University of Notre Dame

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1 University of Notre Dame Notre Dame Rocketry Team Critical Design Review NASA Student Launch Competition Roll Control and Fragile Object Protection Payloads Submitted January 13, Fitzpatrick Hall of Engineering Notre Dame, IN 46556

2 Table of Contents 1 Summary of CDR report Team Summary Launch Vehicle Summary Payload Summary Roll Control Payload Fragile Material Protection Payload Changes made since PDR Vehicle Criteria Changes Payload Criteria Changes Roll Control Payload Fragile Material Protection Payload Project Plan Changes Vehicle Criteria Design and Verification of Launch Vehicle Mission Statement Mission Success Criteria Mission Requirements and Verification Vehicle Description Overview of Vehicle Design Component Design Review Nose Cone Airframe Fins Couplers Subsystem Design Review Roll Control Payload Integration Fragile Object Protection Payload Integration Recovery Subsystem Integration Parachutes Harnesses Bulkheads Attachment Hardware CRAM Mount Propulsion Motor Mount and Retention Fin Integration and Placement Ballast Integration Integrity of Design Fin Shape and Style Materials Motor Mounting Mass of Launch Vehicle Construction and Assembly Verification of Vehicle Design Risks and Mitigations Subscale Flight Results

3 3.3 Recovery Subsystem Hardware Components Parachute Harnesses Bulkheads Attachment Hardware Electrical Components Altimeters CRAM Electrical Components Mission Performance Predictions Validity of Analysis Performance Prediction Program Wind Tunnel Tests Apogee Approximations Stability Kinetic Energy Wind Drift Launch Procedures Vehicle Test Plan Safety Checklist of Final Assembly and Launch Procedures Personnel Hazard Analysis Failure Modes and Effects Analysis (FMEA) Environmental Concerns Project Risks Payloads Roll Control Payload Design and Testing of Payload Equipment System Level Design Review Roll Fins Servomotor Gearbox Electronic Control System Roll Control Algorithm Power Control System Structural Support Measures Ground Station Testing Manufacturing and Assembly Integration Plan External Structure Internal Structure Instrumentation Precision and Repeatability Payload Electronics Electronic Components Arduino Roll Calculations Power Management Methods Payload Safety and Failure Analysis Roll Controller Failure Modes and Mitigation

4 5.1.2 Payload Concept Features and Definition Creativity and Originality Significance Challenge Science Value Objectives Success Criteria Fragile Object Protection Payload Project Plan Testing and Requirement Verification Plans Budget Funding Plan Timeline Appendix A. CAD Drawings Appendix B. Performance Prediction Programs Appendix C. Stability Prediction Program Appendix D. Wind Tunnel Testing MATLAB Codes Appendix E. Launch Procedure Checklists Appendix F. Personnel Hazard Analysis Appendix G. Failure Modes and Effects Analysis (FMEA) Table Appendix H. Environmental Effects on Rocket Appendix I. Safety Concerns for Environment Appendix J. Project Risks Appendix K. Team Timeline Appendix L. Budget Table of Figures Figure 3.1. Vehicle Sections and Subsections Figure 3.2. Final Launch Vehicle Design Figure 3.3. Launch Vehicle-Exploded Model Figure 3.4. Sketch of the Nose Cone Figure 3.5. Picture of the Nose Cone Figure 3.6. CAD Model of the Nose Cone and its Integration Figure 3.7. Airframe Dimensions Figure 3.8. Final Fin Design Figure 3.9. Fin Alignment Mechanism

5 Figure Roll Control Payload Exterior Figure Roll Control Payload Integration Figure Overall Roll Control Payload Dimensions Figure CAD Model of the FOP and its Integration Figure View of the Upper Half of the Vehicle with the FOP Figure Schematic of Recovery System Figure Thrust Curve for Loki L Figure Thrust Curve for AeroTech L Figure Thrust Curve for CTI 3300-L3200-VM Figure Motor Mount Component and Integration Figure Previous Motor Retention Designs Figure Motor Retainer Installation Figure Fin Alignment Mechanism Figure Fin Placement on Fin Can Figure CAD Model of Ballast Container Figure Drawing of Final Fin Design Figure Diagram Showing Both Subscale Designs Figure Dimensions and Components of Launch Vehicle with Roll Induction Fins Figure Dimensions and Components of Launch Vehicle without Roll Induction Fins Figure Dimensions of Main Subscale Fins Figure Recovery System Layout Figure Recovery System PCB Schematic Figure CRAM PCB Layout Figure Subscale Model Wind Tunnel Setup Figure Drag Transducer Calibration Figure CD as a function of Reynolds Number (Without RC Fins) Figure CD as a function of Reynolds Number (RC Fins at 45 ) Figure CD as a function of Reynolds Number (RC Fins at 0 ) Figure Flight Profile of Launch Vehicle with 10 mph Winds Figure OpenRocket Stability Figure RockSim Stability Figure Vertical Descent Speed Calciulations Figure Wind Drift Calculations

