the primary and backup altimeter s DROG leads. In order to allow for the easy removal of the electronics sled between flights, these connections were made impermanent using 4-pin Molex connectors soldered between the altimeter leads and the terminal blocks. These connectors can be seen in Figure 9 and Figure 10, while the entire electronics sled assembly is shown in Figure 11. Figure 11 - Electronics sled assembly 20 P a g e

34 After construction of the coupling tube and its associated systems was complete, the two permanent recovery mounting points in the body tubes were created. Each mount was created by epoxying a steel U-bolt through an aluminum bulkhead of 0.25-inch thickness and 5.35-inch OD using the same hardware described for the coupling tube bulkheads. Each recovery mount was secured by first applying a small amount of epoxy around the inside of the body tube where the bulkhead would be located. After pressing each bulkhead into its final location, a small fillet of epoxy was applied around its circumference, followed by a larger fillet once the first had dried. The aft-side recovery mounting point is shown in Figure 12 after being epoxied into place. Figure 12 - Aft recovery mounting point Main Payload The assembly for the main payload began with cutting the 5.36 OD Blue Tube Coupler to 11 inches. Holes were then drilled into the coupler at a spacing of 3 inches apart starting 3 inches from the bottom of Cylinder 2, as seen in Figure P a g e

35 Figure 13 - Cylinder 2 pin holes 3 inch spacing The payload was also designed with two bulkheads, created with the CNC from 0.25-inch aluminum. The first bulkhead was epoxied into Cylinder 2 and had fifteen 0.2-inch diameter holes were milled and threaded for the bolts used to attach the base springs. Five 0.5-inch holes were milled to center the springs and a 1-inch diameter hole was milled out of the center to reduce the weight of any moving parts within the rocket. The final bulkhead can be seen in Figure 14. Figure 14 - Bulkhead Cylinder 2 22 P a g e

36 The second bulkhead mentioned was milled similarly, however instead of a 1.5-inch diameter hole in the center for weight reduction, it had two 0.3-inch diameter holes for the bolts used to connect to the recovery bulkhead located immediately below in the rocket. Once the bulkheads were machined, the surface around the edge of the first was roughed up with a file to increase surface area for the epoxy to hold to. This can be seen in Figure 15. Figure 15 - Rough finish on bulkhead It was then epoxied into Cylinder 2 and set to dry. Once dry, the first and second bulkheads where mounted to the five base springs via the bolt and washer assembly spoken about in the Payload section of this report. After the base springs were mounted, the wire rope isolators were prepared, small 1-inch by 1-inch squares of 0.1-inch aluminum sheet metal were cut to be epoxied to the 3D printed Cylinder 1 to prevent any failure in tension due to weaknesses in 3D printed material inch holes were drilled into the aluminum squares. The surface of each square was roughed up with a file and then soaked in Acetone to clean prior to epoxying. Cylinder 1 then had inch holes drilled into it at the same spacing as Cylinder 2 however starting 2 inches from the base of Cylinder 1 to allow for a maximum oscillation of 1-inch within Cylinder 2. Pins were cut using a hack saw from standard deck nails that happened to be the correct size as the thru holes in the wire rope isolators. Each wire rope isolator was epoxied to a 2-mm long pin and then epoxied to the aluminum square, as shown in Figure P a g e

37 Figure 16 - Wire rope isolator pin and aluminum square assembly After the epoxy cured on the wire rope isolators, the pins were inserted into the holes drilled into Cylinder 1, the aluminum plate and pin assembly was epoxied to the plastic Cylinder. After 3 hours, Cylinder 1, now attached to all 12 wire rope isolators was inserted into Cylinder 2. Epoxy was placed on all exposed pins and outer faces of the wire rope isolators, to adhere to the inner diameter of Cylinder 2. The pins were inserted into the holes in Cylinder 2 and set to dry. Prior to launch, two bolts were screwed in aft of the recovery bulkhead to secure the entire payload assembly in the rocket. The final assembly can be seen in Figure 17. Figure 17- Final Assembly 24 P a g e

38 Electronic Payload The electronics payload consists of the altimeter, the battery, the mount, and the ballast attach points. The mount consists of four components: the O-ring, base plate, vertical mounting plate, and the battery holder. All four-mount components are machined from Aluminum 6061 and were milled on a 3-axis CNC mill. Once milled, the base plate and vertical mounting plate were tig welded to form one assembly shown in Figure P a g e

39 Figure 18 - Altimeter mounting assembly The O-ring serves as a permanent mounting point for the base plate. The base plate attaches to the O-ring via 4 manually threaded holes. The O-ring was permanently fixed in the nosecone using Rocketpoxy. The attached O-ring can be seen in Figure P a g e

40 Figure 19 - Mounted O Ring Figure 20 shows the battery holder that was designed to attach to the assembly. The battery holder was designed as a separate piece because the original assembly was to hold the battery already. Mounting the battery under the altimeter would not allow the altimeter to accurately measure altitude. Due to the inaccuracy, a new battery holder was designed to securely attach on to the vertical mounting plate. Holes were made and tapped on the backside of the altimeter mount to be the attach points. 27 P a g e

41 Figure 20 - Battery Holder To allow adjustments based on the test flights actual performance compared to simulation, the team needed the ability to add ballasts to the launch vehicle. Weights will be mounted to the aft end of the base plate. Ballast mounts were designed in the base plate. Using the CNC mill, four ballast holes were cut and then tapped. Mounting pins were coated in epoxy and screwed into the holes. When the epoxy dried, remaining was the four mounting pins for ballasts to be attached, as shown in Figure P a g e

42 Figure 21 - Mounting Pins Clear plastic tubing was run from the bottom of the base plate to the base of the nosecone to allow the nosecone compartment to properly pressurize during vehicle flight. The opposite end of the tubing was attached to the outer wall of the nosecone shoulder using a PVC fitting. The PVC fitting was attached using Rocketpoxy. The mounting assembly with the attached altimeter is shown in Figure P a g e

43 Figure 22 - Altimeter Mounting Assembly Recovery The first recovery event is the deployment of a Fruity Chutes CFC-24 parachute at apogee. This 24-inch-diameter ripstop nylon parachute will serve as the launch vehicle s drogue parachute, resulting in an initial descent velocity of 76.5 ft/s. The second recovery event is the deployment of a Fruity Chutes IFC-96 parachute at 750 above ground level. This 96-inchdiameter ripstop nylon parachute serves as the launch vehicle s main parachute, resulting in a final descent velocity of 14.5 ft/s. Two 35 lengths of 1-inch tubular nylon are used as recovery harnesses to tether the three independent sections of the launch vehicle together. To secure the harnesses to rocket, a loop is 30 P a g e

44 stitched at the end of each harness using Kevlar thread. An additional loop is stitched into each harness 5 from one end, which serves as an attachment point for each harness s parachute. Attachment hardware consists of 5/16 steel U-bolts secured to the bulkheads with lock washers and steel backing plates to distribute loading during recovery events, as shown previously in Figure 9. The recovery events are controlled by two PerfectFlite StratoLogger CF altimeters. These altimeters utilize a pressure transducer to determine the altitude of the launch vehicle. The Stratologger CF is relatively simple, yet effective. It has the ability to fire two igniter signals: one at apogee with an adjustable delay time and the other at a fixed altitude. This configuration is ideal for dual-deployment. Using a software transfer kit, altitude and temperature data can be obtained for up to 16 stored flights. Separate 9-volt lithium-ion batteries are connected to the power terminals of each altimeter. The Stratologger CF also has dedicated terminals for connecting a power switch. Using these terminals, a rotary locking switch is connected and used to toggle power to each altimeter. QuickBurst QBECS igniters are connected to the drogue and main output terminals of each altimeter. These low-current igniters ensure reliable, complete ignition of the black powder ejection charges. Redundancy of the recovery system is achieved by utilizing two identical sets of components with completely separate electrical circuits. In this way, if either circuit were to be shorted accidentally or experience an altimeter malfunction, the other circuit would remain unaffected. In addition to the redundant circuitry, each igniter is inserted into its own separate ejection charge well with the appropriate amount of black powder. The result is two black powder explosions for 31 P a g e

45 each ejection event. To avoid over-pressurization of the parachute compartments, the ignition signals from the backup circuit are delayed using the altimeter s built-in software. The first signal for the drogue parachute is fired at apogee and the backup signal is fired 2 seconds later. The first signal for the main parachute is fired when the launch vehicle reaches an altitude of 750 feet and the backup signal is fired when it reaches 650 feet. In both scenarios, a successful ignition at the primary signal results in the backup ejection charge exploding harmlessly into the atmosphere. Conversely, if a main charge fails to ignite for any reason, the backup signal causes ignition and subsequent parachute ejection due to pressurization of the parachute compartment. A block diagram of the redundant recovery electrical systems is provided in Figure 23. Figure 23 - Block diagram of recovery system 32 P a g e

46 The scoring altimeter uses the 470 MHz frequency bands to transmit the GPS and live feed from the rocket. The GPS has an operational altitude limit of 50,000 meters. The scoring altimeter requires Watts to run during the flight. Project ACE recognized that interference from the scoring altimeter to the recovery system is possible. To ensure that the interference would not compromise the recovery system, all ejection tests were done with the scoring altimeter on and near the body of the rocket. Keeping the scoring altimeter near the body would allow any interference to affect the recovery system. In doing so, there was no noticeable interference or change to the recovery system. In addition to the ejection testing, all altitude tests with the drone had all three altimeters mounted in the same location. With all drone testing, there was no noticeable interference with the recovery system. 33 P a g e

47 Mission Performance Predictions Mission Performance Criteria The main mission performance objective for the team is to reach an altitude between 5,125 and 5,375 feet. The goal gives the team a range of 250 feet which is an accomplishable goal for a first-year team. Another goal for the launch transport a piece of fragile material and safely return it back to Earth after reducing kinetic energy to less than 75 pound-force. The altitude range and the fragile payload are but a few of the goals set forth by NASA and the team see the Requirements Compliance section for more. The goals were then measured through testing of the full-scale rocket. The team used three altimeters, one located in the nosecone that can measure acceleration, velocity, and altitude, and two located in the recovery bay measuring just the altitude. An accelerometer was used to measure the fragile material payload bay force reduction and the accelerometer in the nose cone is used to calculate the energy of the rocket as it lands back on Earth. Flight Simulations and Altitude Predictions The full-scale rocket was tested three times with three different configurations. Both ballast weight and quick link (in the recovery section) style was altered. The configurations can be seen in Table 3. The different configurations were simulated in OpenRocket using the conditions of the launch day to mimic actual conditions as closely as possible. 34 P a g e

48 Table 3 - Simulation Summary Different Launch Configurations Launch Day Simulated Simulation Quick Links Ballast (lb) Conditions Apogee (ft) 1a Baseline No Heavy 2.0 5,005 Flight 1 Yes Heavy 2.0 4,967 Flight 2 Yes Heavy 0.0 5,322 Flight 3 Yes Light 1.5 5,326 For a baseline, the rocket was simulated at standard temperature and pressure (70 degrees F and 1 atm) with no launch rail angle. Figure 24 shows the full-scale flight profile of the rocket under these conditions. The maximum altitude that was predicted was 5,005 feet. Ballast was still considered for the first flight based off of the baseline simulation because stability was a concern for the rocket. Another concern was the accuracy of the simulation. A few simulations before the recorded simulation, the apogee was around 5,600 ft which brought some concern for the believability of the software. Because of the high apogee, on the simulation before the baseline, the 2 lb of ballast was used for the baseline and the first flight. 35 P a g e

49 Altitude (ft) Time (s) Figure 24 - Full-Scale Simulation To make the other flight simulations more like the actual launch day, Figure 25 shows the launch day flight conditions. These conditions were applied to all three flights that were flown on the launch day for the full-scale launch. Figure 25 - Flight Simulation Input Data 36 P a g e

50 Altitude (ft) The launch rail sat in the base at an angle of -5 degrees because of the base not being fit for the University of Evansville s rail. For the first flight simulation, a configuration of large quick links (in the recovery section) and a 2-pound ballast was used for the flight. This configuration was used to give a baseline on how to modify the rocket for the following flights. For the first flight, a maximum actual apogee of 4967 feet was reached. The low apogee prompted Project ACE to remove ballast for the second flight. The low apogee could have been because of the weight from the ballast and the heavy quick links, or the launch angle. Figure 26 shows the flight profiles for all three different simulations for the different configurations Flight 1 Flight 2 Flight Time (s) Figure 26 - Simulated Flight Configurations 37 P a g e

51 The second flight simulation reached an altitude of 5,322 feet, which was within the team s goal of a range between 5,200 and 5,400 feet. However, when this configuration was launched, the rocket came off of the rail at about a 14-degree angle. The angle was because of overlooking the stability margin of the rocket coming off of the rail after changing the quick links. When the team removed the ballast from the nosecone, the remaining added weight of the rocket brought the stability of the rocket below 2 calipers off the launch rail. With the second flight showing the team that mass was needed in the nose cone, the third configuration of the smaller quick links and 1.5-pounds of ballast were used in the flight. The third flight simulation reached an altitude of 5,326 feet. The apogee is within the goal the team wished to achieve as a first-year team. The motor thrust curve is given in Figure 27. Appendix D has all component weights for the different flight configurations. 38 P a g e

52 Figure 27 - Anticipated Motor Thrust Curve from OpenRocket One of the NASA requirements was for the rocket to have a minimum rail exit velocity of 52 feet per second. The goal for the team was to have a rail exit velocity of at least 60 feet per second. The difference in being about 8 feet per second higher than the requirement was to mitigate the risk of falling below. Using the same simulations as for the previously described, Table 4 has the predicted rail exit velocities for each flight. Table 4 - Rail Exit Velocity on Different Flights Time to Exit Rail (s) Velocity at Rail Exit (feet per second) Simulation Simulation Simulation P a g e

53 Based on Table 4, the rail exit velocity is well above the requirement and the team goal. Because the first flight was the heaviest, the rail exit velocity was lower than the other two flights. Based on the data, the rail exit velocity for the rocket will be 66.9 feet per second. One of the other requirements for NASA was the Mach number being less than 1. The team again, set a goal of being well below the NASA requirement. The team goal was being below a Mach number of 0.6. Table 5 - Mach Number on Different Flights Mach Number Simulation Simulation Simulation Table 5 shows the predicted Mach Numbers for each of the full-scale flights. Based on the data in the table, the team goal was met being well below a Mach Number of 0.6. Again, because the first flight was the heaviest configuration, the Mach Number would be lower. Based on the flight simulations, the Mach Number of the rocket is Another factor that impacts altitude is the wind speed. Using the same flight conditions, five simulations were conducted with varying wind speeds from 0 to 20 miles per hour. Table 6 shows the change in the altitude at varying wind speeds. 40 P a g e

54 Table 6 Impact of Wind Speed on Altitude Third Flight Configuration Wind Speed (mph) Predicted Altitude (ft) 0 5, , , , ,297 The change in wind speeds plays an important part in the altitude of the rocket. There is a change in height of about 50 feet due to the variance in the wind speed. Based on the team s rocket design, a wind speed of 0 miles per hour would be preferred, while all the wind speeds allow the rocket to be in the range of the team s altitude goal. Validity Assessment An in-depth analysis comparing subscale flights to OpenRocket simulations can be seen in the CDR. It was determined that there was a 5% percent error between OpenRocket and actual flight data. A similarly thorough approach to measuring component weights and dimensions was used for the full scale simulations. A full list of component weights can be seen in Appendix D. Figure 28 through Figure 30 graphically shows the OpenRocket and actual flight data for the full scale flights. A description of the differences and error between the simulations and actual flight data can be seen in the Flight Analysis section. Table 10 (located in the Flight Analysis section) 41 P a g e

55 Altitude (ft) compares the predicted and actual apogees for all three flights. Lastly, in regards to Figure 30 and Flight 3, the main parachute deployed shortly after the drogue parachute. The early ejection accounts for the large discrepancy between the OpenRocket Simulation Data and the Actual Data. A further explanation of this can be found in the Flight Results section Actual Altitude (ft) OpenRocket Altitude (ft) Time (s) Figure 28 - Flight 1 Actual vs OpenRocket Data 42 P a g e

56 Altitude (ft) Altitude (ft) Actual Altitude (ft) OpenRocket Altitude (ft) Time (s) Figure 29 - Flight 2 Actual vs OpenRocket Data Actual Altitude (ft) OpenRocket Altitude (ft) Time (s) Figure 30 - Flight 3 Actual vs OpenRocket Data 43 P a g e

57 Pre-flight, two simulated vehicle factors were validated using empirical data. First, actual stability was measured in order to assess the validity of the OpenRocket stability values (further information can be seen in the Actual Stability Margin section). Second, the launch vehicle was weighed in full to validate final OpenRocket vehicle weight. Due to the low percent difference in predicted altitude for Flights 1 and 3 (approximately 1%), few changes to the predictive models were made post-flight. Coefficient of drag for the drogue and main parachutes were empirically determined from actual flight data (further information can be seen in the Coefficient of Drag section). Acceleration data from the Altus Telemega was used to empirically calculate the forces acting on the launch vehicle during parachute deployments (further information can be seen in the Testing section). Launch day conditions were also used to increase the validity of the flight simulations. See the Flight Analysis section for more detail on the accuracy of the simulation. Actual Stability Margin Stability is a metric (measured in calipers) used in rocketry to help determine a rocket s ability to maintain its speed and direction. This makes stability vital in designing and testing a rocket. When considering stability, NASA dictates a minimum stability of 2 cal to ensure that the rocket would be stable during flight to maintain constant velocity to the target altitude of one mile. In solving for the stability factor, the following equation was used: Stability = (C p C g ) D (1) In this equation, C p is the Center of Pressure, C g is the Center of Gravity, and D is the diameter of the body tube of the rocket. The diameter of the body tube is 5.5 inches, and the 44 P a g e

