University of Mississippi Rocket Rebels

Transcription

1 University of Mississippi Rocket Rebels Project Presidium NASA Student Launch Initiative Flight Readiness Review March 6, 2017 Center for Manufacturing Excellence 1784 University Circle University, MS 38677

2 Table of Contents List of Figures... 6 List of Tables General Information Faculty Advisor/Adult Educators Safety Officer Student Team Leader Team Participants NAR Section Project Summary Team Summary Launch Vehicle Summary Payload Summary Changes Made Since CDR Changes to Vehicle Criteria Changes to Payload Criteria Changes to Project Plan Vehicle Design Design Revision Details Structural Elements and Design Features Airframe Motor Mount Assembly Fin Design Motor Retention System Centering Rings Motor Coupler Tubes Bulkheads, Attachment Hardware, and Ballasts Nose Cone Electrical Elements and Design Features Avionics Retention GPS Tracker Flight Reliability Confidence Mission Success Criteria UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 2

3 4.4.2 Flight Reliability Vehicle Construction Airframe Construction Nose Cone and Fin Procurement Coupler Subassembly Avionics Configuration Parachute Packing Pre-launch Assembly Recovery Subsystem Structural Elements Avionics Sled Attachment Hardware Recovery Harnesses and Packing Electrical Elements and Design Features Altimeters Key Switches Wiring Scheme Redundancy Features Parachute Sizing and Descent Rates Transmitters System Sensitivity Mission Performance Predictions Mission Performance Criteria Flight Simulation Data Drag Assessment Scale Modeling Results Stability Kinetic Energy Drift Calculations Mass Statement Full Scale Flight Data Predicted vs. Actual Flight Data Comparison Error Analysis Drag Coefficient Comparison Payload Criteria Payload Objective and Mission Success Criteria UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 3

4 5.2 Structural Elements Payload Construction Dimensional Drawings and Schematics Payload Component Interaction and Integration System Design Requirements Relevance of Expected Data and Accuracy/Error Analysis Precision of Instrumentation and Repeatability of Measurement Safety Personal Hazard Analysis Failure Modes and Effects Analysis Environmental Hazard Analysis Launch Operations Procedures Launch System Launch Pad Launch Ignition System Launch Procedures Hardware List General Vehicle Assembly Launch Pad Setup Recovery Subsystem Payload Subsystem Motor Subsystem General Field Supplies Launch Location & Setup Procedure Launch Pad Assembly Aft Vehicle and Drogue Parachute Integration Avionics Preparation and Integration Main Bay Integration Payload Preparation and Integration Rocket Motor Installation Launch Vehicle Final Integration Launch Checklist Post-Launch Checklist Project Plan Testing Vehicle Component and Subsystem Testing UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 4

5 Carbon Fiber Stress Testing D-Printed Components Stress Analysis Static Deployment Testing Payload Testing Requirements Compliance Team Derived Requirements Budget Funding Plan Timeline Appendix Appendix A: MSDS Samples Appendix B: NAR High Power Rocketry Safety Code Appendix C: Minimum Distance Table Appendix D: Range Safety Regulation Form UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 5

6 List of Figures Figure 1: Team Hierarchy Figure 2: OpenRocket Side-view Design Model Figure 3: Motor Mount Assembly 2D Drawing Figure 4: Printed Fin Design Figure 5: Aeropack 75mm Motor Retainer Assembly Figure 6: Preliminary 3D Print of Fin Sleeve w/ Centering Ring Figure 7: Loki Research L840 Motor Thrust Curve Graph Figure 8: Ballast Construction and Final Component Figure 9: Final As-Built Nose Cone Figure 10: Completed Debulk Bag Construction Figure 11: Debulked Airframe w/ Defect Figure 12: Final Bagging and Seal Check Before Cure Figure 13: Post Cure Layup Sections with and without Caul Plates Figure 14: Segment of Extracted Carbon Fiber Tubing Figure 15: Final Surface Prep and Powder Coating Operations Figure 16: Finished Airframe Sections Figure 17: Dimensional Drawing of Fin Sleeve and Centering Ring Part Figure 18: CAD Model of Fin w/ Dimensional Data Figure 19: Nose Cone Dimensional Drawing Figure 20: Water-Jet Cutting Procedure for Full-Scale Bulkheads Figure 21: Completed Coupler Tube Subassembly Figure 22: Printed Avionics Sled with Altimeter and Battery Attachments Figure 23: Avionics Bay Bulkhead Figure 24: Rotary Switch, Battery, and Altimeter Configuration Figure 25: StratoLoggerCF Altimeter Schematic Figure 26: Completed Avionics Setup Figure 27a: Main Parachute Canopy Folding Figure 27b: Main Parachute Suspension Line Fold Figure 27c: Main Parachute Canopy Packing Figure 27d: Main Parachute Suspension Line Packing Figure 27e: Main Parachute Final Bagging Figure 27f: Riser Length Inspection UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 6

7 Figure 28: Fin Slot Cutting Tool Dimensional Drawing Figure 29: Recovery System Events Flowchart Figure 30: Avionics Sled CAD Drawing Figure 31: Recovery Package and Shock Cord Configuration Figure 32: Inflated Annular Parachute Design Figure 33: StratoLoggerCF and RRC3 Altimeters Figure 34: Rotary Key Switch Figure 35: Rotary Switch Installation Figure 36: Avionics Bay Electrical Schematic Figure 37: Recovery System Parameter Calculations Figure 38: ABLEGRID Mini Spy GPS Tracker Figure 39: Altitude, Velocity, and Acceleration vs. Time Simulated Flight Data Plot Figure 40: Current Projected Drag Coefficients vs. Mach Number Figure 41: Drag Force vs. Time Figure 42: Comparison of Scaling Factors Figure 43: Stability Margin, CP, and CP vs. Time Figure 44: Flight Side Profile for Maximum Drift Figure 45: As-Flown Full-Scale Launch Vehicle Figure 46: Actual Altitude and Velocity vs. Time Figure 47: Altitude and Temperature vs. Time Figure 48: Altitude and Battery Voltage vs. Time Figure 49: Projected Vertical Motion vs. Time Flight Data Figure 50: CDR Drag Coefficient versus Mach Number Figure 51: FRR Drag Coefficient versus Mach Number Figure 52: Drag Force vs Time Figure 53: FMPA Machining Setup and Stencil Pattern Figure 54: Completed FMPA Figure 55: FMPA 2-D Drawing w/ Dimensions Figure 56: FMPA Exploded View Figure 57: Motor Bulkhead Exploded and Assembled View Figure 58: Assembled L840 Motor Cross-Section Figure 59: Tensile Stress Setup and Point of Failure Figure 60: Stress vs. Strain Curve of Sample Carbon Fiber Material Figure 61: Lower Fulcrum Setup for Short Beam Shear Test UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 7

8 Figure 62: Short Beam Shear Test Action Snapshot Figure 63: Max Drag Force Analysis of Nose Cone Figure 64: Max Motor Thrust Analysis of Nose Cone Figure 65: Max Drag Force Analysis of Fin Figure 66: Max Motor Thrust Analysis of Fin Figure 67: Full Scale Static Deployment Test Snapshots Figure 68: Project Expenditure Pie Chart Figure 69: PDR, CDR, and FRR Budget Comparison Figure 70: Managerial Timeline Figure 71: Subscale Construction Timeline Figure 72: Full Scale Construction Timeline Figure 73: FMPA Construction and Testing Timeline UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 8

9 List of Tables Table 1: Team Members and Responsibilities Table 2: General Vehicle Parameters Table 3: Body Tube Section Lengths and Fiber Orientation by Layer Table 4: Nose Cone Dimensional Data Table 5: Airframe Ply Lengths and Orientations per Mandrel Table 6: Fin Configuration and Dimensional Data Table 7: Nose Cone Specifications Table 8: Parachute Sizing and Projected Descent Rates Table 9: ABLEGRID TK102B GPS Tracker Specifications Table 10: Simulation Flight Data Table 11: Noteworthy Data Values from Flight Table 12: Current Projected Drag Coefficients Table 13: Kinetic Energy of Independent Sections Table 14: Drift Calculations with Change in Wind Speed Table 15: Mass Statement Table Table 16: As-Flown Full-Scale Launch Vehicle General Specifications Table 17: Launch Conditions Data Table 18: Chronological Full-Scale Flight Critical Event Data Points (Retrieved from Altimeter) Table 19: Comparison of Simulated and Actual Flight Data Table 20: CDR Drag Coefficient Projections Table 21: 1 st Test Flight Drag Coefficient Calculations Table 22: FMPA Ring Dimensional Data Table 23: Likelihood and Severity Table Table 24: Personal Hazards Analysis Table 25: Failure Modes and Effects Analysis Table 26: Environmental Hazards Analysis Table 27: Tensile Stress Limit Calculations Table 28: Shear Stress Limit Calculations Table 29: Max Stress and Max Displacement of 3D Printed Parts Table 30: Drop Test Data Table Table 31: Verification Plan Table 32: Team Derived Verification Plan UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 9

10 Table 33: Final Detailed Expenditures Table 34: Donated Material List UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 10

11 1 General Information 1.1 Faculty Advisor/Adult Educators Dr. Jack McClurg Cody Hardin Phone: (662) Phone: (662) Safety Officer Will Thomas Safety Officer 1.3 Student Team Leader Dillon Hall Project Manager & Chief Engineer Phone: (662) Team Participants The members and their subsystem duties are as follows: Dillon H. Blake H. Will T. Olivia L. Barrett F. Ryoma T. Branden L. Peter D. Caroline R. David J. David T. Mac K. Kyle P. Matt W. Garrett R. David B. Student Member Table 1: Team Members and Responsibilities Roles & Responsibilities Chief Engineer/Project Manager Structures Propulsion/Safety Officer Structures Propulsion/Outreach Payload Payload Structures Recovery/Outreach Propulsion Recovery Avionics Outreach/Payload Structures Avionics Propulsion UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 11

12 Caroline R. Dillon H. Will T. Outreach Manager Chief Engineer Safety Officer David B. Peter D. Kyle P. Garrett R. David T. Propulsion Lead Structures Leads Payload Lead Avionics Lead Recovery Lead David J. Will T. Barrett F. Olivia L. Matt W. Blake H. Ryoma T. Branden L. Mac K. Caroline R. Propulsion Team Structures Team Payload Team Avionics Team Recovery Team Figure 1: Team Hierarchy 1.5 NAR Section The Ole Miss Rocket Rebels coordinate with the Mid-South Rocket Society (MSRS) NAR chapter #550. When applicable, the team will collaborate with the Music City Missile Club NAR chapter # Project Summary 2.1 Team Summary The Rocket Rebels are a newly-formed rocket team formed as a part of the Center for Manufacturing Excellence (CME), which is a state-funded program based out of the University of Mississippi in Oxford, MS. The Project Faculty Advisor is Dr. Jack McClurg, Professor of Manufacturing Engineering. The Safety and Rocketry Mentor is Cody Hardin (NAR Level III Certified, NAR #87214). The mailing address of the CME is as follows: 2.2 Launch Vehicle Summary Center for Manufacturing Excellence 1784 University Circle University, MS Figure 2: OpenRocket Side-view Design Model UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 12

13 Table 2: General Vehicle Parameters Structure and Propulsion Total Length 105 in Outer Diameter 6.2 in Dry Weight lb Wet Weight lb Motor Choice Loki Research L840CT Motor Impulse 3888 Ns (874 lbf-s) Average Thrust 844 N (190 lbf) Burn Time 4.6 sec Recovery System Drogue Chute Design: Annular Main Chute Design Annular w/ Pilot Chute Drogue Chute ID-OD 26 in 3 in Main Chute ID-OD 5.93 ft 9.7 ft 2.3 Payload Summary As a new team, the Rocket Rebels have focused its payload efforts to one simplified mission objective. The payload project has been named Project Presidium, which is a Latin term for protection, and consists of designing and constructing a fragile material payload apparatus (FMPA) that can protect a fragile cargo throughout the entirety of a rocket flight aboard the Rocket Rebels vehicle. To fulfill the requirements of the NASA SLI Handbook, the team has designed the apparatus with the capability to easily open and close for quick installation of the cargo before a launch as well as the ability for the apparatus to be easily installed within the vehicle s payload bay. 3 Changes Made Since CDR 3.1 Changes to Vehicle Criteria Since the full-scale test flight, it was found that original flight projections were significantly under actual flight altitude. The discrepancy has since been identified and resolved in the computer model. To compensate the altitude gap from the target of 5,280 feet, two ballasts have been incorporated in the booster and payload bulkheads to reduce altitude, yet maintain vehicle requirements. A 2 nd test launch scheduled for March 25th will confirm the effectiveness of the ballasts. A minor material change has been made to the fin material. Due to a maintenance issue, the primary 3D printing machine used for most of the prints (Fortus) was substituted for a smaller 3D printer that used ABS plastic of similar structural strength. This material was slightly lighter than the ASA used in the Fortus machine and was only utilized on the fins. 3.2 Changes to Payload Criteria Due to miscommunication with Ace Controls, a new vendor has been selected that has delivered similar force dissipation factors (spring and damper coefficients) in shock absorbers at a much cheaper price. However, these shock absorbers equilibrium position is at full damper extension. Therefore, the piston system can only operate in one direction. This required all shock absorbers to be moved to the bottom of the apparatus to mitigate launch loads (the highest source of shock to the cargo). 3.3 Changes to Project Plan Verifications have been added to each safety hazard mitigation in the FMPA, PHA, and Environmental Hazard Analysis to ensure that every mitigation is incorporated and no risk is left unchecked. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 13

14 4 Vehicle Design 4.1 Design Revision Details Since CDR, some component features have been modified with lightening hole grids and redesigned with lighter material to correct a projected weight problem which resulted in the projected flight data being under the target altitude of one mile by about 500 feet. Modified components include the fins and centering rings. Since the first full scale test launch, it was found that the as-built vehicle was significantly under-weight, having a confirmed altitude reading of 6224 feet. Therefore, ballasts have been incorporated into the design. The specifications of these changes are described in detail in the following subsections. 4.2 Structural Elements and Design Features The primary objective for each design feature on every component of the vehicle is to further the mission objective while also ensuring maximum safety to the vehicle, nearby personnel, and environment. The following subsections detail how each element in the vehicle meets these two criteria and in what way. A sampling of MSDS data for building material used is available in Appendix A Airframe The vehicle s body tubes utilize a pre-impregnated carbon fiber-resin weave plies which was laid up around a 6-inch aluminum mandrel. Carbon fiber was chosen for its superior strength to weight ratio, allowing for a lighter design without compromising structural integrity. 95 total inches of circular tubing is made for the full-scale vehicle, consisting of two inner.028 carbon fiber plies and a.014 carbon fiber outer ply. After curing, this resulted in an airframe wall thickness of about.1. The thinner outer layer allows for a sleeker and less porous surface finish so that a powder coat finish can be evenly applied across the airframe. An additional ply of.028 carbon fiber is lain over the actual airframe layup, separated by a peel ply layer, and then removed post-cure to enhance the smoothness and uniformity of the outer airframe surface. Below are the lengths for each section of the airframe and the layup configuration for each ply. Table 3: Body Tube Section Lengths and Fiber Orientation by Layer Section Lengths Fiber Orientation Section Length (inches) Ply Layer Orientation Payload Bay 18.5 Ply #1 (28 mil) 0 Degrees Main Bay 34 Ply #2 (28 mil) +45 Degrees Avionics Ring 2 Ply #3 (14 mil) 90 Degrees Drogue Bay 24 Booster Motor Mount Assembly The Rocket Rebels decided to explore a design option that was not used as often as traditional means of motor mounting. This subassembly relied more heavily on additive manufacturing to deliver a precisely dimensioned design that boasted light weight yet extreme durability. This was achieved by 3D printing the fins and centering rings, joined together by a central sleeve which is inserted over a length of blue UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 14

15 tube which serves as the actual motor mount tube. The following figure gives a good reference for the design features of the subassembly. Each component is discussed in more detail within the following subsections Fin Design Figure 3: Motor Mount Assembly 2D Drawing A minimal aerodynamic drag profile was the goal for the vehicle s fin design, while also maintaining its stiffness and rigidity. Simple research in fluid mechanics revealed that fins with an airfoiled crosssectional shape carried the least amount of drag on the vehicle. To achieve this shape over a much less aerodynamic square cross-section, the corners on the leading edge of the fins were rounded off while those on the trailing edge were tapered using a chamfer feature in CAD modeling software. This simple design change altered the apogee of the as-built flight vehicle by 800 feet. To prevent excessive bending due to launch loads and unexpected turbulence, the fins were widened to ¼ and were given sufficient length and height so that each fin could generally keep its rigidity when moderate bending pressure was exerted to them by hand. Weight considerations were also a factor, since more weight towards the aft of the vehicle generally made the vehicle more stable. This made 3D-printed fins a more attractive option for the team to increase vehicle performance. ABS material was a durable 3D plastic filament that had almost the same strength as aluminum but less than half the weight, making a prime candidate for material. Dimensional data found in section provides the exact specifications for the vehicle s fins. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 15

16 Motor Retention System Figure 4: Printed Fin Design The motor retention system consists of a small subassembly on the far rear of the rocket that prevents the motor from exiting the motor housing. The component is essentially a threaded ring with a cap that places a lip over the back of the motor to support it against the centering ring and airframe. The 75mm retainer is implemented to mate with the Loki L840CT motor that has been selected. The Aeropack retainer is mounted to the rear propulsion system centering ring via Allen screws that were inserted into holes created in the centering ring with a tap and die. The system ensures that the force of the motor is transferred to the airframe through normal forces that can be more easily sustained, as opposed to shear forces were the motor to be mounted via surface adhesion. Figure 5: Aeropack 75mm Motor Retainer Assembly UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 16

17 Centering Rings The centering rings and related interfaces had undergone significant design changes to more efficiently prioritize material. There was a need in the final design phase to reduce weight, and the centering rings were identified as an area where material could be removed while maintaining structural performance. Originally, a G10 fiberglass composite ring system was implemented. 1/8 rings mated to the OD of the motor housing tube and the ID of the airframe body tube. Since then, ABS plastic filament has substituted G10 fiberglass for its similar strength properties and smaller density. The new design features thicker rings (3/8 ) that have been gridded with lightening holes to create a spoked wheel style. The 6 spokes of each centering ring are either coincident with fin slots in the fin sleeve that align the fins correctly or are reinforced with shoulders to further prevent the centering rings from snapping. The 6 spokes of the centering rings are.375 thick, almost 3 times the thickness of the original, to better withstand bending moments and induced bending stress between the ID and OD. The fin sleeve allows for an incredibly strong bond to the motor housing tube by maximizing contact surface area. The rings and fin sleeve are printed as a single part, nominally preventing disconnects between the motor housing and centering rings. The material is an extruded ABS material which provides sufficient strength (~4000psi) and great flexibility in structural geometry design. This design allows a comparable strength at a much lower weight. The estimated mass savings over the G10 rings are significant. The whole system was proven in the first full scale test launch with no failures of any kind relating to the centering rings or propulsion assembly Motor Figure 6: Preliminary 3D Print of Fin Sleeve w/ Centering Ring The motor selection is obviously one of the most crucial parts of the propulsion assembly. The final motor choice for the Presidium vehicle is a Loki Research L840CT reloadable rocket motor. This L840 offers a max thrust of 1190N with an average thrust of 855N. Considered with the current build weight of UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 17

18 the full-scale at 31 lb, the thrust to weight ratio is 6.12, which is well above the threshold T/W ratio of 5. The burn time of 4.6 seconds is long enough to keep the maximum acceleration on the rocket and payload within an acceptable range. The L840 thrust curve can be seen below. The geometry of the motor is such that an acceptable fit is made with the motor mount system Coupler Tubes Figure 7: Loki Research L840 Motor Thrust Curve Graph These components are made from cuts of Blue Tube, a lightweight vulcanized fiber wrap that has exceptional tensile stress capacity. These couplers are capped with bulkheads made from carbon fiber plate material like the fins. The couplers are optimized to allow adequate space for the parachute and payload packages, but also be sized to approximately the diameter of the vehicle on either side of connection points to diminish the risk of bending stress on the vehicle. Blue tube is a shatterproof airframe tube that is extremely resistant to abrasion, brittleness, and cracking. However, the material is still capable of being machined using typical hand tools Bulkheads, Attachment Hardware, and Ballasts To create the bulkheads, 2-D drawings of the shape and sizing for the bulkheads were made in AutoCAD. For the vehicle design, the bulkheads were cut to 5.97 in diameter and then epoxied to the outside of coupler tubes. More specifically, four bulkheads are incorporated within the vehicle: one on the forward section of the booster coupler, one on either side of the avionics bay, and one on the aft end of the payload coupler tube. All bulkheads were epoxied to the coupler tubes with Loctite EA9394 epoxy except for the forward bulkhead on the avionics bay. This specific bulkhead was attached via two 3/8 threaded rods that ran through the avionics bay and on either side of the avionics sled within. The threaded rods ran through both bulkheads and were secured to the forward avionics bulkhead with lock nuts. The forward bulkhead, which is the mount for the avionics sled, was secured to the threaded rod with wing nuts. This allowed the team to access the avionics sled easily when opening the avionics bay. Attachment points in the rocket that are not epoxy bonded are fastened together with 6-32 machine screws. This includes the joining point of the nose cone to the payload body tube and the avionics bay to the drogue and main bay body tubes. 3 nylon screws (4-40) are used at each separation point of flight. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 18

19 Ballasts have been incorporated at the booster and payload sections of the vehicle to weigh the vehicle down so that it will accurately reach one mile. This component was found to be necessary after the first full scale test flight where a confirmed altitude of 6224 feet was reached. The ballasts are made of sections of a cast iron rod. These sections have been shaped to 2 ½ wide, allowing two sections to fit within the inner diameter of the vehicle and mount to the bulkheads of the booster and payload sections of the vehicle with thick machine screws. These ballasts are shaped using hand lathes available on the CME factory floor Nose Cone Figure 8: Ballast Construction and Final Component The Nose Cone is installed on the end of the payload section and was created in a Fortus fused deposition molding (FDM) 3D printer in the CME s manufacturing center. Subscale and full-scale flight test have confirmed that the material is robust enough to withstand launch loads and the duration of flight. The hollowed component is ¼ thick and made of standard ASA plastic filament, optimizing minimal weight and exceptional durability. This FDM material has substantial strength and thermal stability, making it ideal for advanced tooling and prototyping applications. The design of the nose cone utilizes a basic ogive shape to minimize drag coefficient. To prevent unnecessary drag from scarring and damage from testing, one nose cone print will be used for the duration of hardware check and static recovery testing while a separate identical nose cone will replace it right before launch. Dimensional data for the nose cone is provided in the table below: Table 4: Nose Cone Dimensional Data Nose cone shape Tangent Ogive Nose cone material ASA plastic filament UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 19

20 Length 8 in Base Diameter 6.2 in Wall thickness.25 in Shoulder Length 1.5 Figure 9: Final As-Built Nose Cone 4.3 Electrical Elements and Design Features Avionics Retention Though there are little electrical components to the vehicle outside of the avionics bay (discussed in section 4.7.2), it is still important to determine how the subsystem is retained within the vehicle. The avionics package is retained via an avionics sled that incorporates tabs at the bottom of the sled. Machine screws fasten the tabs to the aft bulkhead and prevent the sled from shaking during flight. Smaller nylon screws and nuts fasten the altimeters and battery packs to the sled to prevent them from shaking loose. Loctite EA9394 epoxy bonds are made to the interfaces between the bulkhead and coupler tube and machine screws integrate the subassembly to the airframe sections GPS Tracker The ABLEGRID TK102B GPS Tracker will locate the vehicle at 20 second intervals throughout the duration of flight. This ensures the electromagnetic waves emitted cause the least risk of interference to the avionics bay while keeping a constant signal with the tracker. It is positioned in the nose cone where it can transmit frequencies and has several layers of carbon fiber (an excellent EM wave blocker) between itself and the avionics. This device is discussed in more detail in section UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 20