6 Figure 4.1. Risk Assesment Matrix Figure 4.2. Failure Mode Classification Figure 5.1. CAD Rendering of Roll Fin Figure 5.2. CAD Rendering of Bevel Gearbox Figure 5.3. Servomotor Position Control Diagram Figure 5.4. Servomotor Input Diagram Figure 5.5. Servomotor Example Inputs Figure 5.6. Payload Schematic Figure 5.7. Payload Board Figure 5.8. Ground Station Schematic Figure 5.9. Ground Station Board Figure CAD Rendering of Ground Station Figure Graph of Altitude and Rotation Angle as Functions of Time Table of Tables Table 1.1. Launch Vehicle Summary... 7 Table 3.1. Vehicle System Level Verification Progress Table 3.2. Launch Vehicle Dimensions Table 3.3. Description of Launch Vehicle Sections and Sub-sections Table 3.4. Summary of Fin Dimensions Table 3.5. Description of Recovery System Components Table 3.6. Motor Characteristics for Loki L930, AeroTech L1150, and CTI 3300-L3200-VM Table 3.7. Launch Simulations Table 3.8. Motor Mount Dimensions Table 3.9. Summary of Material Composition Table Weight of Various Rocket Components Table Mass Verification Plan Table Vehicle Design Verification Table Launch Risks and Mitigations Table Motor Characteristics for AeroTech F Table Launch Day Conditions in Three Oaks, MI (12/10/16) Table Characteristics of Subscale Vehicle with Roll Induction Fins Offset at 45 (OpenRocket) Table Characteristics of Subscale Vehicle with Roll Induction Fins Offset at 45 (RockSim)

7 Table Characteristics of Subscale Vehicle without Roll Induction Fins (OpenRocket) Table Characteristics of Subscale Vehicle without Roll Induction Fins (RockSim) Table Launch Results Compared to Simulation Predictions (No RC Fin Configurations) Table Launch Results Compared to Simulation Predictions (With RC Fin Configuration) Table CD for Launch with no RC Fins Table CD for Launch with RC Fins Table Recovery System Components Table Featherweight Raven 3 Altimeter Specifications Table Deployment Settings for Primary and Backup Altimeters Table OpenRocket Simulation at 5 mph Winds Table OpenRocket Simulation at 10 mph Winds Table OpenRocket Simulation at 15 mph Winds Table OpenRocket Simulation at 20 mph Winds Table RockSim Simulation at Low Winds (0-2 mph) Table RockSim Simulation at Medium Winds (3-7 mph) Table RockSim Simulations at High Winds (8-14 mph) Table Locations of CG and CP, and Stability Margins for Both Simulations Table Kinetic Energy at Key Launch Phases Table Vehicle Test Plan Table 5.1. Roll Controller Failure Safety Table

8 1 Summary of CDR report 1.1 Team Summary Scoring Experiment Option: Option 2: Roll induction and counter roll Additional Experiment Option: Option 3: Fragile material protection Team Name: Notre Dame Rocketry Team 365 Fitzpatrick Hall of Engineering Notre Dame, IN Faculty Advisor: Dr. Aleksandar Jemcov, Professor Department of Aerospace and Mechanical Engineering Team Leader: Jonathan Spraul Safety & Reliability Officer: George Porter NAR Mentor: Dave Brunsting, NAR/TAR Level 2 dacsmema@gmail.com NAR/TRA Section: TRA #12369, Michiana Rocketry 1.2 Launch Vehicle Summary Table 1.1. Launch Vehicle Summary Size Motor Recovery System Launch Rail Length: in Diameter: 5.5 in Fin Span: 7.2 in Loaded Weight: 552 oz Weight without Motor: 428 oz Loki Research L1482 Peak Thrust: lbf Average Thrust: lbf Total Impulse: lbf*s The recovery system is a dual-deploy operation with a drogue parachute released at apogee. The main parachute will deploy at 500 ft. At main deployment, the rocket will separate into three tethered sections. Deployment takes place via black powder ejection charges controlled by redundant altimeters. The GPS module transmits its location for ease of recovery. 10 ft. long, 1.5 in. wide 7