58 Center of Pressure is a value determined by simulation from Open Rocket. C g is a value that changes for each flight configuration and has to be determined separately each time the weight in the rocket is shifted (i.e. a ballast is added to the nosecone). The C g for each flight configuration was determined by hanging the rocket by a rope and balancing. C g for each flight configuration is located at the balance point. This data can be found in Table 7 (The C p and C g are both measured from the tip of the nosecone). Table 7 - Actual Stabilities C g (inches) C p (inches) Stability (cal) Flight Flight Flight The rocket was test launched three times and each time a static stability of above 2 was calculated, which was above our minimum objective. This shows that the rocket should be stable in good launch conditions. A sketch of the rocket showing the locations of the C p (in red) and C g s (in blue). Figure 31 - Actual Cp and Cg locations 45 P a g e

59 Kinetic Energy It is crucial to ensure that the kinetic energy of the launch vehicle is managed throughout flight, especially during the final descent. The launch vehicle reaches its maximum kinetic of 173,100 ft-lbf during the ascent, just before motor burnout. To reduce kinetic energy through the initial descent, the drogue parachute is deployed at apogee and achieves a predicted initial descent rate of 76.5 ft/s. This gives the heaviest section a kinetic energy of 1249 ft-lbf during the initial descent. Upon landing, the kinetic energy of any section of the launch vehicle cannot exceed 75 ft-lbf. The kinetic energy of each section at landing can be predicted using the mass of each section and the vehicle s final descent velocity as predicted by an OpenRocket simulation. These predicted values are shown in Table 8. The maximum kinetic energy upon landing is 41.0 ft-lbf, which is experienced by the nose cone and payload. Table 8 - Predicted kinetic energy of launch vehicle sections Section Kinetic Energy (ft-lbf) Nose Cone & Payload 41.0 Coupling Tube Booster 33.9 Drift In order to predict the drift distance of the launch vehicle at landing, five OpenRocket simulations were conducted for wind speeds of 0, 5, 10, 15, and 20 mph. For each simulation, the launch angle was set to zero degrees. The resulting drift distances are shown in Table 9. These 46 P a g e

60 results verify that the launch vehicle will meet the requirement of limiting drift distance to no more than 2500 ft even for high wind speeds. Table 9 - Predicted drift distance for selected wind speeds Wind Speed (mph) Lateral Distance (ft) Full Scale Flight Launch Day Conditions The full-scale launch took place at Elizabethtown, Kentucky on Saturday, February 18 th. It was overcast with a chance of rain throughout the day. It was average wind speeds between 4-8 mph with the cloud layer changing altitude during the day also. The temperature and wind speed changed throughout the day because of an incoming rain shower. The temperature fell as the day progressed, however, only the first flight temperature and wind speed was recorded. For the first launch, it was 59 degrees F at 1 atm pressure with high cloud layer altitude. For the second launch, the weather conditions changed. It started to rain, but not heavy enough for the launch to be cancelled. The rain was believed to have an effect on the rocket, but the result of the effect was uncertain at the time of the launch. The rain could have weighed down the rocket lowering the altitude, and the humidity could have also caused a change in the actual apogee. For the last launch, the rain had stopped, but the cloud layer altitude dropped. 47 P a g e

61 Flight Analysis Comparison with Prediction Three flights were completed for the FRR. A summary of these flights can be seen in Table 10. Flight 1 will be used for the OpenRocket prediction analysis. Flight 2 resulted in an unstable flight with a maximum tilt of 41. For this reason, Flight 2 was not used for the prediction analysis and will not be flown at competition. Flight 3 resulted in the main parachute being deployed prematurely near apogee. For this reason, Flight 3 was not used for the prediction analysis. Each of the flights will be discussed in further detail in the Flight Results subsection. Overall Weight (lb) Table 10 - Actual Flight vs Predicted Flights Summary Ballast (lb) Predicted Apogee (ft) Actual Apogee (ft) Percent Difference (%) Flight ,967 4, Flight ,322 4, Flight ,326 5, A plot of the actual and predicted altitudes for Flight 1 can be seen in Figure 28 on page 42. Graphically, it can be deduced that the actual and predicted flight were very similar. Unfortunately, differing time steps do not allow a direct percent error (Equation (2)) comparison between the actual and OpenRocket flights. To counteract this issue, a 6 part piecewise regression line was created based on the actual flight data. This regression line was then evaluated on the time step of the OpenRocket flight. Error between the best fit line and the actual 48 P a g e

62 Altitude (ft) flight data was calculated, as well as error between the best fit line and the OpenRocket flight data. % Error = theoretical experimental theoretical 100% (2) Actual Altitude (ft) OpenRocket Altitude (ft) Time (s) Figure 32 - Actual Altitude vs OpenRocket Altitude Actual altitude and predicted regression altitude are plotted on Figure 33. Figure 33 also displays the percent error between these altitudes. The percent error assumed the actual flight data as the accepted value and the regression data as theoretical. Percent error remained below 11% between 0.55 seconds and 124 seconds. This is the maximum domain that the regression may be used for when comparing with the OpenRocket data. 49 P a g e

63 Altitude (ft) Percent Error (%) % Regression Altitude (ft) Actual Altitude (ft) Percent Error 10% 8% % % % 0 0% Time (s) Figure 33 - Actual Data vs Regression OpenRocket altitude and predicted regression altitude can be seen graphically on Figure 34. Figure 34 also displays the percent error between these altitudes. The percent error assumed the OpenRocket data as the accepted value and the regression data as theoretical. Percent error remained below 11% between 0.55 seconds and 114 seconds. 50 P a g e

64 Altitude (ft) Percent Error (%) Regression Altitude (ft) Percent Error 12% 10% % % % % % Time (s) Figure 34 - OpenRocket Data vs Regression Although Figure 33 and Error! Reference source not found. show error, it should be remembered that this is not error between actual and OpenRocket data but rather error between these and the regression line. It can be concluded from Figure 33 and Error! Reference source not found. that the largest errors occur at liftoff, main parachute deployment, and low level turbulence. As this is consistent between both figures, it can be concluded that this large is error exists due to the regression used to bridge differing time steps. From this, Project ACE has decided to accept the OpenRocket simulations as a valid prediction method. Error The sources of error can be separated into 4 major types. First, there is the inherent error in the modeling software. Both OpenRocket and Rocksim have documented error within the program that does not allow for perfectly accurate predictions. To counter this, both programs 51 P a g e

65 are used, in order for each to validate the other. Error within the programs was discussed extensively in the PDR. Secondly, there is systematic error in inputs to the modeling software. For example, the accuracy of lengths is limited to the accuracy of the ruler used to measure them. Alternatively, some parameters could not be measured and were thus based on research. For instance, the surface roughness of carbon fiber was not measured, but was instead based on research. Third, there is random error. Similar to the fluctuation of a needle on a gage, there will be a certain variance in the apogee of the rocket from one flight to the next. Lastly, there is error in the best fit curve created to compare OpenRocket data to actual flight data. This is mentioned and described in the previous section. Coefficient of Drag The coefficient of drag is simulated by OpenRocket from liftoff until drogue deployment. A plot of this can be seen in Figure 35. At drogue deployment, the coefficient of drag is assumed to be equal to the manufacturer specified coefficient of drag of the drogue parachute. At the time of the main parachute deployment, the coefficient of drag is assumed to be equal to the manufacturer specified coefficient of drag of the main parachute. 52 P a g e

66 Altitude (ft) Coefficient of Drag Altitude (ft) Drag coefficient Time (s) Figure 35 Predicted Coefficient of Drag During Flight Coefficient of drag was calculated based on experimental values. The Atlus Telemega altimeter located in the nosecone of the launch vehicle records acceleration and velocity data. Post motor burnout, only drag and weight act on the launch vehicle. Using summation of forces, drag can be calculated using the following equation (where acceleration and gravity both act in the positive direction): D = m(a g) (3) Following the calculation of drag force, coefficient of drag can be calculated using the following equation: C D = D ρ v2 2 A c (4) 53 P a g e

67 These equations are valid post motor burnout and pre-drogue deployment. An average of acceleration and velocity were taken over this range. The experimental coefficient of drag was calculated to be Compared to the average OpenRocket coefficient of drag over the same range (0.449), this is a 12% difference. Flight Results The team launched three times with success on each of the launches. Table 10 shows the apogee results from each of the three flights. For the first flight, the team utilized a launch rail provided by the University of Louisville. The launch vehicle was equipped with 2 lb of ballast, and was mostly successful; the vehicle came straight off the launch rod, recovery events occurred at the correct times, and no damage was observed. However, the recorded apogee of 4913 ft was well under the team s minimum goal of 5200 ft. The first flight at the launch site showed the first simulation and the baseline simulation were correct on OpenRocket. The flight with the 2 lb of ballast was ran because of the worry that the OpenRocket simulation was incorrect. The reasoning behind the worry of the simulation not being correct was because a few simulations before the final, the apogee was around 5,600 ft. To mitigate any worry with the simulations being incorrect, the 1 st configuration was ran for a starting point of the ballast optimization and to double check the OpenRocket simulations. In order to increase the apogee of the second flight, the 2 lb of ballast were removed. The second launch was less successful with only the recovery being a success. The problems that occurred in this flight were the altitude being too low, the drift distance was too far, and the stability too low. The second flight reach an altitude of 4,795 feet, which was lower than the first flight. The reason for the low altitude was an unexpectedly low stability. When the ballast in the 54 P a g e

68 nosecone was removed from the rocket, the center of gravity was lowered toward the aft of the rocket lowering the stability. The rocket came off of the launch rail at an angle of 14 degrees from vertical and ended around 40 degrees from vertical a few hundred feet above the ground. The angles were acquired from the altimeter located in the nosecone. The angles led to a lower apogee because of the trajectory of the rocket traveled. Also, due to the angle that the rocket launched, it landed further way from the launch site and in a tree. No parachutes were torn or any part of the rocket harmed when it landed in the tree or when it was removed. The third configuration flown was 1.5 lb of ballast with the smaller diameter quick links mentioned in the Changes Made to Vehicle Criteria section. The last configuration shifted the center of mass further toward the bow, increasing the stability of the launch vehicle to fix the issues observed during the second flight. The third flight was the most successful with respect to the apogee achieved, however unexpected performance of the recovery system resulted in a large drift distance. The apogee of 5291 ft satisfied the team objective to reach within 200 feet of one mile. However, the main parachute deployed early, just after deployment of the drogue parachute. This resulted in a velocity of 15 ft/s for the entirety of the descent. The wind then carried the launch vehicle to just over one mile from the launch site. While the drift distance was greater than the acceptable maximum, the vehicle was able to be recovered without damage. The early deployment of the main parachute was likely due to over-packing of the ejection charges for the third flight; the scale available on-site was not as precise as the one used to measure the ejection charges for the first two flights which was prepared in advance and as a result a larger amount of black powder was used. The larger charge likely caused the bow body tube to separate at a high velocity, pulling the coupling tube out of the aft body tube prematurely. 55 P a g e

69 Despite the excessive drift distance of the third flight, the team has selected this configuration to be flown at competition. The ejection charges will be measured precisely prior to the competition to ensure that the recovery system performs as expected. The recorded apogee for this configuration should fulfill the team s goal. Payload Criteria After initial drop testing proved that neither welding nor epoxying the springs to the bullheads would suffice, the design was changed to a bolt and washer mounting assembly seen in Figure 36. Figure 36 - Final Design Assembly (new bolt and washer mounting) The design change used 30 washers and bolts threaded into both bulkheads to secure the bottom and top layers of each base spring. Each spring had 3 washers on either side allowing one to tighten or loosen one of the bolts to assure the spring was at a constant 90-degree angle to avoid buckling. Due to the addition of 30 bolts, the team repurposed the old bulkheads by adding 15 more threaded holes to each. The drawing of each bulkhead can be seen in Appendix A. The exploded view of the entire payload assembly can be seen in Figure 37 focusing on the base spring assembly. 56 P a g e

70 Figure 37 - Exploded view of payload assembly (annotation following) (1) Represents the U-bolt that is attached to the main parachute which screws into the recovery bulkhead (3). This is epoxied directly to the ID of the rocket s main body tube. The recovery side bulkhead for the payload (5) is attached to the recovery bulkhead (4) via two bolts shown in the figure as (2). (4) is a clear spacer to separate the two recovery side bulkheads. (6) shows the 30 washers used to hold the 5 base springs, labeled (8), in place by inserting them above the last two coils in each spring. (7) is the 30 bolts used to tighten the washers, (6), and base springs, (8), into place. (9) shows the bulkhead epoxied in Cylinder 2, (10). Finally, (11) shows Cylinder 1, a 3D printed canister mounted within Cylinder 2, (10) via the 12 CR1-400 wire rope isolators labeled (12). A second view of the exploded assembly showing how the washers and bolts attach the base springs to each bulkhead can be seen in Figure 38. Figure 38 - Exploded view base spring attachment 57 P a g e

71 Another design change that was implemented due to a failure during initial testing was the addition of thin aluminum squares epoxied to Cylinder,1 as well as pins used in all mounting points of the wire rope isolators. During initial tests, some of the epoxy failed due to shear stress, causing the wire rope isolators to break free from the ID of Cylinder 2. The solution applied was to drill holes into cylinder 2 and epoxy pins inserted in the thru hole of each end of the spring. Both cylinders had holes were drilled to add more strength and reduce total shear stress felt by the epoxy. This solved the epoxy s adhesive failure, but Cylinder 1 experience 2 cases of cohesive failure where the 3D printed plastic was ripped apart due to a weakness in tension. To combat this, thin.1-inch-thick aluminum squares 1x1 inch were epoxied to Cylinder 1 to spread the force over several layers of material. The wire rope isolator with epoxied pins and aluminum plate can be seen in Figure 39. Figure 39 - CR1-400 wire rope isolator pin and plate assembly (1) shows the 0.1-inch thick aluminum plate used to disperse the force along several layers of the 3D printed plastic of Cylinder 1. (2) shows the CR1-400 wire rope isolator and (3) shows epoxied pins in the thru holes of the isolator that would go on to be inserted into Cylinders 1 and 2. The pins were cut from a standard carpentry nail. The pins used in attaching the wire rope 58 P a g e

72 isolators not only add strength by transmitting some of the shear force into the pin, but make assembly easier by fitting the pins into holes in both Cylinders 1 and 2. Safety The University of Evansville s first and foremost priority throughout the duration of this project has been and will continue to be a focus on safety. This consideration and team-wide emphasis on safety has been paramount in this project, as it has allowed the UE s SLI team to stay on schedule and create a safe and successful launch vehicle. Throughout the duration of this project, in order to create the safest possible working and testing atmosphere, risks were identified and mitigations were developed before material handling, fabrication operations, or testing was completed. In addition to this, all team members have been, and will continue to be, educated on the risks associated with all areas of the project. This is significant because, education allows team members to fully understand the risks associated with the operations/items that team members are coming in contact with, and details the proper procedure to take in order mitigate these risks. In the following tables, various hazard and failure mode analyses of the launch vehicle will be considered in order to present possible risks associated with the project, and detail mitigation tactics and verification plans that will be used to alleviate these risks. In order to generate these continually updated tables, the team first began by brainstorming the possible risks associated with each individual section of the rocket from fabrication, to handling of materials, and launch operations. As the project progressed from the design phase to the fabrication phase, and ultimately to the testing phase, the team was able to further identify other unforeseen risks as 59 P a g e

73 well as develop and conduct verification tests in order to mitigate various risks. In the hazard and failure modes analysis tables, the impact and likelihood of each risk was assessed and quantified using the definitions provided in Table 11. Table 11 - Definitions for Hazard and Failure Mode Analyses Severity Definition 1-Catastropic 2-Critical 3-Marginal 4-Negligible Extreme reduction in safety; potential complete loss Substantial reduction to overall safety or functionality Minor reduction to overall safety or functionality Little to no reduction in overall safety of team members or component functionality Likelihood A-Frequent Definition Occurrence of the event is expected B-Probable Occurrence of event is likely, but not guaranteed C-Occasional Chance of occurrence is possible, but not significant D-Remote Minor change of occurrence E- Improbable Occurrence of event is extremely unlikely Following this categorization, mitigations and verification plans were proposed in order to decrease both the significance of the risk as well as the change of occurrence. Lastly, the risk was then reevaluated in order quantify the impact of the mitigations methods. Personnel Hazard Analysis A personnel hazard analysis was conducted to identify hazards, effects, likelihood of occurrence, and impact of individual factors associated with project. Safety practices and 60 P a g e

74 protocols were created to make team members aware of potential hazards, and reduce the chance of risk or injury during the course of the project. The personnel hazard analysis is summarized in Table 12 through Table P a g e

75 Table 12 - Personnel Hazard Analysis - Epoxy Risk/Hazard Epoxy Fumes Epoxy Contacting Skin Spill/Leak of Epoxy Epoxy Burning Through Container Root Cause/Effect Open containers of epoxy during fabrication operations leading to inhalation of toxic fumes, accidental ingestion, or contact with skin leading to potential for irritation or rash Mishandling of epoxy during application leading to skin irritation Mishandling of epoxy resulting in epoxy hardening on the working area, potentially ruining lab equipment or various parts of the launch vehicle Mishandling of epoxy leading to potential damage to user, lab, or equipment Severity/ Likelihood Mitigation and Control Post-Control Severity/Likelihood 4A Work in well ventilated spaces 4C 4A 4C 2E Individuals handing epoxy must be wearing Proper PPE, such as gloves, pants, and closetoed shoes when handling epoxy to prevent contact with the skin. In the event that epoxy does come in contact with the skin, wash it off at the sink Handle the epoxy carefully during mixing or transport. In the event that any epoxy does spill, wipe up the excess with a cloth and dispose of it properly, and clean the dirtied area. Never leave epoxy unattended. Monitor the heat of the epoxy as you mix it. If epoxy does get excessively hot, remove sample from lab and let it cool before disposing of it properly 4C 4D 2E 174 P a g e