21 4.4 Flight Reliability Confidence Mission Success Criteria The Rocket Rebel s vehicle mission is to design, manufacture, launch, and successfully recover a reusable high powered rocket. This rocket must successfully carry a fragile payload without damaging it throughout flight and landing. Also, the rocket must reach an apogee altitude of as close to one mile as possible. For this mission to be a success the following criteria must be achieved: 1. The launch vehicle must successfully lift off the pad at a minimum rail velocity of 52 ft/s. 2. The launch vehicle must attain an altitude of 5280 ft. AGL ±150ft. 3. The drogue/main chute must deploy within 2.0sec after apogee is reached. 4. The landing speed of the rocket after drogue deployment shall be approximately 70 fps to minimize drift. 5. The landing energy of the heaviest section shall be less than 75 ft-lb. 6. All parts of the rocket are recovered and are launch ready within three hours. 7. The FMPA and the fragile material enclosed must survive the entirety of the flight and remain intact upon retrieval Flight Reliability Given the overall design, current testing, and flight tests of the full-scale vehicle, the robustness and reliability of the design is very high. The following points can back the overall reliability: Suitability of Fin Shape and Mounting Techniques The mission requires that the rocket be recovered and launch ready within three hours as well as reach a target altitude of one mile, which means that the vehicle must have such a robust design that in can be flown again. This means the fins must be extremely aerodynamically efficient, creating the least amount of drag possible while still maintaining a stability caliber greater than 2.0 so that the vehicle can ascend as vertically as possible while withstanding all incurred launch loads. The current fin design meets all the above criteria. The overall stability is 2.71 and simulations show that the launch vehicle reaches a projected altitude of close to 5200 feet. The fins are epoxied in 6 different locations: twice to the motor tube, twice to each centering ring inside the body tube, and twice on the outside surface of the airframe. This technique has been tested in subscale and full scale test flights and has been shown to meet all flight criteria. This method of mounting creates rigid fins that resist fluttering and ultimately stable flight. Proper Use of Materials All materials have been handled and used as directed by the manufacturer, as well as being handled per MSDS sheets located in all shops and labs. Proper Assembly Procedures The Rocket Rebels Level 3 certified mentor, Cody Hardin, has overseen the construction of the subscale and full scale rocket, and will continue to oversee all operating procedures to be done in the future. The team will verify that all materials are assembled per the workmanship required to ensure that the rocket is built soundly. Sufficient Motor Mounting and Retention The team has decided to utilize the 76-mm flanged Aeropack retainer assembly mounted in the 3/8 aft centering ring at the bottom of the rocket to retain the motor. The thickness of the centering ring provides sufficient surface area to ensure that the force of the motor is transferred to the airframe through normal forces, as opposed to shear forces were the tube to be mounted via surface adhesion. This system of motor mounting has been UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 21

22 tested in a full scale flight with a Level L motor. Motor retention was completely successful and operated as expected. Design Simulations A multitude of simulations have been performed for the current rocket configuration, placing it under multiple environmental conditions. All flight results have proven to meet competition requirements and are further detailed in Section 4.8. Test flights The design configuration has been subjected to a subscale and full scale test described in detail in Section 4.9. Both flights were deemed successful. 4.5 Vehicle Construction The construction procedures of the Rocket Rebels vehicle, dubbed Presidium, have been compiled in as detailed a manner as possible so that even the most novice of rocketeers can create a replica of Presidium s likeness, given the Rocket Rebels equipment and materials were all at hand. The construction procedure is described in chronological order; however, it is possible for some steps to be completed simultaneously amongst the different subsystems of the vehicle Airframe Construction Body tube construction was completed using composite manufacturing techniques that consisted of three main stages: mandrel prep, layup, and part cure. Before the airframe construction was started, a detailed inventory was taken of all material required and team members verified that enough material was available to complete construction within the deadline requirements. Two 6 wide aluminum tubes (1/8 wall thickness) were used as mandrels for the body tube construction. The tubes were 8 feet and 4 feet in length, respectively. Keeping safe composite layup practices, 6 inches from each side of both mandrel tubes was not covered in layup. Therefore, the 8-foot mandrel would yield a maximum of 7 feet of airframe tubing, while the 4-foot mandrel would yield a maximum of 3 feet of airframe tubing, totaling 10 feet of airframe tubing per layup cycle. Table 5: Airframe Ply Lengths and Orientations per Mandrel Long Mandrel Ply 1 Ply 2 Ply 3 Ply 4 Material Thickness 28 mil 28 mil 14 mil 28 mil Fiber Orientation Length 84 in 84 in 84 in 84 in Width in in 19.2 in in Short Mandrel Ply 1 Ply 2 Ply 3 Ply 4 Material Thickness 28 mil 28 mil 14 mil 28 mil Fiber Orientation Length 36 in 36 in 36 in 36 in Width in in 19.2 in in UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 22

23 The table above shows the dimensions of the carbon fiber plies used to complete the layup. The first three plies of the layup consist of the airframe and are stacked in alternating orientations of 45. This allows any arbitrary element of the airframe tube at any orientation to have nearly uniform normal and shear stress limits. The fourth ply serves as a non-reusable caul plate. Its purpose is to squeeze all the inner layers down to a smooth and uniform surface area to decrease aerodynamic drag. The measurements of the plies were marked out on the available material donated from Orbital ATK, paying special attention to the degree of each ply. An automated Eastman cutting table was used to cut the plies for this launch vehicle, however careful manual cutting with tin snips is also an acceptable means to cut the plies to the correct dimensions. To use the automated cutting table, 2D drawings had to be created and positioned in the cutting plane recognized by the machine. A cutting path was then automatically set by the machine and the machine ran the path file to cut the prescribed ply sizes. The first construction stage was to prep the aluminum mandrels for the hand layup. The main goal of this procedure is to prepare the surface of the material so that the layup does not permanently bond to the mandrel during cure cycle. This is done via several applications of sanding, acetone, and release agent. First, Scotch Brite pads were rubbed with moderate pressure on the outer surface of the mandrel until most or all surface defects had been smoothed. This was followed with a generous application of acetone to clean the surface of all aluminum dust and other impurities. This was done until a paper towel was still clean after wiping it across the entire mandrel. The next operation is the application of chemical products produced and donated by Chem-Trend, the first of which to be applied was called Zyvax SealProof. This mold sealer delivers a durable, visible film that blocks prepreg resin from bonding to the mandrel tube when applied in conjunction with release agent. The first coat was applied to the surface of the mandrel by soaking a paper towel in the sealant and wiping the surface in continuous strokes in one direction. After the coat was applied uniformly, it coat was set aside to cure for 60 minutes. This process was repeated until 3 full coats of sealant had been applied. Zyvax 1034W Mold Release, another chemical produced and donated by Chem-Trend, followed the application of SealProof. A fresh paper towel was wetted with the release agent and wiped across the entire surface of the mandrel in one direction. After two minutes or until the layer could be visibly seen beading up on the surface, a fresh paper towel dried the surface. The coat of release agent was set aside to cure for 15 minutes and the application process was repeated two more times. The mandrel prep stage of the airframe construction was completed after the top coat of release agent was cured. Refer to Appendix A for MSDS of the SealProof and Mold Release Agent. The next stage is the hand layup stage. This stage is crucial in that each ply of the carbon fiber material must be lain as smoothly and uniformly as possible. This ensures uniform characteristics in the end item part after the cure cycle is complete. After mandrel prep is completed, the panels that make up the first ply (28 mil, 0 orientation, in wide) were inspected for any apparent surface defects or foreign object debris (FOD). If satisfactory, the middle of the panels was placed over the mandrel about 6 from the edge and the ends were wrapped completely around the mandrel to verify that they meet square. Then starting from one side, the panel was firmly rubbed down its entire length. All air pockets or wrinkles that may have formed when the material was applied to the mandrel were inspected and smoothed out by hand. After all panels of the first ply on both mandrels was placed, the layup went through a debulk process, which is a shorthand phrase for placing the current layup under vacuum. This procedure allows the outer most ply of the layup to be squeezed more compactly against the mandrel tube, mitigating future surface defects like wrinkles or gaps in the layup. The debulk is done by tightly wrapping the layup in FEP plastic, allowing no air pockets to be visible. This is followed up with a layer of breather cloth, with an extra folded square of breather cloth and the bottom coupler of a vacuum port tapped off near the end of the mandrel. Masking tape was used to secure the FEP and breather cloth in place. Once this was done, UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 23

24 the 6 ends of each mandrel tube were then scrubbed with Scotch Brite pads so that putty tape will stick to the mandrel. Next, a segment of bagging material was cut, roughly 6-7 longer than the layup and about a foot wider than the circumference of the mandrel. Strips of putty tape were applied to the two shorter sides and one of the longer sides of the bagging material previously cut. Placing the mandrel in the center of the bag, the putty tape was firmly pressed to the mandrel, ensuring a firm bond. The bag was then wrapped around the mandrel, applying putty tape to the mandrel while avoiding it from touching the breather cloth. Once the bag met at the top of the mandrel, the last open seam was sealed together, creating a pleat that extends from the mandrel. This pleat will pull excess air away from the surface of the layup creating a smooth outer surface on the ply. Any obvious gaps in the bag were filled to create a seal and a small puncture was made in the bag right over the vacuum port coupler. This allows the top portion of the vacuum port to be installed to the layup. After attaching the vacuum port, the bagged layup was put under about 14.7 psi for 30mins. Once the layup had been under pressure for the required time, the bag, breather cloth, and FEP were removed. The debulked layup was then inspected for any surface defects until satisfactory. Figure 10: Completed Debulk Bag Construction This process continued for the first three plies which would consist of the airframe. After the third ply was debulked, a thin, plastic, peel ply cloth was stretched tightly over the entirety of the layup and UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 24

25 secured with Teflon tape. The 1034W release agent was then applied (two coats) in the same manner as performed during mandrel prep. Both components acted as a mold release to prevent the resin impregnated caul plate sheets from bonding to the layup. After letting the release agent cure, the fourth and final ply was applied to the mandrel in the same manner as the previous plies and debulked. As shown in the following figure, defects occurred during the debulking stage. These gaps in the layup were filled with thin strips of scrap carbon fiber prepreg material and the layup was debulked again to smooth out the surface anomalies. Figure 11: Debulked Airframe w/ Defect The layup is now ready for the curing stage. Curing of the HexPly 6376 prepreg sheets was done by following the instructions provided by the manufacturer as precisely as possible. The layup was prepared in the same bagging configuration used during debulk processes. Per the manufacturer s curing instructions, provided in Appendix A, the layup was to be cured for two hours at 350 F and under seven bars of pressure. The maximum pressure differential provided by the CME s curing oven was approximately two bars. However, the stress limits of carbon fiber cured at this pressure were still far more robust than was required for the projected flight. The curing cycle consisted of four cycles: a purging stage that prepped the oven for high temperature operation, a ramp up stage of 5 F per minute up to 350 F, a steady cure for two hours, and cool down. After curing and ensuring the oven had cooled to safe temperatures, the mandrels were removed from vacuum and placed on a long table. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 25

26 Figure 12: Final Bagging and Seal Check Before Cure The mold release agent applied in mandrel prep allowed the cured body tubes to be easily extracted from the mandrel tube. The caul plate had to be removed to ensure that the outer surface of the body tube was smooth and compact. To remove this ply, a Dremel rotary tool with a high speed cutting wheel was used at approximately 15,000 rpm to slowly cut into the surface of the tube until the peel ply layer was exposed. A flathead screw driver was used to pry under the cut caul plate to help pull it up from the cured layup. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 26

27 Figure 13: Post Cure Layup Sections with and without Caul Plates Once the caul plate had been fully removed, measured cuts were made with a Dremel rotary tool to create the rocket section lengths designed through the OpenRocket model. To ensure a square cut, the body tube sections were placed back onto the mandrel tubes with one end nearly flush with the square cut of the mandrel end. A die-grinder was then used to sand the body tube ends flush to the end of the mandrel tube. Small pits in the fibers and resin matrix created a semi-rough texture on the outer surface. These pits were filled by applying a hearty coat of finishing putty to the entire surface of each body tube and then sanding the excess away, leaving a smooth outer surface on the body tubes. The final body tube operation was powder coating. Glossy white paint powder was generously sprayed onto the body tubes under a vent hood and the tubes were then transferred to the curing oven again. The painted body tubes were cured at 400 F for 10 minutes and removed after allowing to cool. Figure 14: Segment of Extracted Carbon Fiber Tubing UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 27

28 Figure 15: Final Surface Prep and Powder Coating Operations Figure 16: Finished Airframe Sections UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 28

29 4.5.2 Nose Cone and Fin Procurement Both components were made through additive manufacturing by using the Fortus 3D Production System, a fused deposition modeling printer. The components were made with ASA print filament and utilize a.01 tolerance to all dimensional specifications set. CAD models were made for each of the components shown in the following figures. To create the fin configuration, a sleeve and centering ring subassembly had to be created first that would fit the outer diameter of the motor housing and could accommodate slots that would fit the fin tabs. This part was epoxied between the interfaces of the inner diameter of the sleeve and the outer diameter of the motor housing as well as between the outer diameter of the centering rings and the inner diameter of the booster body tube. As can be seen in the figure below, the centering rings were made in a skeleton style to lighten the weight of the vehicle. Supports were added to spools of the centering rings that were not aligned with fin slots. This was to give added shear support to the centering rings during flight. Figure 17: Dimensional Drawing of Fin Sleeve and Centering Ring Part Fins were designed to create the least amount of drag resistance and increase the stability margin to acceptable thresholds. A 45 wire orientation was automatically selected by the machine, relative to the root chord. The leading edge of the fins were chamfered to give an airfoiled shape, while the trailing edge was given a tapered edge. This is the ideal fin shape for all subsonic flights to minimize roll and drag coefficients. Dimensional data and CAD drawings can be seen below: Table 6: Fin Configuration and Dimensional Data Fin Style Trapezoidal Number of Fins 3 UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 29

30 Root Chord 9 in Tip Chord 4.5 in Fin Height 7.5 in Sweep Length 2.25 in Thickness 1/4 in Material ASA Figure 18: CAD Model of Fin w/ Dimensional Data The nose cone was designed so that the shoulder would sit flush with the inner diameter of the payload body tube and the outer surface of the nose cone would meet flush with the outer surface of the body tube. This required a gap that was of equal thickness as the wall thickness of the airframe. To reduce weight yet still have adequate structural integrity, the nose cone was hollowed out to ¼ thickness. Below is the dimensional data and CAD model for this component. Table 7: Nose Cone Specifications Nose cone shape Tangent Ogive Nose cone material ASA plastic filament Length 8 in Base Diameter 6.12 in Wall thickness.25 in Shoulder Length 1.5 UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 30

31 4.5.3 Coupler Subassembly Figure 19: Nose Cone Dimensional Drawing This portion of the rocket construction consisted primarily of the blue tube coupler tubes, 1/8 carbon fiber plate material, and steel attachment hardware. The construction of the avionics bay is included in the following section (4.5.4). This subassembly was begun by procuring the correct dimensions for each part of the subassembly. These steps were completed independent of each other Using safety glasses and gloves, a vertical bandsaw was used to cut blue tube with a 5.97 OD into three tube segments of 9 each. The edges were sanded smooth and deburred of rough edges. The dimensional gap between the inner diameter of the body tubes and the outer diameter of the coupler at this point was.06 inches. This was due to the expansion of the mandrel during the body tube cure cycle, which increased the diameter of the body tube by.03. To minimize this gap for a snug fit, a quick fiberglass layup was performed. A layer of fiberglass cloth was wrapped around the coupler tubes and resin was spread evenly throughout the cloth layer. After a 24-hour cure at room temperature, the outer diameter of the rocket was found to be At the same time, a large sheet of 5K plain weave carbon fiber plate material was placed on a Water-Jet cutting machine. A 2-D CAD drawing modeled in AutoCAD was transferred to the Water-Jet with a nest of four circular bulkheads. One was dimensioned to be 5.97 and the other three were dimensioned to be The slightly smaller bulkheads were to fit flush with the inner diameter of the blue tube while the UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 31

32 slightly larger bulkhead would fit flush with the outer diameter. The sheet of plate material was clamped down to the cutting surface and the safety windows were locked in the upright positon before the automated cutting operation was started. Figure 20: Water-Jet Cutting Procedure for Full-Scale Bulkheads Two 3/8 U-bolts were then selected and positioned atop two of the smaller bulkheads which were then designated as the booster and payload bay bulkheads. The U-bolts were positioned to be situated in the middle of the bulkheads and the two threaded ends of the U-bolt were marked. These marks were drilled with a drill press, using a ¼ bit as a pilot hole and finishing with a 3/8 drill bit. The U-bolts were inserted through the holes and lock nuts were threaded on each end to secure them permanently to the bulkheads. Two coupler tubes were then designated as the payload and booster coupler tubes with one bulkhead designated to each tube. A half batch of the 2-part Loctite EA9394 epoxy was prepared in a small mixing cup, which consisted of 50 grams of part A and 8 grams of part B. A 90-minute timer was set to keep track of the pot life and the curing oven was preheated to 150 F. Using a wooden popsicle stick, a thick UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 32

33 bead of epoxy was applied to the first ½ of the inner diameter of the booster and payload coupler tubes. A bead of epoxy was also applied to the outer diameter of the designated bulkheads. The bulkheads were inserted through the opposite end of the coupler tubes that contained epoxy and were pushed to the other end. Slight adjustments were made so that the bulkheads sat flush with the end of their respective coupler tube. The interfaces of the bulkheads to the coupler tubes were inspected for gaps and filled with epoxy as necessary. To ensure adequate surface bond, a thin fillet was applied to the inner surface of the coupler tube and the inside surface of the bulkhead. The coupler tubes were then transferred to the curing oven, cured for one hour, and then removed. After the cure cycle, the coupler tubes were marked at the halfway point of 4.5 and another half batch of epoxy was made. A thick bead of epoxy about 2.5 was applied to the inner diameter of the payload and booster body tubes. A bead was also applied to about 2.5 of the outer diameter of the coupler tubes, opposite the side with the bulkhead attached. Each coupler tube was then inserted, with the bulkhead up, into the body tube until the halfway mark was flush with the body tube. The subassembly was then placed in the oven again, cured for one hour, and removed Avionics Configuration Figure 21: Completed Coupler Tube Subassembly Full assembly of the avionics bay was performed pre-flight to ensure appropriate altimeter behavior was achieved. The first subcomponent to be created was the avionics sled. A 3-D model of the sled (shown in section ) was created through the PTC Creo modeling software and was converted to an STL file. This ensured the most detailed dimensional data that would be transferred to the print. After installing the file to the Fortus 3D production system, the automated machine printed the model as designed machine screws included with the battery holders secure the holders to the plate on the back side of the sled. Standard 9V batteries easily clip into the seat of the holders. The spacer orientation of the two altimeter mounts on the other side of the sled make it evident of where each altimeter is to be installed nylon screws were mounted through the altimeters and into the spacers printed on the sled and are secured on the backside with nuts. To ensure the sled is stable within the avionics bay, a tab was created on the UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 33

34 bottom side. Holes sized for 4-40 machine screws were drilled and tapped through the tab and the aft avionics bulkhead so that machine screws could be screwed through and fastened to the bulkhead. Figure 22: Printed Avionics Sled with Altimeter and Battery Attachments The coupler tube was then marked so that a 2 segment was positioned in the middle of the outer diameter of the coupler tube. A half batch of epoxy was prepared like that described in the previous section and a bead was spread ½ along the inside of one end of the avionics coupler tube, designated as the forward end. A bead was also spread over the outer diameter of the last smaller bulkhead, designated as the forward avionics bulkhead. The two interfaces were joined and a small fillet was added to the inside interface of the bulkhead and coupler tube. A thick bead was then spread on the 2 segment of the outer diameter of the coupler tube and along the entire inner diameter of the avionics ring. These two interfaces were then bonded and any gaps found were filled with small amounts of epoxy. This subassembly was then placed in the oven, cured for one hour at 150 F, and then removed. The following pattern of 3/8 holes were drilled through both avionics bay bulkheads that would house the rest of the attachment hardware. Two symmetric holes, equally spaced 1.5 from the center, were drilled in each avionics bulkhead and 3/8 threaded rod was fitted through them. This rod would connect the free aft bulkhead with the avionics sled to the fixed forward bulkhead that is bonded with epoxy to the avionics bay. Lock nuts secured the threaded rod on the forward end, while wing nuts secured the threaded rod through the aft bulkhead so that it sealed the avionics bay closed throughout flight. Two more 3/8 holes were drilled symmetrical to the center of the bulkheads and oriented 90 degrees from the previous hole pattern. U-bolts were fitted into these holes and tightened down with lock nuts. Two 1 holes were then drilled towards the outside of the bulkheads symmetric with the center and PVC adapters were threaded through and secured with threaded caps. A small hole was drilled through the cap. This element allowed for ejection charges to be snugly fitted to the bulkhead and easily be sealed from the avionics bay. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 34

35 Figure 23: Avionics Bay Bulkhead 20-gauge, multi-stranded copper wire was taken in two strips of about 10 inches. One end of each segment was stripped bare and soldered to the connection points corresponding to the 110 position of a two-port rotary switch. This was representative of the ON position. The two wires were then twisted around one another to form a single continuous wire. The ends of this wire were stripped and connected to the switch ports of the StratoLoggerCF altimeter. The same process was repeated for the RRC3 altimeter. Pre-installed leads from the battery holders were then installed into the battery ports of each altimeter. A series of beeps from each altimeter verified each circuit was closed and continuous. After the altimeters finished pre-launch announcement codes, the rotary switches were moved to the 220 position representing OFF. Figure 24: Rotary Switch, Battery, and Altimeter Configuration UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 35

36 The wires in the switch ports were removed from both altimeters and the 150 and 210 azimuth of the avionics ring were designated. These locations were selected under the assumption that the launch rail would rest against the 0 azimuth of the vehicle. Using a 7/16 drill bit installed in a drill press, holes were created in the avionics rings at the two designated locations. The key switches were snugly threaded into these holes and nuts were threaded from the inside of the avionics bay. Four fingers of a latex glove were cut and placed aside. On a scale, the appropriate amount of 4F black powder for the drogue charge was measured and placed in one of the glove fingers. An e-match was inserted within the black powder charge and the glove finger was twisted closed and fastened with electrical tape. The opposite end of the e-match was then fed through one of the PVC attachments made in the aft bulkhead and excess wire was trimmed. The ends of this wire were stripped and installed into the drogue ports of the StratoLoggerCF altimeter. The same process was performed with the RRC3 altimeter through the other hole in the aft bulkhead. The main ejection charge was then measured and assembled like the drogue charge, however, the leads of the e-match were fed through the PVC attachments of the forward bulkhead and installed within the main ports of the altimeter. 3M Fire Block Sealant was then used to fill the PVC attachment up to the ejection charge. This seals the avionics bay from the blast of the ejection charges and prevents unwanted pressure changes in the avionics bay. Figure 25: StratoLoggerCF Altimeter Schematic With care, the avionics sled was installed to the aft avionics bulkhead and slid into the avionics bay. The threaded rods were fitted through the bulkhead and tightened with wing nuts. Figure 26: Completed Avionics Setup UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 36

37 4.5.5 Parachute Packing Drogue parachutes were folded with the suspension lines stacked and folded over the middle of the folded canopy. Then the canopy was folded in an S-shaped style and secured with a small piece of masking tape. The main parachute recovery system was lain completely outstretched to begin packing it within the main bay. The parachute, suspension lines, and risers were inspected for knots and entanglements before folding. A small weight was placed at the end of the riser furthest from the parachute to fix one end of the recovery system. From the location where the suspension lines meet the riser, the furthest right suspension line was isolated and traced up to the parachute. This section of the parachute was placed furthest to the right than the remainder of the parachute. The next closest suspension line was then isolated and placed on the first suspension line placed, causing the section of parachute between the two lines to fold in half over itself. This process of stacking suspension lines and folding the adjacent parachute segments together was repeated until half of the suspension lines (8) were stacked together. The remainder of the parachute was loosely flipped over the top of the parachute while maintaining the stack created previously. The suspension line immediately to the left of the first stack was placed over the suspension line stack and the adjacent parachute segment was folded over itself to the left of the suspension line stack. The stacking process was repeated on the left side for the remaining 8 suspension lines. The following figure shows the completed stack of the semi-folded parachute. Figure 27a: Main Parachute Canopy Folding The suspension line stack was then inspected for the point where the line insertions were located. Pinching the stack at this point and halfway between the canopy, the stack segment was folded in half and placed in the center of the canopy stack. Figure 27b: Main Parachute Suspension Line Fold 1/3 of the canopy was then folded over the suspension line segment on both sides. The cord that ties the deployment bag to the harness was then folded over the top of the canopy stack and the canopy was folded in half on top of it. Holding the stack together with a knee or an assistant, a 12-inch length of cord was fed through the red ribbon located on one of the flaps of the deployment bag. The canopy stack was folded into an S shape to about the length of the deployment bag and then snugly installed within the deployment bag. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 37