9 1.3 Payload Summary Roll Control Payload The Roll Control Payload is an experimental payload intended to fulfill the specific requirements posed by NASA to induce two full revolutions about the main axis of the rocket, and then to cease rolling motion and mitigate any further roll brought about during the rocket s ascent. These requirements will be met through the use of roll control fins which will be canted by a servomotor in order to utilize aerodynamic lift forces to create a moment about the main rocket axis. The rotation of the rocket will be controlled by an Arduino running a sensor-feedback algorithm. The entire payload will be controlled by an extensive electronic control system and can be monitored and further controlled through the use of a ground station Fragile Material Protection Payload The primary goal of the fragile object payload is to contain and protect a fragile object provided on launch day. The design for the payload includes a dual cylinder system; the fragile object will be placed in a viscous fluid within an inner cylinder made of aluminum. This cylinder will be insulated by foam and contained in a larger cylinder (made of phenolic tubing). The outer cylinder will be the coupler connecting the first and second body tubes. Both cylinders will have screw tops, which will allow for easy installation and removal of the fragile object. The viscous fluid will prevent the fragile object from hitting the sides of the inner cylinder, while the foam padding surrounding the inner cylinder will absorb most of the shock experienced during launch, flight, and landing. The payload will be integrated into the rocket by a series of blocks and screws. Four blocks will be placed 90 degrees from one another inside the coupler and approximately 1 inch below the coupler s screw cap, and a bolt will be screwed into the rocket, through the body tube and coupler, into each block. This will allow the forces experienced during parachute deployment to be evenly distributed, thus preventing the payload from being pulled out of its bay during flight. 2 Changes made since PDR 2.1 Vehicle Criteria Changes Since the presentation of the Preliminary Design Review, the Vehicle Sub-team of the Notre Dame Rocketry Team has made several changes to its design. The coupler material was changed to phenolic from fiberglass to ease the process of acquiring the material. The couplers are not structurally significant; therefore, the fiberglass was considered unnecessary. The thickness of the launch vehicle was changed from 0.05 to Again, this was from a logistics standpoint. The manufacturer of the body tube was unable to reach an inner-diameter of 5.4 inches. The team could not spend its time sanding the inside of the body tube. This increased the weight slightly. In addition, a communication error had been made and the Fragile Object Protection Payload had more weight than expected. Epoxy and paint consideration were made. The weight thus increased to 600 ounces from 543. A new motor was thus chosen. Lastly, FEM analysis was downgraded on the team s priorities. Load analysis by calculations on critical areas like the motor mount produced high margins of error. Outsourcing much of the work such as body tube construction eliminated many sources of error that FEM analysis was no longer considered crucial. It has been started, and will be finished if time permits. 8

10 For the Recovery system, there have been two primary changes made to the recovery system. The first change is a reduction in drogue size. The drogue will now measure 24 inches in diameter in order to minimize wind drift during descent. The second change is an edit to the altimeter programming. As recommended by the NASA team during the PDR presentation, the altimeters will no longer include a velocity constraint for deployment. This is to ensure that parachutes deploy even in the case of a rogue launch. 2.2 Payload Criteria Changes Roll Control Payload The only major conceptual change for the Roll Control Payload is the way in which the aerodynamic forces acting on the payload are utilized. Initially, the plan was to rotate the fins to an extreme angle of 45 degrees and utilize the pressure forces on the fins to rotate the rocket. However, simulations should an increase in drag that was not manageable, as well as concern that flow separation over the roll fins would cause serious instability over the main fins. For this reason, the rotation angle was reduced to 4 degrees, and instead, flat-plate lift forces are being used to rotate the rocket instead. This will minimize flow separation and will have a much lower drag component. Three design changes have been made since PDR. The first is that the roll fins will now be printed in Carbon fiber instead of manufactured using steel sheet and rods and welding them together. This is mainly to reduce weight while still ensuring that the fins are strong enough to withstand the forces. Second, the microcontroller was switched from an Arduino Uno to an Arduino Feather. This change has two major advantages. First, the Feather is smaller which saves space and weight inside the payload, and the second is that it has radio transmission capabilities, which allows for the third design change: an implementation of a ground station. This station is being used to successfully monitor the payload on the pad, get real-time GPS and altitude data, as well as ensuring that all aspects of the payload are functional before launch Fragile Material Protection Payload A primary design modification from the PDR is that there will no longer be springs connecting the curved edges of the inner cylinder to the outer cylinder. The main problem with using circumference springs is that there is limited space separating the exterior of the inner container from the interior of the outer container (coupler), proving it difficult to adhere the springs to both surfaces while also allowing for sufficient displacement of the springs to mitigate movement of the inner container. The difficulty of this assembly process would be compounded with the fact that the springs would be required to connect to a curved surface on both ends. In order to avoid the risk of attaching the springs through a difficult, delicate process, foam will replace the springs in order to provide an effective cushion that is less likely to suffer from errors in construction. Therefore, a high density foam tube will be added on the sides of the inner container in place of springs to cushion it against the walls of the phenolic coupler. The final design will now feature springs on only the top and bottom surfaces of the inner container attached to bulkheads above and below. The springs on the top and bottom have been split up so that the forces will be divided into two smaller springs on the top and the bottom. The sets of springs will be oriented perpendicular to each other to keep the axis of the inner canister parallel to the axis of the main fuselage as much as possible. The use of two springs at the top and two at the bottom of the inner container halves the previously estimated spring constant of 40 psi to 20 psi per spring. Having more springs lends to improve the structural stability of the inner container by distributing the force of the springs on the top and bottom surfaces. An additional result benefit of splitting up the springs is that it allows for the center of the bottom bulkhead to be clear of the springs. This is important for the interfacing of the Fragile Objects Payload with the rest of the rocket. The center area of the bottom bulkhead needs to be clear so that the eyebolt that the shock cord runs through can be adequately bolted through the bulkhead. In the coming weeks, time will be dedicated to further 9