76 Table 13 - Personnel Hazard Analysis - Launch Operations/Post-Launch Inspection Risk/Hazard Root Cause/Effect Severity/ Likelihood Mitigation and Control Post-Control Severity/Likelihood Debris In Team Member's Eye Particles flying through the air during fabrication operations leading to potential scrape or cut to user's eyes 2C Wear proper PPE, such as safety glasses during launch and fabrication. In the event that debris does enter the eye, the eyewash station will be used to cleanse the eye of the debris. 3D Sharp Edges on Fins and Nosecone Improper sanding or fabrication of fins and nose cone resulting in potential splinters or cuts to team members 4D Team member will be required to wear proper PPE, such as gloves, close-toed shoes, and pants during testing operations and inspection procedures to prevent direct contact with fragments of the rocket 4E Cracks or Chipping in Body Tube Improper fabrication operations or faulty components putting team members at risk for splinters or cuts when coming in contact with rocket 3C Team member will be required to wear proper PPE, such as gloves, during testing setup and inspection procedures to prevent direct contact with fragments of the rocket 4E 174 P a g e

77 Materials Experience Explosive Breaking When Opening for Inspection Failure of component durability or subsystem resulting in a range of possible injuries to team members from minor to severe depending on the intensity of the explosion 1E Team members will wait in designated safe launch zone until rocket is deemed safe for retrieval by RSO. Safety officer will retrieve rocket, wearing proper PPE and keep face directly out of line of launch vehicle. 2E Direct Contact With Hot Material Oversight or ignorance when approaching hot materials for handing, yielding to varying degree of burn to team members 2D Proper PPE, such as gloves or aprons, must be worn at all time when handling hot objects 4D Improper storage of flammable All flammable objects will be kept in proper components or inappropriate locations away from sparks and open flames. Materials Catching Fire fabrication operations/tool usage leading to potential severe injury 1D In the event of a small fire, a fire extinguisher will be used to put out the fire. In the event of 2E or burns to team members, a large fire, the team will evacuate the equipment, or work space building and the fire department will be called 64 P a g e

78 High exposure to black powder Black powder is stored in portable fireproof Black Powder Fumes when handling and preparing samples of this toxic gas can result in coughing, dizziness, and 2D case to keep away from fire and high temperatures. When handling substance recovery subsection lead will measure 3D fainting samples in well ventilated areas Per the motor preparation checklist, the motor Rocket will be transported from an offsite location to Propellant Improper transportation and the launch location in a protective, waterproof Comes In configuration of motor subsystem 2C casing. Upon installation, the propulsion team 2D Contact With leading to irritation and burns lead will prepare the motor according to Skin manufacturer specification while wearing proper clothing, shoes and PPE 65 P a g e

79 Table 14 - Personnel Hazard Analysis - Testing Risk/Hazard Root Cause/Effect Severity/ Likelihood Mitigation and Control Post-Control Severity/Likelihood Per malfunctioning electronics Unsafe working conditions or lack troubleshooting checklist, electronics of care when handling electronics subsection lead will inspect faulty instrument Electrical Shock leadings to electrocution resulting 3D for improper connection. Care has been take 3E in burns, significant injury, or to ensure nothing with exposed or fraying death wiring is being used in fabrication. Electronics will be stored in dry, secured area Inexperienced Test Personnel Improper handling of shop tools or machining operations leading to personal injury or destruction of equipment 3C Only authorized individuals have run tests. Multiple team members are present during testing to report and issue if one should occur 3D Fractured Particles During Testing Failure of various components leading to potential splinters or cuts to team members 3B Team member have been required to wear proper PPE during testing setup and inspection to prevent direct contact with fragments of the rocket 4E 174 P a g e

80 Table 15 - Personnel Hazard Analysis - Fabrication Risk/Hazard Root Cause/Effect Severity/ Likelihood Mitigation and Control Post-Control Severity/Likelihood Proper clothing, shoes, and PPE must be worn Allergic Handling of materials team at all time when handling materials. Allergies Reaction to Building member is allergic to resulting in an allergic reaction in the form of 2E of all team members are kept on file, and members allergic to a specific material will 2E Material skin irritation, rash, or swelling not work with that material while it is fabricated. Improper Heavy Machinery Usage Improper handling of shop tools or machining operations leading to personal injury or destruction of equipment 2C All team members have been trained on how to properly use shop equipment and have passed written and practical tests regarding proper handling and maintenance of shop equipment 2D 67 P a g e

81 All team members have been trained on how Improper Handheld Tool Usage Bruises, cuts or scrapes from mishandling of basic handheld shop tools such as hammer or saw 3C to properly use handheld tools. During fabrication operations, team members have a spotter to ensure proper safety procedures are followed and to monitor surroundings during 3D fabrication operation Improper Tool Storage Tool storage in improper location following fabrication operations leading, or usage by unauthorized individuals leading to damage to equipment or environment. 3C Tools are stored in proper locations to keep team members and work area clean and safe, and prolong life of tool. Periodic checks will be conducted by safety officer to ensure all materials are returned following construction and placed in their proper locations 3D Improper Use of Craft/Exacto Knife Cuts leading to injury as a result of unsafe precision cutting operations on fins or other pieces of the rocket body 2D During fabrication operations, team members have at least one spotter to ensure proper cutting procedures are being followed by cutting away from body. 2E 68 P a g e

82 Improper Work Attire Lack of education on proper clothing or inspection of shop workers leading to potential damage to body or attire 4D Proper clothing, shoes, and hair styles will be required in the lab to ensure safety for all team members. Safety officer will conduct periodic checks of fabrication work attire and PPE. 4E Tripping Hazards Cords or other materials lying on the floor could cause team members to trip, thus resulting in cuts, scrapes, bruises, or broken bones 4B Cords will be plugged in closest to the area in which their machine is being used. The work area will be kept tidy in order to prevent debris from accumulating on the floor 4D Lack of awareness to surroundings Ensure all needed materials are close by Overreaching leading to potential falls, cuts, or 4B before beginning fabrication to avoid 4C scrapes overextension. Keep proper footing/balance. 69 P a g e

83 Table 16 - Personnel Hazard Analysis - Education Engagement Outreach Events Risk/Hazard Root Cause/Effect Severity/ Likelihood Mitigation and Control Post-Control Severity/Likelihood Car Accident Not following driving rules and regulations resulting in a range of potential injuries to team members in the car, or others from minor to severe, and potential property damage 1E Seat belts are worn at all times by all members inside the vehicle. All individuals in the vehicle will also sign a waiver releasing the team of liability in the event of an accident. The driver must follow all federal driving laws including have a valid license and insurance 2E Child Using Tools Inappropriately Lack of oversight of individuals managing event or disobedient participants could lead to child experiencing a range of injuries depending of the tool being used at the event 2D Age appropriate tools will be given during educational outreach events. Strict supervision will be used to monitor all activates to ensure all children are safe and know what they are supposed to do. 3D Child Not Following Instructions Lack of oversight of individuals managing event or disobedient participants could lead to child could experience a range of injuries depending operation/event 3D All children will be closely monitored in order to ensure they are doing what they are supposed to. If they continue to be defiant, they will be removed from the activity. 4D 174 P a g e

84 Failure Modes and Effects Analysis In order to analyze the functionality and safety of the rocket and all of its components, a failure modes and effects analysis was created. In this analysis, presented in Table 17 through Table 23, verification plans, referencing various pre-launch checklists or data obtained from tests conducted on individual components, are stated in order to verify mitigation tactics to reduce risks. Then, post-control severity and likelihood was then reevaluated to see the impact that the mitigation tactics and verification checks had on the risk. 174 P a g e

85 Table 17 - Failure Modes and Effects Analysis - Design/Fabrication Risk/Hazard Root Cause/Effect Severity/ Likelihood Mitigation and Control Verification Plan Post-Control Severity/Likelihood Faulty components or inappropriate fabrication Care has been taken to ensure all In accordance with the final operations lead to failure of parts have been fabricated assembly checklist, each Cracking or rocket upon launch or testing according to specification. Parts subsection lead will inspect their Chipping of operation. Potential to splinter 2D will be stored in appropriate section of the rocket for any 3D Fabricated Parts and cause significant damage to containers and holders within compromises in structural other sections of the rocket or locked room to prevent integrity as a result of fabrication lead to failure of subsequent accidental damage. operations components Lack of Precision When Fabricating Parts Fabrications not completed according to specifications leading to potential inability to assemble components of rocket properly and have secure attachment, resulting in failure of rocket upon launch or testing. 2D Only trained individuals are allowed to operate any machinery during the fabrication and construction process. Other team members will verify work to ensure it meets standards set for in the design In accordance with the final assembly checklist, each subsection lead will inspect their section of the rocket to ensure all parts are fabricated to specified dimensions 3D 174 P a g e

86 Gaps Between Connecting Pieces Fabrications not completed according to specifications resulting in inability to assemble components of rocket properly and have secure attachment potentially leading to failure of rocket upon launch or testing operation. 2C All components of the rocket have been measure after fabrication in order to ensure they meet the dimensions specified in the design. The prelaunch safety checklist will be used to ensure team members visually inspect connections of components prior to launch In accordance with the final assembly checklist, each subsection lead will inspect their section of the rocket to ensure all parts are fabricated to specified dimensions. In the event that there are gaps between adjoining sections, the troubleshooting checklist will be followed to remedy the issue 2D Insufficient Epoxy Lack of attention to security of connection causing inability of components to hold together leading to separation and potential failure 3D Epoxy has been mixed in accordance with instructions in order to ensure a good adhesive mixture. Components will be tested prior to launch to ensure a secure, water-tight seal In accordance with the secure attachment inspection within the final assembly checklist, each subsection lead will inspect their section of the rocket to ensure attachment between adjoining subsections 3E 73 P a g e

87 Only trained individuals are Lack of knowledge by team allowed to operate any Wrong Equipment Usage for Fabrication Operation members during fabrication operations leading to potential damage to component(s) of rocket and individual harm. Also potential to generate flawed component that is not suitable 4D machinery during the fabrication and construction process. If a component is corrupted, fabrication will be done to salvage as much of the material as possible without All team members will pass a practical and written test on proper usage of shop equipment before they are allowed to use equipment. 4E for usage compromising safety of the launch vehicle and operations Improper storage or handling of Team members will operate in a Materials Catch Fire materials causing potential damage to component(s) of rocket and individual harm resulting in minor to major loss of equipment, workspace, or components, or compromise of structural integrity of launch 1D safe manner to prevent the start/spread of fire. In the event of a small fire, it will be extinguished using the fire extinguisher in the energy systems lab. For larger fires, 911 will be called and the team will Team members will be trained on how to properly use fire extinguisher in the event of a small fire. Safety officer will periodically test fire extinguisher to ensure it is fully functional. 2E vehicle retreat to a safe distance. 74 P a g e

88 Improper Storage of Materials/Equipment Lack of knowledge as to where to put supplies when fabrication operation is completed leading to potential damage to materials/equipment resulting in compromise of the structural integrity of various components, or inability to use equipment for further fabrication operations 4A The energy systems lab is cleaned after each work period. Checklists have been created to ensure that all materials and equipment being used are returned to their proper locations before everyone can leave. The parts checklists will be used to sign tools and materials in and out. Additionally, the safety officer will periodically check the supply cabinets to ensure all tools are returned and in their proper locations following fabrication operations 4B Parts checklists will be used to Degradation of Epoxy Oversight of connection security during inspection process or improper storage leading to lack of adhesion between parts resulting in separation and potential failure 3E Epoxy is stored in the in the lab at room temperature according to specification listed by the manufacturer. check out tubes of epoxy so that the safety officer has all supplies accounted for. Furthermore, each subsection of the rocket will be inspected via the final assembly checklist to ensure proper connection and adhesion 3E between adjoining sections. 75 P a g e

89 Table 18 - Failure Modes and Effects Analysis - Payload Risk/Hazard Root Cause/Effect Severity/ Likelihood Proposed Mitigation Verification Plan Post-Control Severity/Likelihood Failure to properly prepare ignition system according to Black powder is stored in a dry checklist leading to a range of area at room temperature in Premature Ignition Charge failure modes from minor damage to the payload system or components to catastrophic 2E accordance with manufacturer specifications. Testing has been done to ensure premature See ejection testing summary and results in project plan section of FRR 3E failure due to premature ignition does not lead to separation and parachute recovery failure deployment Testing and research has been Failure of Motor Faulty motor or improper storage leading to inability for rocket to ascend off launch pad. Potential damage to payload or other components upon misfire 3D completed in order to ensure the proper motors for each size of launch vehicle created is being used. Rocket motors will be kept in a dry area at room temperature in accordance with Motor has been tested via fullscale launch operations. For further detail see full scale testing section of FRR 3E manufacturer specifications. 174 P a g e

90 Improper storage of black Black powder is stored in a dry powder sample or compromised area at room temperature in Failure of Black Powered Charge sample resulting in an inability for launch vehicle to separate leading to potential catastrophic 1D accordance with manufacturer specifications. Testing has been done to ensure proper amounts See ejection testing summary and results in project plan section of FRR 1E damage to payload and rocket of black powder is used for failure of recovery system ignition. Ejection test was completed in Deployment of Black Powder Change Resulting in Damage to Payload Holding Container Inability to input proper amount of black powder into launch vehicle as determined by testing resulting in over pressurized capsule causing minor to catastrophic damage to payload, holding container, or springdamper system 1E Black powder is stored in a dry area at room temperature in accordance with manufacturer specifications. Testing is done to ensure proper amounts of black powder is used for ignition. order to determine the proper amount of black powder to be used to pressurize the launch vehicle and deploy the parachutes. Impact tests were also completed on the payload container to determine how much force is felt by the fragile material while within the 2E dampening system 77 P a g e

91 Failure to properly account for Tensile and impact tests were forces experienced by launch Testing has been conducted in completed on the payload Bending/Breaking of Spring-Damper System vehicle during flight or improper assembly resulting in compromise in the structural integrity of the spring-damper 3A order to minimize forces on payload and ensure fragile payloads of all types will be kept safe and secure during launch container in order to measure it's tensile and ability to dampen direct impact. For further detail, see MTS testing summary and 3D system and damage to the fragile and recovery operations results in project plan section of payload FRR Crack in Payload Holding Container Failure to properly account for forces experienced by launch vehicle during flight or improper inspection prior to launch resulting in compromise in the structural integrity of the springdamper system and damage to the fragile payload 3D An inspection has been completed in accordance with the pre-launch safety checklist in order to ensure the structural integrity of the payload systems is not in any way compromised prior to launch operations In accordance with the final assembly checklist, the payload container will be inspected for cracking or any other structural imperfections that could have been acquired during fabrication or transport prior to launch 3E Dampening material is unable to Testing was completed in which absorb impact and restrict the movement of the payload Inability to Keep Payload Static within Holding Container movement leading to potential minor to catastrophic damage to fragile payload and damage to other components of the launch 2D within its holding container is measured in order to ensure it does not experience collision with surrounding walls or See MTS testing summary and results in project plan section of FRR 3D vehicle. Potential failure to meet anything else that could cause mission objective fracture or damage. 78 P a g e

92 Payload Damage Upon Impact With Ground Upon Decent Inability to slow speed of launch vehicle during decent leading to damage to fragile payload resulting in repair, replacement, or failure to meet mission objective 2E Tests were conducted in order to validate that the materials used for the payload security container can withstand the forces experienced by the fragile material without damage to its structural integrity See MTS testing section of FRR for results on impact testing of payload container 2E 79 P a g e

93 Table 19 - Failure Modes and Effects Analysis - Payload Integration Risk/Hazard Root Cause/Effect Severity/ Likelihood Proposed Mitigation Verification Plan Post-Control Severity/Likelihood In accordance with the final Lack of Space in Body Tube for Payload Container Fabrication operations not completed according to specifications leading to inability to properly assembly launch vehicle resulting in removal of payload container and failure to meet mission objective 3C Throughout the fabrication process all complete parts have been measured and verified by the team lead in order to ensure they meet the proper dimensions laid out in the design. This will allow for proper fit and connect within the launch vehicle assembly checklist, each subsection lead will inspect their section of the rocket to ensure all parts are fabricated to specified dimensions. In the event that there are gaps between adjoining sections, the troubleshooting checklist will be followed to 3E remedy the issue Failure to properly inspect Payload Container Not Properly Mounted in Body Tube payload subsection system prior to launch causing potential damage to the payload or its housing container. Could compromise the structural integrity of various components or lead to failure of other 2D Prior to launch the pre-launch checklist will be used to verify that payload is mounted correctly in place and all connections are secure to ensure safe launch operations In accordance with launch procedures checklist, payload will be reviewed for flight readiness and proper mounting prior to launch operations 3D operations 174 P a g e

94 In accordance with the secure attachment inspection within the Weak Attachment Between Payload Container and Recovery System Failure to properly inspect connection between adjoining subsections leading to possible cracking or separation between the payload and recovery systems and inability to return fragile material 2D Prior to launch the pre-launch checklist will be used to verify that payload is mounted correctly in place and all connections are secure to ensure safe launch operations final assembly checklist, the connection of adjoining subsections will be checked by the safety officer to ensure proper connection. In the event that there are gaps between adjoining sections, the 3D troubleshooting checklist will be followed to remedy the issue Failure to fabricate payload All fully fabricated parts have Inability to Fit Given Payload Into Container container according to specifications given by NASA resulting in potential inability to meet mission objective of safely launching and returning a fragile 3E been measure and compared to design requirements in order to ensure they meet the proper dimensions, thus ensuring that the fragile payload fits within the Payload container has been design in accordance with the envelope of fragile material provided by NASA 3E material on our launch vehicle envelope of the container 81 P a g e

95 Inability to Fill Payload Container with Material that Dampens Force Felt by Payload Improper material used to dampen forces or unreliable impact testing data causing damage to the payload that could result in failure to meet mission objective 2C Testing has been done in order to measure the force felt by the payload during launch and landing operations See MTS testing section of FRR for results on filler material's ability to dampen impact 3D Inability to Fill Improper material used to Testing has been done in order Payload Container with Material that Restricts Payload Movement During restrict movement or unreliable impact testing data resulting in damage to the payload that could result in failure to meet mission 3C to measure the movement of the payload within the container in order to ensure it will not be damaged as a result of striking See MTS testing section of FRR for results on filler material's ability to restrict movement 4D Flight objective interior walls 82 P a g e