38 Figure 27c: Main Parachute Canopy Packing The small length of cord and red ribbon were then fed through the grommet of the bag flap opposite them. The length of cord was removed while placing a pin through the red ribbon to hold the ribbon through the grommet. The small cord was wrapped around the suspension line stack and fed through the ribbon again, pulling about 2 inches of the suspension line stack through the ribbon with it. The cord was removed again and wrapped around the suspension line stack. The cord and suspension lines were then fed through the first rib sleeve of the deployment bag, without pulling any of the 2-inch suspension line segment from the ribbon. The suspension lines were folded through all the following rib sleeves on the flap. Figure 27d: Main Parachute Suspension Line Packing Once the second set of insertions were reached during packing, the remaining suspension line was folded in the same manner as the previous segments but left outside of a rib sleeve. The entire flap was then folded over itself in an accordion style into the bag and the riser section was folded in an S-pattern over itself up to the pull pin insertion. The small length of cord was fed through another red ribbon located at the top of the folded rib sleeve flap. The ribbon was threaded through the three remaining flaps of the deployment bag and the metal pin on the riser insertion was inserted into the ribbon to fasten the bag together. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 38

39 Figure 27e: Main Parachute Final Bagging Before packing the bag into the vehicle, the length of riser cord that led to the pin was inspected to ensure that it was shorter than the segment of riser protruding from the bag to the insertion point. This ensures the force of the pilot chute deployment properly pulls the pin from the ribbon and allows the deployment bag to unfurl the parachute. Figure 27f: Riser Length Inspection UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 39

40 4.5.6 Pre-launch Assembly For the rocket to be at full flight-ready configuration, small miscellaneous steps had to be completed prior to beginning the launch procedures detailed in section 7.2. Fin slots had to be machined away from the booster section and rail buttons had to be installed before any bonding could be performed. To help cut straight fin slots, a cutting tool was printed from the Fortus machine that served as a guide for a Dremel tool along the body tube. This tool is shown in the following figure. Figure 28: Fin Slot Cutting Tool Dimensional Drawing Once three fin slots had been cut 120 radially and 9 long, the holes for the 1515 rail buttons were then drilled using a standard power drill. The lower rail button was installed by drilling a 3/16 hole in the booster section at 0 azimuth and half the distance up the fins, or at 4.5. The upper rail button hole was drilled two feet above this point along the rocket stack in the drogue body tube. The upper rail button was installed into this hole. The lower rail button was not installed until after the fin sleeve had been inserted and bonded into the booster. At this point, all booster components were procured and prepped for subassembly. A batch of two-part Loctite 9394 Slow Cure Epoxy was prepped by using popsicle sticks to spoon out 100 grams of part A into a cup zeroed on a small kitchen scale, followed by 17 grams of part B (117 total grams in the cup). The components were mixed together until uniform in texture and a timer for 90 minutes was started to UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 40

41 keep track of pot life. A bead of epoxy was spread across the entire bottom and side surface of all three fin tabs and along the fin slot of the fin sleeve installed on the motor mount tube. The fins were fitted together into the slots and fillets were made over the concave edges between the fins and the fin sleeve. A think bead was then spread over the entire outer diameter of both centering rings and the first and ninth inch of the booster body tube (both sides of the fin slots cut). The motor mount and fin subassembly was fitted within the booster body tube until the fins were flush with the ends of the fin slots. An extra bead of epoxy was spread over the interface between the edge of the body tube and centering ring. Outer fin fillets were then constructed by laying Teflon tape around the fins with about a ½ gap between the fin and tape. The remaining batch of epoxy was spread along this gap to cover the concave edge between fins and body tube. The fillets were evenly spread smooth and the Teflon tape was slowly removed to leave clean and even fillets. The completed subassembly was then placed in the oven to accelerate cure for one hour at 150 F. The booster section was removed and, after cooling, the lower rail button was installed into the hole made previously. To mount the rocket in its complete stack on launch day, machine screw and shear pin holes had to be drilled in the correct body tube locations. This procedure was started from the bottom to make it easier to detect alignment errors and make corrections. The interfaces between the booster coupler tube and the drogue bay body tube, as well as the interfaces between the payload coupler tube and the main bay body tube were constrained via 4-40 nylon shear screws. The holes drilled in the body tubes were sized to be 90% of the thread diameter of the nylon screws so they could be worked into the hole without bending or shearing while still maintaining a snug fit. Machine screw attachment points included the avionics bay coupler tube to both the drogue bay and main bay body tubes as well as the payload bay body tube to the nose cone shoulder. These holes were drilled to 75% total diameter to preserve material for the screws threads to interface with. A Phillips-head screwdriver was used to tighten the screws to the body tubes. After all bonds for the vehicle had been completed, spray paint was applied to the 3-D printed features of the rocket to create a secondary color scheme for the rocket. These parts included the nose cone and fins. The nose cone was capable of being unscrewed and painted independently while the fins had to be taped off from the rest of the rocket to prevent paint being applied to the body tubes. 2 coats of paint primer were applied, followed by 2 coats of powder blue spray paint, and 1 layer of clear gloss topcoat. These layers were applied 30 minutes apart from each other to ensure enough drying time between each layer. To minimize pre-launch operation time, the shock cord is folded the day before each launch. To fold the shock cord properly, the two length of tubular nylon are laid flat on the floor. Starting at the looped end farthest from the parachute attachment loop, each shock cord is folded over itself in flat lengths of about 12 inches to form a stack. After the drogue parachute attachment loop is reached, the stack is held in place with a small strip of masking tape. Another stack is created on the other side of the attachment loop and held in place with a strip of tape. 4.6 Recovery Subsystem The recovery system for the launch vehicle involves a two-phase parachute deployment with the drogue parachute being released at apogee and the main parachute being released at roughly 900 ft. Deployments involve the detonation of black powder charges initiated by a StratoLoggerCF altimeter, which collects altitude, temperature, and battery voltage data throughout the flight. A Missileworks RRC3 Xtreme altimeter serves as a backup altimeter, triggering its own set of detonation charges in the unlikely event that the primary altimeter should fail. The final recovery system design has not experienced any drastic modifications over the course of the vehicle design improvement process. The dual parachute deployment system has been deemed capable of UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 41

42 safely recovering the launch vehicle and has been determined as the best available recovery option. The drogue parachute deployment provides an acceptable velocity of 70 ft/s for most of the descent, ensuring that the rocket will not drift more than 1100 ft from the launch site, presuming that the wind speed is no more than 20 mph. The second portion of the descent, beginning 900 ft AGL, is governed by the main parachute, which will drift the rocket to the ground at approximately 20 ft/s. The following block diagram shows the recovery events of the vehicle in chronological order. Launch is sensed by altimeter Rocket reaches apogee Primary drogue ejection charge fires. Backup drogue ejection charge fires 2 seconds later. Booster coupler and drogue bay body tube seperation Drogue parachute deploys Rocket descends to specificed main chute altitude Primary main ejection charge fires Backup main ejection charge fires 100 feet lower. Payload coupler and main bay body tube seperation Main parachute deploys Main parachute descent to ground Touchdown Figure 29: Recovery System Events Flowchart Physical separation is achieved via controlled detonation of pyrotechnic charges located on the forward and aft avionics bay bulkheads to create pressure inside the parachute bays and push them apart from the separation points. Nomex blankets are situated between the charges and the parachute bundles to prevent them from burning or melting before escaping their respective bays. The pyrotechnic charges consist of 4F black powder charges specifically sized through calculation to create about 213 lbs of separation force per charge. This force is adequate to shear 3 #4-40 nylon screws at each separation point. Each deployment event detonates a primary and backup charge, each capable of independently forcing separation and controlled by their own altimeters. These will be fired 100 feet in altitude apart in the unforeseen occurrence of a misfire or other anomaly in the primary altimeter. The following equations are used to determine ejection charge and static port sizes. Static deployment testing was performed before all vehicle launches to ensure the calculated black powder charge masses were enough to induce rocket separation. This testing is elaborated in section Black powder exhaust gases behave as ideal gases so the ideal gas law is used to determine the amount of black powder charge needed: PV = nrt lb in R = 266 lbm R T = 3300 R The black powder exhaust gases can be simplified to exert force only on the rocket s bulkheads so that the pressure exerted can be calculated: UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 42

43 Area = πd2 4 = π(5.8)2 = 26.4 in 2 4 Pressure = P = F A 213 lb = = 8.06 psi 26.4 in2 Volume can be calculated as a function of the area and the chamber length of each parachute bay. The volume of the parachute, shock cord, U-bolts, and blast caps can be ignored. V drogue = A L drogue = (26.4 in 2 )(16 in) = in 3 V main = A L main = (26.4 in 2 )(26 in) = in 3 The mass of 4F black powder needed for each chamber can then be calculated using the ideal gas law: And PV drogue = n drogue RT (8.06 psi)(422.4 in 3 lb in ) = n drogue (266 ) (3300 R ) lbm R n drogue = lbm = 1.77 g PV main = n main RT (8.06 psi)(686.4 in 3 lb in ) = n main (266 ) (3300 R ) lbm R n main =.0063 lbm = 2.85 g As good practice, the actual amount of black powder used is the calculated amount rounded up to the nearest gram and +1 gram to ensure a successful separation. Thus, the drogue charge will contain 3 grams of black powder and the main charge will hold 4 grams. The altimeter static ports also must have a specific diameter to assure accurate barometric readings for the altimeter and accurate parachute deployment events. The rocket incorporates four static ports which are each.192 inches in diameter. This value was calculated by the equation: D N = D T L N = ( )(5.775) 9 =.192 in 4 In this equation, D T is the inner diameter of the avionics bay, L is the length of the bay, and N is the number of static ports desired. The final value is rounded up to.219 (7/32 ) to accommodate conventional drill bit sizes Structural Elements Avionics Sled The most unique and essential part of the recovery system is the avionics sled, which mounts and retains the altimeters and batteries for the vehicle. The sled is made of durable ASA plastic filament through the Fortus 3D Productions System. CAD models were created which dimension the sled to perfectly fit the altimeters and battery mounts onto printed spacers or mounts. The ASA material has been proven to withstand launch loads as part of the nose cone during subscale and full scale test flights, furthermore proving to survive flight as the material of the avionics sled as well. Below is the CAD model and general dimensional data for the sled. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 43

44 Attachment Hardware Figure 30: Avionics Sled CAD Drawing 3/8 U-bolts are used to attach 7/8 shock cord to each vehicle section that is to separate. One U-bolt is fastened to each bulkhead of the avionics sled, to the payload bulkhead, and to the booster bulkhead. The shock cord is attached to these U-bolts via ¼ quick links and swivel eye bolts attach shock cord to the risers of the parachutes. Avionics bay bulkheads are retained with two 3/8 threaded rods that run through the avionics bay and fasten on opposite sides of either bulkhead. Locknuts secure the aft side to permanently fix the bulkhead to the avionics bay. Wing nuts are used on the other side so that the avionics sled can be accessed before and after flight, yet the system still be reliably retained within the avionics bay Recovery Harnesses and Packing Each parachute is encased in an 18 x 18 Nomex blanket that prevents the chute material from being burned by hot ejection gasses. The drogue parachute attaches to a 30 ft, 1 tubular nylon shock cord that is secured to the rocket sections by 3/8 aluminum U-bolts; one on the booster section and one on the avionics section. The shock cord tethers the drogue parachute to the aft end of the avionics bay and the forward end of the booster bay. Materials used are 0-3 cfm, 1.1 oz PIA-C-44378D type IV Rip-Stop Nylon for the canopy, type 3 nylon tape for reinforcement, Dynema for suspension lines, tubular nylon for risers, and size E nylon thread. These materials are lightweight but also strong enough to safely return the rocket to the ground. The main parachute incorporates a para-pak cloth deployment bag that is attached to a 30 ft shock cord which are secured to the aft end of the payload bay and the forward end of the avionics bay via 3/8 aluminum U-bolts. A polyconic pilot chute with Cd of.85 is located at the leading end of the main UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 44

45 recovery package and is attached directly to the top of the deployment bag. As the rocket splits and the pilot chute is in the stream of air, it inflates and starts pulling the deployment bag. The pilot chute pulls the deployment bag releasing the suspension lines in an organized manner. Before the canopy is extracted from the deployment bag, a closing loop secured with a tight loop of suspension lines ensures proper line stretch before the canopy feeds out. When the canopy comes out it wants to immediately start inflating. This could result in a high intensity force exerted on a small portion of the parachute which could case failure so inflation is slowed by a tie from the apex of the canopy to the deployment bag/pilot chute. The pilot chute gives a final pull on the canopy to elongate it. The main then starts inflating and a slider consisting of grommets and rip stop nylon travels slowly down the suspension lines to control the speed of deployment. The main parachute design is unique in that a vent opening is situated at the center of the parachute. This design has one of the highest drag coefficients of all parachutes and one of the softest openings. The main recovery package also incorporates Dynema suspension lines, 1.1oz rip stop nylon and a slider as previously mentioned. As with the drogue chute, the main parachute is covered in a Nomex blanket to ensure it is unscathed during the flight and deployment. Figure 31: Recovery Package and Shock Cord Configuration UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 45

46 Figure 32: Inflated Annular Parachute Design Electrical Elements and Design Features Altimeters The PerfectFlite StratoLoggerCF is the designated primary altimeter for the vehicle and the MissleWorks RRC3 Xtreme is the backup altimeter in case of unforeseen deployment errors of the primary. The primary altimeter is programmed so that the drogue parachute will deploy at apogee and the backup altimeter will deploy its drogue charge 2 seconds after apogee. At 900 feet, the primary altimeter fires the main parachute, followed by the backup altimeter firing at 800 feet in case of misfire. These altimeters, shown below, are commonly trusted devices used in the rocketry field and are certified to operate at the projected altitudes the vehicle is to reach. Figure 33: StratoLoggerCF and RRC3 Altimeters UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 46

47 Key Switches Two rotary switches, one for each altimeter, control the operation of the avionics from outside the vehicle. The switches are mounted through a small 2 ring that is set in the middle of the avionics bay and are each retained by a threaded nut that fastens them to the wall of the ring. The switches have two operating positions that represents the ON and OFF position. Electrical wire is soldered to the thin metal prongs that correspond to the ON position, which allows for a continuous circuit for the altimeters to operate. A flathead screwdriver is used to rotate the switch between each operating position from outside the vehicle in launch-ready configuration. Below is a figure of the rotary switch and its installed configuration. Figure 34: Rotary Key Switch Wiring Scheme Figure 35: Rotary Switch Installation 20-gauge electrical wire is used to transfer current through the avionics system. Below is the complete electrical schematic for the avionics bay. Altimeter 1 and 2 refer to the StratoLoggerCF and the RRC3, respectively. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 47

48 Figure 36: Avionics Bay Electrical Schematic Redundancy Features The avionics bay contains two independent circuits, one for each altimeter installed. The circuit that each altimeter is a part of consists of a small LED that is connected in series with the altimeter. When each altimeter is armed, the altimeter sounds a series of beeps, denoting pre-launch status, and the LED illuminates. Each circuit contains an independent battery source and a designated key switch that allows for each altimeter to be armed outside the launch vehicle Parachute Sizing and Descent Rates The following shows a detailed dimensions and equations sheet used to determine the as-flown parachute sizes and the projected descent rates for those dimensions. It is followed by a simplified table of the parachute sizing and descent rates. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 48

49 Drogue Parachute Configuration Figure 37: Recovery System Parameter Calculations Table 8: Parachute Sizing and Projected Descent Rates Hemispherical Main Parachute Configuration Drogue Parachute Diameter 26 in Main Parachute Outer Diameter Drogue Suspension Line Length 39 in Main Parachute Vent Diameter Shock Cord Length 30 ft Main Parachute Canopy Height Shock Cord Material 7/8 Tubular Nylon Main Suspension Line Length Annular with Pilot Chute 9.7 ft 5.93 ft 2.95 ft ft Drogue Descent Rate 64 ft/s Main Descent Rate 20 ft/s The projected descent rate is adequate to slow the rockets descent for an easier deployment of the main parachute while still minimizing the potential for excessive drift of the vehicle. Slowing the rocket to 20 UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 49

50 ft/s after main parachute deployment slows the rocket significantly enough to prevent a hard impact upon landing Transmitters The launch vehicle incorporates one location transmitting GPS tracker placed inside the nose cone to locate the rocket at any time during flight. The model to be used is a data puller type device called an ABLEGRID TK102B GPS Tracker. The device uses GPS technology to determine the exact location of the vehicle and can always be queried when powered on. First, it receives a signal from a mobile phone conveyed as a radio wave at a predetermined frequency or via SMS messaging. The tracker s signal is broadcasted using a different frequency through the built-in frequency converter. Thus, by transmitting and receiving different frequencies at the same time, it can also be detected simultaneously. This device can be located within an area as precisely as 10 feet in diameter and then transmits the data as geocoordinates that are sent via SMS messaging to any authorized phone. The device also has the capacity to continuously track the device and send new coordinates every 20 seconds of flight to retrieve real-time location information. See the following table concerning the specifications for the device. Frequency Wattage Range Table 9: ABLEGRID TK102B GPS Tracker Specifications MHz 2.96 Wh m System Sensitivity Figure 38: ABLEGRID Mini Spy GPS Tracker The avionics section alone is sensitive to radio or electromagnetic waves that could excite the circuit and interfere with the altimeter flight codes. To prevent the onboard GPS tracker from causing this interference, vehicle bulkheads are made of carbon fiber which is an excellent shielding material for EM waves. Two 1/8 bulkheads separate the avionics from the GPS tracker placed inside the ASA plastic nose cone. This configuration prevents EM interference while still allowing the tracker to send and receive frequencies within the vehicle without the carbon fiber components blocking its signal. Also, the UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 50

51 tracker only transmits location every 20 seconds, which further reduces the chance of interference occurring. 4.7 Mission Performance Predictions Most of the predictions for the leading design were computed through OpenRocket s flight simulation software. Hand calculations were performed when possible. The following data represents samples of the predicted flight patterns for the current full-scale vehicle design that is planned to be utilized on competition day, highlighting critical variables and verifying that the vehicle is robust and safe enough to launch Mission Performance Criteria The Rocket Rebel s vehicle mission is to design, manufacture, launch, and successfully recover a reusable high powered rocket. This rocket must successfully carry a fragile payload without damaging it throughout flight and landing. Also, the rocket must reach an apogee altitude of as close to one mile as possible. For this mission to be a success the following criteria must be achieved: 1. The launch vehicle must successfully lift off the pad at a minimum rail velocity of 52 ft/s. 2. The launch vehicle must attain an altitude of 5280 ft. AGL ±150ft. 3. The drogue/main chute must deploy within 2.0sec after apogee is reached. 4. The landing speed of the rocket after drogue deployment shall be approximately 70 fps to minimize drift. 5. The landing energy of the heaviest section shall be less than 75 ft-lb. 6. All parts of the rocket are recovered and are launch ready within three hours. 7. The FMPA and the fragile material enclosed must survive the entirety of the flight and remain intact upon retrieval Flight Simulation Data Table 10: Simulation Flight Data Event Time (s) Altitude (ft) Vertical velocity (ft/s) Vertical acceleration (ft/s 2 ) Stability Margin Launch Launch Rail Exit Burnout Apogee & Drogue Deployment Main Deployment Ground Hit UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 51

52 Figure 39: Altitude, Velocity, and Acceleration vs. Time Simulated Flight Data Plot Table 11: Noteworthy Data Values from Flight Max Altitude (ft) 5279 Velocity off rod (ft/s) 63.7 Max Velocity (ft/s) 595 Time to Apogee (s) 18.7 Max Acceleration (ft/s 2 ) 206 Flight Time (s) Average Thrust (lbf) 190 Velocity at Main Deployment (ft/s) 79.7 Thrust to Wet Weight Ratio 5.12 Velocity at Ground Impact (ft/s) 17.2 The mission performance predictions show that the current model will reach an altitude just below the target, at 5,279 ft. This would be in stark contrast to the first full-scale test, wherein the rocket traveled nearly 1,000 feet above the target. Since the first full-scale test flight, steps were taken to increase the weight of the vehicle, yet keep the same configuration and dimensions. This was due to the unforeseen aerodynamic advantage that airfoiled fins provided, causing the vehicle to travel 800 feet over projected apogee. To bring max altitude back towards the target, a combined weight of 6.5 lbs was added to the rocket through two ringed-shaped ballasts located on the payload and booster bulkheads. One minor anomaly to acknowledge is the descent speed which is predicted to be around 79.7 ft/s until main parachute deployment. It is very probable that the descent speed will be approximately 64 ft/s, based off other calculations presented in Recovery System, section With this data, along with the successful flight of the full-scale vehicle, the stress data of the carbon fiber building material and epoxy provided in the Appendix, and the expected loads given by the motor thrust curve, drag coefficients (section 4.7.3), and projected weight of the rocket, it is valid to assume that the vehicle design is robust enough to withstand the expected load of flight. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 52

53 4.7.3 Drag Assessment The following data represents the projected drag coefficients and subsequent drag forces that are expected during flight of the final launch vehicle as of FRR submission. There is little difference in design from the vehicle flown in the first test flight outside of a major weight addition that serves to lower apogee closer to the target altitude. However, it is justifiable to reevaluate the drag assessment of the current vehicle design to reveal any unforeseen changes in drag force due to this design change. Table 12: Current Projected Drag Coefficients Figure 40: Current Projected Drag Coefficients vs. Mach Number UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 53

54 Figure 41: Drag Force vs. Time The drag assessment provided shows that very minimal drag coefficient changes exist between the flight data in February (section 4.8) and the projected flight data of the currently revised design. The figure above shows the maximum drag force to be expected is 40 lbf, which is well within the launch load limits of all structural components within the vehicle Scale Modeling Results The transition from the subscale to the full-scale vehicle involved manipulating certain variables and scaling dimensions to accurately portray the same aerodynamic profile between the two vehicles. However, this meant that some variables would remain constant. The strength of the rocket was not required to be scaled if the structural integrity of the material could be proven to withstand the stresses incurred on both vehicles. Therefore, the strength properties of the carbon fiber airframe were not altered. Many of the shapes and geometric ratios were left the same as well. For example, the stability margin was maintained between the two vehicles stayed between 2 and 3.5 so that all vehicles exhibited low tendencies to move with or against crosswinds. The ratio of rocket length to diameter was also kept constant to help maintain the stability margin. This is not to be confused with the component values of this ratio, which were scaled to produce a larger rocket. Flight properties like max velocity, rail exit velocity, and max acceleration were kept constant as well when possible to recreate similar flight loads for the full-scale vehicle. On the other hand, there were many other factors that were scaled to either increase the vehicle s point of apogee to the target altitude and maintain a stable flight profile. To create this scale, the subscale vehicle was reverse engineered to closely recreate the desired full-scale model dimensions. This meant that each subscale component s weight and volume was measured and then recreated in OpenRocket modeling software. A scaling feature was then used to maintain the computed densities of each component and scale up the entire model by 50% (the subscale vehicle is a 67% scale of the full-scale). This gave the projected full-scale vehicle model that was to be built and was the team s point of justification for changes to the scaled design. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 54