11 modeling the dynamic response to avoid spring responses that coincide with the natural frequencies as well as conducting physical tests on models of the payload to decide on a final spring constant. 2.3 Project Plan Changes The only significant change in the project plan was moving the subscale launch date from November 20 th to December 10 th due to bad weather. 3 Vehicle Criteria 3.1 Design and Verification of Launch Vehicle Mission Statement The mission is to successfully build, test and fly a reusable rocket carrying 2 payloads to an altitude of 5,280 feet. At apogee, the vehicle will separate into two tethered sections and the drogue parachute will deploy. The upper section will separate into two additional sections at an altitude of 600 feet when the main parachute deploys. The vehicle shall carry two payloads. The first shall be a Fragile Object Protection (FOP) Payload. The second shall be a Roll Control Payload (RCP). The RCP will induce a roll during the coasting phase of the flight path that will last for a few seconds and return the rocket to non-rolling status. The launch vehicle shall then be recovered and prepared for re-launch Mission Success Criteria Several conditions must be met for the mission to be considered a success. The high level design of the launch vehicle was designed with these in mind and are always referenced with every change in design. Verifications methods are based on these. The full-scale launches will be determined as successful based on these criteria. The dominant criteria for a successful mission are: 1. Altitude: The vehicle must reach an apogee of as close to 5280 feet as possible. Success on this criterion will be determined based on readings from an altimeter onboard the rocket. A desirable altitude range is 5280 ± 100ft, or ft. 2. Stability: The rocket must maintain an acceptable degree of stability for the duration of its flight. Stability is determined theoretically with OpenRocket and RockSim models. An off-rail stability coefficient of 2.0 is considered acceptable. 3. Structural Integrity: The vehicle must remain intact for the duration of its flight. Each component of the rocket from the nose cone to the fin can to internal bulkheads must survive the flight without compromise so that the rocket can be re-used if necessary. 4. Recovery: The vehicle must be re-usable upon recovery without requiring repairs. Success in recoverability is predicted by the kinetic energy of each section upon landing, based on simulation data. Recoverability of the rocket will be determined based on the condition of each component after the rocket lands. 5. Roll Control Payload: The vehicle must successfully perform two induced revolutions about its roll axis and reduce its rotational motion. Success of the roll control payload will be determined based on data from an accelerometer located within the payload. With respect to the launch vehicle, the payload shall remain undamaged and will not damage its immediate surroundings for it to be considered successful. 6. Fragile Object Payload: The fragile object payload must successfully protect the fragile object during launch, flight, and landing. Success of the fragile object payload will be determined based on whether 10

12 the object is intact upon vehicle recovery. With respect to the launch vehicle, the payload shall remain undamaged and will not damage any of its surroundings in order to be considered successful Mission Requirements and Verification NDRT has put together design requirements that have guided the launch vehicle s design and construction. These are the overarching goals that NDRT strove to meet as the launch vehicle was designed. These verifications attempted to meet the mission success criteria of the overall launch vehicle. Table 3.1 lists these goals as well as the method by which they will be met. They were listed in Section of the Notre Dame Rocketry Team Preliminary Design Report, and are repeated here with the progress. Table 3.1. Vehicle System Level Verification Progress Requirement Requirement will be met by: Method of Verification Progress Launch Vehicle shall rise to an apogee of 5280 feet AGL. - Choosing a motor that provides enough thrust to overcome the mass of the vehicle. - Ensuring smooth surface to minimize drag. - Subscale Test will test the accuracy of the performance prediction software. Subscale launched on Dec 3 rd. Results used to more accurately utilize results from OpenRocket and RockSim Full Scale Tests will verify the effect of the roll control payload to the apogee Construction to be started in Jan Wind Tunnel Tests will test the accuracy of simulated drag coefficients Wind Tunnel test results are described in Section Launch vehicle and payloads thereof must have the capability to be launched again the same day without needing repairs or modifications - Ensuring that every sub-system undergoes extensive individual test. - Ensuring that all subsystems fit together with minimal interference - Designing the launch vehicle to withstand more than one flight in Extensive tests for each sub-system as elaborated in Section An extensive launch vehicle checklist shall be followed before every Full Scale Test and on the day of In Progress according to Section Launch Vehicle Checklist developed Launch Vehicle 11