96 Table 20 - Failure Modes and Effects Analysis - Recovery System Risk/Hazard Root Cause/Effect Severity/ Likelihood Proposed Mitigation Verification Plan Post-Control Severity/Likelihood Testing have been done in order to verify the packing method and Various parachute packing Improper packing method give the student practice with methods have been researched leading to failure in parachute to packing the parachute into the and tested in order to determine Parachute is Not Packed Properly deploy properly resulting in launch vehicle experiencing 1E body tube. On launch day, the pre-launch checklist will be an optimal method. Tests have been conducted with various 3E more force than planned upon followed to ensure proper packing styles. For further detail, when landing packing of the parachute in see parachute deployment force accordance with standard testing subsection of FRR practices. Tear in Parachute Failure to properly inspect parachute prior to launch or faulty parachute resulting in launch vehicle descending at a faster rate than planned in an uncontrolled manner causing potential damage to components or total loss 2D Various tests have been completed in order to verify the strength of the parachute. The parachute will be inspected prior to launch using the pre-launch checklist in order to verify it does not have any tears, pulls, rips, or other imperfections that could result in failure In accordance with the recovery preparation checklist, all parachutes will be inspected prior to launch for tears, snags, or any other imperfection that could result in recovery failure 3E 174 P a g e

97 Tear in Shock Cord Failure to properly inspect shock cord prior to launch or faulty component resulting in launch vehicle descending at a faster rate than planned in an uncontrolled manner potentially damaging components or resulting in a total loss 2E Tests have been run in order to verify the strength of the shock cords. The cords will be inspected prior to launch using the pre-launch checklist in order to verify it does not have any tears, pulls, rips, or other imperfections that could result in failure In accordance with the recovery preparation checklist, all shock cords will be inspected prior to launch for tears, snags, or any other imperfection that could result in recovery failure 3E Testing has been done in order Incorrect shock cord for force to verify the strength of the Shock Cord Cannot Withstand Force of Parachute Deployment experienced during deployment causing separation of rocket into multiple pieces, some of which will not be attached to the parachute, causing damage and 1D connection between the shock cords and the main rocket body tube. The connection between these two will also be inspected prior to launch with the pre- See parachute deployment force testing subsection of FRR for details regarding shock cord strength and durability for various black powder charges 2D potential harm to by standards launch checklist in order to ensure a secure attachment 84 P a g e

98 Failure to properly pack parachute or use correct amount Drogue Parachute Deployment Failure of black powder for pressurization leading to uncontrollable decent until the opening of the main parachute resulting in the launch vehicle landing with a greater impact, thus causing damage to 1E Tests have been conducted in order to verify the precise amount of black powder that will need to be used to pressurize the parachute and allow for proper deployment. See parachute deployment testing and full-scale testing subsections of FRR for details regarding proper quantity of black powder for launch vehicle pressurization and drogue parachute deployment 3E components or endangering spectators Failure to properly pack Main Parachute Deployment Failure parachute or use correct amount of black powder for pressurization leading to decent at a quicker rate than expected and potential drift of the launch vehicle off course resulting in damage to the components or 1E Various tests have been completed in order to verify the precise amount of black powder that will need to be used to pressurize the parachute and allow for proper deployment. See parachute deployment testing and full-scale testing subsections of FRR for details regarding proper quantity of black powder for launch vehicle pressurization and main parachute deployment 3E endangerment of spectators. 85 P a g e

99 Wind Blows Rocket Off Course Launch of rocket in excessively windy conditions resulting in an inability for parachutes to open at proper altitudes or launch at excessive wind speeds leading to potential for rocket to become lost or components becoming damaged if landing occurs outside of the space given. 3D The rocket will only be launch in proper conditions, therefore minimizing the chance for wind gusts to blow the rocket off course. In the event that this does occur, the launch vehicle will be retrieved using the GPS tracking system in the altimeter UE SLI team in conjunction with NASA and local rocket clubs will monitor wind speeds and make launch related decisions accordingly 4E Testing has been done in order Parachute Deploys at Incorrect Time Incorrect packing of parachute, or faulty electronics leading to potential for uncontrollable decent, damage to components or compromises to structural integrity of the launch vehicle 2D to verify the precise amount of black powder that will need to be used to pressurize the parachute and allow for proper deployment at the correct time. The recovery system will also be tested prior to launch in accordance with the pre-launch See altimeter testing and fullscale testing subsections of FRR for details regarding proper quantity of black powder for launch vehicle pressurization and parachute deployment at proper times 3D checklist. Interference from the Scoring Altimeter Causes System Failure Operating on the same, or close to the same frequency resulting in a failure in the recovery system. 1D Testing the recovery system with the scoring altimeter on and nearby to determine if there is any interference with the system. Verify that all electronics are working according to the Pre- Launch checklist. 1E 86 P a g e

100 Table 21 - Failure Modes and Effects Analysis - Testing Risk/Hazard Root Cause/Effect Severity/ Likelihood Proposed Mitigation Verification Plan Post-Control Severity/Likelihood Lack of knowledge or All tests have been done by Inexperienced Student Running Tests experience by test personnel resulting in damage to lab equipment, facilities, components of the launch vehicle, or other team members 3D supervisors of each group. Multiple team members will be present during testing in order to ensure proper protocols are followed and safety precautions Subsection team leads will be presents thought the duration of their area's testing to ensure proper usage of lab equipment. 4D or students are taken Wind Tunnel Operation at Excessive Speeds Lack of knowledge or experience by test personnel resulting in damage to the testing apparatus or components of the rocket being tested in the wind tunnel 4D All tests have been done by supervisors of each group. Multiple team members will be present during testing in order to ensure proper protocols are followed and safety precautions are taken Aerodynamics subsection team leads will be presents thought the duration of wind tunnel testing to ensure proper usage of lab equipment in accordance with manufacturer specifications 4E Failure to properly clean and Proper PPE and eye protection inspect wind tunnel prior to must be worn at all times in the Safety officer will monitor Debris in the Wind Tunnel testing resulting in potential damage to testing apparatus, components of the rocket or 4B lab. In the event that debris does fly out of the wind tunnel during testing, multiple students will be testing to ensure proper PPE, such as gloves, ear and eye protection are being worn during 4D harm of individuals running the present to assist in the clean-up testing operations. test of the debris. 87 P a g e

101 Overuse of Wind Tunnel Excessive testing beyond apparatus capabilities causing damage to the testing apparatus that could result in inability to conduct future tests 3D Wind tunnel testing will be scheduled in advance with breaks in between tests to allow the engine to properly cool. The tunnel will also be inspected before testing to ensure proper conditions. Aerodynamics subsection team leads will be presents thought the duration of wind tunnel testing to ensure proper usage of lab equipment in accordance with manufacturer specifications 3E Improper storage of black Black Powder Fails To Ignite powder leading to no separation or deployment of parachute, thus creating a potential for catastrophic damage to launch 2D Secondary charges can be used in order to ensure that if one change fails another can engage to deploy the parachutes. See ejection testing subsection of FRR regarding proper black powder handing 3D vehicle or injury to spectators Failure to properly measure Excess of Black Powder Used in Testing correct amount of black powder for sample resulting in full separation of rocket leading to damage to various components and potential failure of other 2D Manufacturer specification are followed in order to determine how much black powder is need to pressurize the rocket based on its weight See ejection testing subsection of FRR detailing the amount of black powder to be used for complete and optimal separation 2E systems and debris 88 P a g e

102 Improper connection inspection Failure to Properly Secure Payload prior to launch leading to potential damage to the payload or its housing container. Could compromise the structural integrity of various components or lead to failure of other 3C Prior to testing the pre-launch checklist were used to verify that payload is mounted correctly in place and all connections are secure to ensure safe launch operations In accordance with launch procedures checklist, the payload will be inspected prior to launch for secure attachment and flight readiness 3D operations Multiple tests have been Testing have been done in order conducted in order to verify the to verify the packing method and Incorrect packing procedure or packing method and give the give the student practice with Parachute is Not Packed Properly for Testing method used resulting in failure in parachute to deploy at proper altitude resulting in launch vehicle experiencing more force 1C student practice with packing the parachute into the body tube. On launch day, the pre-launch checklist will be followed to packing the parachute into the body tube. On launch day, the pre-launch checklist will be followed to ensure proper 3D than planned upon landing ensure proper packing of the packing of the parachute in parachute in accordance with accordance with standard standard practices. practices. 89 P a g e

103 Table 22 - Failure Modes and Effects Analysis - Launch Support Equipment Risk/Hazard Root Cause/Effect Severity/ Likelihood Proposed Mitigation Verification Plan Post-Control Severity/Likelihood Faulty component(s) or failure to In accordance with set up on account for various changes to Prior to launch operations, the launch pad checklist, the launch rocket made between launches guide rails will be inspected by pad and guide rail will be Instability in Guide Rail can cause launch vehicle to deviate from its deal path 2C the lead safety officer in accordance with the pre-launch inspected for structural flaws or bowing that could lead to 3D potentially leading to checklist in order to ensure safe instability in launch and endangerment of spectators or operations deviation of the rocket from its damage to components ideal flight path Lack of care when handing The launch vehicle will be In accordance with final launch vehicle or components transported in its specially made assembly checklist, prior to Improper Transport of Launch Vehicle resulting in potential damage to launch vehicle and/or compromise of structural 2C container which will provide support for all fragile areas of the rocket, while protecting it launch, all subsection will be inspected by their team lead for cracks, chipping, or other 3D integrity of individual from slipping, vibrations or other structural flaws that could have components potentially damaging impacts. been acquired during transport 174 P a g e

104 The launch vehicle is stored in a specific container in a locked In accordance with the parts Lack of care and safety when room during fabrication and in checklist, the safety officer will storing launch vehicle or between tests. Only team leads periodically inspect storage Improper Storage of Launch Vehicle components leading to damage to launch vehicle and potential 2D and safety officers have access to the room to prevent the rocket cabinets as well as launch vehicle holders to ensure all 3D compromise of structural from being mishandled. The supplies have been returned integrity of components room is kept at room following use, and are being temperature to not adversely stored in their proper place affect any components The rocket motors will be Improper Transport of Rocket Motor Handling of motor not in accordance with specifications leading to potential damage to payload or other essential components upon launch 3C transported in a fireproof case that will prevent moisture for getting into the motor. The case will also protect the motors against slipping, vibrations, and other potentially damaging In according with motor preparation checklist, the rocket motor will be stored off site and will be transported to the launch location in a protective, waterproof case 3E impacts. 91 P a g e

105 The rocket motors are stored in a Improper Storage of Rocket Motor Potential damage to payload or other essential components upon launch due to improper placement of motor in potentially compromising locations 2D fireproof case. This case is kept in a locked cabinet in order to prevent from other team members handing the motors. This cabinet will remain at room temperature and dry in order to not allow heat or moisture In according with motor preparation checklist, the rocket motor will be stored off site and will be transported to the launch location in a protective, waterproof case 3E adversely affect the motor. In accordance with motor Improper Handling of Rocket on Launch Pad Handling of launch vehicle not in accordance with guidelines listed on pre-launch checklist leading to endangerment of spectators, and minor to catastrophic failure of the rocket and its subsystems 2D Only trained and essential team members will handle the rocket during launch operations. Prelaunch safety checklists will be used in order to ensure everting is safe for launch preparation checklist, the leader of the propulsion subsection will retrieve the rocket motor from its protective, waterproof casing, and will ready the motor for ignition after all subsequent subsection inspections have been 2E completed 92 P a g e

106 In accordance with the set up on Instability of Launch Pad Faulty component or failure to properly inspect launch pad prior to flight leading to potential for rocket to deviate from ideal flight path endangering spectators and causing failure of components or drift 2B Prior to launch operations, the launch pad will be inspected by the lead safety officer in accordance with the pre-launch checklist in order to ensure safe operations launch pad checklist, prior to launch operations, the launch pad will be retrieved from NASA will be inspected by the lead safety officer in conjunction with the RSO for any structural flaws that could lead to 2C instability in launch operations Faulty Ignitor Clips Improper handling or storage of component causing rocket to be unable to ascend off of launch pad 3C Prior to launch operations, the ignitor clips will be inspected by the lead safety officer in accordance with the pre-launch checklist in order to ensure safe operations and successful launch In accordance with ignition checklist, ignitor clips will be inspected prior to attachment to the launch vehicle by propulsion subsection lead. 3D 93 P a g e

107 Table 23 - Failure Modes and Effects Analysis - Launch Operations Risk/Hazard Root Cause/Effect Severity/ Likelihood Proposed Mitigation Verification Plan Post-Control Severity/Likelihood Improper storage of launch Prior to launch, the body tube vehicle or transportation leading will be thoroughly inspected for In accordance with the launch Cracking in Main Body Tube to compromise in the structural integrity of the rocket leading to potential damage to other 2D cracking, splintering, or fatiguing in according with the procedures listed in the pre- procedures checklist, the main body tube will be inspected by the safety officer prior to launch 3D components or failure of other launch checklist in order to for any structural imperfections subsystems ensure safe launch operations Gaps Between Connecting Pieces Failure to fabricate subsections according to specifications yielding an inability to assemble components of rocket properly with secure attachment potentially leading to failure of rocket upon launch or testing operation. 2C All components of the rocket were measured after fabrication in order to ensure they met the dimensions specified in the design. The pre-launch safety checklist will be used to ensure team members visually inspect connections of components prior to launch All components of the rocket have been measure after fabrication in order to ensure they meet the dimensions specified in the design. The prelaunch safety checklist will be used to ensure team members visually inspect connections of components prior to launch 3C 174 P a g e

108 Test have been run in open areas to ensure no overhanging trees, roofs, or other items could Failure for parachutes to deploy impede the flight operations. Collision with Object in Sky (Tree, Bird, Etc.) at proper altitude resulting in damage to launch vehicle and compromise in structural integrity of impacted 2C Furthermore, subsequent tests have been completed in order to determine the strength of the body tube and nosecone so that See scale model testing subsection of FRR report regarding impact testing of launch vehicle 2D components in the event the launch vehicle does strike a bird it can withstand impact and return safely Failure of team members to account for ballast adjustments In accordance with post-flight Instability During Flight made prior to launch resulting in change of center of gravity and leading to inability of the rocket 1B Maintain safe distance from launch pad inspection, all team members will wait until rocket has landed in a safe location before leaving 2C to maintain its projected flight safe launch zone path 95 P a g e

109 Altimeter or Other Electronics in Avionics Bay Malfunction/Fall Off Failure to properly inspect security of attachment between electronics, or test functionality prior to launch causing potential short circuiting or harm to spectators below 3B Verify all electronics work properly before launch and are firmly attached to the rocket In accordance with launch procedures checklist, all electronics will be tested for functionality prior to launch operations. For further detail on electronics testing, see altimeter testing section of FRR 3C In accordance with launch Coupler Excessively Tight Parts not fabricated according to specifications resulting in potential failure of parachute to deploy leading to damage to rocket 2D Run multiple tests to ensure proper amounts of black powder is used to allow rocket to separate procedures checklist, the ability of adjoining sections to separate will be tested. For further detail on proper ejection charges for separation, see ejection testing 3D subsection of FRR 96 P a g e

110 Environmental Considerations Additionally, when considering the safety and impact of the rocket, considerations must be given to how the vehicle will impact the environment, and how the environment will impact the vehicle. These considerations are represented below in the environmental hazard analysis, shown in Table 24. Table 24 - Environmental Consideration Hazard Analysis Risk/Hazard Root Cause/ Effect Severity/ Likelihood Mitigation and Control Vehicle Effects on Environment Post-Control Severity/Likelihood Fumes released during Epoxy Fumes construction resulting in hazardous working conditions created for team members 4A Work in well ventilated spaces and dispose of waste properly 4D as a result of toxic air Failure to follow proper disposal Let epoxy fully Epoxy Not protocols leading cure before Disposed of to potential fire 4C disposal in order 4E Properly hazard and damage to prevent fire to lab or hazard equipment 174 P a g e

111 Fabrication operations Wear mask when producing small sanding to avoid dust particles from inhaling dust sanding or particles and try Dust Particles machining 4A to contain dust 4D operations are when sanding released into the opposed to freely environment which releasing it into can result in surrounding air. breathing problems Failure to properly secure motor, Rocket Motor Ignition therefore, upon ignition, when motor reaches high temperatures and hot exhaust is released, the motor could become displaced or burn the areas where the 2D Place flame resistant material beneath the launch pad to avoid burning the immediate surroundings or starting a fire 3D rocket is launched or lands 98 P a g e

112 Ensure fully Component failure functioning leading to parachutes before fragments of the launch via pre- rocket breaking off launch recovery during flight or preparation upon landing checklist and impact and check to make Debris from Rocket becoming irretrievable, 3D sure all components of 3E leading to minor the rocket and environmental payload are harm due to accounted for inability to upon return in decompose and accordance with toxicity of post-flight component inspection checklist 99 P a g e

113 Water Environmental Effects on Vehicle Improper storage or launching in unfit conditions can lead to water exposure, which Avoid launching can cause rocket in wet malfunctioning of conditions and electronics within store rocket in 2E the avionics bay, proper stand in a or damage to the dry area for body of the rocket, storage and which will be transport constructed out of Blue Tube that is not 100% water resistant 3E Launch into excessive wind speeds leading to deviation from Avoid launching launch vehicle's rocket on days of Wind ideal flight path 3C high speed winds 3E thus leading to or unpredictable, damage to the strong wind gusts rocket and potential harm to spectators 100 P a g e

114 Improper storage of components resulting in potential corrosion Store rocket in a and weakening of dry area to avoid Humidity/Moisture various materials used to construct 3D moisture entering the rocket over 4E the rocket. This time via humid moisture can also air negatively impact on-board electronics Launch during times of low cloud coverage resulting Avoid launching in inability to track rocket on days Visibility the rocket thus 4C with low cloud 4E leading to debris coverage and not being retrieved poor visibility and damaging the environment 101 P a g e