55 The following figure compares the various aspects of the as built subscale rocket to the projected fullscale model at CDR, the as-flown full-scale vehicle, and the projected full-scale after design revisions have been made. Figure 42: Comparison of Scaling Factors As the figure shows, the vehicle generally maintained a constant scale where expected to, specifically in stability, rail exit velocity, max acceleration, max velocity, and length-to-diameter ratio. Motor impulse was very over-scaled, however, the chosen motor presented the most impulse available in an L motor while delivering the least amount of average thrust. This decision was made to reduce the loads exerted on the payload section, thus the scaling of the motor impulse was ignored for this project Stability To find the stability margin of the rocket, the center of gravity and center of pressure must be known. The center of gravity is given by using the following equation for each of the components of the rocket: CG = d 1w 1 + d 2 w 2 + W Where d is the distance of a component from the forward end, w is the weight of that component, and W is the total weight of all components analyzed. The Center of pressure for this vehicle is controlled by the center of pressure for the fins and nose cone. The center of pressure coefficient for a conical nose cone (C Nn) is equal to 2 and X N is the center of pressure for the nose cone, which for an ogive nose cone with length, L N is equal to X N = 0.466L N The center of pressure coefficient for the fins (C Nf) was calculated by: UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 55

56 C Nf = [1 + R S + R ] 4N ( S d ) 2 2 C R + C ) T ( 2L f [ For this equation, R is the radius of the body at the aft end, S is the fin semi span, N is the number of fins, d is the diameter at the base of the nose, L f is the length of the fin mid chord line, C R is the fin root chord, and C T fin tip chord. The center of pressure for the fins (X f) is calculated using: X f = X B + X R 3 (C R + 2C T ) (C R + C T ) [(C R + C T ) C RC T (C R + C T ) ] Where X B is the distance from the nose tip to the fin root chord leading edge and X R is the distance between the fin root leading edge and fin tip leading edge, parallel to the body. From these calculated variables, Barrowman s Equation is used to determine the center of pressure for a rocket without any conical transitions. X = C NnX N + C Nf X f C Nn + C Nf Now that the center of gravity and pressure are known, the following data and equations are used to determine stability caliper: Center of Gravity in Center of Pressure in Rocket Diameter Stability = C P C G D = 6.2 in ] = 3.15 Per data from OpenRocket, the stability margin increases from 2.56 at rail exit to 3.31 at motor burnout. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 56

57 4.7.6 Kinetic Energy Figure 43: Stability Margin, CP, and CP vs. Time The kinetic energy was determined through the given equation for each tethered section of the vehicle: KE = 1 2 mv2 The velocity variable (v) in this equation was determined through private algorithms with the parachute manufacturer to be 64 ft/s under drogue and 17.2 ft/s under main. The three tethered sections are listed in the following table in order from the tip of the rocket to the bottom. The values for weight in the table represent the dry weights (no propellant) of the tethered sections. Based off the calculations performed for each section, the heaviest section will hit with the hardest kinetic energy and be within 75 ft-lbf. Data is also included for kinetic energy of each component during drogue descent. Note that weights were converted to slugs (divided by 32.2 ft/s 2 ) before being applied to the above equation. Section Booster Section (postburnout) Table 13: Kinetic Energy of Independent Sections Weight (lbf) Kinetic Energy during drogue descent (ft-lbf) Recovery Section Payload Section Drift Calculations Kinetic Energy on Ground Impact (ft-lbf) Excessive wind accompanied with poor symmetrical design considerations will cause undue drift from the launch pad, making it more difficult to track the vehicle s descent and potentially causing personal UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 57

58 injury or property damage if the vehicle happens to drift outside the launch site radius. Drift calculations were conducted through OpenRocket to predict the likelihood of this occurrence with the project vehicle. The following table shows the anticipated drift distance for 0, 5, 10, 15, and 20 mph winds. Since 25 mph winds warrant aborting launch, drift simulations above this threshold were not analyzed. Drift must be limited to half the expected altitude which would mean that drift cannot exceed 2,640 lateral feet from the launch pad. That data shows the maximum lateral drift is within this range. Assumptions implicated the rocket will be launched straight up (perpendicular to the ground) and a low turbulence intensity of 10%. The following figure shows a side view flight track, tracking the displacement of the rocket with a 20- mph wind. Table 14: Drift Calculations with Change in Wind Speed Wind Speed (mph) Lateral Drift Distance (ft) Projected Altitude (ft) Mass Statement Component Figure 44: Flight Side Profile for Maximum Drift Table 15: Mass Statement Table Weight (oz) Component Weight (oz) Nose Cone 19.4 Drogue Bay Body Tube 33.6 UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 58

59 Payload Body Tube 25.2 Drogue Shock Cord 19.6 Payload Coupler Tube 10 Drogue Parachute Package 1.4 Payload Bulkhead 8.6 Recovery Section 190 FMPA Apparatus 16 Booster Ballast 8 Payload Ballast 96 Booster Bay Bulkhead 8.6 Payload Section 175 Booster Coupler Tube 9.8 Main Parachute Package 23.5 Motor Mount Tube 8 Main Shock Cord 19.6 Booster Body Tube 24.9 Main Bay Body Tube 47.7 Fin Sleeve w/ Centering Rings (Single Printed Part) Avionics Coupler Tube 10.1 Fin Set (Three Fins Total) 15.9 Forward Avionics Bulkhead 8.6 Loaded Motor Casing 132 Aft Avionics Bulkhead 8.6 Retainer Assembly 4.9 Avionics Ring 2.6 Booster Section 226 (Dry- 153) Avionics Sled and Inner Hardware 4.8 Full Scale Flight Data 14.7 Total 591 oz Dry 518 oz The full-scale test flight described in this section was conducted to validate the aerodynamic profile and confirm the projected flight patterns of the newly constructed full-scale launch vehicle. The test flight provided the team with invaluable data about the tendencies of the rocket during flight, an opportunity to rehearse drafted launch procedures, and address any errors in launch operations concerning the vehicle or launch pad assembly. The following model and specifications table shows the most accurate representation of the vehicle that was flown for the test launch. Figure 45: As-Flown Full-Scale Launch Vehicle Table 16: As-Flown Full-Scale Launch Vehicle General Specifications Structure and Propulsion Total Length 105 in Outer Diameter 6.2 in Dry Weight lb Wet Weight lb UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 59

60 Motor Choice Loki Research L840CT Motor Impulse 3888 Ns (874 lbf-s) Average Thrust 844 N (190 lbf) Burn Time 4.6 sec Recovery System Drogue Chute Design: Hemispherical Main Chute Design Annular w/ Pilot Chute Drogue Chute Diameter 26 in Main Chute ID-OD 5.93 ft 9.7 ft The launch was conducted in Memphis, TN off Gardener Road in the Shelby Farms area on February 18, The temperature forecast for the day varied only slightly throughout the time the team was at the launch site, ranging from 61 F to 68 F. Wind speeds were projected to be around 4 mph to the North and slightly West. The following input table reflects all the launch conditions of the test launch. Table 17: Launch Conditions Data The following table shows the critical points of the flight data retrieved from the primary altimeter of the avionics sled in chronological order. This is followed by a graph generated by the altimeter data transfer software that creates a graph of the data gathered by the StratoLoggerCF altimeter. Additional data recorded by the altimeter, including temperature and battery level readings, are provided as well. Table 18: Chronological Full-Scale Flight Critical Event Data Points (Retrieved from Altimeter) Event Time (s) Data Value Recorded Max Acceleration (ft/s 2 ) Max Velocity fps Apogee (Drogue Deployment) ft Main Parachute Deployment Altitude ft Ground Hit Altitude ft UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 60

61 Figure 46: Actual Altitude and Velocity vs. Time Figure 47: Altitude and Temperature vs. Time UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 61

62 Figure 48: Altitude and Battery Voltage vs. Time Predicted vs. Actual Flight Data Comparison The following flight data was procured from OpenRocket simulations made a day prior to the launch when the model best represented the as-built vehicle. This data is compared to the actual flight data listed previously. Figure 49: Projected Vertical Motion vs. Time Flight Data UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 62

63 Table 19: Comparison of Simulated and Actual Flight Data Simulated Flight Data Actual Flight Data Percent Difference Apogee (ft) % Max Velocity (fps) % Max Acceleration (ft/s 2 ) % Flight Time (s) % Drogue Descent Speed (fps) % Main Descent Speed (fps) % Error Analysis The compared data shows that most critical data points were accurate to the projected values generated by OpenRocket by within 8%. The vertical motion data (apogee, velocity, acceleration) was extremely accurate, implying that the vehicle models dimensional and weight data is a near perfect representation to the as-built model. Two values that interrelate to each other, flight time and drogue decent speed, exceeded this percent difference and justify error analysis. Persistent with the error found during CDR, the OpenRocket software does not have the capacity to simulate descent rates with annular style parachutes. This gives very inaccurate readings for the size that parachutes must be to properly control the descent of the vehicle. Therefore, the parachute manufacturer computes the data for the team by using proprietary algorithms to calculate the descent speeds. These calculations are consistently closer to actual descent speeds from the two previous flights. Therefore, they will continue to be the primary means of descent rate projections. However, OpenRocket descent speeds will be still be heeded as rough estimates for future design changes. The only independent anomaly in the actual flight data worth addressing is the sharp, momentary decline in altitude that occurs at both parachute deployment events. This spike is caused by an increase in pressure, which would mean that the gases of the ejection charges are rushing into the avionics bay at some point. This means that the avionics bay was not fully sealed when flown. This problem was encountered during subscale flight and actions were taken to seal the bay for the full-scale test flight. For future flights, the bay will be heavily inspected for gaps and will be exceptionally filled as necessary. It is also important to evaluate the bay temperature and battery voltage data gathered by the altimeter. It shows that the temperature steadily decreased as the flight computer increased in altitude and then spiked up by.25 F at apogee. The avionics bay temperature steadily decreased again and then spiked again by.5 F at main parachute deployment. The temperature then slowly decreased until ground hit. This data is synonymous to what was expected, considering that the ejection charges would transfer some amount of heat through the bay and increase the bay temperature slightly. This occurred for both ejection charges and the outside temperature cooled the bay back down intermittently. The battery voltage data retrieved by the altimeter revealed only a slight anomaly. As shown in the graph, the voltage optimally dropped by about.5 V when firing the drogue parachute ejection charge. However, the main ejection charge resulted in the battery to drop to nearly half its original voltage, momentarily losing 4 V. This could have potentially caused a momentary brownout of the altimeter which would have resulted in a small gap where the altimeter did not record altitude. This problem could have originated from an unsealed avionics bay and the rush of atmospheric pressure during main parachute deployment could have slightly affected the electrical current. For future launches, all holes and connection points will be adequately sealed before flight. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 63

64 4.8.3 Drag Coefficient Comparison The predicted full-scale drag coefficient is calculated through the OpenRocket components analysis page and shows the drag coefficients of each component. The following tables and graphs show the maximum aerodynamic drag coefficients projected as of CDR in mid-january and the drag coefficients calculated from the as-flown model replicated in OpenRocket. These values are generated for when the launch vehicle has reached its maximum velocity, which was at 651 fps (.59 Mach) in the CDR design and is currently 595 fps (.54 Mach) in the FRR design. Table 20: CDR Drag Coefficient Projections Table 21: 1 st Test Flight Drag Coefficient Calculations As can be seen in the tables, the total C D of the launch vehicle decreased by.05 since revisions past CDR were implemented. The main differences occur in the C D of the booster bay and fins. Since CDR, the entire booster has gone through a moderate design revision that was an attempt to lighten the aft end of the rocket while maintaining certain aerodynamic properties. By shortening the booster section, this component s fiction C D decreases. Implementing an airfoiled fin design with lighter material decreases the pressure C D of the fins while only slightly increasing friction C D. Minor design revisions occurred that have increased the drag coefficients like increasing the main bay body tube length and using a thicker fin set. The following figures graph the change of each drag coefficient and the total drag coefficient of the vehicle as it approaches max velocity. The red line on the graphs show the total drag coefficient expected along the range of Mach speed that will be experienced on the rocket. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 64

65 Figure 50: CDR Drag Coefficient versus Mach Number Figure 51: FRR Drag Coefficient versus Mach Number The following graph shows the drag force overtime as the rocket reaches max velocity at about 4.5 seconds. As can be seen, the max expected drag force is about 40 lbf. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 65

66 5 Payload Criteria Figure 52: Drag Force vs Time 5.1 Payload Objective and Mission Success Criteria The Rocket Rebels are researching a fragile material protection apparatus (FMPA). This system s primary function is to hold a small fragile cargo and protect it from the loads transferred to it through launch, parachute deployment, and recovery. Success criteria for this objective to be fulfilled are as follows: 1. The FMPA can be installed within the payload bay. 2. The fragile cargo can be easily loaded into the FMPA. 3. The FMPA cargo survives the entirety of a successful vehicle flight and is recoverable. 4. The FMPA is recoverable and reusable. 5.2 Structural Elements The Fragile Material Protection Apparatus (FMPA) is created as one integrated system of 10 shock absorbing pistons. The inner capsule consists of a screw-on cap that allows for the fragile material to be easily inserted into the rocket at any given time. This capsule is fixed to a carbon fiber plate ring that is epoxied around the lower portion of the capsule s circumference. This ring is attached to the rods of the piston system which is installed to a larger fixed ring that lines the inside diameter of the rocket. The fixed system includes a set of 10 pistons arranged in an equal radial manner. This material is exceptionally light-weight while providing stiff structural integrity to the system and is machinable with the tools available. The inner ring is epoxied to the rods of the shock absorbers while the outer ring is drilled to fit the outer diameter of the pistons. The capsule can fluctuate up and down via the inner ring to divert force, and the rings cradle force by direct attachment to the pistons. The inner capsule is made of high-strength polymer and accommodates the maximum volume requirements for the competition. The fragile material will be filled with packing peanuts upon installation to ensure it is snugly secured, UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 66

67 regardless of its volume. The capsule is fitted with a sealed screw-on top allowing easy access to the inside of the payload. Two rings are made for the apparatus: one that is flush with the inner diameter of the rocket (epoxied to the inner diameter of the payload bay) and one that is slightly smaller which moves freely up and down the payload bay with the inner capsule. This inner capsule will be filled around the fragile cargo with a loose bed of foam packing peanuts. The FMPA incorporates Losi shock pistons, commonly used for suspension systems on RC cars. Ten will be positioned 36 radially below the capsule. The pistons measure slightly over 2 each, which is ideal for limited payload bay volume and utilize reactionary forces from a spring and damper fluid combination. 5.3 Payload Construction The construction of the payload system is fairly straightforward and simple. The inner capsule is removed of all labels and sanded heavily with 220 grit sandpaper. This is to give two-part epoxy a rough surface that is more susceptible to holding heavier loads. The waterjet cutter is used to cut the inner and outer rings from a flat piece of cured 1/8 thick carbon fiber plate material. Two ring-shaped drawings were created in AutoCAD and transferred to the waterjet cutter computer. The dimensional data is provided in the table below: Table 22: FMPA Ring Dimensional Data Outer Diameter (in) Inner Diameter (in) Inner Ring Outer Ring After pattern cutting, the extracted fins were sanded smooth and holes for shock absorbers were drilled in the rings. The tops of the shock absorbers are.22 and fit in the inner ring. Thus, the holes in the inner ring were drilled to the next biggest drill size available (15/64 ). The bottoms of the shock absorbers were measured as.32 and fit in the outer ring. Thus, the drill bit size used for the bottom hole was 11/32. Since the shock absorbers must align vertically between the rings, the rings were secured together with duct tape. A stencil with a laser-etched pattern was created to help position the holes as they were being drilled. The following figure shows the setup of the rings and stencil before drilling. Figure 53: FMPA Machining Setup and Stencil Pattern UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 67

68 After the holes were drilled, a half batch of two-part Loctite 9394 Slow Cure Epoxy was prepared. A small popsicle stick was used to spread a bead around the inner diameter of the inner ring and the outer diameter of the lower portion of the capsule. A dab of epoxy was spread into every hole in both rings and the larger side of the shock absorbers were installed into the outer ring holes. The inner ring was installed onto the capsule and a fillet was made on both sides of the interface between the ring and capsule. Finally, the inner ring was placed over the outer ring and the shock absorbers were carefully aligned and inserted through the smaller holes of the inner ring. A loose bed of packing peanuts were placed in the capsule and the epoxy was allowed to cure for 24 hours. The following figure shows the final assembly. Figure 54: Completed FMPA UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 68

69 5.4 Dimensional Drawings and Schematics Figure 55: FMPA 2-D Drawing w/ Dimensions Figure 56: FMPA Exploded View UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 69

70 5.5 Payload Component Interaction and Integration The principle of this system is to dissipate impulse from launch loads, ejection charge forces, and impact from the ground. Since the greatest of these is the impulse delivered from launch loads, the design of the FMPA system will be prioritized to dissipate this load. Impulse from this event is transferred across the rocket body to the payload bay where the FMPA capsule rests. The capsule tends to travel in the direction of the force towards either the nose cone or the aft payload bulkhead respective to the force applied. Hard impact with either of these could potentially damage the fragile cargo. To prevent this, force is transferred from the body tubes to the outer carbon fiber rings. These outer rings are integrated to inner rings that are fixed to the FMPA capsule through a hybrid shock absorber. This system uses a damper and resistive spring combination to detract from the energy transferred to the inner capsule. This system allows for energy to be dissipated by two main factors: how quickly and how far the shock absorbers contract or extend. The analytical analysis performed to calculate the effectiveness of the chosen design is described in section System Design Requirements The FMPA system must complete the following system requirements along with the requirements provided by NASA for the payload: System must be capable of recovery and reusability FMPA must easily house a fragile cargo within the inner capsule without damaging the cargo. FMPA piston system can maneuver the inner capsule via the shock absorber system. FMPA piston system can be adjusted in a configuration to maximize launch load dissipation at beginning of flight. FMPA can protect fragile cargo from all other stresses incurred throughout duration of flight. System can be easily accessed after flight to retrieve fragile material. The team derived requirements verification plan is listed in Section 8.3 where the method of fulfilling these requirements is described in detail. 5.7 Relevance of Expected Data and Accuracy/Error Analysis The FMPA s accuracy and error analysis will be governed largely by the number of pistons used and the length of the piston rods that will be used for this project. If the lengths are too short or not enough pistons are installed, the launch loads could compress/extend the shock absorbers to their extreme, negating any shock dispersion from the inner capsule that houses the fragile cargo. This error would be the most catastrophic to the cargo, yet would not cause any noticeable damage to the rest of the vehicle nor largely affect its launch telemetry. Dimensional analysis was performed to find the most effective shock absorbers that will fit within the payload bay. 5.8 Precision of Instrumentation and Repeatability of Measurement There are no electronics used in the Rebel Rockets payload. The precision and repeatability data will be retrieved through trial and error. The initial calculations used will be based upon drop test trials to test the durability of the payload structure. Tests were performed with the full-scale model to record the amount of stress endured during a flight and how the apparatus controls the applied stresses. The removable attachment of the capsule allows for constant changing of either the fragile testing material, or the safety materials used within the capsule. Any further adjustments will be based on newer data obtained from future flights. The modified payload will then be subject to additional launches and/or drop tests, repeating the process. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 70

71 Severity 6 Safety Rocket Rebel s safety officer is William Thomas. As the Safety Officer, William is responsible for briefing the team on proper procedure and risks associated with the design, construction, testing and storage of materials and equipment. When working on the shop floor, William will coordinate with the shop technicians to ensure compliance with all shop rules. In addition to William, the team mentor Cody Hardin will oversee each stage of the design and build process to verify NAR requirements are met and maintained. The faculty advisor, Jack McClurg, is available to verify designs are structurally sound. Table 23: Likelihood and Severity Table Likelihood Rare Unlikely Moderate Probable Extremely Likely Negligible 1A 2A 3A 4A 5A Minor 1B 2B 3B 4B 5B Moderate 1C 2C 3C 4C 5C High 1D 2D 3D 4D 5D Extreme 1E 2E 3E 4E 5E This table will be used for risk identification and mitigation purposes. Likelihood is ranked on a scale from 1 to 5 with 1 being rare and 5 being extremely likely. Severity is ranked on a scale from A to E with A being negligible and E being extreme. Green boxes indicate an overall low risk to the project. This is due to either the risk severity being negligible or the chance of the risk occurring is extremely rare. Risk assigned this rating will be monitored. Yellow boxes indicate a moderate amount of risk to the project. These risks can delay the project or create a need for additional funding. Moderate risks should be monitored by all team members. Red boxes indicate a high amount of risk to the project. These risks can cause major delays, cause serious harm to the rocket or team members, or create a need for a massive increase in funding. High risk procedures and activities should be closely monitored by all team members and the risk will receive a separate briefing addressing the risk and proper mitigation. 6.1 Personal Hazard Analysis The primary dangers to team members come from handling chemicals, machining parts, handling explosives, handling electrical components, and preforming the launch. In addition to the team, bystander stander safety is primarily concerned with dangers at the launch site. The team safety officer will make every effort to brief all team members and bystanders of all relevant safety procedures and risks associated with each stage of development, build, assembly, and launch. Also, the outreach team will follow all relevant safety procedures associated with the task at hand. Each team member has read and agreed to abide by safety standards (included in appendix) enforced on the shop floor. All machining processes will involve 2 team members as well as be overseen by a shop technician. This procedure helps to mitigate any risk associated with the machining of the rocket parts. The team safety officer will give safety briefings before any machining, assembly or launch procedures. During the safety briefing, the team safety officer will provide the relevant MSDS for the planned procedures. The team will also be informed of the proper PPE required for each procedure. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 71

72 Likelihood & Severity Likelihood & Severity after Mitigation Control To reduce risk, all flammable chemicals will be stored in a designated metal cabinet and all high-energy materials will be properly stored in a designated magazine. These materials will be secured until the head engineer, the team safety officer, or a factory technician removes them for the intended purpose and will be returned by the person who originally removed the material. This procedure ensures that no material is accidentally left out and helps verify that anyone handling or working with the material has been properly informed of hazards and procedures associated with working with the material. NAR guidelines have been thoroughly reviewed by the team. The team mentor is level 3 NAR certified. In addition to this certification, some team members may pursue NAR certifications during the build and subscale testing process. The following is a table addressing common personal hazards associated with building and flying high power rockets. The table addresses the hazard, examines the causes, lists the potential effects, and then makes a recommendation as to how to mitigate the risk. Finally, the table includes a likelihood and severity rating for post risk mitigation. All risks should be accounted for with special attention paid to risks with a moderate or high rating. Table 24: Personal Hazards Analysis Hazard Cause Effect Mitigation Control Mitigation Verification Acetone Prolonged exposure to hazardous fumes Skin and eye irritation. Possible sickness or lightheaded-ness 4B Goggles, masks, and appropriate protective clothing will be worn when handling chemical. Adequate ventilation will serve to remove fumes. 2C Ensure PPE and ventilation is available and PPE is equipped before using chemical. HexPly 6376 Prepreg Carbon Fiber Contact with rough edges and fine, fibrous sheets Fibers could be lodged in skin, inhaled through nose or mouth. Could cause long term skin rashes or respiratory problems 4C Thick nylon/leather gloves and long sleeve clothing will be worn when handling carbon fiber material. Material will be held away from face. 2C Ensure PPE is available and equipped before handling cured Carbon Fiber West System 105 Epoxy Resin Exposure to skin, eyes, or respiratory system Eye, skin and respiratory irritation 3C Material will be used in ventilated area with chemical resistant gloves and eye protection 2B Ensure PPE and ventilation is available and PPE is equipped before using epoxy UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 72

73 West System 206 Slow Hardener Exposure to skin, eyes, or respiratory system Eye, skin and respiratory irritation 3C Material will be used in ventilated area with chemical resistant gloves and eye protection 2B Ensure PPE and proper ventilation is available ensure that PPE is equipped before using chemical Cesaroni Pro54 Profire Igniter Combustion, burn risk Burn to skin, face, or other appendages. Medical care may be required 2C Use outside, with face and body protection. Have fire-fighting method on hand. Keep away from heat. 1B Ensure PPE and fire extinguisher is available ensure PPE is in use before handling igniter ensure team members know proper use of fire extinguisher Cesaroni Pro54 Rocket Motor Reload Kit Combustion, burn and explosion risk Burn to skin, face, or other appendages. Hospitalization may be required 2E Use outside, with face and body protection. Have fire-fighting method on hand. Keep away from heat. 1D Ensure PPE and fire extinguisher is available ensure PPE is in use before handling igniter ensure team members know proper use of fire extinguisher Chem-Trend SealProof Exposure to skin or eyes; ingestion Eye, skin and respiratory irritation 3C Material will be used in ventilated area with chemical resistant gloves and eye protection 2B Ensure PPE is available and proper PPE and ventilation is in use when handling the chemical Chem-Trend Zyvax 1034W Mold Release Exposure to skin or eyes; ingestion Eye, skin and respiratory irritation 3C Material will be used in ventilated area with chemical resistant gloves and eye protection 2B Ensure PPE is available and proper PPE and ventilation is in use when handling the chemical UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 73