13 one day competition Checklist to be used during test The rocket will be evaluated after each Full Scale Test to determine that it is able to launch again Full Scale Test scheduled for Feb 2017 Launch vehicle will have commercially available solid motor using APCP - Choosing a motor that fulfills the NAR Safety Code. - Ensuring that the team stays in contact with its mentor regarding the motor choices as design changes Inspection Team s mentor shall handle motors during Full Scale Tests and on the day of competition A motor has been chosen and ordered Motors have been ordered by the team, but handled by Mentor Dave Brunstig Launch Vehicle will have a stability margin of at least Placing the Roll Control Payload above the CP in a way that allows a stability margin of 2.5 and lets the forces on the fins move the CP up to avoid overstability when the propellant has burned - Ensuring that necessary ballasts are spread across the launch vehicle so as not to affect the stability margin OpenRocket and RockSim statistics from simulations will show the changes in CP and CG across flight Subscale and Full Scale Tests will verify if the stability of Launch Vehicle is sufficient OpenRocket and RockSim show that the stability shall remain above 2. Subscale tests with stabilities of 2.43 and 2.68 proved themselves very stable Full Scale test scheduled for Feb Launch Vehicle will have at least 52 fps for velocity at rail exit - Choosing a motor that provides an impulse that achieves this velocity. - Ensuring that the launch vehicle s interaction with OpenRocket and RockSim simulations will verify that this velocity is achieved. OpenRocket and RockSim predict a velocity above 55 fps 12

14 Launch Vehicle performance shall be verified by Wind Tunnel Testing Payload affecting flight shall be verified before launch at competitions Launch Vehicle s general design and construction plan shall be established by PDR Launch Vehicle s subsystems shall have finished the design phase; Subscale will have been launched by CDR KEY: Not started; In Progress; Finished the launch pad is uninterrupted and smooth by carefully calculating interruption angles for rail buttons. - By relating Reynolds number for fluid scaling and relating Reynolds number to the coefficient of drag of both subscale and Full-scale - Confirming the effect of Roll Control Fins on flight path compared to lack thereof - Confirming the structural strength of Payload Integration - Confirming the efficacy of Payload in a practical sense - Working with different Payload and Recovery subsystems in order to gain understanding on what is possible and what is not Ensuring each subsystem fits in with the overall system and can be edited at short notice. - Designing and launching a subscale that will verify the accuracy of our software prediction for confidence. Evaluation following Full Scale Tests shall affirm this velocity in practice before the team launches at competition The acquired coefficient of drag from wind tunnel testing shall be compared against all the performance predictions software and programs as well as Subscale Test results. Subscale Flight Full Scale Flights The overall vision of the launch vehicle has been established. The PDR has been finished. Different options considered and described in PDR have been narrowed down to create one robust launch vehicle with robust sub-systems Subscale vehicle to verify the software prediction and aid the completion of theoretical design Full Scale test scheduled for Feb 2017 Wind tunnel tests were performed on Dec 16 and are described in Section Subscale flight proved that the presence of extra fins decreases apogee Full Scale Flight scheduled for Feb 2017 PDR turned in on the 4 th of November The theoretical designs have been finished; drawings for every sub-system are in Appendix X Subscale launched Dec 3 rd. 13

15 3.1.4 System Level Design Review System level CAD drawings can be found in Appendix A Vehicle Description The launch vehicle, designed by the Notre Dame Rocketry Team, and described here-in, will be capable of reaching an altitude of feet and will carry two payloads. The launch vehicle will be recoverable and will have the capability to be launched again on the same day as the first launch. The length and weight of the rocket are two of the most important considerations when evaluating the stability and projected apogee of a rocket. With that in mind, the total length and weight of the rocket were reduced compared to last year. This will increase both the apogee and the stability, allowing for greater flexibility in dealing with unforeseen problems or changes to the mass. The rocket, as listed, is broken down into three main sections. Section I consists of the nose cone and the payload bay. This Fragile Object Protection Payload bay is housed in a coupler. The section is detachable from section II, which is comprised of the main and the drogue parachute and the CRAM (Compact Removable Avionics Module), which controls the ejections of the parachute. Section III is made up of Roll Induction Payload and the motor mount. This section is the fin can. The section also houses the electronics for the payload fins, as well as both the payload fins and the main fins. During flight, section I will break from II and III at apogee and release the drogue parachute. At deployment of the main parachute, section II and III will split. If at any point it becomes necessary to add extra length to the rocket, it will most likely be at the parachute bay, as there is no worry of too much movement, and it will give more room for the parachutes to be packed Overview of Vehicle Design The launch vehicle is will have a diameter of 5.5 inches, and is inches in total length. The rocket will weigh approximately 615 ounces. There are 4 fins with a trapezoidal shape, with a 7 inch tip and root chord and a 7.2 inch height. At launch, the rocket has a stability margin of 2.42 calibers. Table 3.2 and Table 3.3. Description of Launch Vehicle Sections and Sub-sections show these dimensions. The rocket consists of 5 sub-sections, the nose cone, the Fragile Objects Payload (FOP), the parachute bay, the Roll Induction Payload, and the fin can. The nose cone is 13 inches long with a 4 inch shoulder protruding into the body of the rocket, and is made of polypropylene. The second section, the FOP, contains the materials required to protect the fragile object given to the team. The parachute bay contains the Compact Removable Avionics Module (CRAM) and both the main and drogue parachutes. The Roll Induction Payload contains 4 secondary fins and a bevel gear system to turn the secondary fins. The fin can contains the motor and the 4 main fins. These 5 sections are sub-sections of the main 3 sections. These sections are broken down into how they separate during the descent of the rocket. Section I separates from II upon apogee, and the drogue is ejected. At 600 ft AGL, section II separates from III and the main parachute deploys. Section I consists of the nose cone and the FOP, Section II is the parachute bay, and Section III is the Roll Induction payload and the fin can. Figure 3.1 (a) and (b) shows the sections and subsections that the rocket is broken down into. Further, the CAD drawings in Figure 3.2 and Figure 3.3 illustrate the overall vehicle more clearly. 14