115 General Risk Assessment Finally, a general risk assessment, shown in Table 25, was conducted in order to account for various extraneous risks not accounted for in previous sections, such as time, resources, the budget, scope, and functionality. Table 25 - General Risk Assessment Risk/Hazard Root Cause/ Effect Severity/ Likelihood Mitigation and Control Post-Control Severity/Likelihood Being a first year Limited Resources team with a small budget could lead to a lack of quality design or fabrication material and to failure to meet mission objective or overall poor 2C The team has work with faculty members as well as local rocketry club members in order to gain a better understanding of rocketry and develop a functional rocket. 2C performance 102 P a g e

116 Tight or Minimal Budget Being a first year team with minimal established funding, the team could be forced to use parts that are not optimal, or be unable to replace parts of the rocket that are broken during testing 3A The team and its adult educators have applied for and been given grants in order to fund parts of the project. Additionally, the team has held fundraisers to provide the team with a flexible budget beyond the normal amount of money allotted to the project by the school 3B Inability to Mismanagement of Time manage project on a weekly basis could potentially lead to major delays resulting in the quality of work lacking, or the rocket not being completed by 1E Team members have and will continue to fill out weekly time cards and log their hours in the task breakdown in order to ensure everyone remains on schedule 2E competition 103 P a g e

117 Inability to budget There has been and will time properly continue to be constant could leave the communication Underestimation project running amongst all team of Scope of behind schedule 2E members and with 3E Work and various facets NASA project leads to of the rocket not ensure the scope of being completed in work is clear and a quality manner everyone stays on task Increase in Safety Regulations Failure to meet proper FAA and NASA safety regulations could lead to team to be forced to add material to the rocket in order to increase safety, which will result in an increase in expenses 2D The team has designed and downselected with safety as the foremost priority, and will clearly identify all safety measures before all operations so that additional, last-minute safety measures do not have to be taken that will inflate the budget. 2E 104 P a g e

118 Launch Operations Procedures Parts Checklist In order to ensure safe and uninterrupted transportation of components and launch day procedures, parts checklists were developed for each of the following subsections: propulsion, aerodynamics, main payload, avionics bay/electronics payload, recovery, and safety and education, as well as a miscellaneous checklist to account for various extraneous items that the team will need for launch day operations. These checklists, which can be seen below in Table 26 through Table 32, were developed by each subsections respective team lead in conjunction with the safety officer to ensure all vital parts of the launch vehicle, as well as supporting materials, are accounted for and available for use on launch day. Table 26 - Parts Checklist - Propulsion Initial Part Quantity Liner for Motor Case 1 Motor Case 1 Aft Closure 1 Bow Closure 1 Forward Seal Disc 1 Reload Kit 1 Grains 3 Ignitor 2 Retention System Cap P a g e

119 Initial Part Quantity Water - Rags 3 Pocket Knife 1 Flat head screw Driver 1 Super Lube Synthetic 2 Grease Wire Strippers 1 Box Cutter 1 Table 27 - Parts Checklist - Aerodynamics Initial Part Quantity 3/36" Hex Key 1 3/16" in Hex Bolts 6 Nose Cone 1 Bow Body Tube 1 Aft Body Tube 1 Body Tube Holders 2 Spare Fin 1 JB Weld Tube 2 Extra Rail Buttons 4 Launch Rail P a g e

120 Table 28 - Parts Checklist - Main Payload Initial Part Quantity CR1-400 Wire Rope 12 Isolators 5.36" Blue Tube (Cylinder 1 2) Base Springs (#866) 5 Spacer (clear acrylic) 1 Recovery Bolts 3/8" x 1.25" 2 Length Spring Fastening Bolts For 30 3/8" x 16 Bolt, 1/4" Height Spring Fastening Washers 30 Bulkheads 2 Pins inch diameter 24 Aluminum Squares 1x1x inches 3D Printed Cylinder 1 (Cylinder 1) 3D Printed Cap P a g e

121 Table 29 - Parts Checklist Electronics Payload/Avionics Bay Initial Part Quantity Atlus Metrum TeleMega 1 Starter Pack 1 Arrow Yagi Antenna 1 SMA to BNC Adapter /16" O-Ring Bolts /8" Altimeter Bolts 4 O-Ring 1 1" Long, 0.25x40" Studs for 4 Ballast 108 P a g e

122 Table 30 - Parts Checklist - Recovery Initial Part Quantity Coupling Tube 1 Electronics Sled 1 Ejection Charge Igniter 8 Plastic Bags 4 Flat Washers 4 Lock Washers 2 Wing Nuts 2 1/4" Hex Nut 1 Shear Pins 12 1/4" Quick Links 6 35' Recovery Harness 2 Nomex Sleeves 2 Nomex Squares 2 24" Drogue Parachute 1 96" Main Parachute 1 Roll of Masking Tape 2 Black Powder Assurance Recovery Fiber Sheets P a g e

123 Table 31 - Parts Checklist - Safety and Education Initial Part Quantity First Aid Kit 1 Fire Blanket 1 Fire Extinguisher 1 Safety Glasses 15 Ear Plugs 15 Dust Mask 5 Table 32 - Parts Checklist - Miscellaneous Initial Part Quantity Folding Table 1 Chairs 4 Quick Dry Epoxy Tub P a g e

124 Final Assembly Checklist Below, in Table 33 through Table 40, are final assembly checklists for each subsection that were used for full-scale rocket assembly prior to launch to ensure safe and successful operations. For each checklist, the leader of the subsection is required complete each check-off point, in the order that they appear on the list, and then present the list to the safety officer for approval and sign-off. After this, the next checklist can be completed. It is important to note that each checklist is to be completed one at a time, in the order that they appear in this document, and not in parallel with other checklists currently in progress. In the event that any point on the checklist cannot be completed, the subsection team lead should immediately notify the safety officer so that the problem can be dealt with according to the procedures listed in the troubleshooting tables (Table 46 through Table 49). After all pre-launch checklists and inspections have been completed and approved by the safety officer, launch operations may commence. Table 33 - Final Assembly Checklist - General Set Up Initial Check-Off Point Set up table for launch vehicle preparation and pre-launch inspection Equip all personnel handling the launch vehicle with proper PPE equipment Inspect all members for safety glasses, gloves, and proper attire before handling any launch vehicle-related supplies Unpack all supplies and boxes from the truck Separate supplies by subsection Remove launch vehicle from transport case and transport to housing on inspection table 111 P a g e

125 Table 34 - Final Assembly Checklist - Comprehensive Structural Inspection Initial Check-Off Point Visually inspect body tube for cracks, bumps, abrasions or any other imperfections that could have been acquired during transport that could adversely affect the flight of the rocket Physically inspect rocket tube for structural integrity and flight readiness Inspect fins for any structural imperfections or bowing that could have been acquired during transport Physically inspect nosecone for cracks, chipping, or any other damage that could have been acquired during transport and handling Examine thrust plate and couplers for solid connection and structural integrity Table 35 - Final Assembly Checklist - Electronics Initial Check-Off Point Inspect avionics bay for flaws or damage to ensure nothing was broken or disconnected during transport Ensure proper connection of all electrical wires by inspection and comparison to wiring diagrams Test avionics for proper functioning Assemble avionics bay and check for proper connection to shock cord Test GPS tracking device and altimeter to ensure proper functioning Secure avionics bay using proper fasteners 112 P a g e

126 Table 36 - Final Assembly Checklist - Payload Initial Check-Off Point Examine payload housing container for any structural imperfection that could have been acquired during transport Inspect wire rope isolators for fraying or fatigue Visually examine springs on payload housing container for structural integrity Ensure proper filling of dampening material to protect payload Check for secure connection between fragile material protection apparatus and recovery section Table 37 - Final Assembly Checklist - Recovery System Initial Check-Off Point Inspect drogue parachute and shock cord for any imperfections or tears that could lead to error in recovery operations Examine connection between drogue parachute shock cord and main body section Examine connection between drogue parachute shock cord and drogue parachute Fold and pack the drogue parachute 113 P a g e

127 Initial Check-Off Point Wind excess drogue parachute shock cord to ensure proper deployment of drogue parachute Inspect main parachute and shock cord for any imperfections or tears that could lead to error in recovery operations Examine connection between main parachute shock cord to main body section Examine connection between main parachute shock cord and main parachute Fold and pack the main parachute Wind excess main parachute shock cord to ensure proper deployment of main parachute 114 P a g e

128 Table 38 - Final Assembly Checklist - Motor/Ejection System Preparation Initial Check-Off Point Inspect individuals preparing motor for proper PPE, including glasses, gloves, and mask Remove black power container from storage case Check black powder to ensure no moisture has compromised the sample Measure and pour 2 grams of black powder into charge cup to be used for drogue parachute Measure and pour 3 grams of black powder into charge cup to be used for main parachute Inspect motor casing for any structural imperfection acquired during transport Remove motor from storage container Examine motor and casing to ensure it is not wet or containing any moisture that could cause misfire or deviation from ideal flight path Assemble motor following manufacturer specifications Install motor into launch vehicle 115 P a g e

129 Table 39 - Final Assembly Checklist - Secure Attachment Inspection Initial Check-Off Point Check for secure attachment between motor and casing Examine nosecone for level and secure attachment with main body tube Inspect electronics bay within nose cone for proper fastening Inspect for proper connection between nosecone and payload bay Check for secure attachment between main payload and recovery system Inspect all exterior connections and assemblies on the rocket for proper fitting Table 40 - Final Assembly Checklist - Launch Pad/Pre-Launch Inspection Initial Check-Off Point Transport launch vehicle to Range Safety Officer for inspection Continuity test igniter clips for proper functioning with launch controller Inspect launch rail for bowing or imperfection that could cause the rocket to launch in an unplanned direction Connect the ignitor clips to the motor ignitor 116 P a g e

130 Motor Preparation In order to prepare the motor for ignition and launch operations, the following checklist, shown below in Table 41, was used. Table 41 - Motor Preparation Checklist Initial Check-Off Point Remove motor from protective, waterproof casing Assemble motor according to manufacturer specifications Remove the top of the screw on the retention system Place motor into inner tube with the nozzle facing the rear of the rocket in the open-air Examine placement of motor in inner tube to ensure secure fit Screw top of retention system back into place Place cap around nozzle to ensure a moisture does not enter the grains Motor is ready for ignition 117 P a g e

131 Recovery Preparation To prepare the recovery system for launch operations the recovery preparation list, displayed below in Table 42, was used. Table 42 - Recovery Preparation Checklist Initial Check-Off Point Test each battery with a multimeter to ensure that it is fully charged to 9 volts Reconnect each battery to its respective altimeter Insert the mounting sled into the coupling tube by sliding it over the threaded steel rods Connect mating female molex plugs with their male counterparts from the altimeters Electrical connections for the drogue and main ejection charges are established Attach aluminum bulkhead with lock washer and wing nuts Assemble the coupling tube Open end of coupling tube is now sealed Measure two 2.00 g black powder samples to be used for the drogue charges Place sample into small plastic bag with an ignitor Measure two 3.00 g black powder samples to be used for the main charges Place sample into small plastic back with an ignitor Twist each bag to compress the black powder around the tip of the ignitor Insert each ejection charge into ejection well Insert foam insulating material to hold each charge in place 118 P a g e

132 Initial Check-Off Point Seal each ejection well using masking tape Strip electrical leads Clamp electrical leads to terminal block Attach recovery harnesses Secure quick links on the end of each harness to U-bolts on the body and coupling tubes Wrap each harness in a spiral form Insert the wrapped harness into the body tube Wrap main parachute in Nomex flameproof fabric Insert main parachute into launch vehicle Wrap drogue parachute in Nomex flameproof fabric Insert drogue parachute into launch vehicle Insert the coupling tube into the aft body tube Secure the coupling tube and aft body tube using two nylon shear pins Fit aft body tube onto top of the coupling tube Secure the aft body tube and coupling tube using two nylon shear pins Activate altimeter alarming switches through exterior holes in the coupling tube 119 P a g e

133 Setup on Launch Pad After all subsections of the rocket had been properly configured, Table 43 was used in order to ensure proper safety procedures were followed when transporting the launch vehicle to the launch pad and when preparing the rocket for launch operations. Table 43 - Launch Pad Configuration Checklist Initial Check-Off Point Obtain launch pad for official competition from NASA Set launch box down at safe viewing distance Inspect the launch rail for any structural flaws that could cause the rocket to deviate from its ideal course of travel Lower launch rail height for safe rocket insertion Transport launch vehicle to launch pad with approved team members Place the launch vehicle on the launch rail Insert launch rail onto base of launch pad Secure launch rail to base of launch pad with two threaded bolts Adjust launch pad to vertical setting using the design feature on the base of the launch pad All non-level two members retreat to safe launch zone Complete ignitor installation checklist Arm rocket for launch Remaining members retreat to safe launch zone 120 P a g e

134 Ignitor Installation After the launch vehicle was properly configured on the launch pad and non-level two members of the team had retreated to the safe viewing area, the ignitor was installed in accordance with the checklist in Table 44.Table 44 - Ignitor Installation Checklist Table 44 - Ignitor Installation Checklist Initial Check-Off Point Strip ignitor wires 2 inches to allow for more surface contact with the composite for ignition Remove paper around the end of the ignitor from the composite Insert ignitor into motor Inspect ignitor to ensure entire ignitor is within the grains of the motor Pinch ignitor wires where end of the wires reach the end of the motor Remove pinched wire from the motor Measure pinched wire length Check to ensure that pinched wire length is the matches up with the length of grains in the motor Replace measured wire back into motor Attach stripped wires to ignition system Wrap stripped part of the wires around the system to allow for proper surface contact Inspect continuity of system Connect ignitor leads to launch controller 121 P a g e

135 Initial Check-Off Point Ignitor and ignition system is set-up and ready for launch 122 P a g e

136 Launch Procedures Following the installation of the ignitor, the rocket was armed and ready for launch. The launch procedures checklist, below in Table 45, contains all of the necessary checkpoints that must be met in order to ensure a safe and successful launch. To ensure the safety of all team members as well as spectators, equipment, and facilities, all check-off points listed in the final assembly checklist and launch procedures checklist must be initialed by subsection leaders in order for launch operations to commence. Table 45 - Launch Procedures Checklist Initial Check-Off Point Ensure a safe working area before transporting rocket to the launch pad Check the safety and readiness of team members and bystanders by ensuring proper PPE and safety glasses are worn by all individuals transporting the rocket Carefully transport rocket to launch pad Visually inspect the rocket main body tube for any structural imperfections Visually inspect the fins for any structural imperfections Inspect launch vehicle for proper connections between all sections of the rocket Test nosecone and body tube's ability to separate Examine main body tube for flight readiness Inspect fins for flight readiness Inspect nosecone for flight readiness Review payload to ensure flight readiness 123 P a g e

137 Initial Check-Off Point Test electronics (GPS, camera, altimeter, etc.) to ensure they are armed and functional prior to launch Inspect launch pad and guide rails for readiness Place rocket on launch pad Have non-level two team members move away from the launch pad back to the safe-viewing area Arm the rocket motor for ignition Disarm all safeties on the rocket Have remaining team member retreat to safe-viewing distance to watch launch Check with Range Safety Office to ensure all codes and rules ae met and the rocket is clear for launch Initiate rocket ignition Check for proper ignition Watch flight so that launch vehicle sections do not get lost Recover payload and main body section after landing Disarm altimeter and any unfired charges Disassemble launch vehicle Inspect launch vehicle for any cracks, breaks or fatigue as a result of testing Record altimeter data 124 P a g e

138 Troubleshooting Table 45-Table 49 below, detail troubleshooting tactics that can be used to address common problems that could be encountered during the pre and post launch subsection inspections. Table 46 - Troubleshooting - Cracking in Main Body Tube or Subsection Initial Check-Off Point Replace cracked part if spare part is available Evaluate severity of structural compromise Determine if cracked piece is load bearing If not load bearing, epoxy part If cracked part is critical and load bearing, postpone launch until replacement part can be obtained or manufactured Table 47 - Troubleshooting - Insecure Fit Between Adjoining Subsections Initial Check-Off Point If too large, sand oversized subsection down until secure fit is reached If too small, replace with spare part If spare part is unavailable and part is too small, add layers or tape to increase diameter until secure fit is reached 125 P a g e

139 Table 48 - Troubleshooting - Unresponsive or Malfunctioning Electronics Initial Check-Off Point Inspect wiring to see if there is any disconnect or break in the circuit Test battery to ensure it is operating at the proper voltage Inspect wiring switch Examine wiring terminals for crossed wires or insertion into incorrect ports Replace unresponsive/malfunctioning electronic piece Table 49 - Troubleshooting - Insecure Connection Between Launch Rail and Launch Pad Initial Check-Off Point Inspect launch pad for debris that could be limiting proper connection Inspect launch rail for bowing that could be limiting proper connection Screw threaded bolts further into launch pad to create more secure connection If connection is still not secure, drill new holes to screw threaded bolts into 126 P a g e

140 Post-Flight Inspection Following flight operations and retrieval of the rocket, all areas of the rocket will be inspected in order to determine the success of the team s testing and design, as well as individual component suitability to be reused on a subsequent flight. In order to complete this post-flight inspection, Table 50 is used. Table 50 - Post-Flight Inspection Checklist Initial Check-Off Point Wait until rocket has landed in a safe location before leaving safe launch zone If the rocket is not deemed safe for retrieval by RSO, stay in safe launch zone and have proper individuals retrieve rocket If the rocket is deemed safe for retrieval by RSO, have the safety officer approach launch vehicle for retrieval Retrieve launch vehicle and return to working area for inspection Remove motor casing once it reaches a temperature that is cool enough to handle Inspect motor casing for cracking or other structural flaws Clean motor casing Disassemble the rocket into individual subsection Remove altimeter from the rocket Record the official altitude of the launch vehicle following flight operations as measured by the altimeter 127 P a g e