74 3-M Fiberglass Cloth Contact with rough edges and fine, fibrous sheets Fibers could be lodged in skin, inhaled through nose or mouth. Could cause long term skin rashes or respiratory problems 4C Thick nylon/leather gloves and long sleeve clothing will be worn when handling carbon fiber material. Material will be held away from face. 2C Ensure proper PPE is available and in use when handling the cloth ABA Plastic Filament Contact with rough edges Mild to severe lacerations, abrasions, or puncture wounds. Infection possible. Member may need medical care. 3C Thick gloves will be worn when handling material. Rough edges will be deburred with sander. 1B Ensure proper PPE is available and in use when handling the filament Skinirritating materials Exposure of acetone, epoxy, or cleaner to skin or eyes Skin or eye irritation, burning, or corrosion. 4B Long sleeve pants and shirts will be worn with thick latex gloves and goggles when handling or applying potentially hazardous liquids or gels 2C Ensure proper PPE is available and in use when handling hazardous materials Bridgeport Milling Machine Inattention to moving drill or running machine Mild to severe lacerations, abrasions, or puncture wounds. Infection possible. Member may need medical care. 3C Proper PPE and tooling instruction will be given prior to every construction operation by technical staff or experienced team members. 2B Ensure proper PPE is available and in use when working with the mill. Ensure all team members using the mill have received instruction on proper usage Roll-In Horizontal Swivel Band Saw Inattention to moving blade or running machine Mild to severe lacerations, abrasions, or puncture wounds. Infection possible. Member may need medical care. 3C Proper PPE and tooling instruction will be given prior to every construction operation by technical staff or experienced team members. 2B Ensure proper PPE is available and in use when working with the saw. Ensure all team members using the saw have received instruction on proper usage UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 74

75 Stationary Vertical Band Saw Inattention to moving blade or running machine Mild to severe lacerations, abrasions, or puncture wounds. Infection possible. Member may need medical care. 3C Proper PPE and tooling instruction will be given prior to every construction operation by technical staff or experienced team members. 2B Ensure proper PPE is available and in use when working with the saw. Ensure all team members using the saw have received instruction on proper usage Drill Press Misuse, lack of attention to detail or failure to follow safety procedures Member is seriously injured and could need hospitalization 2D The team will only operate machinery after being cleared by the shop technicians. The shop technicians will review all planned operations and be present for supervision during all operations. 1D Ensure proper PPE is available and in use when working with the press. Ensure all team members using the press have received instruction on proper usage Eastman cutting table Inattention to moving parts or running machine Mild to severe lacerations, abrasions, or puncture wounds. Infection possible. Member may need medical care. 3C Proper PPE and tooling instruction will be given prior to every construction operation by technical staff or experienced team members. 2B Ensure proper PPE is available and in use when working with the cutting table. Ensure all team members using the cutting table have received instruction on proper usage Walk-in curing oven Improper use including: Opening doors during cure cycle or standing inside oven during cure cycle Mild to severe burns. Hospitalization may be necessary 2D Oven will only be operated with technical staff present. No personnel will be allowed in the oven during or within an hour of a cure cycle. 1C Ensure proper PPE is available and in use when working with the oven. Ensure all team members using the oven have received instruction on proper usage. Ensure all team members have exited oven UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 75

76 before activating heat Black Powder Inhalation, Eye/ Skin Contact, Injuries sustained during accidental Ignition Lung Irritation, Eye/ Skin Irritation, Minor to severe burns 4D Wear masks when handling black powder to prevent accidental inhalation. Wear eye protection to shield eyes from black powder dust, wash eyes thoroughly if exposed. Discharge static electricity before working with black powder, keep flames away from area when black powder is being handled 2B Ensure proper PPE is available and in use. Ensure a fire extinguisher is available and team members know proper use of fire extinguisher Tripping over loose cords Long power cords/wires being run across the shop floor Personal injury 3C Power strip will be installed near all machines/workspaces that may require a power outlet. 1C Verify power strips are available and installed near operations area. All other cords will be situated off walkways and operating areas Loctite EA9394 Inhalation, Skin/ Eye contact Lung irritation, Eye/ Skin Irritation 4B Handle/ work with epoxy in well ventilated areas. Wear latex gloves when working with epoxy 2A Ensure proper PPE and ventilation is available ensure PPE is in use when working with Loctite Scotch Brite Pads Inhalation of particles Lung Irritation 4B Wear mask when using, only use in well ventilated areas 2A Ensure proper PPE and ventilation is available ensure PPE is in use when working with Scotch Brite Powder Coating Inhalation of Powder, Accidental Ignition of powder Lung Irritation, Minor to Sever Burns 4D Wear masks when working with material. Ensure part is well grounded to minimize discharge sparks, ensure that air flow does not 2B Ensure proper PPE and ventilation is available ensure PPE is use during the UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 76

77 allow for build-up of explosive concentration powder coating process Fortus 450- MC 3D Production System Burns from improper use Minor to sever burns from hot oven 3B Wear safety gloves and long sleeves when removing parts, working inside oven 2B Ensure proper PPE is available and in use when working with the Fortus machine Water Jet Cutter Injury from improper use, Eye Exposure, Hearing Damage Severe Lacerations, Crushed appendages from moving parts. Eye irritation due to exposure to abrasives mixed in water. Hearing irritation/loss due to loud noise involved with water jet cutting 4E Verify that all safety guard are in place before turning on machine, verify all personnel present are away from the cutting area. Wear safety goggles when working with machine. Wear hearing protection when working with the machine 2A Ensure proper PPE is available ensure PPE is in use when working with water jet ensure water jet operator is instructed on proper usage Electrical Sanders Lung Exposure, Eye Exposure, Moderate to Severe Skin Abrasion, Pinch Points Lung irritation due to inhalation of dust. Eye irritation due to exposure of dust. Skin abrasion due to Improper use. Appendages/ clothing caught in moving parts 4C Wear protective mask when working with sanders. Wear protective goggles when working with sanders. Wear gloves when working with sander, ensure no one touches a moving sandpaper. Avoid wearing baggy clothing, keep excess hair tied back, keep appendages away from moving parts 2B Ensure proper PPE and ventilation is available ensure PPE is in use when working with sanders MIG welder Careless use or negligent behavior with welding rod Severe skin burning, retinal damage, respiratory damage 3E Welding helmet and clothing will be worn. Technical staff will overview operation before using welder 2D Ensure proper PPE and ventilation is available ensure PPE is in use when working with Loctite ensure weld operator is instructed on proper usage UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 77

78 Aluminum Mandrels Inattentive transportation, rough edges, fine powder residue from sanding Member could be hit by mandrel, cut by rough edges, or have respiratory problems from fine powder inhalation 3C Path will be cleared for mandrel prior to transport, rough edges will be deburred, mandrel will be rinsed and wiped with water immediately after sanding 2B Ensure proper PPE is available and in use when deburring ensure all personal are informed before mandrel is moved Hand tools Careless or negligent use of portable tools that could lead to injury Mild cuts, or other wounds. Infection possible. Member may need medical care. 4B Proper PPE and tooling instruction will be given prior to every construction operation by technical staff or experienced team members. 3B Ensure proper PPE is available and all hand tool users are properly instructed on hand tool usage Hot materials Touching recently machined or baked materials; mishandling a welding tool; exposure to exothermic epoxy Mild to severe burns could require members seek medical attention 2E Hot materials will be set aside to cool after any heat related operation or after machining materials that are very heat conductive. 2C Ensure proper PPE is available and in use ensure proper cooling area is established Burns from high energy materials Member is burned by an ejection charge or ignited motor at launch Mild to severe burns. Exhaust can also irritate eyes and lungs. 2E The team safety officer will brief the team on all relevant safety procedures when handling the highenergy material. Handling, use and transport of the rocket motor will be supervised by the team mentor. High energy materials will be appropriately stored when not in use. 1E Ensure proper PPE is available and in use when handling high energy materials ensure team mentor is aware of high energy material transfer Electrical Components Mishandling of ungrounded circuits, batteries, launch boxes or electronics prior to launch Discomfort, chance of unconsciousness or heart attack in extreme cases. Possibility of hospitalization 3C Members working on batteries and circuits shall have training on proper handling and wiring techniques and ensure that all components are grounded and use caution. 2B Ensure proper PPE is available and in use when handling electrical components ensure team members handling electrical components UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 78

79 have been properly trained Free-floating dust particles Machining metals and cured carbon fiber without proper dust collection system Difficulty breathing, coughing, other respiratory problems. Can lead to long term respiratory disease 3C Members will wear face masks and always have a dust collection system active when machining carbon fiber and metal materials. 2B Ensure PPE is available and used during operations. Verify dust collection system is active before machining. Sharp drill bits Mishandling when changing out bits with hand drill or drill press. Can cause lacerations and lead to bleeding and infection 4B Members will wear thick gloves when handling drill bits. Precautions will be taken to carefully secure drill bits into machinery or in drill bit cases. 1A Ensure gloves are available. Test launches Member is burned by rocket motor during test Member is harmed and could require medical attention 2E As required, there shall be supervision during testing. Members shall be behind a secure and shielded testing frame. 1C Ensure all personal are properly briefed before test launches Launches A fire is started from the ignition of the rocket, late separation, improper rocket mounting, etc. Member could be burned and require medical attention 2E A fire extinguisher and safety radius around rocket shall be maintained at launch. All launches will be in full compliance with NAR safety codes. Team mentor will be present at all launches. 1D Ensure all personal are properly briefed before launches 6.2 Failure Modes and Effects Analysis The following is a table addressing failure modes and effects associated with each subsystem. The table categorizes the data, identifies the point of failure, examines the causes, defines the possible effects, and provides risk mitigation procedures. In addition, each point of failure is given a likelihood and severity rating. The team will give special attention and oversight to any risks bearing a moderate or high rating. Table 25: Failure Modes and Effects Analysis UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 79

80 Likelihood & Severity Likelihood & Severity after Mitigation Control Risk Cause Effect Mitigation Control Mitigation Verification Improper assembly of motor Incorrect spacing between propellant grains; motor case improperly cleaned; end caps improperly secured Motor failure, unstable flight, target altitude not reached, damage or loss of rocket or motor casing 3E Ensure proper training and supervision by team mentor for motor assembly by student safety officer 2C Verified mitigation in Consultation of manufacturer s assembly and installation instructions. Premature propellant burnout Improper motor packing; faulty propellant grain Altitude estimate not reached; main parachute may not deploy 3E Proper motor assembly; obtain motor only from reputable source 1C Manufacturer s instructions will be followed for motor preparation and installation. Transportation/ handling damage Improper protection during transportation and handling Unusable motor; potential damage or loss of rocket 3E Motor will be stored in cushioned and sealed box overseen by team leader, safety officer, and team mentor 1D Safety officer will verify casing and grains are secured in a safe storage container during transport. Mandrel Tube Defects Improper layup procedures, creating deep cuts or scratches in the layup mandrel Poor layup finish. If severe, cured part could be unusable. 3C Ensure great care in mandrel prep operations, use fine sandpaper if needed instead of high speed tools 2C Ensure mandrel is properly prepared for Carbon Fiber layup Body tubes Unforeseen layup anomalies, poor extraction or machining practices Poor quality body tube sections produced. Could compromise vehicle during flight, potential loss of vehicle 3D Careful extraction procedures will be followed. Post-cure machining will be minimal and meticulous. Layup will be performed with special care. 2C Verified design in Ensure team members are aware of proper extraction techniques Motor retaining ring Poor motor assembly or poor retainer rings assembly. Motor falls out of the rocket once the propellant is spent. 2E Specify assembly procedures in final launch checklist and double check retainer ring assembly. 1D Verified in Ensure assembly procedure is adhered to UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 80

81 Centering Ring Inadequate epoxy strength, inadequate epoxy heat tolerance. Motor launches through rocket, unstable flight. 2D Verify epoxy used to join centering rings to rocket is strong enough to withstand the forces applied to it. 1D Verified ensure assembly procedures are followed Electric Igniter Failure Defect in igniter, faulty wiring of igniter. Motor fails to launch. 3A Verify the igniter is working properly prior to launch date. Verify wiring is correct. 1A Verified in 4.5.4, 4.5.5, and 4.6 series ensure proper assembly procedures are followed Propellant ignition Damage to motor, poor propellant packing, bad propellant. Motor fails to launch. 2D Verify the motor is properly packed for transport, team mentor oversees motor preparation. 1D Verified in 4.5.4, 4.5.5, and 4.6 series ensure proper assembly procedures are followed CG/CP Poor rocket design causes the CG and CP to be closer than the length of 2 radii. Unstable Flight. 2D Hand weigh rocket to find center of gravity with simulation center of pressure to verify stability margin is acceptable. 2B Verified in 4.7.5, and Nose Cone Failure Damage incurred from previous flights, improper handling, or transportation. Unstable Flight. 2D Proper nose cone selected. Create a spare if funds allow. 1D Verified in 4.2.5, and Fin Failure Poor adhesion between the fins and the motor housing. Unstable Flight. 2D Proper epoxy used to adhere the fins to the motor housing. 1D Verified in , and Tube Coupler Shearing Poor adhesion between tube coupler and rocket section. Rocket parts separate during flight, damage to or loss of rocket. 1E Verify the correct adhesive is used to join the tube coupler to the rocket sections. 1E Verified in Bulkhead failure Buckling due to bulkheads lacking the ability to withstand force produced by motor. Rocket incurs damage. 1E Proper selection of materials, Proper manufacturing techniques. 1C Verified in UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 81

82 Lacking Team Participation Lack of communication between subgroups, poor understanding of overall rocket design. Rocket cannot be built. Rocket is incorrectly built. 4E Encourage intergroup communication. Overview the entire design before the build is commenced. 1D Ensure adequate team members are aware of meetings and builds Incorrect Preflight assembly Rocket is not preassembled in correct configuration. Negligent of launch ops procedures Unpredicted trajectory and possible rocket loss 3C Label aft, forward, and 0 azimuth locations on all components before assembly 1C Ensure team members are present for dry run of assembly and procedures are followed 7.2 Premature or late rocket separation Faulty wiring, altimeter malfunction. Rocket separates before intended (i.e. while motor is still burning), leads to loss of rocket, potential danger to personnel and property. 2E Ensure the wiring to the separation charge is correct, ground test altimeter and wiring configurations. 1D Verified in 4.6 Launch Pad misassembled Personnel error, rail lugs not properly installed, wind. Rocket does not take the desired launch path 2D Verify the rocket is properly placed and aligned on the launch rail, verify rail lugs are installed correctly, do not fly in high winds. 1B Verified Inadequate parachute space Poor communication between subgroups on space allocation. Lack of space to properly pack parachutes within the rocket. 2C Promote clear communication about the needs and wants of each subgroup. 1B Verified in Compactness of ejection charge and recovery packages Negligent assembly and packing of parachutes and shock cord without Nomex blankets Severe damage to recovery system. Potential damage or loss of vehicle 3E Ensure Nomex blankets are installed to shock cord and are wrapped around parachute package during installation and vehicle assembly. 1D Verified in launch procedures UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 82

83 Parachute Deployment Drogue or Pilot chute drag force is not sufficient to deploy parachute. Rocket goes into free fall, potential danger to property and personnel. 2E Fit between couplers and body tubes will not be excessively tight, recovery package will be properly prepped for deployment 1E Verified in 4.6 Tangles in suspension lines and/or risers Improper packing. The parachute does not perform as intended, damage to or loss of rocket or payload. 2D Verify the line is free from tangles before packing the parachute, properly pack the parachute to minimize the chance of tangles. 1C Verified in Slow main parachute decent Parachute is too large for desired decent rate. Rocket is blown out of the landing area. 4B Verify calculations for parachute size and execute appropriate subscale testing. 2B Verified in Fast main parachute decent Parachute is too small for desired decent rate. Damage to or loss of payload or rocket. 4D Verify calculations for parachute size and execute appropriate subscale testing. 2C Verified in Excessive excitation of avionics bay from GPS tracker EM waves transferred by GPS tracker could intercept the avionics circuitry, inducing varied currents Recovery system does not deploy main or drogue parachute during vehicle; potential loss of vehicle 2E EM shielding bulkheads will deter any interference to the avionics circuit. GPS tracker will always be situated in the nose cone, away from the bay. 1D Ground tests will confirm GPS operation doesn t interfere with avionics in launch configuration Premature ejection charge ignition Ejection charges cause unintentional ignition in the rocket. Damage to recovery system, loss of avionics, loss of rocket. 1E Isolate ejection charges from flammable material via packing with flame retardant material. 1D Verified in Parachute melting Poor insulation between parachute and separation charges. Hole in parachute or shock cord; potential separation of rocket section or increased descent speed 2E Ensure the parachute is properly insulated from separation charges, verify with ground testing. 1E Verified in UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 83

84 Separation failure Excessive shear pin holding strength; inadequate amount of black powder charge; altimeter failure Recovery system malfunction; potential loss or damage of rocket 3E Verify calculations used to calibrate separation charges, verify ground testing shows adequate separation. 1E Verified in 4.5.3, and through full scale testing Tears in Parachute Parachute is improperly packed increasing chance of snagging on deployment, damage incurred form transportation. The parachute does not perform as intended. Decent velocity is greater than expected, leads to damage to or loss of rocket. 2E Verify parachute is packed correctly, inspect the parachute before packing for any tears or other defects. 1D Verified in Damaged Shock Cord Defective or fatigued cords from improper storage or damages incurred form transportation. Rocket parts separate, damage to or loss of payload or rocket. 2E Verify shock cord is stored properly, inspect shock cord before packing. 1E Verified in Avionics power loss Disconnection in wiring scheme Loss of flight data and failure to deploy parachutes 3D Fresh wire will be used for every launch and soldering will be performed with care to avoid disconnection. 2C Verified in proper wiring assembly will be followed Failure of NASA marked altimeter Damage due to negligent handling, water, or unintended electrical charge Altimeter does not power on or malfunctions in flight. Flight data not recoverable 2D Altimeters will be designated to a securely closed and sealed plastic box which protects from impact, water, and static charge. 1C Verified in 4.3, and UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 84

85 Piston Bind-up Pistons are misaligned with outer and inner rings Pistons lose range of motion and could become stuck, increasing risk of losing cargo. 3D Use extreme caution when bonding the pistons to mitigate misalignment. Verify all pistons are working as intended after assembly. 1C Verified through proper construction of payload system Wildlife interference Launch vehicle flies through flock of birds or is trampled by wildlife after ground hit. Disturbance or damage to flight vehicle 1D Skies will be verified as clear before flight. 1B Consultation of RSO to ensure clear launch field prior to flight Volatile Gas Failure to protect payload could allow corrosive gases from the motor or separation charges to damage the fragile payload. Payload is damaged or destroyed. 3D Verify the payload is properly protected before flight. 2B Verified in Parachute sections come apart Inadequate parachute design; poor stitching between sections Catastrophic failure of recovery system; damage to/loss of rocket and payload cargo 3E Manufacturer performs deployment test prior to shipment to the team 2C Deployment test report given by manufacturer before parachutes are delivered to the team. Suspension lines and/or risers become unattached Weak stitching or materials Catastrophic failure of recovery system; damage to/loss of rocket and payload cargo 3E Manufacturer performs deployment test prior to shipment to the team to ensure suspension line security. 2C Deployment test report given by manufacturer before parachutes are delivered to the team. Avionics bay not properly sealed Holes in rocket body or avionics bay; gaps between sections Incorrect detonation of drogue and/or main ejection charges; failure to reach desired altitude; dame to/loss of launch vehicle; errored altitude sensing 3C Teflon tape used to seal edges between avionics bay body tubes. 1D Ensure Teflon tape is available. Static deployment tests will reveal where seals are needed. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 85

86 Structural failure despite stress analysis Stress model underestimates true flight stress incurred and overestimates factor of safety Damage to vehicle or components; flight failure 3E Verify the model with multiple iterations; use conservative safety factors 1E Compression tests and validation from test launches will confirm stress model predictions Rocket comes loose from launch pad Rail buttons not securely mounted; extreme wind Rocket breaks off launch rail before rail exit; potential loss of vehicle and damage to bystanders or property. 4D Rail buttons ensured to be snugly secured to vehicle. Careful alignment when installing rocket on launch pad 2C Design analysis to align rail buttons; validation from test flights Vehicle component not brought to launch site Component not packed Rocket does not launch 1E Ensure all components are brought to launch event. Follow checklist. Bring extra components where possible 1C Verify launch operations checklist is followed and is correct Carbon fiber element failure Weak structural integrity; low material strength in concentrated area Unstable flight; damage to/loss of vehicle. 3D Follow proper procedure for carbon fiber layup, bagging, and curing 2D Consultation of fabrication protocol for composite carbon fiber operations Excess friction between rocket and launch rail Misalignment or poor installation of rail buttons; misalignment of launch rail Rocket gets stuck on launch rail; rocket does not reach desired altitude 3B Rail button alignment ensured upon installation. Proper care of rail and correct assembly of launch pad 2A Verification from subscale and fullscale launch; preflight inspection of alignment of rail buttons. 6.3 Environmental Hazard Analysis The team will perform all due diligence to mitigate the environmental impacts of flying high power rockets. The launch area will have trash receptacles to prevent litter and the team will police the launch area after all subscale launches to remove any trash. In addition, the launch pad will include a blast shield to mitigate the effects of the rockets launch. The area will also be cleared of all flammable materials for the designated radius. This mitigates the risk of starting fires at the launch site. The following table identifies environmental risks associated with the project. Risk areas are addressed, causes examined and effects listed. Mitigations are recorded to reduce the environmental impact of the project. Each risk is given a Likelihood and Severity rating. All risks bearing a moderate to high rating will be observed by the team. Table 26: Environmental Hazards Analysis UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 86

87 Likelihood & Severity Likelihood & Severity after Mitigation Control Risk Cause Effect Mitigation Control Mitigation Verification Disposing Batteries Improper disposal, malfunction leading to caustic chemicals in environment. Batteries can leak and cause corrosive damage to vegetation. Could be ingested by animals. 4B Verify batteries can be safely disposed of or stored with normal trash, make sure proper equipment is present on launch day to remove malfunctioning batteries safely. 3B Ensure proper waste disposal facilities are available Uncontrolled Black Powder ignition Improper disposal, accidental ignition Accidental ignition could cause fires on the launch site 3D Verify static electricity is discharged prior to working with black powder, verify any black powder left over is properly stored, verify that black powder that is not useable is properly disposed. 2C Ensure proper waste disposal facilities are available ensure fire extinguishers are available at the launch site Extreme weathercocking Excessive wind Unexpected vehicle trajectory; altitude not reached; vehicle drifts past launch site radius. 4D Vehicle s stability margin minimizes risk of nonstable or overstable flight. 1B Design analysis of fins and mass distribution of vehicle verified to procure acceptable stability margin. Verify with RSO that winds are minimal for launch. Wildlife interference Contact with animals, causing misalignment or damage to vehicle Unexpected vehicle trajectory, damage to vehicle or wildlife 2E All sensitive components secured inside the vehicle body; wildlife managed away from launch site 1C Verify with local land owners and RSO that wildlife are not present and launch is GO. Chemical poisoning of ponds or ground water Leakage of battery fluid or excess fuel; slag discharge from motor casing Harm to environment from motor slag and battery acid. 2C All debris and hazardous materials will be accounted for and removed quickly from the field by the team before leaving. 2C Ensure proper waste disposal facilities are available UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 87