16 (a) (b) Figure 3.1. Vehicle Sections and Subsections Table 3.2. Launch Vehicle Dimensions Property Dimension Total Length (in) Diameter (in) 5.5 Number of Fins 4 Tip / Chord Length (in) 7 15

17 Fin Height (in) 7.2 Fin Width (in).125 Weight with Motor (oz) 599 Weight without Motor (oz) 474 Stability Margin with Motors 2.71 Stability Margin without Motors

18 Figure 3.2. Final Launch Vehicle Design 17

19 I Figure 3.3. Launch Vehicle-Exploded Model Table 3.3. Description of Launch Vehicle Sections and Sub-sections Section Sub-Section Label Composed of Description Hollow nose cone, 13 in height and 5.5 in Connected to the fragile materials payload bay Nose Cone A diameter, made of below polypropylene Fragile Materials Payload Bay B 12 of carbon fiber tubing body tube, and 12 inches of phenolic coupler Holds fragile materials protection payload II Parachute Bay C 40 carbon fiber body tube Holds CRAM (Compact Removable Avionics Module), as well as a main and drogue parachute. 18

20 Roll Control Payload Bay D 12 phenolic coupler. Holds second set of fins for controlling roll during flight and electronics to read and compute proper spin III Fin Can and Motor Mount E Carbon Fiber tube and carbon fiber fins. Hold motor and motor mount and carbon fiber fins Component Design Review Nose Cone The various options for the nose cone material were fiberglass, carbon fiber, and polypropylene. Fiberglass was considered for its strong structural properties and easiness to work with. It was not selected because of its high density and thus heavy weight. Carbon fiber was also considered because of its extremely high strength-to-weight ratio. It would have been the ideal choice, since the body tube for the team s rocket is made of carbon fiber. It was not chosen, though, because of its difficulty to manufacture. Also, purchasing a premade carbon fiber nose cone was too expensive for our budget. We decided to purchase a LOC/precision rocketry polypropylene nose cone from Red Arrow Hobbies. Polypropylene was the ideal material because of its low cost, and purchasing a premade nose cone will help to keep our team on time while ensuring reliability and quality. Because the nose cone does not experience heavy loading or stresses during flight, the material for the nose cone does not necessarily have to have as strong structural properties as the material for other parts of the rocket. The mass of the nose cone is 10 oz. Its dimensions are given below in Figure. The bottom of the nose cone will have a 4 shoulder that inserts into the body tube of the rocket. Two bolts spaced directly across from each other on the rocket will connect the body tube and nose cone, holding it in place. For security, the nuts holding the two bolts will be epoxied to the inside of the rocket body tube. Figure 3.5 shows the physical nosecone, and Figure 3.6 is the CAD model for the integrated nose cone, including nuts and bolts. We decided not to be put ballast in the nose cone because of the difficulty of adhering regular epoxy to polypropylene. 19

21 Figure 3.4. Sketch of the Nose Cone Figure 3.5. Picture of the Nose Cone 20

22 (a) (b) Figure 3.6. CAD Model of the Nose Cone and its Integration Airframe In the past, phenolic tubing has been utilized to construct the airframe of the rocket because it is inexpensive, durable, and resistant to water and heat. Additionally, no pre-processing was needed when phenolic tubing was used for 21