141 Initial Check-Off Point Aerodynamics team inspects main body tube, fins, and couplers for cracking or structural flaws acquired during flight Main payload team inspects payload for structural integrity and security of fragile material Electronics payload team inspects altimeter and avionics bay for proper functioning and any damage to electronic systems as a result of flight operations Recovery team inspects all components of the recovery subsection Safety officer completes overall inspection of all subsection inspections Receive all good from RSO 128 P a g e

142 Project Plan Testing The testing plan outlined in the CDR has almost been completed. All but two tests have been completed, and each can be seen in greater detail below. Table 51 summarizes each test and its results. Table 51 - Test Results Test Data Taken Status and Results All three altimeters Altimeter Testing precisely measured altitude. Altitude, GPS tracking, The GPS tracking and live and live feed feed worked properly. Complete. MTS Bulkhead Testing The epoxy failed, not the The force required for carbon fiber or aluminum. failure of the assembly. Complete. Ejection Testing If separation is achieved, the amount of black powder needed. Separation was achieved 10 times for each body tube. Complete. Parachute Force Force of the parachute Deployment Testing deployment. In Progress. Wind Tunnel Testing Strain from a strain gauge. Incomplete. Scale Model Testing Full System Test Two Successful Flights. Complete. Payload Spring Testing Spring Constant Check Complete spring constant matches. Complete. 129 P a g e

143 Test Data Taken Status and Results Full Scale Testing Full System Test Three Successful Flights. Complete. Payload impact testing Modified Charpy Impact Test Complete Altimeter The main scoring altimeter and the recovery altimeters were re-tested with the drone as they were for the subscale launch. The process for this test can be found in the CDR. Attaching to the drone allowed all three altimeters to be tested to ensure the GPS, altitude reading, and live feed all worked correctly. The GPS tracking and live feed is only on the main scoring altimeter and both worked correctly. The altitude data from the drone flight averaged two feet higher than the three altimeters, however the altimeters measured the same altitude. The difference in height makes sense because the altimeters were suspended below the drone two to three feet depending on which test number it was. The drone test was repeated five times with the lengths of the rope being measured after the altimeters were tied off. Along with the altitude being tested, the flight data from the recovery altimeters also showed where the parachutes would have been deployed. The deployment altitudes demonstrates that not only is the altimeter reading altitude, but both recovery altimeters have been correctly set to deploy at the proper altitude. 130 P a g e

144 MTS (Bulkhead) Tensile testing with the MTS Machine was designed to determine, which component would fail and how much force is required to cause failure. Knowing how much force will cause failure will verify the manufacturers specifications. The test also shows that nothing should fail in flight because all components have been designed and selected to withstand more stress than what will be endured in flight. The assembly was manufactured with a small piece of the body tube, two spare recovery bulkheads, and two identical U bolts. The two bulkheads were epoxied into the body tubes, with one at each end. These bulkheads are identical to what will be used in the full-scale flight. The U bolts were attached to the bulkheads in the same manner as the full scale. Two pieces of fracture mechanics clevis grip were used to mount the U bolts in the MTS machine. Figure 40 shows the clevis grips attached to the U bolts before mounting into the MTS machine. Figure 41 shows the assembly mounted into the MTS machine. The MTS test was repeated twice, on two identical assemblies. To ensure the data was as consistent as possible the angle of the bulkhead was measured while it was attached to the MTS machine, the first angle measured 7.4 degrees and the second measured 7.9 degrees. With both tests the epoxy failed first, which is called adhesive failure. The test is considered to be successful because the epoxy failed first and at a force greater than what it will endure in flight. 131 P a g e

145 Figure 40 - Fracture mechanics clevis grip attached to U bolts Bulkhead assembly for MTS testing 132 P a g e

146 Figure 41 - The Assembly Mounted into the MTS Machine The test was determined to be successful if the failure is a higher force than what will be experienced during flight. If the MTS machine reached the maximum travel distance, and the assembly did not fail, then the maximum force put onto the assembly would determine if the test was a success. Table 52 shows the results from the MTS test. The OpenRocket simulation showed that the parachute ejection should put a force of 400 lbf onto the rocket body. Using data from the full scale flight, the actual force felt on the rocket was 206 lbf. With both MTS tests, the bulkheads withstood a significantly higher force than what will be experienced in flight. 133 P a g e

147 Table 52 - MTS Test Results Maximum Force lbf lbf Component Failure Epoxy Epoxy The bulkheads were tested in order to make sure that extra inspections are done at the point of failure before and after the flight. Safety is the primary consideration and locating the most likely point of failure allows the team to ensure safe flights. The maximum force for the first test was lbf, and the maximum force for the second test was lbf. Both assemblies used were made from the same materials, however keeping the exact same amount of epoxy is impossible. On the second test more epoxy was used to better represent the actual amount on the recovery bulkhead in the rocket. On the second test before epoxying the bulkhead into the carbon fiber, both pieces were roughed up with a file. Roughing up each piece allows the epoxy to adhere better compared to two smooth surfaces. The increase in epoxy, along with the rougher surface area, is what caused the higher force needed to fracture the assembly. The procedure used to run the MTS Machine and perform the tensile test on the assembly can be found in Appendix K. Ejection Testing To be sure that the entire recovery system would function as designed, multiple ground ejection tests were performed for each body tube and altimeter. A successful ejection test consists of complete ignition of the black powder charge and separation of the body tube from the coupling tube. Satisfactory performance of each altimeter signal is attained through two 134 P a g e

148 successful tests. These tests ensure that the wiring of the recovery electronics is sound and that the parachute compartments are sufficiently airtight. Additionally, the test will test the shearing of the nylon pins holding the body tubes together. The size of the ejection charges were determined using equations available through the website of the Nevada Aerospace Science Association. Based upon the guidance of the NAR members at Mid-South Rocket Society, the mass of black powder used in each charge will be double the calculated mass. Before the test can begin, an ejection charge must be packed according to the procedure described in the Recovery Preparation section. After inserting the coupling tube into the body tube to be tested, the coupling tube was braced between two sandbags. This ensured that the coupling tube remained stationery during the test, preventing damage to the electronics. The body tube was rested on an adjacent sandbag. To slow the body tube after ejection and protect it from external damage, a series of cloth dampers were hung in the path of the body tube s motion away from the coupling tube. The USB data transfer kit was connected to the altimeter and a test signal was fired. The ejection tests were entirely successful with the exception of a single backup circuit test of the drogue parachute charge, which led to the re-soldering of a disconnected wire. An overview of the tests is given in Table P a g e

149 Table 53 - Results of ejection testing Signal Number of Tests Number of Failures Notes Primary Main 2 0 Primary Drogue 2 0 Backup Main 2 0 Backup Drogue 3 1 Wiring Issue The overwhelming success of ejection testing indicates that the recovery electronics are reliable and that the ejection charges are suitable sized for each body tube. Altogether, the system can be relied upon for triggering recovery events at the appropriate times. Parachute Deployment Force Testing The force experienced by the launch vehicle during recovery events was determined by analyzing acceleration data from the Altus TeleMega. Computing these forces was important for understanding how the fragile material payload would respond under such conditions and provided assurance that critical mounting hardware was not in danger of failure. Using the measured mass of the nosecone/payload section of the launch vehicle and accelerometer data from the full-scale test flights, it was possible to determine the force exerted on this section by the recovery harness during various stages of flight. For each flight, the maximum force occurred during one of the recovery events. These maximum values are given in Table P a g e

150 Table 54 - Maximum force on launch vehicle during descent Flight # Maximum Force (lbf) The forces recorded above are well below the minimum breaking strength of the tubular nylon recovery harness (4000 lbf) and the tested minimum breaking strength of the recovery mounting points (insert value here). These results indicate that the launch vehicle is wellequipped to handle the forces associated with parachute deployment. Wind Tunnel Testing Introduction The wind tunnel is an important instrument used for studying the airflow across solid specimens. Using a scale model of the rocket inside the wind tunnel for testing helps simulate the effects of air resistance, or drag force, during the actual flight. The drag coefficient must be determined in order to best predict the shape, the performance, and the altitude of the rocket. The experimental drag coefficient will be used to empirically validate simulated CFD and OpenRocket drag coefficient values. Testing Apparatus Components Table 55 shows the apparatus used for performing the wind tunnel experiment. 137 P a g e

151 Instrument Make/Model Table 55 - Testing Apparatus Components Model number Diameter (in) Length (in) Width (in) Strain gage Strain indicator Vishay -TN Strain Gage Vishay 3800 Wide Range Strain Indicator Scale Model 0.5 Air fan Wind Tunnel Test section Motor Differential pressure 1295L108A Honeywell - SSCSNBN010NDAA transducer beam Cantilever 6061 rectangle Aluminum beam Conditioner m-prep conditioner A Neutralizer m-prep neutralizer 5A P a g e

152 Instrument Make/Model Model number Diameter (in) Length (in) Width (in) Carbide paper 320 and 400 grit silicon carbide paper Degreaser m-prep CSM Figure 42 through Figure 46 are the components used to set up the experiment. Inside the wind tunnel, there is an attached electric fan that functions to flow air through the testing area. When the air crosses the test section, the air pressure increases due to the decrease in cross sectional area. A pitot tube connected to a differential pressure transducer will be used to measure the velocity of the air inside the wind tunnel. Figure 42 Variable Frequency Drive 139 P a g e

153 Figure 43- Strain Gage (From Vishay website) Figure 44 - Strain Indicator Figure 45 - Air Fan 140 P a g e

154 Figure 46 - Wind Tunnel Figure 47 - Example of wiring strain gage to strain indicator 141 P a g e

155 Figure 48 - Wiring Diagram (strain gage to strain indicator) The strain indicator will be positioned near the testing area wired with the mounted strain gage on the 6061 aluminum rectangular beam (refer to Figure 48). There will be a hole in the test section allowing the operator to insert the beam. The strain gage will be mounted at the base of the clamped beam. When the air crosses through the test section, the scale model rocket will resist drag causing deflection in the beam. When the beam is deflected, the strain indicator will display the strain readings. - To see how the strain gage wired to the strain indicator, refer to Figure 47 and Figure 48. Procedure 1. Strain gage installation. 142 P a g e

156 1.1.Surface preparation for 6061 Aluminum rectangular beam. 1.2.Degreasing 1.3.Abrading 1.4.Burnishing. 1.5.Conditioning 1.6.Neutralizing. 1.7.Gage bounding 1.8.Apply catalyst 1.9.Apply adhesive Soldering strain gage Prepare the leadwire Tin the copper CSA strain gage tabs Trim the lead Position the lead wire for soldering Solder the lead wire to the tabs Remove all flux residue Apply protective coating. 2. Wire the strain gage to the strain indicator. 2.1.Refer Figure 48 to see how strain gage is wired to strain indicator. 2.2.Turn on the strain indicator. 2.3.Set the excitation voltage to be 5 volts 3. Set up the wind tunnel: 3.1.Push the button 143 P a g e

157 3.2. Operate the tunnel at airspeeds of 20 mph( 351 in/s). 3.3.Use the differential pressure transducer to measure the velocity. 4. Use lab view to get the voltage readings. 5. Use equation (4), Equation (5) and Equation (6) to indicate the velocity. 6. Set up the 6061 aluminum rectangular beam: 6.1.Clamp the A 6061 rectangular aluminum beam to the support 6.2.Insert the beam through the test section. 7. Make sure the strain gage is wired to the strain indicator. 8. Position the model rocket inside the test section. 9. Record the reading on the strain indicator readings. 10. Turn off the wind tunnel. 11. Disconnect the strain gage. Analytical method From the wind tunnel testing, measured quantities such as velocity and strain will be used to calculate the drag coefficient. There are two assumptions made before calculating the expected drag coefficient. 1. The velocity is constant. 2. Air density is constant. Equation (1) defines the aerodynamic drag coefficient of an object due to air resistance. C D = F d ρ U2 2 A c (1) 144 P a g e

158 Where Fd is the drag force(lbf), ρ is the air density ( lbm/ in 3 ), U is the velocity of the air wind (in/s), and Ac is the cross sectional area of the scale model rocket. Equation (2) shows the relation between the strain and the cantilever beam that determines the drag force. F d = ϵewt2 6L (2) Where ϵ is the strain (in/in), w is the width of the beam (in), t is the thickness of the beam (in), E is the modulus of elasticity of the beam (lbf/in 2 ), and L is the length where the bounded gage is positioned (in). Equation (2) is only valid for a rectangular beam. By substituting equation (2) into equation (1): C D = Eϵwt2 6Lρ U2 2 A c (3) Another measured quantity is the velocity of air. The velocity will be calculated using the differential pressure transducer. The differential pressure outputs only voltage. Therefore, there will be at least two related equations for indicating the velocity. P = (V out 2.5)5 (4) Where P is the difference in pressure (in H2O), Vout is the output voltage (volts). In order to solve for velocity, a unit conversion of the pressure is required. P ( lbf ft 2) = P (in H 2 O) (5) Therefore, the velocity is calculated using Equation (6). 145 P a g e

159 U = 2 P (lbf ft 2 ) (6) Uncertainty Table 56 shows the mean, systematic uncertainty and random uncertainty used to predict the total uncertainty of the drag coefficient at a velocity of 352 (in/s). Since the testing is not performed yet, the strain was calculated using the predicted drag coefficient from the CFD and OpenRocket simulation. The uncertainty analysis is done for the best case where no precision error is involved. For the best case uncertainty, the expected total uncertainty for the drag coefficient was expected to be ± The Pareto chart (Figure 49) shows the factor that contributed most to the uncertainty analysis, which is the velocity. Detailed calculations are provided in Appendix H. Table 56 - Inputs for Uncertainty analysis Symbol Description Units Mean Systematic (Bias) Random (Precision) Uncertainty Uncertainty L Length in b Width in E Young's Modulus (6061 ALM) psi t Thickness in u Velocity in/s Strain in/in E-06 0 ρ Density lb/in P a g e

160 Percent (%) Ac Area of the subscale rocket. in Systematic (Bias)(%) Random (Precision)(%) Figure 49 - Pareto Chart Test Status The wind tunnel test has not been performed yet. A thorough uncertainty analysis was performed before proceeding with the test procedure to address concerns of accuracy. Complicating factors included: Test section size (length and diameter of the scale model must be reduced due to minimal size constraints of the test section in the wind tunnel) Surface finish of the 3D model (surface roughness on minimal scale could affect coefficient of drag) Cantilever beam assumption validity 147 P a g e

161 Uncertainty of equipment Validity of results as check for CFD Scale Model Testing The sub-scale model was tested in December. It had a goal to reach an apogee of 2,500 feet. The model was launch twice successfully. Although the first flight reached an apogee of 2592 feet, we learned that we were not using the correct black powder or enough black powder. Changing the black powder for the second flight resulted in no issues on the second sub-scale flight. The second flight was closer to our target and reached an apogee of 2498 feet. For a more detailed breakdown, refer to the CDR report. Payload Testing Before the entire payload assembly was tested, the spring constant for the 5 base springs given by the manufacturer was tested to verify that the values used in the math model were accurate, (for math model refer to CDR).. No variable can affect the tests except the change in weights. After these criteria where met, the test was deemed successful if the spring deforms, measurements are accurately taken, and the weight is properly recorded. The procedure for the spring constant test is as follows. 1. Fasten the spring to a mounting plate and turn apparatus upside down so that weights can be suspended from it 2. Fashion a hook and attach it to the end of the spring so that weights can be attached 3. Measure and record the un-stretched length of the spring 4. Attach a weight to the hook 148 P a g e

162 5. Record the mass of the weight and change in length of the spring 6. Repeat steps 4 and 5 until enough data has been collected 7. Calculate the spring constant using Hooke s Law, Equation 1. F = kx Equation 1: Hooke s Law The results of this test are summarized in Table 57. Table 57 - Spring Constant Test Values Mass (kg) Weight (lbf) Spring Displacement (m) Spring Displacement (in) k (kg/m) k (lbf/in) Average Uncertainty ±.16 (kg/m) ±.23 (lbf/in) 149 P a g e

163 The manufacturer specified spring constant is 15 confirming that the tested springs were reasonably close to the specified values. The weights picked for the spring constant test started at 2 pounds and went incrementally until the spring failed to get the entire range of forces. After confirming that the spring constant was near what the manufacturer had specified, a drop test was performed on the payload assembly while mounted in a mock rocket body tube created out of Blue tube. The test was a vertical drop from three stories or 30 feet high. This was designed to simulate forces worse that actual flight and to calculate the force endured by the fragile material within Cylinder 1 through the use of an accelerometer. After the first test however, three of the base springs epoxy and welds broke causing the springs to buckle and five of the wire rope isolators failed, three had adhesive failure and two had cohesive failure. One reason for the failure was that the math model simulated the force as a purely longitudinal force along the length of the rocket, however during the drop test, the tube hit the ground at approximately a 45 degree angle. The first drop test was meant to determine the amount of freefall time to properly calculate the impact force and acceleration that the payload experienced. However, due to the failure of the springs and a malfunction with the accelerometer, no data was gathered. Due to the drop test s lack of repeatability, a substitute test was designed with maximum variable control providing more accurate data. This test was the modified Charpy Impact test. The modified Charpy impact test employed for lack of sources of error and repeatability. The test was set up by placing the payload assembly in the mock rocket body tube and positioning it in the impact zone of the hammer on the Charpy Impact test machine. The U-bolt used to attach the parachute was also used to be the connection point where the hammer transmitted its force to the payload. A frame with sheets draped over it was set at the end of the 150 P a g e