88 Improper Motor Propellant disposal Negligent disposing of toxic grains Motor Propellant can poison vegetation 3B Verify any unused motor propellant is properly stored and that any wasted propellant is properly disposed. 2A Ensure proper waste disposal facilities are available ensure fire extinguisher is available at launch site and team members are instructed on proper use Launch Pad Fire High temperature or exposure to sunlight Fire damage to launch area; overheating of electronics; warping of structural components 3E A fire extinguisher will be present at the launch site. The launch area will be cleared by the appropriate radius. Rocket assembly performed in tent when possible. 2C Ensure fire extinguisher is available at launch site and team members are instructed on proper use. Field time minimized. The rocket s structure is compromised The rocket lands in a body of water, rough terrain such as rocks, or in a tree. The rocket s integrity and ability to be reused will be compromised. 3D The rocket shall be launched from the place which provides the largest radius of safety from environmental hazards. The launch shall not occur if there are strong winds to deter the rocket from the expected trajectory. 2C Ensure the terrain is suitable for rocket launch ensure winds are with in an acceptable speed before launching Onboard electronics damage Electronics are damaged due to overexposure to the sun, excessive humidity, or unexpected rain The launch is not possible. Electronics will need repair or replacing. Data may be unrecoverable. 3E Weather will be checked prior and during launch. Electronics will be kept in water resistant containers until needed for launch 2C Ensure electronics are sealed in a water proof container prior to launch. Ensure the weather will not interfere with electronics on launch day Catastrophic motor failure Motor explodes on launch pad Debris from vehicle, casing, propellant, etc. is spread across launch field and contaminates environment 1E Commercially tested motors that are proven for safety will be used and fire extinguishers will be present. 1D Ensure the rocket motor is of a quality commercial grade UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 88

89 Rocket is destroyed on impact. Premature or incorrect rocket separation Nonlevel launch pad Rocket experiences ballistic recovery. Shock cord breaks upon separation, other section comes loose, altimeter misfires ejection charge at incorrect altitude Pad sinks due to soft ground; pad not perpendicular to surface Materials from the rocket become lodged in ground causing environmental harm and possible contamination. Ballistic recovery of a section of the rocket. Hazardous material may spread across area of launch site, causing environmental damage Unexpected vehicle trajectory 2D 2D 3C Tests of rocket components shall be performed prior to launch. Redundant systems shall be used to ensure recovery. If the risk does occur, members shall sift through impact crater to remove small debris. Tests of rocket components shall be performed prior to launch. All components, primarily hazardous ones, shall be accounted for before departure. Plywood boards installed under pad to prevent pad legs from sinking 1C 2C 2B Verified through test launches and Open rocket simulations Verify that the rocket separates as desired through test launches and Open rocket simulations Level brought to verify proper launch pad setup and alignment. Untethered hardware lost Excessive wind, careless handling of hardware Final assembly made more difficult or not possible; personnel hazard 2C Securement of hardware in labeled boxes 1A Design and consultation of launch hardware list. Cleaning Supplies Mid-air explosion of rocket or components Improper disposal methods or littering Internal failure of payload or main engine Cleaning chemicals can harm local flora and fauna ecosystems Widespread scattering of vehicle debris 4C 1E Proper chemical disposal methods will be implemented. Trash bags and launch site litter check will be used. Motor is sourced from a reliable, commercial supplier 2B 1B Ensure proper waste receptacles are in place on launch day Consultation of motor assembly instructions Injury to wildlife Animals contact launch pad or vehicle; rocket impacts wildlife during flight or landing Animal wounds or death 1D Launch conducted in area clear of wildlife or obstables 1A Adherence to launch operations procedure; visual monitoring of launch pad UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 89

90 throughout launch day Vehicle strikes people, animals, or structures outside planned launch area Excessive wind; delayed parachute deployment; no parachute deployment Vehicle drifts outside launch field 2C Large launch field chosen and system hardware checks done prior to launch 1B Simulation testing to predict drift under nonideal conditions; launch procedures followed in choosing launch field. Litter Improper disposal due to lack of trash receptacles or littering General litter around launch site can create choking hazards for animals or remain in the environment long after launch is over. 4B The team will ensure enough trash bags are present at the launch site, the team will verify all trash has been removed from the launch area before departure. 2A Ensure proper waste receptacles are in place on launch day 7 Launch Operations Procedures 7.1 Launch System Launch Pad The Rocket Rebels launch pad has been constructed in-house using simple welding techniques and materials, yet has the weight and strength to endure the stresses incurred from launch. It features a simple tripod configuration with hollow sleeves in the central bracket. Angled leg mounts are inserted here to raise the pad above the ground and allow for easy transport. A 1/8 stainless steel blast plate deflects exhaust away from the ground to prevent fires. The hinged and locking rail erector allows for easy mounting of the vehicle and rigid locking of the 80/20 rail back to its apex. The pad has been confirmed to satisfactorily operate for rocket launches through its operation Launch Ignition System The launch ignition system consists of a 12V battery connected to a safety switch that is only active through a key insert. The safety switch leads through a 300 ft electrical wire long enough for operators to be outside of the minimum distance requirement set by the NAR to the motor igniter. The safety switch ensures that no voltage differential is delivered to the electric match igniter until a key is inserted into the safety switch and a button is pressed. 7.2 Launch Procedures The team will review system preparation requirements well in advance of the launch. Safety oversight and launch procedures are managed by the launch leader. The student safety officer, Will, will oversee all UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 90

91 operations where danger to people or rocket might exist. Team leaders will supervise the preparation of their respected areas. Formal integration of systems requires the launch leader s oversight. The team mentor will oversee the assembly of the propulsion system. This is a quality control measure to ensure that mission essential systems are properly prepared for launch. All launch procedures are divided into systems along with the personnel required Hardware List General Vehicle Assembly Portable Handheld Power Drill Phillips screw driver One box of 6-32 machine screws One box of 4-40 nylon shear pins Flathead screwdriver Masking Tape Duct Tape Launch Pad Setup 5/8 pivot bolt, two washers, and nut 3/8 locking bolt and pin 1 ½ steel tube legs 1 ¾ central bracket and rail mount 1 ½ launch rail 2x4 wood blocks Electrical wire (100 yds) Recovery Subsystem Avionics Bay w/ installed key switches Main Bay Body Tube Drogue Bay Body Tube Avionics Sled w/ StratoLoggerCF and RRC3 altimeter Forward Bulkhead Wing Nuts Two fresh 9V batteries Two drogue and two main ejection charges Main Parachute w/ Deployment Bag and Pilot Chute Drogue Parachute Shock Cords Nomex Blankets Teflon Tape Quick Links Extra electrical wiring Payload Subsystem Payload Body Tube GPS Tracker Nose Cone FMPA UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 91

92 Fragile Material (egg) Packing Peanuts Motor Subsystem Booster Body Tube Motor Case with thrust ring Forward Bulkhead Graphite Nozzle Nozzle Washer Retaining / Snap Rings, (2) large (motor case), 1 small (bulkhead) Liner Tube (black phenolic) Propellant Grains (2) Primary O-Rings (1/8 thick) (4) Tracking smoke O-Ring (3/32 thick x 1.25 OD) Liner shoulder O-Ring (3/32 thick x OD) Tracking Smoke Element Snap ring pliers Grease Aeropack outer retainer ring General Field Supplies Two tables Vehicle Assembly Stilts General Toolbox Grease Pop-up tent Garbage bags Electronics box Tape Measure Scissors Chairs Box Cutter 5-minute epoxy resin and hardener Mixing Cups Popsicle Sticks Latex Gloves Electrical Tape Film Canister of 4F Black Powder Scale Small Funnel Launch Location & Setup Procedure Unload equipment, toolboxes, and rocket components from vehicles. Setup tent, tables, and trash collection bags. WARNING: Uncollected trash is an environmental hazard. Place rocket components on tables. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 92

93 Place avionics/electronics on separate table. Ensure desired launch pad location is flat and will provide a sturdy base for rocket launch. WARNING: If ground conditions are not satisfactory, abort launch. Place launch pad components near desired launch location Launch Pad Assembly Insert the three angled legs into the central bracket sleeves. Align holes and install bolts with nuts through sleeve and leg to lock in place. Place one washer through the pivot bolt and slide the launch rail into the rail bracket. Install the pivot bolt through the 5/8 holes aligned in the rail bracket and launch rail. Place another washer on the pivot bolt. Slide the pivot bolt through the upper mount on the central bracket and screw on nut. Install the locking pin through the 3/8 hole aligned through the rail bracket and central bracket. Place a level flush with every side of the vertical launch rail and ensure that every side of the launch rail is oriented perpendicular to the launch area. o If launch rail is not aligned correctly, place small wooden planks under pad legs to adjust the rail orientation. Re-evaluate with the level to ensure correct orientation. WARNING: This step is critical for straight flight. Remove the locking pin from the pad and pivot the launch rail to the horizontal position for rocket integration. Inspect fully assembled launch pad for anomalies or damage. WARNING: If unrepairable damage exists, abort launch Aft Vehicle and Drogue Parachute Integration Required Personnel: At least two team members Inspect all vehicle components for damage or other anomalies. WARNING: If any damaged component is found to be damaged beyond repair or irreplaceable, abort launch. Clip quick links into both loops of one length of folded shock cord. Notice the parachute attachment loop located somewhere along the length of the shock cord. Clip a quick link into the attachment loop of the shock cord and into the riser loop at the end of the drogue parachute riser. Ensure the drogue parachute and suspension lines are folded cleanly, without apparent knots. WARNING: If drogue parachute seems improperly folded, refold the parachute package in accordance to the manufacturer s packing instructions. Fold the riser over the shock cord stack until a small length exists between the folded shock cord and the suspension lines. Hold in place with another piece of masking tape. Clip a Nomex blanket into the quick link of the second stack. Run the first shock cord stack s quick link through the drogue bay body tube and attach to the U- bolt of the booster section coupler. A small length of shock cord unraveling is acceptable. WARNING: Assistant needed for next steps. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 93

94 Have an assistant hold the first stack, drogue parachute, and second stack to prevent unraveling. Integrate the drogue bay over the booster section coupler and install three 4-40 nylon shear pins through the aligned holes. WARNING: Failure to use three shear pins may result in premature separation in flight. Insert the first stack, drogue parachute, and second shock cord stack into the drogue bay, respectively. Tuck the Nomex blanket around the side of and on top of the installed shock cord. WARNING: Ensure the shock cord is completely covered. Exposed shock cord could burn from ejection charge fumes Avionics Preparation and Integration Required Personnel: Avionics Lead and Team Mentor Inspect the altimeters, key switches, and battery mounts. WARNING: If any unrepairable or irreplaceable damage is found, abort launch. Verify that the altimeters and battery mounts are correctly and securely fastened to the avionics sled. WARNING: If the altimeters or battery mounts are not correctly and securely fastened to the sled, access for repair. If unrepairable or irreplaceable, abort launch. Have the avionics lead begin wiring the altimeters by first installing the battery leads into the marked positive and negative ports of each altimeter. Install two fresh 9V batteries. Install the key switch leads to each altimeter, ensuring that both key switches are in the 220 position, representing OFF. Inspect the wiring scheme of the altimeters. WARNING: If any damage or incorrect wiring is observed repair and correct. If damage is unrepairable, abort launch. Turn key switches into the on position and ensure that each altimeter is receiving power and is working correctly, evident by a series of beeps. WARNING: Each altimeter must be powering on and signaling all is well before proceeding with the assembly. WARNING: The following ejection charge preparation steps require members to be supervised by team mentor. Turn the key switches off. Retrieve two pre-made drogue charges (labeled with a D) and two premade main charges (labeled with an M) inspect them for leaks or damage. WARNING: If any ejection charges are damaged, make new ejection charges with available materials. If new ejection charges need to be made, complete the following: o Weigh out the predetermined amount of black powder needed for each ejection charge with a lightweight kitchen scale and place it into a cut off latex glove finger. o Place each charge over the correct e-match and tightly pack the charge and e-match into the tip of the glove finger. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 94

95 o Wrap each charge tightly with electrical tape ensuring no black powder is leaking from the charges. WARNING: If any black powder is leaking from the charge remake the charge and secure it, making sure no black powder leaks. Feed the leads of the drogue charges through the designated openings on the aft bulkhead of the avionics bay and attach the leads of one ejection charge to the drogue ports of each altimeter. Feed the leads of the main charges through the designated opening on the forward bulkhead of the avionics bay and attach the leads of one ejection charge to the main ports of each altimeter. Place Teflon tape over the wire of the four ejection charges and over the entire hole of the bulkhead that the ejection charge wire leads through. Carefully slide the avionics sled inside the avionics bay, feed threaded rods through the forward bulkhead, and securely fasten it to the avionics bay with wing nuts. Wrap a long strip of Teflon tape over the interface between the end of the avionics bay and forward bulkhead. Place the aft end of the avionics bay into the forward section of the drogue bay until the body tubes meet flush. Align the pre-drilled holes and secure with four machine screws Main Bay Integration Required Personnel: At least two team members Clip quick links into both ends of the second length of shock cord. Notice the parachute attachment loop located somewhere along the length of shock cord. Attach the main parachute riser loop to a quick link and clip it into the parachute attachment loop of the shock cord. Inspect the parachute package for evident damage, knots, or packing errors. WARNING: If the parachute package seems improperly packed, unpack and refold the system in accordance with the manufacturer s packing instructions. Layout the assembled bottom half of the rocket and the main bay tube horizontally on the table. WARNING: Assistant needed for next step. Feed the smaller stack of shock cord through the main bay body tube and clip the quick link into the forward end U-bolt of the avionics bay. A small length of shock cord unraveling is acceptable. Clip a Nomex blanket into this quick link. Slide the body tube onto the forward avionics coupler until the body tube is flush with the avionics ring. Ensure the Nomex blanket is tucked completely inside the body tube. Install four machine screws into the aligned holes. Slowly insert the main parachute, pilot chute, and second shock cord stack into the main bay body tube, respectively. Ensure the main parachute is inserted before the pilot chute attachment when packing. WARNING: If the pilot chute is not packed closer to the main chute separation point, the main parachute may not deploy correctly Payload Preparation and Integration Turn on GPS Tracker and request GPS coordinates to verify operation. Inspect the FMPA for apparent damage or defects. WARNING: If irreparable structural damage or defects are present, abort launch. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 95

96 Retrieve the fragile material to be flown. Unscrew the cap of the inner capsule and fill about 1/3 full with small packing peanuts. Gently place the fragile material over the bed of packing peanuts. Fill the rest of the capsule with packing peanuts. WARNING: Do not attempt to stuff peanuts into the capsule. Screw on the inner capsule cap. Attach the quick link of the main parachute shock cord to the aft coupler U-bolt on the payload subassembly. Slide the coupler inside the main bay body tube until it is flush with the payload body tube. Install three #4-40 nylon shear pins into the aligned holes. WARNING: Failure to use three shear pins may result in premature separation of payload section in flight. Retrieve nose cone and inspect for surface defects. WARNING: If nose cone defects are extensive, use replacement nose cone if available or abort launch. Place the GPS Tracker into a latex glove and use duct tape to install and secure the covered GPS tracker inside the nose cone. Fit nose cone on top of the payload body tube and install 4 machine screws into the aligned holes Rocket Motor Installation Personnel Required: Team Mentor, Team Leader, and Propulsion Lead WARNING: Perform entire rocket motor installation procedure under supervision of team mentor. Inspect all motor components to ensure that no damage occurred during transportation or handling that could result in such failures. WARNING: Failure to transport motor safely and securely may result in cracking of propellant grains and inflight motor failure. If such damage has occurred, abort mission and safely dispose of faulty motor under the supervision of the safety officer. If intact replacement reload kit is available, use this. Check that all hardware pieces are clean and free of grease and soot. Check and clean both retaining ring grooves and the delay cavity in the bulkhead. Run your finger around the inside ends of the motor case. Feel for any nicks or sharp raised metal that may cut or tear the O-rings. WARNING: If found, remove them with a sharp knife or small file prior to motor assembly. Rub a thin layer of grease on the inside of the bulkhead, each end of the case and over all O-rings. Temporarily place the O-rings on a clean surface. Place the small 1.25 OD by 3/32 thick O-ring onto the tracking smoke grain. Apply a light film of grease to the back side of the smoke grain. Loosen the head bolt at the top end of the bulkhead. Push the smoke grain into the smoke well, being careful as the O-ring is compressed into the well. Push the grain all the way in and then tighten the head bolt down. Install small snap ring in the smoke well to retain the smoke element. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 96

97 Install all O-rings: the larger black 1/8 thick O-rings onto the bulkhead, the smaller 3/32 O-ring onto the bulkhead shoulder, and the orange 1/8 thick O-ring onto the nozzle. WARNING: Check the motor liner to make sure one end is sanded and the inside corners are chamfered smooth. Lightly grease this inside end of the liner. This will be the head end of the motor liner. Next, install each of the blue and red propellant grains. The placement of grains in this reload are specific. All Blue propellant grains with the small core are to be placed at the forward end of the motor. All Red propellant grains with the large cores are to be place at the nozzle end of the motor. WARNING: If the blue and red grains are not placed properly at the correct ends, the motor will fail. If the grains are marked with lines, arrange them so that they are lined up with each other and insert them into the motor in the same position. Insert the recessed shoulder of the nozzle into the nozzle end of the liner tube and slide together (liner tube first) into the thrust ring end of the motor case (see diagram below). Place the stainless-steel nozzle washer behind the nozzle and install the retaining ring using appropriate retaining ring pliers. Install the assembled bulkhead into the top of the motor and carefully push it in straight until the bulkhead shoulder O-ring is seated into the end of the liner. Secure with the second retaining ring. WARNING: If there is a gap between the bulkhead and retaining ring, pull the bulkhead up flush against the retaining ring. There will be a small amount of empty space in the case and the grains may rattle. This is normal. Make a final physical inspection to ensure that both retaining rings are fully seated in their grooves. The motor can now be installed in the rocket. Insert the assembled motor into the motor mount tube. Verify that the positive screw cap retention ring is securely fastened to the rocket. Tighten the outer retainer ring cap onto the retainer assembly and verify it is securely fastened. Perform a final inspection of the motor mount assembly for correctness. Figure 57: Motor Bulkhead Exploded and Assembled View UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 97

98 7.2.9 Launch Vehicle Final Integration Figure 58: Assembled L840 Motor Cross-Section Required Personnel: Team Mentor, Team Leader, and Propulsion Lead Carry rocket assembly to the launch pad. Line up the airfoiled rail buttons on the rocket to the launch rail slot. Very slowly slide the rail buttons into the rail slot of the launch rail and lead the rocket down the rail until it hits the stop screw towards the bottom of the rail. WARNING: Failure to slowly slide rail buttons onto the rail may result in detaching the rail buttons and flight failure. Once the launch vehicle is seated, slowly raise the launch rail back into a vertical configuration and slide the locking pin into the aligned bracket holes. 7.3 Launch Checklist With a flat head screwdriver, turn both key switches counterclockwise one quarter turn to the ON position. Listen for the correct prelaunch sequence of beeps. After pre-launch beep sequences, both altimeters should repeat a series of three short beeps to indicate continuity of the main and drogue parachutes. WARNING: If altimeters do not turn on or malfunction, dismount vehicle from launch pad and dismantle to inspect avionics. Consult the following troubleshooting list. If irreparable damage or unknown anomaly exists, abort launch. o o o If either altimeter emits a series of one short beep, that altimeter only has continuity with the drogue parachute charge. Inspect the main bay charge for disconnects. If either altimeter emits a series of two short beeps, that altimeter only has continuity with the main parachute charge. Inspect the drogue bay charge for disconnects. If either altimeter emits no series of beeps, that altimeter has no continuity with either the drogue or main charge. Inspect both elements for disconnects. When correct altimeter operation is confirmed, unroll length of electrical wiring from the launch pad to at least 100 yards away. Install launch controller at far end of the length of electrical wire with the key out. WARNING: Ensure the launch key is not in the launch controller while final integration is still in place. Slide the igniter through the nozzle all the way up until it touches the smoke element. Secure in place with tape. Attach the leads of the igniter to the long length of electrical wiring. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 98

99 Have all personnel clear the launch area at least 100 yards away. WARNING: Do not launch if any personnel are closer to the pad than the launch controller. Install the launch controller to a 12V battery. Ensure connections are secure. Loudly announce that the range is hot and insert the launch key into the launch controller. Watch for the bright red LED on the launch controller signaling a continuous circuit. WARNING: If LED does not illuminate, remove launch key and check all connections and wiring for damage. Ensure skies are clear of aircraft and birds. Begin initial countdown to launch at T-minus five seconds. At T-minus 0 seconds, press and hold the red button on the launch controller to ignite the motor. WARNING: If the motor does not ignite, consult the following troubleshooting list. If the problem persists, abort launch. o o o o o Sixty seconds after launch attempt, approach the vehicle, turn the key switches to the OFF position, and slowly remove the igniter from the motor. Check for indications of ignition. If the igniter is slightly burned, replace the igniter and reinstall. Ensure that the launch button is held in place to allow current to light the igniter. If the igniter is not burned, check the igniter leads and the electrical wiring for damage or disconnects. Follow RSO instruction to inspect battery sources for continuity. Restart the instructions of this section again from the beginning to reattempt vehicle launch. Immediately after launch, remove key from ignition system. Disconnect launch controller leads from power supply. 7.4 Post-Launch Checklist Visually track the rocket throughout the flight. Once the main parachute has opened, begin to predict the landing position. As soon as the launch vehicle lands during a full-scale test flight begin heading towards the launch vehicle. o On competition day, as soon as a NASA official gives the OK to go recover the launch vehicle begin to head towards the launch vehicle. WARNING: Ensure a NASA official is present to record apogee when recovering the vehicle. Record apogee as measured by the altimeters by listening to the audible beeps produced by the StratoLoggerCF altimeter. Remove machine screws from the nose cone and remove inner payload capsule cap. Check for damage to fragile material. Carry the sections of the rocket back to the staging area. Mark and discard the 9V battery. Check for structural damage on the airframe. Check for rocket and payload parachute damage. Discard the spent engine casing. Check for fractures in the avionics section. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 99

100 Debrief the launch, including: motor used, rocket configuration, altitude achieved, avionics onboard, and vehicle and payload recovery. 8 Project Plan 8.1 Testing Vehicle Component and Subsystem Testing The following subsections describe the tests conducted on components of the vehicle to verify the structural integrity and robustness of the overall design Carbon Fiber Stress Testing A critical material testing procedure to conduct is tensile and shear stress analysis of the carbon fiber airframe. For obvious reasons, it is important to know how much stress the carbon fiber layup can withstand in any dimension and direction. These two tests were conducted using a Material Test System (MTS), which allows the user to manipulate samples of material with a gradually increasing load to determine ultimate shear and tensile stress limits. To determine the maximum tensile stress, a double sided hydraulic wedge grip was configured to grip a 1 by 12 sample of carbon fiber. It is important to note that all samples used were constructed and cut out from a flat layup that was of identical ply number and orientation so that strength values gathered here would be a good comparison of the actual body tube material. Figure 59: Tensile Stress Setup and Point of Failure UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 100

101 Once a sample was loaded into the wedge grips of the stress machine, a program was run to slowly pull the sample apart until failure. After failure, the max load applied was given from the computer software operating the machine. This value, along with the area of the cross section stretched, was used to calculate the ultimate tensile stress of the carbon fiber, given by: σ = F A The following table gives the stress limits calculated for a set of three nearly identical samples. The stress vs. strain curve of one of the sample pieces is shown below. Table 27: Tensile Stress Limit Calculations Sample Area (in 2 ) Max Load (lb) Ultimate Shear Stress (psi) Average Figure 60: Stress vs. Strain Curve of Sample Carbon Fiber Material Launch loads incurred during motor burn are expected to produce about 140 psi of compressive stress across the cross section of the body tubes, given by: σ = F A = 4(267.5) π( = 140 psi ) This stress is well below the stress capabilities of the carbon fiber, confirming the normal strength to be robust enough to endure flight. To conduct shear stress, a method called the Short Beam Shear Test was conducted. In this procedure, a ¾ by ¼ sample of carbon fiber was cut from the flat layup made for testing. The small samples were UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 101

102 then loaded onto a rack consisting to two curved fulcrums. This rack was placed on the MTS machine and a third curved fulcrum point was attached to the upper hydraulic port. When the machine was activated, the upper curved fulcrum pressed down on the area of the sample between the lower two fulcrum points. This induced a shear stress in the material that was gradually increased until failure. Figure 61: Lower Fulcrum Setup for Short Beam Shear Test Figure 62: Short Beam Shear Test Action Snapshot UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 102