23 the body tube and couplers. However, the team has decided to use carbon fiber for the some of the rocket s airframe components this year. The body tubes and fins will be made out of carbon fiber, while the fin can bulkheads, motor mounts, and couplers will still be made out of phenolic tubing. Despite the additional costs and time investment required to process and make the carbon fiber body tubes and fins, carbon fiber is extremely versatile and is expected to provide additional support and protection during launch and landing. Figure 3.3 displays the body tubes that will be made from carbon fiber and how the various carbon fiber components will be attached during airframe construction. Since the team does not have any experience laying and rolling carbon fiber, the body tubes and fins will be ordered (with the slots for the fins pre-cut in the appropriate tubes). The team will perform the rest of the construction on the rocket. The vehicle airframe will be verified through full scale testing. The full scale test will involve constructing the rocket that is to be used during the NASA launch (Figure 3.2). The team will utilize OpenRocket to estimate how the given design components (such as fin shape and size, airframe materials, and locations of the center of pressure and center of mass) affect the apex of the rocket s flight. After the full scale test launch, the team will examine the rocket to see how the carbon fiber airframe supported the rocket during flight. A thorough examination will be conducted to search for severe dents, scratches, and cracks along the rocket s frame. A dent, scratch, or crack is labeled severe if it will significantly affect the outcome of a future flight. A severe marking is one that can add additional drag to the rocket, requires the affected section or component to be replaced and/or rebuilt, or moves the center of pressure or the center of mass to a location that results in the static stability margin to be less than the allowed value. The team will discuss the results of the full scale launch to determine which components need to be improved and if carbon fiber will be safe and stable enough to utilize during the competition. Figure 3.7 shows the dimensions of the airframe used for the launch vehicle Fins Figure 3.7. Airframe Dimensions Several fin configurations were considered for our rocket s main fins: an elliptical fin shape, a square fin shape, a rectangular fin shape, and a parallelogram fin shape. Each fin configuration was considered with either three total fins or four total fins. Plywood, fiberglass, and carbon fiber were considered materials for each fin configuration. Elliptical fins are only effective at low Reynolds numbers when the rocket is deflected because it is only then when the forces acting on the fins are large enough to cause them to be effective in straightening out the flight of the rocket. A parallelogram fin shape was chosen because of its high effectivity at low Reynolds Numbers. Conveniently, 22

24 this fin shape is easy to make and replicate and because all of the fins have the same airfoil shape, there is no drag that would arise from asymmetrical fin shapes. In the past, plywood fins were used because they are easy to construct, inexpensive, and easy to attach to the rocket. This year, carbon fiber fins will be used. Carbon fiber fins are light and strong thereby increasing apogee while sustaining structural robustness. In order to maintain flight in the vertical direction, fins were chosen that maximize stability and minimize drag thereby also maximizing apogee. The fin is shown in Figure 3.8. Furthermore, with this fin configuration, the team can make changes quickly when necessary. The team can decrease or increase the height of the fins for apogee changes after the first Full Scale Test if necessary. That is, decreasing the height of the fins increases the apogee slightly, while increasing it decreases it slightly. The team ensured that the stability is sufficient so that these changes don t under-stabilize the launch vehicle too much. Figure 3.8. Final Fin Design The fins will be ordered as carbon fiber plates. They will be cut appropriately and will have the edges sanded to reflect airfoils in order to ease airfoil and decrease drag. The fins edges will be epoxied lightly so that the edges do not dull. Rocketpoxy will be used to attach the fins to the body tube. Figure Figure 3.9 shows the mechanism shown to align the fins at exactly 90 degrees from each other. There will be two of these plates on the leading edge and trailing edge of the fins and the fins will be epoxied. They will wait overnight to dry. Table 3.4 shows a summary of the fin dimensions. Table 3.4. Summary of Fin Dimensions Characteristic Dimension 23

25 Length (in) 7 Height/Span (in) 7.2 Sweep (degrees) 28.3 Thickness (in) Figure 3.9. Fin Alignment Mechanism Couplers Couplers were chosen to extend the length of the rocket past a single body tube and allow the rocket to be separated into multiple sections both on the ground and at parachute deployment. Multiple ideas were considered for achieving access to the systems inside the rocket, but couplers were chosen for their overall convenience and ease of use. An alternative would be the use of access doors on the rocket. However, the pressure change occurring inside the rocket would cause the access doors to open, unless they were bolted shut, which would add drag to the rocket. Access doors would also only allow access to the inside of the rocket while on the ground, but would not help keep the rocket together during launch. Because the rocket design has a two stage deployment, it makes the most sense to couplers to access the inside of the rocket, and connect the body tubes which will eventually separate at the deployment of each parachute. The couplers used will be made from 0.08 inch thick Kraft Phenolic tubing, instead of the carbon fiber used for the main body of the rocket. Phenolic is durable enough to not bend or break under the stresses of flight and is much easier to work with than carbon fiber. Although this material will be different from carbon fiber, the two materials can 24

26 still be epoxied together when needed. The coupling tubes will also have an outer diameter in line with the body tube inner diameters. This tight fit will help to hold the rocket together when needed and work to support the carbon fiber during flight. The coupling tubes will be purchased commercially at the correct diameter (5.20 inner diameter and 5.36 outer diameter) and size to be sure of the uniformity throughout the tubes. The rocket will have two couplers. One contains the fragile object protection payload, which also connects the payload bay body tube to the main body tube, which contains the CRAM and both parachutes. The other is attached directly to the roll control body tube and connects the main body tube to the fin can. In this second coupler is where the roll control payload will be placed. These two couplers can be seen respectively in Figure 3.10, Figure 3.11, Figure 3.12 and Figure Figure Fragile Object Protection Payload Coupler (External and Section) 25