164 testing apparatus to catch the payload when it was launched from the machine. Several tests were conducted to determine the reduction in acceleration, the optimal fill material, and overall performance of the payload. The testing data can be seen in Table 58. To be able to calculate the percent reduction in force and acceleration, the accelerometer was first mounted to the outside of Cylinder 2 or on the mock rocket body to get a base acceleration value. This was later used in comparison to the accelerometer values within Cylinder 1 showing the percent reduction. Table 58 - Charpy Impact Acceleration Test Data Acceleration (counts/g) average of x,y and z directions. Percent reduction from base Fill Material Cotton Filling Shredded Paper Paper/Cotton mix Base Value N/A N/A N/A Initially, the fill material selection was going to be based off the accelerometer values and percent reduction given from those. However, as can be seen in Table 58, the acceleration values were very different between the tests using shredded paper and cotton filling. The values were determined by the graphs found in Appendix I Payload Accelerometer Graphs. By the time this had been discovered, the testing housing had been disassembled and the payload was already in use in the full-scale rocket, so further testing could not be done to determine the reason for such a large difference in acceleration. The Base Value seen in the table represents the acceleration values in the x, y and z directions for the accelerometer mounted to the mock body tube receiving acceleration reducing effects of the springs or fill material. This was used as the basis for comparison. For the shredded paper and cotton fill tests, the accelerometer was placed inside Cylinder 1 to mimic what the fragile material will endure. One source of error and possible 151 P a g e

165 explanation of why the accelerometer values were so different could be the rate of data logging. The maximum rate of data logging for the model of accelerometer used was 4 hz. This means if the impact occurred in a small enough time step, the entire event could have been missed and not logged, which is most likely what happened in all 3 tests. The systematic uncertainty for the accelerometer used is given as: Nonlinearity (x,y,z)=±0.5% FS ; Where FS= 32 g Equation 2 Nonlinearity (x,y,z)=± = ±0.16g Equation 3 Zero-g Offset level accuracy: X and Y-Axis = ±150 mg= ±0.15 g. Equation 4 Z- Axis = ±250 mg= ±0.25 g. Equation 5 Overall systematic error = B = BOIE= e L 2 + e z 2 Equation 6 BOIE(x,y) = e L 2 + e z 2 = (0.16g) 2 + (0.15g) 2 =± g Equation 7 BOIE(z) = e L 2 + e z 2 = (0.16g) 2 + (0.25g) 2 =± g Equation 8 Although the systematic uncertainty demonstrates accuracy in the accelerometer measurements the decision was made to base fill material choice off of the survival of the sample fragile material specimens shown in Table 58. However, this means that no numerical data can prove that the payload reduces the maximum force and acceleration felt by the fragile material by at least 50 percent. 152 P a g e

166 The fragile material tested in the Modified Charpy Impact test was tested first with the frame and draped sheets to catch the payload after impact. The testing data can be seen in Table 59. Table 59 - Fragile Material Sample Testing Break y/n? Fill Material Cotton Filling Shredded Paper Paper/Cotton mix no fill 2 large incadescent bulbs no no no no 2 candelobra bulbs no no no yes Fragile material glass sheet no no no yes egg 1/2 power swing no no no yes egg full power swing no no no yes egg full power swing double impact no no no yes The original test matrix in the CDR included several other fill materials to test, however other materials were omitted due to volume density considerations. Each fragile material was first tested with the hammer of the Charpy Impact Tester at 90 degrees or parallel with the floor, and all fragile materials survived in each of the fill materials. After no fragile material had broken, each fragile material was then tested with no fill material. From the testing with no fill material, the egg was determined to be the most fragile of all materials. The same egg was re-tested 2 times with the hammer on the Charpy Impact Tester raised to the maximum as in every test. However, this time a plywood board supported with cinder blocks was placed 2 feet from the payload so that immediately after impact with the hammer, the payload would impact with a 153 P a g e

167 sturdy wall. The wall allowed a simulation of compression as well as tension in all springs for one test. This test was performed only with the egg as the fragile material selection and both times the egg survived un-cracked. Since both the shredded paper and cotton fill worked in protecting the fragile objects, a combination was selected for the final payload design. Shredded paper will be placed in the top and bottom of the cylinder to crumple and provide axial cushion while the fragile object will be wrapped in cotton fill to project a majority of side impact and keep the material centrally located in Cylinder 1. The final test the payload endured was the full-scale flight tests. The rocket was flown three times, and each time an egg was placed in the payload with the shredded paper and cotton fill mix. During the first two flights, the accelerometer was placed in Cylinder 1 with the egg to try to obtain the maximum force experienced by the fragile material. Accelerometer data for Flight 1 is seen in Figure P a g e

168 Acceleration (counts/hz) Time Step 4 Hz X Y Z Figure 50 - Accelerometer Data Full-scale Flight 1 Flight 1 was the only full scale flight that the accelerometer recorded data for due to the battery malfunctioning. The graph shows the maximum acceleration in the x direction which is the direction of flight of the rocket. This value is counts/hz and converting to acceleration values is ft/sec 2. The altimeter data from the scoring altimeter shows the maximum acceleration of the entire rocket as being ft/sec 2. The uncertainty for the both acceleration values are under ±10ft/sec 2 the proving that the accelerometer used malfunctioned during recording. This also helps prove the random values that occurred during the Charpy Impact Tests. Also through all 3 test flights, the same egg was used and each flight the egg survived unscathed. The team considered this to be evidence of successful performance of the fragile material payload. 155 P a g e

169 Full Scale Testing The full-scale rocket was tested on February eighteenth. A successful flight was defined by, the fragile material payload needed to survive the entire flight, and the apogee needed to be within 5,125 feet and 5,375 feet. Table 60 shows a summary of the results from the full scale flights. A full review of the full scale test can be found in the Full Scale Flight Analysis section above. Table 60 - Full Scale Flight Results Apogee Did the Payload Survive? Flight feet Yes Flight feet Yes Flight feet Yes Requirements Compliance In order to be succeed in the competition, and follow all rules and regulations set forth by NASA, the team will abide by both NASA & team-created requirements. These requirements involve various facets of the project from rocket design parameters, to launching procedures, and safety protocols. Each individual NASA requirement is listed in Table 61, sorted by the corresponding USLI Handbook number. Within this table, each requirement is summarized and a verification plan is given to ensure compliance with all NASA requirements. Additionally, information has been added pertaining to the status of each item, as well as where further information can be found in this report. 156 P a g e

170 Table 61 - NASA Requirement Compliance Handbook Number Summarized Requirement NASA Requirements Verification Method(s) Description of Verification Plan Status & Location The vehicle shall deliver the science or engineering payload to an apogee altitude of 5,280 feet above ground level (AGL). The vehicle shall carry one commercially available, barometric altimeter for recording the official altitude used in determining the altitude award winner. All recovery electronics shall be powered by commercially available batteries. The launch vehicle shall be designed to be recoverable and reusable. Reusable is defined as being able to launch again on the same day without repairs or modifications. Test Analysis Inspection Inspection Test Inspection The rocket team will utilize OpenRocket, RockSim, CFD, & test flight data to achieve an accurate prediction of altitude. The rocket will house a Atlus Metrum TeleMega altimeter in the nosecone to record the official altitude used in determining the altitude award winner. Batteries & altimeter will be purchased from online rocketry sources. The rocket is reusable in design because the team is using a motor that has refuels that can be reloaded into the motor under supervision. Full scale test completed with apogee of 5,291 feet. See Full Scale Flight for more detail. Three altimeters meeting requirements were flown for full scale; all producing valid altitudes. See Validity Assessment for more detail. Recovery altimeters powered by Energizer 9V Lithium Batteries. See Line Item Budget in Appendix F. Three test flights were conducted on the February 18 th. No repairs were made to the rocket, making it reusable. See Full Scale Flight. 174 P a g e

171 NASA Requirements Handbook Number Summarized Requirement Verification Method(s) Description of Verification Plan Status & Location 1.5 The launch vehicle shall have a maximum of four (4) independent sections. Inspection The launch vehicle will have 3 independent sections: the aft body tube, the bow body tube and nosecone, and the coupler. The launch vehicle has 3 independent sections: the aft body tube, the bow body tube and nosecone, and the coupler. See Vehicle Criteria. 1.6 The launch vehicle shall be limited to a single stage. Inspection Demonstration The launch vehicle shall be a single stage. Only one L850W is used, as seen in Vehicle Criteria. 1.7 The launch vehicle shall be capable of being prepared for flight at the launch site within 4 hours. Test The team will conduct multiple tests on full-scale test day and measure re-launch times. The team was able to prepare the rocket in 32 minutes on February 18 th. See Full Scale Flight for more detail on multiple launches that day. 1.8 The launch vehicle shall be capable of remaining in launchready configuration at the pad for a minimum of 1 hour without losing the functionality of any critical on-board component. Test The launch vehicle design will ensure all components have a life of greater than 1 hour without loss of functionality via a fullscale launch pad test. All systems remained on during the full scale test for over 2 hours while the rocket was stuck in a tree. See Full Scale Flight for more detail. 1.9 The launch vehicle shall be capable of being launched by a standard 12-volt direct current firing system. Inspection Test The ignition system will use a 12- volt direct current firing system. The ignition system and igniters used during the full-scale test is 12V. See the Line Item Budget in the Appendix for exact ignitor specifications. 158 P a g e

172 Handbook Number Summarized Requirement NASA Requirements Verification Method(s) Description of Verification Plan Status & Location The launch vehicle shall require no external circuitry or special ground support equipment to initiate launch (other than what is provided by Range Services). The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR). Pressure vessels on the vehicle shall be approved by the RSO. The total impulse provided by a University launch vehicle shall not exceed 5,120 Newtonseconds (L-class). The launch vehicle shall have a minimum static stability margin of 2.0 at the point of rail exit. Inspection Inspection Inspection Inspection Test Analysis There will be no external circuity for the ignition system because it will be a ground based ignition system being placed underneath the rocket before launch with 300 ft of cord between the igniter and the controller. The motor being used is a solid fuel motor from AeroTech. The motor is the L850W. No pressure vessels will be used. The motor will produce an impulse of 3695 N-s which is below the specified total impulse that is allowed. Using OpenRocket, Rocksim, and Test Data determine rail exit velocity and then stability. The ignition system comprised of only 1 ignitor that runs off of 12V. This setup was successful during the full scale test. See the Line Item Budget in the Appendix for exact ignitor specifications. The team has purchased and flown on an Aerotech L850W, see the Line Item Budget for further information on the motor. As of final design, no pressure vessels are used. See Design and Construction of Vehicle. The motor details can be found via Aerotech s website. See Line Item Budget for specific motor look-up information. The flight configuration for competition has an actual flight stability of P a g e

173 Handbook Number Summarized Requirement NASA Requirements Verification Method(s) Description of Verification Plan Status & Location The launch vehicle shall accelerate to a minimum velocity of 52 fps at rail exit. All teams shall successfully launch and recover a sub-scale model of their rocket prior to CDR. All teams shall successfully launch and recover their fullscale rocket prior to FRR in its final flight con- figuration. Any structural protuberance on the rocket shall be located aft of the burnout center of gravity. Analysis Test Test Test Analysis The rocket team will utilize OpenRocket, RockSim, CFD, & test flight data to achieve an accurate prediction of minimum velocity at rail exit. The current value is 66.9 fps. A sub-scale model with comparable weights, lengths, and masses will be launched prior to the CDR. The project schedule will ensure a full-scale rocket launch occurs before the FRR. No structural protuberances will exist bow of the burnout center of gravity. The Full-Scale flight was a success. See Section Flight Simulations and Altitude Predictions The Sub-Scale test was successful and has been completed. See the CDR s Sub-Scale Flight section. Launch both complete and successful on February 18 th. See Full Scale Flight for more detail. The rocket has 3 bolts holding the nosecone to the bow body tube. These are located bow of the burnout center of gravity but has been cleared by NASA. No other structural protuberances exist bow of the burnout center of gravity. 160 P a g e

174 Handbook Number Summarized Requirement NASA Requirements Verification Method(s) 1.19 Vehicle Prohibitions Inspection Vehicle must deploy a drogue parachute at apogee, followed by a main parachute at a much lower altitude. A successful ground ejection test for both parachutes must be conducted prior to sub- and fullscale launches. No part of the launch vehicle may have a kinetic energy greater than 75 ft-lbf at landing. Recovery electrical circuits must be independent of payload circuits. Test Test Analysis Demonstration Inspection Description of Verification Plan The launch vehicle will follow all prohibitions laid out in section 1.19 of the 2017 SL NASA Student Handbook. Dual-deployment altimeters are programmed to fire ejection charges at apogee and at 750 feet. Multiple ejection tests conducted prior to sub- and full-scale launches. Parachute sizes are optimized to minimize kinetic energy at ground impact. Recovery electronics are housed in a separate compartment. Status & Location The launch vehicle has followed all prohibitions laid out in section 1.19 of the 2017 SL NASA Student Handbook. See Vehicle Criteria for full design. Full-scale test flights resulted in successful recovery events. See Full Scale Flight & Recovery for more. Sub-scale and full-scale test ejections were successful 8 consecutive full scale test ejections. See Ejection Testing section. Full-scale test flights resulted in kinetic energy below the maximum allowable. See Recovery section. Coupling tube constructed completely independent of other electronics. See Recovery section for more detail. 161 P a g e

175 Handbook Number Summarized Requirement NASA Requirements Verification Method(s) Description of Verification Plan Status & Location Recovery system must include redundant, commercial altimeters. Motor ejection cannot be used for primary or secondary deployment. Each altimeter must be armed by a dedicated switch accessible from the rocket exterior. Each altimeter must have a dedicated power supply. Each arming switch must be lockable to the ON position. Removable shear pins must be used to seal the parachute compartments. Inspection Demonstration Inspection Inspection Inspection Inspection Inspection Two PerfectFlite Stratologger CF altimeters will be used. Black powder ejection charges are used to deploy parachutes. A separate switch accessible through pressure sampling holes is used to arm each altimeter. Separate 9-Volt batteries are wired to the power leads of each altimeter. Locking rotary switches are wired to the switch leads of each altimeter. Three #2 nylon shear pins are used to seal each parachute compartments. Redundant ejection charges observed for each recovery event during full-scale test flights. See Recovery section. Black powder ejection charges successfully triggered recovery events for full-scale test flights. See Recovery section. Rotary switches successfully armed from rocket exterior for full-scale test flights. See Recovery section. Recovery altimeters were powered up for duration of each full-scale test flights. See Recovery section. Recovery altimeters were powered up for duration of each full-scale test flights. See Recovery section. Pins sheared successfully during ejection testing and fullscale test flights. See Recovery section. 162 P a g e

176 Handbook Number Summarized Requirement NASA Requirements Verification Method(s) Description of Verification Plan Status & Location Tracking device(s) must transmit the position of any parts of the launch vehicle to a ground receiver. Recovery system electronics must not be adversely affected by any other on-board electronics. Design container capable of protecting an unknown object of unknown size and shape. Object must survive duration of flight Test Demonstration Inspection Test Inspection Testing Testing All parts of the launch vehicle are tethered together; position will be transmitted via a flight computer in the nosecone. Recovery electronics located in separate compartment. Math model is used to develop spring system in conjunction with a concentric cylinder model to provide sufficient vibration dampening and force reduction. The spring and concentric cylinder design will be tested with a matrix of different support materials as well as testing materials to assure the unknown object(s) can survive the flight during demonstration. All sections of launch vehicle remained tethered during fullscale test flights. Position data was successfully transmitted throughout each flight. See Line Item Budget for exact GPS specifications. Recovery altimeter data showed no signs of interference after full-scale test flights. See Recovery section. Full scale flights resulted in safe return of an egg with ability to adjust to multiple eggs. See the Payload Testing section for more detail, including a % reduction in force. Full scale flights resulted in safe return of an egg. See the Payload Testing section for more detail, including a % reduction in force. 163 P a g e

177 Handbook Number Summarized Requirement NASA Requirements Verification Method(s) Description of Verification Plan Status & Location Once the object is obtained, it must be sealed in its housing until after the launch and no excess material may be added after receiving the object. Each team shall use a launch and safety checklist Each team shall identify a student safety officer who shall be responsible for the safety of the team and ensure all proper rules and guidelines are followed The team safety officer shall monitor team activities with an emphasis on safety throughout the design, construction, and testing of the rocket by maintaining MSDS sheets and hazard analyses. Test Inspection Demonstration Inspection Inspection During full scale flight, verify that an object can be contained using no excess material. Final assembly and pre-launch checklists will be created and reviewed at the appropriate time to ensure safe launch of the rocket and all members involved in the launch The team has appointed a safety officer to monitor the safety of the team throughout the project and ensure all federal rules and laws are met. The team safety officer will monitor the progress of the project emphasizing the proper safety procedures for the current stage of the project. Using only material already in the rocket, this setup was tested on February 18 th for the full scale flight and passed. See Payload Testing section for more. Launch and safety checklist used for full-scale test flight. See Launch Operations section for more detail. Safety officer Bryan Bauer oversaw both fabrication and testing phases to ensure safe and successful operations. Safety officer has monitored the full-scale testing, fabrication and launch in order to ensure safe operations. 164 P a g e

178 Handbook Number Summarized Requirement NASA Requirements Verification Method(s) Description of Verification Plan Status & Location Each team shall appoint a mentor who has certification and is in good standing with the NRA. During test flights, teams shall abide by the rules and guidance of the local rocketry club's RSO Teams shall abide by all rules set forth by the FAA Students shall do 100% of the project excluding motor / black powder handling. Inspection Demonstration Demonstration Demonstration Inspection The team has assigned an school faculty member to mentor the project to provide valuable insight on the rocket design and construction as well as assume full liability of the rocket. Team will converse with RSO at local rocketry club to ensure all of their chapter s rules and regulations are abided by. Team will converse with NASA lead safety officer and thoroughly research all rules and regulations set forth by the FAA to ensure all rules and regulations are abided by. The team will continuously demonstrate an independently managed and executed project. The team lead will routinely monitor this quality. Dr. David Unger is the team mentor; his information can be found on the cover page. Team is in compliance will all rules and regulations set forth by local rocketry club BluesRocks. See Full Scale Flight for more detail. Team is in compliance will all rules and regulations set forth by FAA and NASA. See Full Scale Flight for more detail on the flight. The team has only used mentors for guidance and will continue to do so. 165 P a g e