103 After failure, the max load applied was given from the computer software operating the machine. This value, along with the area of the cross section stretched, was used to calculate the ultimate tensile stress of the carbon fiber, given by: τ =.75 P m A The following table gives the stress limits calculated for a set of three nearly identical samples. Table 28: Shear Stress Limit Calculations Sample Area (in 2 ) Max Load (lb) Ultimate Shear Stress (psi) Average Generally, the max shear stress of a material is normal stress divided by two. For this vehicle, that expected value is 70 lb. Therefore, the shear stress limits are also well above the expected stresses incurred through flight. Based off this testing, the carbon fiber layup was deemed robust enough for flight D-Printed Components Stress Analysis A stress analysis on all 3D components was required to determine their robustness, especially since the ASA and ABS material are not commonly used in sport rocketry. The objective of this testing is to compare the stress and displacement limits of the nose cone and fins at two different points in flight: max drag force and max motor thrust. To conduct this analysis, the Rocket Rebels utilized the 3D-modeling software called PTC Creo Parametric 3.0. This software is the same that was used to create and prep the models for printing in the Fortus machine. The software includes a Simulation and Analysis application that can perform a complete stress analysis of any part given certain parameters. To run the analysis, a material assignment must first be given to the part. The specific material that is used to print the nose cone and fins was not an available material in the Creo database, so PVC was selected as a material with similar material properties. Constraints were then selected for the material to determine the points where the part would be fixed. In the case of the nose cone, the constraints were along the outer surface of the shoulder and the narrow fringe that would meet flush with the wall of the vehicle body tube. For the fins, constraints were made all along the bottom edge where the fin would interface with and be epoxied to the motor mount assembly. The final parameter needed before running analysis is the loads to be exerted on the part. For this analysis, the two main loads to be investigated were the drag force, referenced from the drag force calculations procured from OpenRocket simulations, and motor thrust, referenced from the motor thrust curve published by the manufacturer. These loads had to be given a magnitude, orientation, and surface location. The surface locations were chosen from where the loads were known to interface with the part. For example, 40 lb of total drag force was applied along the entire leading edge of the fin and 83.5 lb of total thrust was applied parallel to the bottom-most surface of the fin, closest to the motor mount tube. In this instance, the forces on the fin during maximum drag force were being analyzed. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 103

104 Once parameters had been determined and the software considered the analysis runnable, a von Mises and Displacement analysis was conducted. The most passes of analysis were selected to be converged into the most accurate and precise value of stress and displacement over every point on the part. This data was displayed into the following easy to read windows. Figure 63: Max Drag Force Analysis of Nose Cone UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 104

105 Figure 64: Max Motor Thrust Analysis of Nose Cone Figure 65: Max Drag Force Analysis of Fin UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 105

106 Figure 66: Max Motor Thrust Analysis of Fin As can be seen in the figures above, the points of max stress and displacement are labeled. The following table consolidates the max stress and displacement of the stress analysis performed: Table 29: Max Stress and Max Displacement of 3D Printed Parts Analysis Part Max von Mises Stress (psi) Max Displacement (in) Max Drag Force Nose Cone e-05 Max Motor Thrust Nose Cone e-05 Max Drag Force Fin e-03 Max Motor Thrust Fin e-04 From the results given, it can be implied that the nose cone can withstand all the stresses incurred during flight. The point of max stress is at the gap transitioning from the shoulder to the outer surface of the nose cone, where the payload body tube meets flush with the nose cone. The data point of most consideration is the max stress incurred on the fin during the point of max drag force. Evidently, the drag force applied to the fins induces a bending moment that is transmitted to the points of constraint in the fin, at the junction between the actual fin and the fin tab. This is where the most stress is concentrated and proves the necessity for strong and functional fin fillets. Without these fillets, there is a risk of deforming the fins and losing the vehicle. Though the stress limits of the ABS material is well above this stress concentration, the fin fillets will add more assurance for a strong flying component. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 106

107 Static Deployment Testing A complete, ground-based test of the launch vehicle recovery system is performed prior to each subscale and full-scale. This is to assure that the electric matches, parachutes, and drogue and main parachute deployment charges meet safety and operational standards. To test the deployments, the launch vehicle is assembled in the horizontal position using mass simulators for excluded payload components. During assembly, 4F Black Powder charges are carefully placed in the designated blast locations on the forward and aft bulkheads of the avionics bay. For these tests, 3 grams of powder was used for the drogue charges and 4 grams was used for the main charges. The vehicle is then constructed as it would be for a true launch. The edges of the bay are sealed with Teflon tape to protect the internal instrumentation from forces experienced during the black powder ignition. Igniters are connected to a custom electrical relay control mechanism that allows for remote, manual ignition. The area is cleared for obstructions and personnel as verified by Safety Officer Will Thomas. Once the range is clear, the drogue blast charges are fired manually. After a successful ignition and separation, the launch control system is wired to the main blast charges and these are fired manually. After verifying both separation events, the Safety Officer assures there is no un-ignited black powder in the launch vehicle which could cause a safety hazard. Once this is verified, the test of the rocket separation is complete and equipment is transported back to the CME for manual inspection for damage. A successful deployment test was conducted before the subscale launch in December and the full-scale test in February. Below are pictures taken of the full scale static deployment tests Payload Testing Figure 67: Full Scale Static Deployment Test Snapshots The main payload testing procedures performed were trial and error drop tests of the completed fragile material protection apparatus (FMPA) with a plausible fragile material, an egg, installed. This will be conducted in the factory floor. The apparatus will be raised to a point where the potential energy from that height with respect to the floor is equal to the amount of energy transferred upon ignition and launch of the rocket from the launch pad. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 107

108 The objective of this test is to test how receptive the FMPA system is to impulse forces that are like that of a rocket motor ignition. Success criteria are as follows: The impulse of the motor and the height at which the FMPA must be dropped are calculated to be similar in magnitude. The fragile material is successfully loaded into the FMPA. The FMPA is dropped at the correct height. The FMPA is not extensively damaged from the drop test. The fragile material is intact from the drop test. As stated in section 5.5, the principle of the system is to dissipate the impulse transferred from the motor to the FMPA. This concept is proven through the equations for resultant energy dissipated by a spring and damper system given below, where k equals spring constant, v equals velocity, x equals displacement, and b equals damper constant: E damper = b v x E spring = 1 k x2 2 The energy produced from the burn of the motor is the greatest energy and is thus the most significant to be dissipated from the payload. To dissipate the impulse delivered from the motor, the thrust curve is consulted to determine the area under the curve which is equal to the impulse via the equation: I = Ft The biggest force dissipation will occur during the 4.6 second burn time of the rocket motor, where a force steadily decreasing from 267 lbf will be transferred to the FMPA up until motor burnout. However, the complexity of this problem lies with the parabolic and ill-defined expression of the thrust curve for the force overtime. To simplify calculations and determine an accurate estimate for the most impulse to be dissipated, the average thrust was integrated over the first three seconds of motor thrust that is delivered to the FMPA which totals as: I = Ft = 189 lb 3 sec = 567 lb s The shock absorbers must be durable enough to dissipate this impulse. Another expression for the total impulse exists which is given as: I = mv This will be the means of testing if the FMPA will be able to dissipate the major launch impulse. The mass and velocity variables given here can be used to solve for the equivalent kinetic energy of the same FMPA system. Knowing the mass of the entire vehicle and expected impulse dissipation, the velocity can be solved as: v = I m = 567 lb s 1.16 slugs = And the mass and velocity of the FMPA determine the equivalent kinetic energy: KE = 1 2 mv2 = 1 2 (. 0233)(488.8)2 = J This kinetic energy can then be converted to a potential energy due to gravity given by: KE = PE = mgh So, the ideal height to drop the FMPA by is given as: h = PE mg = = 3758 ft (. 023)(32.2) UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 108

109 Note that this is the ideal height from which to drop the apparatus in a vacuum. Obviously, these means are not possible to accommodate. Therefore, the test can only partially test the structural integrity of the apparatus. The apparatus will be dropped from incrementally increasing heights of five feet up to about twenty feet. This will provide an adequate measurement for the strength of the FMPA. The main independent variable to test is the stiffness of the shock absorber system. Too many pistons would transfer force through the system to the fragile material. Not having enough pistons would cause force to be transferred to the fragile material after all pistons had fully contracted from the initial load. If more pistons are needed, small increases in stability margin, weight, and total cost would be the main consequences. To conduct this test, a thin rope was tied around the bottom of the cap of the inner capsule and looped around the hand rail of the catwalk that goes over the factory floor. The necessary height was measured vertically with a tape measure hanging from the catwalk. The capsule was raised to the indicated heights in the following table and then released. On the floor, a compact square of wood is positioned just under the FMPA so that the lower outer ring will strike the wood fixture. This allows for the capsule to travel past the outer ring and engage the full stroke length of the shock absorbers. The FMPA is then retrieved and analyzed for damage. The cap is then removed and the cargo is inspected for damage. The results of this experiment are as follows. To determine the height from which to drop the FMPA, the following table was used as a guideline. It has been completed with the results from the test conducted. Height Dropped (ft) Table 30: Drop Test Data Table Projected Impulse Dissipated (N-s) Survival of Cargo? Y N Y N Y N Y N Evident Damage to FMPA? The test proves that the apparatus can withstand and dissipate a portion of the impulse delivered to the apparatus. This doesn t mean that the FMPA can t protect the cargo during a full-scale flight so a 2 nd test launch will be performed to test the full extent of the FMPA s protection capabilities. 8.2 Requirements Compliance The following table lists all requirements, the verification plan chosen, and the execution methods which will be performed to verify the requirements are fulfilled. Status updates as of January 13 th, 2017, are also included to show progress of each requirements compliance. These updates are highlighted based on their completion progress: green means the requirement is completed, readily verifiable, or non-applicable to the Rocket Rebels project, yellow means the execution is still in progress, and red means the verification is planned but has not yet been started. Table 31: Verification Plan Vehicle Requirements Requirement Shall Statement Verification Plan 1.1 The vehicle shall deliver the science or Execution Method Verify apogee is 5,280 through OpenRocket Status Projections show a less than 100- UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 109

110 engineering payload to an apogee altitude of 5,280 feet above ground level (AGL). Analysis Testing simulation and test launches foot difference in target altitude based off the most recent and accurate model simulations. 1.2 The vehicle shall carry one commercially available, barometric altimeter for recording the official altitude used in determining the altitude award winner. A high-quality altimeter will be used to record the altitude for the vehicle. Complete. Altimeters have been purchased, designated, and tested The official scoring altimeter shall report the official competition altitude via a series of beeps to be checked after the competition flight. Inspection Altitude measurement will be performed after the flight through beeps Test flights confirm the post flight altitude report works nominally Teams may have additional altimeters to control vehicle electronics and payload experiment(s). Inspection One backup altimeter will be used to ensure redundancy of the avionics system Test flight confirms that backup altimeter works nominally At the LRR, a NASA official will mark the altimeter that will be used for the official scoring. Inspection NASA official will be informed of altimeter to score Primary altimeter is designated and is proven to work as expected At the launch field, a NASA official will obtain the altitude by listening to the audible beeps reported by the official competition, marked altimeter. Inspection Functionality of altimeter will be verified for ease of altitude measurement Test launch confirms that primary altimeter post-flight report operates as expected At the launch field, to aid in determination of the vehicle s apogee, all audible electronics, except for the official altitude-determining altimeter shall be All altimeters will be off when determining the altimeter s measured altitude via external key switches. Key switches have been verified to work nominally for altimeter cut off. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 110

111 capable of being turned off Team shall receive zero altitude points if the official, marked altimeter is damaged and/or does not report an altitude via a series of beeps after the team s competition flight. Inspection Altimeter will be inspected pre-flight for damage or not operating at all Full scale test launch confirms that altimeter can withstand expected fullscale loads The team will receive zero points if it does not report to the NASA official designated to record the altitude with their official, marked altimeter on the day of the launch. Inspection The NASA official will be accounted for and present during the recording of the altitude Launch operations account for NASA official to be present to record altitude The team will receive zero points if the altimeter reports an apogee altitude over 5,600 feet AGL Testing Test launches will be conducted prior to launch day. Weight will be added or an alternative motor will be selected in case of extreme deviation from expected altitude Off-limits apogee altitude is very unlikely to occur with final motor selection and current vehicle configuration. See section The team will receive zero points if the rocket is not flown at the competition launch site. The vehicle will be flown at the competition launch site Competition launch site and date of launch has been confirmed. 1.3 All recovery electronics shall be powered by commercially available batteries. 9V batteries will be used to power all altimeters and electronics Test launch confirms that commercial batteries are capable for vehicle flight. 1.4 The launch vehicle shall be designed to be recoverable and reusable. Inspection The vehicle is designed to be reloaded and reassembled for launch within the same day Subscale launch and recovery have proven that the vehicle is capable of being reused. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 111

112 1.5 The launch vehicle shall have a maximum of four (4) independent sections. The vehicle will be composed of three sections maximum Final vehicle design consists of only three independent sections. 1.6 The launch vehicle shall be limited to a single stage. The vehicle is designed to be a single stage or single motor vehicle Final vehicle design consists of a single stage. 1.7 The launch vehicle shall be capable of being prepared for flight at the launch site within 4 hours, from the time the Federal Aviation Administration flight waiver opens. Testing Testing will be done post-construction for swiftness of assembly to ensure that the time limit is not exceeded Subscale preflight assembly time trials have confirmed that the vehicle is capable of flight preparedness in less than 2 hours. 1.8 The launch vehicle shall be able to remain in launch-ready configuration at the pad for a minimum of 1 hour without losing the functionality of any critical on-board component. Testing Inspection The altimeter operational life span on a 9V battery will be analyzed to ensure it meets the time threshold The subscale vehicle has been tested to remain in launch-ready configuration for one hour without losing functionality of any critical onboard components. 1.9 The launch vehicle shall be capable of being launched by a standard 12V direct current firing system. The firing system will be provided by the NASA-designated Range Services Provider Testing The vehicle is designed to be ignited through a 12V firing system The subscale test launch confirms that the vehicle is capable of being launched through any system that utilizes a 12V direct current system 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 vehicle will not necessitate special ground support equipment to ensure a successful flight No special ground support equipment is necessary for final vehicle design. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 112

113 1.11 The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) Final motor choices must be made by the Critical Design Review (CDR) Any motor changes after CDR must be approved by the NASA Range Safety Officer (RSO), and will only be approved if the change is for the sole purpose of increasing the safety margin Pressure vessels on the vehicle shall be approved by the RSO and shall meet the following criteria: The minimum factor of safety (Burst or Ultimate pressure versus Max Expected Operating Pressure) shall be 4:1 with supporting design documentation included in all milestone reviews The low-cycle fatigue life shall be a minimum of 4: Each pressure vessel shall include a solenoid pressure relief valve that sees Analysis Inspection Analysis Inspection Analysis Inspection Commercial motors will be used in the making of the vehicle. No motors will be manufactured. Availability of primary motor choice inspected, simulation analysis performed on alternatives in case of unavailability Motor change will be heavily analyzed for safety and functionality if the case arises that the current motor choice is unfit for flight No pressure vessels are present on the vehicle design Final motor choice is a commercially available brand that utilizes APCP propellant Final motor choice has been confirmed. No present need exists to change motor choice. n/a n/a n/a n/a n/a n/a n/a n/a n/a n/a UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 113

114 the full pressure of the tank Full pedigree of the tank shall be described, including the application for which the tank was designed, and the history of the tank, including the number of pressure cycles put on the tank, by whom, and when The total impulse provided by a Middle and/or High School launch vehicle shall not exceed 5,120 Newton-seconds (Lclass) The launch vehicle shall have a minimum static stability margin of 2.0 at the point of rail exit The launch vehicle shall accelerate to a minimum velocity of 52 fps at rail exit All teams shall successfully launch and recover a subscale model of their rocket prior to CDR The subscale model should resemble and perform as similarly as possible to the fullscale model, however, the full-scale shall not be used as the subscale model. n/a n/a n/a n/a n/a n/a Analysis Testing Testing Testing Analysis Testing The rail exit stability is analyzed through OpenRocket and is proven to be above 2.0 at rail exit OpenRocket simulations and designs will be updated to ensure rail exit velocity does not fall below threshold. Timeline management will ensure a sub scale rocket will be launched prior to the due date of the CDR. Weight and dimensional scaling in OpenRocket and small modifications will be made to make the subscale Stability analysis through simulated flight has been determined. See section Rail exit velocity has been simulated to be above threshold. See section 4.8 Subscale flight has been confirmed as a successful launch and recovery. See PDR for flight data Subscale model dimensions closely resembled the full-scale model as a 67%- dimensional scale model. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 114

115 The subscale model shall carry an altimeter capable of reporting the model s apogee altitude. Testing There will be a primary and secondary altimeter, each capable of recording 20 altitudes per second. This will be tested and recorded during the launch of the subscale rocket. The subscale model carried two altimeters that both recorded an apogee altitude of 3940 ft. See section All teams shall successfully launch and recover their fullscale rocket prior to FRR in its final flight configuration. The rocket flown at FRR must be the same rocket to be flown on launch day. The following criteria must be met during the full-scale demonstration flight: Proper launch operations will be followed to ensure a successful launch and timeline management will ensure it is done on time. Full-scale vehicle successfully launched and recovered prior to FRR. This vehicle will be the same vehicle flown at launch day The vehicle and recovery system shall have functioned as designed. Analysis Test flights and simulation models will be conducted to ensure the vehicle and recovery system perform as intended Full-scale test launch confirms that the vehicle and recovery system function as designed The payload does not have to be flown during the full-scale test flight. The following requirements still apply: Testing Analysis Design collaboration will determine whether a subscale payload system or ballast mass will be used in the subscale payload The FMPA payload was not flown during the subscale launch. An on-board camera and ballast with equivalent mass were used instead If the payload is not flown, mass simulators shall be used to simulate the payload mass. Testing If a ballast is used it will be a mass just slightly higher than the entire payload system, to as closely as possible account for the unknown payload. Ballasts were used to simulate the scaled mass of the payload system The mass simulators shall be in the same The ballast will be situated in the payload Ballasts were installed in the UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 115

116 approximate location on the rocket as the missing payload mass. Analysis bay in the same configuration as the apparatus same location as the payload would have been If the payload changes the external surfaces of the rocket (such as with camera housings or external probes) or manages the total energy of the vehicle, those systems shall be active during the fullscale demonstration flight. Testing Analysis All payload changes that could alter total energy will be active during full-scale flights The FMPA, which alters the total kinetic energy of the vehicle as it falls, was installed on the vehicle for full-scale demonstration flight Test flights shall simulate maximum velocity and maximum acceleration of the launch day flight Testing The full-scale motor is intended to be used for test launches Full scale motor and all on-board components will be installed for test flights to ensure maximum acceleration is known The vehicle shall be flown in its fully ballasted configuration during the full-scale test flight. Testing The vehicle will be flown with the same amount of ballast for test launches and launch days The vehicle is not expected to use any ballast for full-scale configuration After successfully completing the fullscale demonstration flight, the launch vehicle or any of its components shall not be modified without the concurrence of the NASA Range Safety Officer (RSO). Inspection Testing Any changes that must be made after full scale test flights will be tested thoroughly through simulation to ensure successful and risk free flight The full-scale vehicle modifications confirmed by NASA before continuation Full scale flights must be completed by the start of FRRs (March 6th, 2017) Testing The timeline reflects flights being completed before March 6 th, 2017 Full scale test flight conducted on February 18, Any structural protuberance on the rocket shall be located Inspection Testing No structural protuberance will be present on the vehicle No structural protuberances exist forward of the burnout center UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 116

117 aft of the burnout center of gravity The launch vehicle shall not utilize forward canards The launch vehicle shall not utilize forward firing motors The launch vehicle shall not utilize motors that expel titanium sponges The launch vehicle shall not utilize hybrid motors The launch vehicle shall not utilize a cluster of motors The launch vehicle shall not utilize friction fitting for motors The launch vehicle shall not exceed Mach 1 at any point during flight Vehicle ballast shall not exceed 10% of the total weight of the rocket Testing Analysis Testing Recovery System Requirements The vehicle only utilizes a trio of main fins in the rear of the rocket The vehicle only utilizes a rear firing motor The motor prepollent data sheet will be inspected to ensure titanium sponges are not present Hybrid motors will not be considered for motor selection A single motor will be used for the vehicle launch Retaining rings and aerospace epoxy are components used to ensure the fit and robustness of the motor housing The max speed will be tested through simulation to ensure max velocity does not exceed Mach 1 Need for ballast will be determined through simulation and test flights. This ballast will not exceed 10% of the vehicle s weight of gravity. See Section 4. Final design does not include the prohibited component. See Section 4. Final design does not include the prohibited component. See Section 4. Final design does not include the prohibited component. See Section 4. Final design does not include the prohibited component. See Section 4. Final design does not include the prohibited component. See Section 4. Final design does not include the prohibited component. See Section 4. Simulation flights confirm that the full scale shall not exceed Mach 1. See section 4.8 No ballasts are included in the full-scale design. See section 4. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 117

118 Requirement Shall Statement Verification Plan Vehicle must deploy a drogue parachute at apogee and a main parachute at a lower altitude. Team must perform a successful ground ejection test for both the drogue and main parachute. The landing of the vehicle can have no more kinetic energy than 75 ft-lbf. 2.4 The recovery system electrical circuits will be completely independent of any payload electrical circuits. 2.5 The recovery system shall contain redundant, commercially available altimeters. The term altimeters includes both simple altimeters and more sophisticated flight computers. 2.6 Motor ejection is not a permissible form of Analysis Testing Testing Analysis Testing Testing Inspection Analysis Execution Method Dual deployment altimeters and two separate recovery parachutes will be used on board the craft that will deploy via ejection charges at both apogee and a lower altitude. Each parachute ejection system will be tested before the first flight. Recovery system designed in relation to the mass of the vehicle to ensure acceptable kinetic energy. No other electrical systems will be present in the vehicle payload or the rest of the vehicle Two altimeters will be employed by the recovery avionics system. One will be the primary altimeter, the other a backup in case of failure. The avionics bay will house separate charges from the motor. These charges will be Status Complete Ejection tests have been performed with success on subscale vehicle. Tests will be conducted for full-scale as well. OpenRocket simulations show that the heaviest rocket section will not create generate excessive kinetic energy before impact. Complete Two simple altimeters have been purchased, designated, and tested in subscale flight. Complete UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 118

119 primary or secondary deployment. Testing activated at the appropriate point of flight by the avionics system. 2.7 Each altimeter shall be armed by a dedicated arming switch that is accessible from the exterior of the rocket airframe when the rocket is in the launch configuration on the launch pad. Inspection A dedicated key switch that slightly protrudes from the avionics ring, for accessibility, will be connected to each of the two altimeters Key Switches have been tested and purchased. 2.8 Each altimeter shall have a dedicated power supply. Testing Avionics system will be designed with dedicated power supplies for altimeters. Fresh 9V batteries used for each flight. 2.9 Each arming switch shall be capable of being locked in the ON position for launch. Testing Inspection Arming switches will be tested to ensure safe locking. They will also be inspected before launch. Key switches have hold on/off switch capacity Removable shear pins shall be used for both the main parachute compartment and the drogue parachute compartment. Testing Analysis Shear pin integration will be designed to ensure they are removable Complete An electronic tracking device shall be installed in the launch vehicle and shall transmit the position of the vehicle or any independent section to a ground receiver. Inspection An electronic GPS tracker shall be used to transmit location data. GPS tracker designated and purchased Any rocket section, or payload component, which lands untethered to the launch vehicle, shall also carry an active electronic tracking device. Testing Analysis Vehicle will be designed with no untethered components so as only one tracker will be required. Complete The electronic tracking device shall Testing Tracking system will be tested prior to flight confirmed UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 119

120 be fully functional during the official flight on launch day. Inspection launch to ensure proper operation during flight. operational integrity of tracking device The recovery system electronics shall not be adversely affected by any other on-board electronic devices during flight (from launch until landing). Identification GPS tracker is separated from altimeters by carbon fiber bulkheads, which will block radio frequency from exciting altimeters Test flight and deployment tests confirm altimeter and GPS operate without interference The recovery system altimeters shall be physically located in a separate compartment within the vehicle from any other radio frequency transmitting device and/or magnetic wave producing device. Analysis Testing The GPS tracker will be housed in the nose cone, which is adequate distance from the avionics bay. Design finalized. Testing will confirm operational integrity The recovery system electronics shall be shielded from all onboard transmitting devices, to avoid inadvertent excitation of the recovery system electronics. Analysis Testing Appropriate shielding materials will be implemented in avionics bay design. Testing will be performed to ensure no interference will occur during flight. Test flights confirm altimeters were shielded The recovery system electronics shall be shielded from all onboard devices which may generate magnetic waves (such as generators, solenoid valves, and Tesla coils) to avoid inadvertent excitation of the recovery system. Analysis Testing Avionics design will ensure no magnetic interference occurs. No subsystems on the vehicle generate magnetic waves. Complete The recovery system electronics shall be shielded from any other onboard devices which may adversely affect the proper operation of the Analysis Testing Avionics bay will be designed in such a way to shield avionics from any interference that may impact performance during flight. Carbon fiber enclosed avionics bay ensures no interference from other electronics. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 120

121 recovery system electronics. Experiment Requirements Requirement Shall Statement Verification Plan Each team shall choose one design experiment option from the following list Additional experiments are encouraged and may be flown, but they will not contribute to scoring If the team chooses to fly additional experiments, they shall provide the appropriate documentation in all design reports so experiments may be reviewed for flight safety Teams shall design a container capable of protecting an object of an unknown material and of unknown size and shape The container shall be able to accommodate multiple of one replicated object The object(s) shall survive throughout the Test Analysis Inspection Test Analysis Inspection Test Execution Method The team will take on the task of Fragile Material Protection There will be no additional experiments to be contributed to the initial experiment. No additional experiments will be flown, no extra documentation needed The container will be designed to protect one or more unknown fragile objects during flight. Each design will be tested to observe their performance in a rocket simulation for many differently sized materials. The testing of each design will be done with multiple objects. The design process will include analysis of having multiple objects to protect during a single flight. The various designs will be analyzed to Status Complete. Complete. Complete. Final design has been tested and confirmed for flight operation. Complete. flights and drop UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 121

122 entirety of the flight. Analysis Inspection protect from impact from all directions. tests confirm confidence of success for launch competition Teams shall be given the object(s) at the team check in table on launch day The team will be sure to check in at the check in table to receive the object(s) Launch procedures ensure object is acquired during setup Teams may not add supplemental material to the protection system after receiving the object(s). Once the object(s) have been provided, they must be sealed within their container until after launch. Inspection The payload will not be filled with any additional material, so that when the object is given the Rocket Rebels will only be required to insert the object inside the capsule and seal the container. Complete The provided object can be any size and shape, but will be able to fit inside an imaginary cylinder 3.5 in diameter, and 6 in height The object(s) shall have a maximum combined weight of approximately 4 ounces. Test Inspection Analysis Test Inspection Analysis Safety Requirements Requirement Shall Statement Verification Plan 4.1 Each team shall use a launch and safety checklist. The final checklists shall be included in the FRR report and used during Testing Flight simulating testing with objects that fit the volume and mass tolerances will be used. The design will be configured around these tolerances. The structure of the payload design was created to sufficiently support 4 ounces of weight. Testing with objects weighing 4 ounces will be best to calculate the strength of each payload design. Execution Method Test runs of launch and safety checklists will be done to ensure soundness and safety Container purchased with accommodating dimensions. Cargo weight taken into consideration of design. Status Full-scale test launch confirmed launch procedures checklist as valid and correct. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 122

123 the Launch Readiness Review (LRR) and any launch day operations. is maximized for launch day operations 4.2 Each team must identify a student safety officer who shall be responsible for all items in section 4.3. A safety officer has been selected and will coordinate safety operations throughout the project Complete The safety officer shall monitor team activities with an emphasis on Safety during design of vehicle and launcher Inspection The design of the vehicle and launcher has shown in simulations to allow the rocket to reach safe speeds when leaving the launch rail. The vehicle has been shown to reach the desired apogee and recover safely. FMEA, PHA, and Environmental Concerns updated as needed to improve safety and reduce risk The safety officer shall monitor team activities with an emphasis on Safety during construction of vehicle and launcher Inspection Construction of the vehicle will take place on the shop floor. During all phases of construction, the team will be briefed on the proper way to use the machines involved. The team will also receive briefings when dealing with potentially harmful materials. When working on the shop floor, the team will be supervised by shop technicians to ensure all procedures are conducted in the safest way possible. FMEA, PHA, and Environmental Concerns updated as needed to improve safety and reduce risk The safety officer shall monitor team activities with an emphasis on Safety during assembly of vehicle and launcher Inspection The team will follow pre-launch assembly checklists to ensure that safety procedures will be followed and FMEA, PHA, and Environmental Concerns updated as needed to improve safety and reduce risk UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 123

124 The safety officer shall monitor team activities with an emphasis on Safety during ground testing of vehicle and launcher The safety officer shall monitor team activities with an emphasis on Safety during sub-scale launch test(s) The safety officer shall monitor team activities with an emphasis on Safety during full-scale launch test(s) The safety officer shall monitor team activities with an emphasis on Safety during launch day The safety officer shall monitor team activities with an emphasis on Safety during recovery activities The safety officer shall monitor team activities with an Inspection Inspection Inspection Inspection Inspection Inspection the rocket will be properly assembled. Prior to ground testing, the team will be briefed on procedures used during the test. Only designated people will be allowed to load the separation charges. Subscale tests will take place on designated NAR launch sites. The team will use procedure checklist for launch day. Full scale rocket tests will be conducted at a NAR officiated rocket launch site. The primary launch site the team will be using is in Memphis, TN. The team will develop a checklist for procedures on launch day. The team will be briefed on what is expected to happen. When recovering the rocket, the team will go in pairs and take a cellphone with them. This ensures that if an injury were to occur, the team can get proper medical help as quickly as possible. The team will not attempt to recover the rocket should it present a personal hazard to the team. Prior to any Educational Engagement FMEA, PHA, and Environmental Concerns updated as needed to improve safety and reduce risk FMEA, PHA, and Environmental Concerns updated as needed to improve safety and reduce risk FMEA, PHA, and Environmental Concerns updated as needed to improve safety and reduce risk FMEA, PHA, and Environmental Concerns updated as needed to improve safety and reduce risk FMEA, PHA, and Environmental Concerns updated as needed to improve safety and reduce risk FMEA, PHA, and Environmental Concerns updated UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 124

125 emphasis on Safety during educational Engagement Activities Activities, the team will coordinate all planned activities with the proper administration. The team will always work in pairs when doing outreach work. Working in pairs mitigates many of the concerns that come with outreach work. as needed to improve safety and reduce risk The safety officer shall implement procedures developed by the team for construction, assembly, launch, and recovery activities Analysis Testing Procedures will be made prior to each build to ensure time and safety management during building Construction, launch, assembly, and recovery procedures have been made The safety officer shall manage and maintain current revisions of the team s hazard analysis, failure modes analysis, procedures, and MSDS/chemical inventory data Inspection The safety officer will have an updated record of all MSDS and chemical inventory data, as well as all relevant safety analysis documentation FMEA, PHA, and Environmental Concerns updated as needed to improve safety and reduce risk The safety officer will assist in the writing and development of the team s hazard analyses, failure modes analyses, and procedures. Analysis The safety officer will lead the effort of creating, updating, and revising all safety analysis and procedures Complete 4.4 Each team shall identify a mentor. A team mentor has been identified Complete 4.5 During test flights, teams shall abide by the rules and guidance of the local rocketry club s RSO. Inspection The RSO has final say in all safety concerns and the team will adhere to his instruction Complete 4.6 Teams shall abide by all rules set forth by the FAA. Inspection All laws relevant to high power rocketry will be acknowledged and obeyed FAA waiver regulations are observed and UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 125

126 General Requirements Requirement Shall Statement Verification Plan 5.1 Students on the team shall do 100% of the project, including design, construction, written reports, presentations, and flight preparation except for those procedures required to be done by the team s mentor. 5.2 The team shall provide and maintain a project plan to include, but not limited to the following items: project milestones, budget and community support, checklists, personnel assigned, educational engagement events, and risks and mitigations. 5.3 Foreign National (FN) team members shall be identified by the Preliminary Design Review (PDR) 5.4 The team shall identify all team members attending launch week activities by the Critical Design Review (CDR) Inspection Inspection Inspection Execution Method Rocket Rebels will do the entirety of the tasks without any outside completion from advisors, professionals, or nonstudent team members. Within Rocket Rebels, each sub-system team will complete their respective tasks with only the members of Rocket Rebels. Rocket Rebels will create a project plan including all necessary components. Each subsystem shall complete their respective portions and those that are nonspecific to a group. The Rocket Rebels team is not comprised of any Foreign National team members. Rocket Rebels will indicate the members attending launch week activities before CDR. An effort will be made to include the maximum number of students actively engaged in the project. obeyed by team during launch. Status Team acknowledges and accepts workload responsibility. Project plan completed as required. Complete. Complete. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 126

127 5.4.1 Students actively engaged in the project throughout the entire year. Inspection Students will be accessed for their participation throughout the project One mentor Inspection One team mentor has been identified No more than two adult educators 5.5 The team shall engage a minimum of 200 participants in educational, hands-on science, technology, engineering, and mathematics (STEM) activities, as defined in the Educational Engagement Activity Report, by FRR. An educational engagement activity report shall be completed and submitted within two weeks after completion of an event. 5.6 The team shall develop and host a Web site for project documentation. 5.7 Teams shall post, and make available for download, the required deliverables to the team Web site by the due dates specified in the project timeline. Inspection Inspection One adult educator has been identified During all outreach events, interactive activities shall be coordinated to reach the maximum number of participants in the most hands-on way possible. Unaffiliated third parties or advisors shall sign off where activities of direct engagement are present. Visual proof, signatures from participants, and signin sheets will ensure the team is held accountable for participants reached. All activities shall be done in relationship to STEM principles. Rocket Rebels has created a website to show the progress of the project and to maintain transparency through documentation. Rocket Rebels will complete and submit all items to be delivered by the specified due dates as per the project timeline. Complete. Complete. Complete. Educational engagement events conducted and activity reports have been submitted for review. Complete Complete UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 127

128 5.8 All deliverables must be in PDF format. Rocket Rebels will convert all files to be delivered in PDF format. Complete 5.9 In every report, teams shall provide a table of contents including major sections and their respective subsections Rocket Rebels will provide a table of contents listing the required sections and sub-sections at the beginning of every submitted report. Complete 5.10 In every report, the team shall include the page number at the bottom of the page. Rocket Rebels will include a page number at the bottom of each page for every report. Complete 5.11 The team shall provide any computer equipment necessary to perform a video teleconference with the review board. If possible, the team shall refrain from use of cellular phones as a means of speakerphone capability. Rocket Rebels will utilize the technology provided by the CME facility to perform video teleconference. A teleconferencing system with internet video conferencing is available for use in the CME. Complete 5.12 All teams will be required to use the launch pads provided by Student Launch s launch service provider Rocket Rebels will utilize a custom pad for test launches but will use the launch pads provided by the Student Launch s launch service provider on launch day Complete Teams must implement the Architectural and Transportation Barriers Compliance Board Electronic and Information Technology (EIT) Accessibility Standards (36 CFR Part 1194) Inspection Rocket Rebels will adhere to the Architectural and Transportation Barriers Compliance Board EIT Accessibility Standards which requires no undue burden defined as a significant expense or difficulty. Complete UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 128

129 8.3 Team Derived Requirements As a new team, these requirements will be very simple design requirements that ensure that certain components of the rocket developed from a new USLI team are robust enough to help coordinate the successful delivery of the payload to a mile apogee and back to ground with minimal stress on the fragile material within the payload apparatus. Requirement Stress analysis will be performed on key vehicular components (i.e. ABS plastic and carbon fiber Carbon fiber body tube will withstand stress values with a safety factor of at least 2 The launch vehicle shall not weight more than 40 lbs. Launch vehicle shall be capable of being prepared at launch site within 2 hours. Center of gravity and center of pressure will be known to the nearest.01 before launch. The descent speed of the rocket after the drogue parachute is released should be greater than 50 ft/s to minimize drift. All team members involved in composite layup will wear latex/vinyl gloves when handling chemicals or carbon fiber sheet. Table 32: Team Derived Verification Plan Verification Plan Testing Testing Inspection Analysis Analysis and Inspection Execution Method Tensile and Shear Stress Tests will be conducted in the UM ME Labs. Creo Parametric 3.0 will simulate launch loads on models of the 3D printed nose cone, fins, and motor mount sleeve. Tensile and Shear Stress Tests will determine limits of carbon fiber samples and be compared to expected stresses. The weights of all components will be summed to verify total vehicle weight is less than 40 lbs Launch procedures will be verified at test launches for timeliness. OpenRocket computation will be performed with constructed model to determine CP and CG locations. Simulation analysis and onboard sensors will monitor speed during flight demonstration and verify drogue descent speed. Safety officer will verify use of proper equipment during operations. Status Stress analysis conducted. Results in 4.9 Stress analysis conducted. Results in Weight below threshold. See Vehicle Overview Vehicle prepared in one hour and 20 minutes during test launch. See Mission Performance Predictions Complete. See Full Scale Flight Data and Mission Performance Predictions See section 6.1 UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 129

130 Eye protection will be worn always while on the factory floor Inspection Safety officer will verify use of proper equipment during operations. See section 6.1 Either two team members or a team member and on-staff technician will be present for hazardous machining or layup operations. Inspection Safety officer will verify use of proper equipment during operations. See section 6.1 Black powder charges will be isolated from all components excluding parachutes and shock cord. Inspection Launch vehicle will be inspected for seal breakage or gaps in avionics components or ejection charges. See Launch Procedures All machining operations will be done with a dust collection system active Inspection One team member will stand by with an industrial vacuum to collect all dust and shavings made from machining operations. Section 6.1 Team members will wear face masks when machining carbon fiber Inspection Safety officer will verify use of proper equipment during operations. See section 6.1 Fresh batteries will be used for each launch performed. Inspection Fresh batteries will be purchased ahead of time and verified during pre-launch status sound off from the altimeter See section 7 Launch controller will be held instead of pressed during motor ignition RSO will verify launch button is being held by team leader during launch. Complete. Payload system must be capable of reusability. Testing Shake tests and drop tests will be conducted to ensure the FMPA has the durability to be reusable. Recoverability is dependent on projected flight telemetry. FMPA constructed and tested. Confirmed to be reusable, reloadable, and resilient to stresses. FMPA must easily house a fragile cargo within the inner capsule without damaging the cargo. Inspection Final design configuration will allow for a user to remove the screw top from the inner capsule and delicately place the fragile cargo within. Inner capsule and FMPA design integrate so that cargo insertion is delicate and easy. FMPA piston system can maneuver the inner Inspection Good manufacturing and precision machining will ensure the pistons can move freely as one system without binding up. FMPA constructed and smoothly operational. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 130

131 Research and Development capsule via the shock absorber system. Optimum FMPA pistons will be used to maximize shock dissipation to as close to 100% Analysis Testing Drop tests will be performed to see if equivalent force of impact to motor thrust is dissipated without damage to the fragile material. Pistons can be added or removed to vary test parameters. Optimum shock absorber number and size verified to protect material. Section FMPA can protect fragile cargo from all other stresses incurred throughout duration of flight. Testing Drop tests and parachute deployment tests will help discover the stress mitigations capabilities of the FMPA. Drop tests completed. Stress verification verified. Section FMPA can be easily accessed after flight to retrieve fragile material. After drop tests are performed, the ease of access to the fragile cargo will be accessed to determine if material strengths can maintain assembly configuration. Inner capsule and FMPA design integrate so that cargo extraction is delicate and easy. 8.4 Budget The Rocket Rebels have created a comprehensive budget to facilitate this year s expenditures. The following table shows the complete and exact costs of all expenditures for the Rocket Rebels throughout the project that were not donated material from its sponsors. A pie chart is provided to show the allocation of funds to each category of the project and a clustered bar graph is available to show the change in projected budget costs per section from PDR, CDR, and the finalized budget costs submitted here. Table 33: Final Detailed Expenditures Item Vendor Part Cost Quantity Total Cost ATHENA3 Rocket Kit LOC Precision $ $87.95 GRADUATOR Rocket Kit LOC Precision $ $69.25 Chute Release Chris Rocket Supplies $ $ G80-7T Single Use Rocket Motor Chris Rocket Supplies $ $22.39 G77-4R Single Use Rocket Motor Chris Rocket Supplies $ $22.39 Subtotal $ UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 131

132 Payload Project Vehicle Subscale Vehicle 1/4" x 1" U-Bolts Fastenal $ $ BC 1/4" Zinc Plated Quick Link, Pack Of 5 National Hardware $ $ "x12" NOMEX Blanket Apogee Components $ $ mm Blue Tube Full Length Coupler Apogee Components $ $ mm Blue Tube Apogee Components $ $23.95 MW RRC3 Altimeter Chris Rocket Supplies $ $69.95 Rotary Switches Chris Rocket Supplies $ $9.00 9V Battery Holder Chris Rocket Supplies $ $4.00 Perfectflite StratologgerCF Chris Rocket Supplies $ $54.95 Mounting Screws and Nuts Fastenal $ $20.00 Fiberglass Sheet (G-10/FR4), 1/8" Thick, 24" x 24" McMaster-Carr $ $46.68 Cesaroni 54-3 Grain Case Chris Rocket Supplies $ $64.85 Cesaroni 54mm Rear Closure Chris Rocket Supplies $ $39.95 Cesaroni J295-16A Classic Reload Kit Chris Rocket Supplies $ $85.95 Motor Retainer Assembly (54 mm) Aero Pack $ $38.00 Airfoiled Rail Button, 10 series, pair Apogee Components $ $7.00 Subtotal $ Airfoiled Rail Button, 15 series, pair Apogee Components $ $ /4" U-bolts and Lock Nuts Fastenal $ $20.00 Sunward 18" Nomex Red/Orange Parachute Protector Apogee Components $ $20.98 Fiberglass Sheet (G-10/FR4), 1/8" Thick, 24" x 24" McMaster-Carr $ $46.68 Motor Retainer Assembly (76 mm) Aero Pack $ $ " x 48" Blue Tube Apogee Components $ $ " x 48" Blue Tube Apogee Components $ $ /8" - 16 Wing Nuts Home Depot $ $1.18 1/2" PVC Male Adapters Home Depot $ $1.92 1/2" PVC Caps Home Depot $ $3.44 Glossy Sky Blue Adhesive Vinyl (pack of 6 12"x24" sheets) Vinyl Ease $ $ /2" x 3/8" U-bolts with Collar Home Depot $ $10.32 Loki Motor Kit Chris Rocket Supplies $ $ First Fire Igniter Chris Rocket Supplies $ $25.98 TK102 Mini Spy GPS Tracker Silicon Electronic $ $19.30 Rotary Switches MissleWorks $ $9.00 Loki L840 Cocktail Chris Rocket Supplies $ $ Subtotal $1, Time Capsule Airtight Container Skyway Products $ $14.95 Losi Front Shocks with Springs (pair) AMain Performance Hobbies $ $79.95 Subtotal $94.90 UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 132

133 Outreach/Travel Facilities and Equipment 96" X 1.00 X 1.00 T-Slotted Profile Rail - Four Open T-Slots 80/20 Inc. $ $ " x 1.5" x 1.5" T-Slotted Profile Rail - Four Open T-Slots 80/20 Inc. $ $41.73 Go Box Launch Controller Apogee Components $ $ x 3/4-Inch - Nylon Machine Screw, 40-Pack National Hardware $ $8.00 Firewire Electric Matches Chris Rocket Supplies $ $ M Fire Block Sealer (10.1 oz) Home Depot $ $5.97 Data Transfer Cable Chris Rocket Supplies $ $ fl. oz. Autobody Icing Polyester Finishing Putty U.S. Chemical $ $ mm Motor Cleaning Brush Chris Rocket Supplies $ $ m, 20 gauge phone cable Sportsmans' Guide $ $29.99 West Systems 105 Epoxy Resin, 1 Gallon West Systems $ $80.46 West Systems 206 Slow Epoxy Hardener, 1 Quart West Systems $ $45.81 Resin Pump Kit West Systems $ $19.93 Epoxy Mixing Cups and Sticks Wal-Mart $ $ " x 8ft OD Aluminum Tubing (0.25" thickness) OnlineMetals.com $ $ " x 12ft OD Aluminum Tubing (0.25" thickness) OnlineMetals.com $ $ " x 0.12" Steel Tube A513 (17ft) Tigrett Steel $ $ " x 0.12" Steel Tube A513 (17ft) Tigrett Steel $ $95.20 Stainless T-304 Annealed 2B Sheet/0.06" (16 ga.) 12" x 12" OnlineMetals.com $ $ /8" Pivot Bolt Fastenal $ $5.00 1/2" Pull Pin Fastenal $ $4.99 Subtotal $ Slush Fund for Part Shipping Costs --- $ $ Poster Board Display Wal-Mart $ $24.99 Competition Lodging Costs Embassy Suites $1, $1, Subtotal $2, Total $5, UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 133

134 Figure 68: Project Expenditure Pie Chart Figure 69: PDR, CDR, and FRR Budget Comparison 8.5 Funding Plan The Rocket Rebels have been given a great opportunity to work with local businesses and be sponsored by them. By displaying each sponsors logo on the vehicle, the team has been granted free building materials, monetary donations, or equipment time at company facilities. The Rocket Rebels also can take advantage of any materials and equipment that the CME already owns and that the team can use. The remaining expenditures and equipment needs for this project are covered by the Center for Manufacturing Excellence. Though most of the high dollar materials are donated, there are still several things that were paid for by the CME. UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 134

135 The donated and pre-owned materials and their estimated market values are included in a separate table of the budget shown in the previous section. These include carbon fiber prepreg rolls and all layup materials needed, Henkel Loctite EA-9394 epoxy, ABA polymer filament spools, and drogue and main parachute recovery packages, along with other low value pre-owned or pre-purchased equipment or materials. These materials were used for both subscale and full scale vehicles. Sponsors include Orbital ATK, GE Aviation, Chem-Trend, and Graybeal Recovery Systems. GE Aviation also donated cure time with their in-facility autoclave in the case that the CME s curing oven was not sufficient for curing. Additional sponsors since PDR include Chem-Trend, which has donated high grade mold release and sealant for fiberglass and carbon fiber layups. These materials ensure that the cured material can easily be removed from the mandrel after curing. Other companies are still being reached out to for potential sponsorships to further reduce budget costs for the CME through monetary or material donations and to give local businesses the opportunity to market their brand. The following table shows the donated material given by these companies and the projected market value. Pre-Purchased/Donated Material Table 34: Donated Material List Item Company Estimated Cost Hexply 8552, 5 Harness Plain Weave Prepreg Orbital ATK $ H Plain Weave Carbon Fiber Plate Stock Orbital ATK $ ABA Plastic Filament CME $80.00 Subscale Drogue and Main Parachute Package BAMA Recovery Systems $ Hexply 8552, 5 Harness Plain Weave Prepreg Orbital ATK $ H Plain Weave Carbon Fiber Plate Stock Orbital ATK $ ABA Plastic Filament CME $ Full Scale Drogue and Main Parachute Package BAMA Recovery Systems $ gal Henkel Loctite EA9394 Adhesive Orbital ATK $ Armalon/Peel Ply GE Aviation $75.00 Breather Cloth CME $50.00 FEP Material GE Aviation $75.00 Vacuum Bagging Material GE Aviation $80.00 Zyvax 1034W Mold Release Agent Chem-Trend $ lb Red Powder Coat Paint CME $ lb White Powder Coat Paint CME $45.00 Chem-Trend SealProof Mold Sealant Chem-Trend $75.00 UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 135

136 8.6 Timeline Figure 70: Managerial Timeline Figure 71: Subscale Construction Timeline Figure 72: Full Scale Construction Timeline UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 136

137 Figure 73: FMPA Construction and Testing Timeline UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 137

138 Appendix Appendix A: MSDS Samples UNIVERSITY OF MISSISSIPPI USLI FLIGHT READINESS REVIEW 138

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