179 Handbook Number Summarized Requirement NASA Requirements Verification Method(s) Description of Verification Plan Status & Location A detailed project plan shall be maintained. Foreign National members shall be identified by the PDR. All team members attending launch week shall be identified by the CDR. The educational engagement requirement shall be met by the FRR. The team shall develop and host a website for documentation. The team shall post & make available for download all deliverables by the specified date. Demonstration Inspection Inspection Inspection Test Inspection Documents for scheduling, budget tracking, outreach, and safety will be continuously updated and reported. The team lead will ensure that any Foreign National members are clearly indicated in the PDR. It will be checked that a list of team members, with indications of those attending launch week, will be included in the CDR. The Educational Engagement lead shall confirm that all documentation has been received and approved by NASA prior to the FRR. Team members will periodically confirm that the website is functioning as intended by opening each posted document. The team lead shall confirm that all documents are posted prior to the specified date. Project plan is updated and can be found in the Project Plan section. Foreign National members have been identified in s with NASA. Team members have been identified in s with NASA, along with completed waivers. Team has completed outreach activities with over 200 students reached. Educational engagement is not discussed in this report. Website has been developed and is being updated. All reports, presentation slides and flysheets have been and will continue to be posted to the team website by the deadline set forth by NASA. 166 P a g e

180 Handbook Number Summarized Requirement NASA Requirements Verification Method(s) Description of Verification Plan Status & Location All deliverables must be in PDF format. A table of contents must be included in all reports. Page numbers shall be provided in each report. The team shall provide videoconference equipment needed for reviews. All teams shall use launch pads provided by the SLS provider. The team must implement the EIT accessibility standards. Inspection Inspection Test Demonstration Test Demonstration Demonstration The team lead shall confirm that all documents posted are in PDF format. The team lead shall ensure that a table of contents is located at the start of each report. Page numbers shall be checked to the table of contents to ensure continuity throughout the report. Videoconference rooms will be reserved and trialed immediately prior to each design review. The team shall design the rocket to utilize launch rail. If software or applications are created (not planned) the team will abide by 36 CFR Part Otherwise, all components containing software will be checked to ensure compliance. All deliverables to the team website are upload in PDF format. See Table of Contents, Figures, and Tables sections. See lower right hand corner of each report. Requirement met, same setup will be used for all future correspondence. The Rail used for subscale launch operated as intended, see Full Scale Flight section. Software not designed by team. See Line Item Budget for exact electronic components. 167 P a g e

181 As mentioned, Project ACE has developed a set of team derived requirements as well. The team requirements can be seen in Table 62. They cover things that were not touched on by the handbook and also add depth to certain handbook requirements. 174 P a g e

182 Table 62 - Team Requirement Compliance Number Requirement Team Requirements Verification Method Description of Verification Method Status & Location All reports shall be compiled at least three days prior to NASA due dates. Each member of the team shall have a working knowledge of each subsystem. Safety shall be made the team s first priority. Altimeters shall be in good working order. Demonstration Inspection Demonstration Test Reports shall be completed, according to team schedule, prior to NASA due dates to allow for revision time and mitigate risk of late submissions. At each team meeting, every subsection lead will review the status of their section with the entire team. The team leader will confirm that the information presented is sufficient. The safety officer will periodically ask team members what the most important aspect of the project is. All altimeters shall be flown on sub-scale and full scale flight tests. Altitude readings will be compared to confirm consistency. The team has completed all reports on time. The dates can be seen in the Schedule portion of the report. This has been maintained. It was recently demonstrated at the full-scale launch where team members had to work on each other s sections. See Full Scale Flight for more details. Safety officer has asked 17 team members what the most important part of the project is and has had 15 safety answers. The two outliers have been reminded of safety. Altimeters have all been extensively tested and have passed all tests. See Altimeter Testing section for more detail. 174 P a g e

183 Team Requirements Number Requirement Verification Method Description of Verification Method Status & Location 5 The tracking system shall be in good working order. Test The tracking system shall be flown on the sub-scale and full scale flight tests. This will be used to find the rockets thus confirming its operation. The tracking system performed perfectly in the sub-scale & full-scale test. During the full scale test, the tracking system located the rocket over 1 mile away. 6 A solid output signal must be given from triggered altimeters. Test Analysis All altimeters will be triggered while voltage is read on the output. This output will be read to confirm it is acceptable. The output voltage is seen in real time at the base station. 7 All circuits shall be checked prior to use. Inspection All circuits will be confirmed at each node to ensure connections. This was completed for both sub-scale and full-scale tests. Continuity and amperage were both inspected. 8 Impulse for the parachute deployment shall be determined experimentally. Test Analysis The main parachute shall have an apparatus (strain gauge) attached to it that enables a force to be read as it opens at high speed. This will cut down in the large ambiguity that exists in estimating an impulse value. Parachute force testing completed using acceleration data on altimeter. See Parachute Deployment Force Testing section. 170 P a g e

184 Number Requirement Team Requirements Verification Method Description of Verification Method Status & Location A spring constant for parachute cords shall be determined experimentally. Payload must reduce force felt by object(s) by 50 % Payload must reduce acceleration of object(s) by 35 % Test Analysis Testing Testing The spring constant shall be determined using forces related to what is experienced with parachute opening. This helps when estimating energy absorption by the cord when the chute opens. From the mathematical model, appropriate springs will be selected to induce oscillation and reduce force. These will be tested by Charpy Impact Tests. From the mathematical model, appropriate springs will be selected to insure acceleration graphs show 35 percent reduction from inputs. Will be tested via Charpy Impact Test. This testing has been bundled into the parachute deployment force test. The parachute as it relates to the body had its acceleration measured, which as a system includes the cords expansion. See Parachute Deployment Force Testing. This testing has been completed and selected springs were also tested to assure spring constants given by the manufacturer were accurate. The test to deduce the percent reduction in force and acceleration was completed however a faulty accelerometer caused data to be useless and therefore percent reduction cannot be found. See Payload Testing section for full explanation. 171 P a g e

185 Number Requirement Team Requirements Verification Method Description of Verification Method Status & Location Electronics must operate in cold temperatures Mach Number will be less than 0.6 The rail exit velocity will be above 60 ft/s Complete a Combustion Analysis on the Motor to obtain Pressure of fuel ignition Complete a Modal Analysis on the Motor Mount System to ensure safety and stability of the rocket Demonstration Testing Simulation Test Simulation Simulation Simulation First, temperature sensitive components will be identified. Then components will be tested in the cold with ejection testing. From simulations, the motor and aerodynamics of the rocket will ensure the rocket has a Mach number of 0.53 From the simulations, the rocket weight and motor section will ensure of having a rail exit velocity of 66.9 ft/s From hand calculations to obtain the temperature and pressures to run the FEA analysis on the motor casing to find the Factor of Safety of the motor casing Used hand calculations to determine the natural frequency of the motor mount and then used Finite Element Analysis to find operational frequency 2 Full Scale ejection tests done in cold (-2 C) environment, test passed. See Ejection Testing section. Full-Scale completed and simulations ran. See section Flight Simulations and Altitude Predictions Full-Scale completed and simulations ran. See section Flight Simulations and Altitude Predictions Combustion analysis complete and has a pressure of 2155 kpa. Located in Combustion Analysis in CDR. Modal analysis complete with an operating frequency not near natural frequency. See Modal Analysis in CDR. 172 P a g e

186 Number Requirement Team Requirements Verification Method Description of Verification Method Status & Location Complete a Shear Stress Analysis on the motor mount to ensure that the epoxy being used will withstand the motor forces Have a Factor of Safety above 2 for the Combustion Analysis and Shear Stress Analysis Reach an altitude between 5,200 and 5,400 feet Design flexibility on full-scale test launch day to raise or lower altitude on a second test-flight. Demonstration Demonstration Simulation Testing Demonstration Simulation Used hand calculations to have a verification of the results found using the Finite Element Analysis to find the Factor of Safety of the motor mount Calculated the Factor of Safety of Combustion and Shear Stress areas. Found Combustion factor of Safety to be 103. Found Shear Stress factor of safety to be 17 Use OpenRocket and Rocksim to simulate the altitude of the fullscale rocket. Test the full-scale to see the actual altitude Using simulation and demonstration of design, the team will prove that on test launch day, small changes can be made to raise or lower altitude for a second flight. Analysis complete with F.O.S. of See Shear Stress Analysis in CDR. Complete with F.O.S. of & 17.07, respectively. See Propulsion Section in CDR. Full scale test apogee of 5,291 feet. See Full Scale Flight section for more detail. Ballast adjusted on February 18 th to change height for three different flights. See Full Scale Flight for more detail on the configurations. 173 P a g e

187 Budgeting and Timeline Budget Project ACE received funding from three primary sources. First, the Indiana Space Grant Consortium generously awarded Dr. David Unger & Project ACE a total of $5, The University of Evansville s Student Government Association (SGA) and University of Evansville s College of Engineering & Computer Science contributed as well, resulting in total funding of $10, seen in Table 63. Table 63 - Sources of Funding Source Amount NASA Grant (INSGC) $5, Student Government Association $2, U.E. College of Engineering & Computer Science $2, Total $10, After obtaining funding, Project ACE created a detailed budget that resulted from a complete parts list (Appendix F). For financial purposes, this budget broke the project into 10 sections. Additionally, a variable contingency fund was built into the budget for each section. The sum of the parts list and variable contingency fund is shown as the Forecasted Amount column in Table 64. Using detailed cost-tracking methods an Amount Expended column was created in Table 64. The Amount Expended figure represents the total amount spent on that section of the project. As of FRR submission, all spending has been completed aside from fuel to/from competition. Fuel costs have been conservatively estimated and are included in the Travel / 174 P a g e

188 Lodging figures. As such, all expended amounts reflect final values. From this a Difference column was created that is the difference between the forecasted and expended amounts. Figures containing parenthesis and a red background indicate a section that went over budget while figures with a green background indicate a section that remained under budget. A visual comparison of forecasted and actual expenses is provided in Figure 51. Table 64 - Sectional Budget Breakdown Section Forecasted Amount Amount Expended Difference Operating $ $ $(270.90) Travel / Lodging $2, $2, $ Launch Pad $ $ $22.41 Aerodynamics (Body) $1, $ $ Propulsion $2, $2, $ Main Payload $ $ $(292.17) Electronic Payload $ $ $15.14 Recovery $1, $1, $(39.49) Scale Model $1, $ $6.95 Educational Engagement $ $74.83 $25.17 Total $10, $10, $ Operating costs were over budget due to the purchase of tools and team polo reimbursements. The main payload went over budget due to the mounting re-design (discussed in Payload section) & while the recovery excess was caused by unforeseen component costs. 175 P a g e

189 Ultimately, the project concluded under budget by $ a total expenditure nearly 5% under the forecast. Figure 51 - Sectional Budget Amounts 176 P a g e

190 Schedule A detailed breakdown of each task, accompanied with all pertinent dates, can be found in the detailed task breakdown in Appendix G. All critical dates for completion of the project are shown in Table 65. Additionally, a broader view of the task breakdown can be seen in Gantt chart form in Figure 52. Despite a few testing delays the project is on schedule as of FRR. Table 65 - Critical Dates Due Date Activity NASA U.E. Team Project Kickoff Aug. 15, General Motor Selection/Data Sept. 30, Sept. 16, 2016 Informal Design Sketches - Sept. 21, 2016 Sept. 19, 2016 Proposal Sept. 30, 2016 Oct. 3, 2016 Sept. 27, 2016 Motor Selection/ Data Oct. 31, 2016 Oct. 7, 2016 Proposal Presentation - Oct. 24, 2016 Oct. 22, 2016 PDR Report Nov. 04, Oct. 26, 2016 PDR Flysheet Nov. 04, Oct. 26, 2016 PDR Presentation Nov. 04, Oct. 28, 2016 Sub-Scale Launch Motor Selection - - Nov. 30, 2016 Sub-Scale Launch - - Dec. 11, 2016 Design Report - Dec. 2, 2016 Nov. 29, 2016 Design Presentation - Dec. 5, 2016 Dec. 2, 2016 Motor Mount Design/ FEA Jan. 13, Nov. 30, 2016 All Structural elements FEA Jan. 13, Nov. 30, 2016 CDR Report Jan. 13, Dec. 9, 2016 CDR Flysheet Jan. 13, Dec. 9, 2016 CDR Presentation Jan. 13, Jan. 11, 2017 Full Scale Launch - - Feb. 12, 2017 FRR Report Mar. 6, Mar. 1, 2017 FRR Flysheet Mar. 6, Mar. 1, 2017 FRR Presentation Mar. 6, Mar. 3, 2017 Competition Apr. 5, Apr. 5, 2017 LRR Report Apr. 6, Apr. 3, 2017 UE Final Report - Apr. 17, 2017 Apr. 12, 2017 UE Final Presentation - Apr. 20, 2017 Apr. 17, 2017 PLAR Report Apr. 24, Apr. 21, P a g e

191 Testing Construction Design Reporting Project ACE Gantt Chart Period Highlight: 27 Plan Duration Actual Start % (Planned) Actual (beyond plan) % (Unplanned) ACTIVITY T/M RESPONSIBLE PLAN START PLAN DURATION Proposal David Preliminary Design Report David PDR Presentation David Interim Design Report David Critical Design Report David CDR Presentation David Flight Readiness Report David FRR Presentation David Project Final Report David 31 2 Launch Readiness Review David 29 4 Post Launch Assesment David 33 2 Budget Creation David Website Creation Bryan Motor Type Selection Andrew G Motor Mount Design Andrew G Rocksim Model Andrew G Body Component Selection Torsten D Rocket Model Torsten CFD Model Torsten Payload A Design Justin Payload B Design Braden Data Acquisition Design David Data Transmission Design David Design of Recovery System Andrew S Design Tracking System Andrew S Design Education Activity Bryan Propulsion Construction Andrew G Body Construction Torsten Payload A Construction Justin Payload B Construction Braden Recovery System Construction Andrew S Data Systems Construction David Scale Model Construction Torsten Scale Model Test Team Bulkhead Testing Rakan Payload Testing Braden Parachute Testing Andrew S Wind Tunnel Testing Feras Recovery Testing Andrew S Educational Engagement Bryan Ongoing Ongoing Preparation for Competition David 31 1 Competition David 32 1 ACTUAL START ACTUAL DURATION PERCENT COMPLETE 100% 100% 100% 100% 100% 100% 100% 100% 0% 0% 0% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 100% 30% 100% 100% 0% 0% (Week 1 ends September 4th, 2016) PERIODS Figure 52 - Gantt Chart 178 P a g e

192 References Autodesk. (2015, December 28). External Incompressible Flow. Retrieved from Autodesk Knowledge Network: explore/caas/cloudhelp/cloudhelp/2014/enu/simcfd/files/guid-4eed9e6e-a d9cc42a5ec2-htm.html Center, G. C. (2016, 08 10) NASA's Student Launch. Retrieved 08 11, 2016, from NASA: Engineering Toolbox. (n.d.). U.S. Standard Atmosphere. Retrieved from Engineering Toolbox: G. Lengellé, J. D. (2004, January). Combustion of Solid Propellants. Research Scientists, Energetics Department Office national détudes et de recherches aérospatiales (ONERA). Lofton, J. (2016, November 29). Mechanical Engineering Professor. (T. Maier, Interviewer) Michael J. Moran, H. N. (2014). Fundamentals of Engineering Thermodynamics. Hoboken: John Wiley & Sons, Inc. NASA. (n.d.) NASA Student Launch: Colleges, Universities, Non-Academic Handbook Niskanen, S. (2009). Development of an Open Source model rocket simulation software. OpenRocket. Helsinki: HELSINKI UNIVERSITY OF TECHNOLOGY. Ring, C. (2016, 9 27). Launch Crue. Retrieved from LaunchCrue.org: P a g e

193 Schmidt, D. P. (2016, October 15). Natural Frequency. Weidong Cai, P. T. (2008). A MODEL OF AP/HTPB COMPOSITE PROPELLANT COMBUSTION IN ROCKET-MOTOR ENVIRONMENTS. Taylor & Francis Group, LLC. 180 P a g e

194 Appendix A Machine Prints Dimensioned Drawings 181 P a g e

195 Figure 53 Aft Body Tube Drawing 182 P a g e

196 Figure 54 - Bow Body Tube Drawing 183 P a g e

197 Figure 55 - Fin Drawing 184 P a g e

198 Figure 56 - Motor Drawing 185 P a g e

199 Figure 57 - Nosecone Drawing 186 P a g e

200 Figure 58 - Launch Vehicle Drawing 187 P a g e

201 Figure 59 Recovery bulkhead drawing 188 P a g e

202 Figure 60 - Payload Main bulkhead residing in Cylinder 2 Figure 61 - Payload assembly general dimensions 189 P a g e

203 Figure 62 - Recovery attachment bulkhead and hardware 190 P a g e

204 Figure 63 - Altimeter Mounting Plate Piece P a g e

205 Figure 64 - Altimeter Mounting Plate Vertical P a g e

206 Figure 65 - Metal O-Ring 193 P a g e

207 Figure 66 Propulsion Section 194 P a g e

208 Figure 67 Inner Tube 195 P a g e

209 Figure 68 - Centering Ring 196 P a g e

210 Figure 69 - Thrust Plate 197 P a g e

211 Figure 70 - Inner Cylinder 198 P a g e

212 Figure 71 - Payload Coupler 199 P a g e

213 Appendix B OpenRocket Simulation Sub-scale OpenRocket Inputs 200 P a g e

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216 Appendix C Best Fit Curve OpenRocket Simulation Piecewise Regression 203 P a g e

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218 Appendix D OpenRocket Simulation Inputs for OpenRocket Flight Simulation and Different Flight Configurations Flight Configuration P a g e

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221 Flight 2 Configuration 208 P a g e

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224 Flight 3 Configuration 211 P a g e

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227 Appendix E Payload Part Specification Payload Part Specification Sheets 214 P a g e

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