27 Figure Roll Control Payload Coupler (External and Section) 26

28 Figure Fragile Object Protection Payload Coupler Dimensions 27

29 3.1.6 Subsystem Design Review Roll Control Payload Integration Figure Roll Control Payload Coupler Dimensions The Roll Control (RC) payload will induce rotation of the rocket about its center axis after burnout, according to the payload requirements as outlined in Section 3.1 in this document. It consists of four separate payload sections also called payload bays which house the internal structure of the payload, including the arduino, servo motor, and gearbox. The payload will be secured in a coupler. Four carbon fiber fins spaced radially around the rocket will be used to induce roll and counter roll. The payload bays are stacked on top of each other and placed within a coupler, which attaches the main airframe body tube to the fin can. Because the payload connects these two parts of the rocket, a plywood bulkhead with an eye bolt is placed directly in front of the payload and another singular rear bulkhead behind the payload. The shock cord used in the recovery system will be attached to the eye bolt on the bulkhead. Four threaded steed rods are 28

30 connected to a different bulkhead in the fin can right above the motor mount. They run through the entire payload and both bulkheads. Lock nuts and washers will be screwed onto the rods above the bulkhead with the eye bolt and below the read bulkhead, holding it in place. For extra security, two bolts spaced directly across from each other on the rocket will connect the payload and coupler to the fin can. The nuts holding the two bolts will be epoxied to the inside of the payload. The roll control payload utilizes four additional carbon fiber fins (RC fins). To accommodate this, four aluminum rods from the roll control payload will extend out of four circular holes spaced radially around the payload. Because of the strength of the carbon fiber body tube, the holes in the body tube and coupler are not a structural concern. The RC fins will be attached to the rods using aluminum friction couplers, one fin per rod. The aluminum rods, aluminum friction couplers, and carbon fiber fins are used because of their high strength-to-weight ratio. Figure 3.14 below depicts a CAD view of the Roll Control Payload exterior, and Figure 3.15 shows how the payload will be integrated and secured into the rocket. Figure 3.16 gives the overall dimensions of the payload. 29

31 Figure Roll Control Payload Exterior Figure Roll Control Payload Integration 30

32 Fragile Object Protection Payload Integration Figure Overall Roll Control Payload Dimensions The fragile object protection payload will be comprised of a concentric cylinder containment system and will be located in the first section of the rocket, directly beneath the nose cone. The outer cylinder of the payload will be the coupler that connects the first and second body tube components; this will allow for the fragile object provided on launch day to be easily installed in and removed from the rocket. The inner cylinder of the containment system will be just large enough to protect the largest possible object, and this cylinder will be suspended in the larger cylinder (the coupler) by foam padding that is responsible for reducing most of the impulses imparted to the payload. Figure XXXXX displays a vertical cross section of this dual cylinder payload containment system without the block and screw integration system. In order to maintain the fragile object s position within the inner cylinder, it will be suspended in a viscous fluid that is able to evenly support the mass of the object. The preliminary viscosity that is desired for the fluid surrounding the fragile object is on the 10 3 to 10 4 order of magnitude in units of centipoise, and the team is considering using corn syrup in order to satisfy this requirement. Corn syrup will be able to support the fragile material and prevent it from shifting violently during launch, parachute deployment, and landing. Additionally, corn syrup was chosen as the viscous fluid because it has a slow absorption rate, will not put the rocket overweight, and will not significantly shift the center of gravity and/or aerodynamic center. Foam was chosen to further insulate and protect the inner cylinder because it is lightweight and easier to manufacture. While the rocket s body tubes and most of its couplers will be made out of carbon fiber, the inner cylinder used to contain the fragile object will be made out of aluminum in order to further prevent the absorption of the viscous fluid. The coupler that houses the payload will be made of phenolic tubing. These materials were chosen for the payload cylinders in order to reduce overall cost and ensure a stable casing for the fragile object. 31

33 When loading the fragile object, the concentric cylinder containment system will be completely removed from its bay in the rocket. The viscous fluid will be added to the inner cylinder, the object will be placed in the fluid at the midpoint of the inner cylinder, and the containment system will be loaded back into its appropriate bay. Both cylinders will be capped on one end with a screw-top lid so that installing and removing the fragile object is as efficient as possible. In order to secure the payload in the rocket, four wooden blocks will be epoxied inside the coupler (Figure 3.17). These wooden blocks will be sanded so that one edge follows the curve of the coupler, and they will be secured approximately 1 inch below the lid of the outer cylinder. The blocks will be approximately 1.5 inches long and will extend only 1 inch into the space between the inner and outer cylinders so that they do not interfere with the vibrations the payload will inevitably experience during all stages of flight. The blocks will be placed 90 degrees apart in order to evenly distribute the stresses that will be felt by the payload and nose cone during parachute deployment. A single screw will be used in conjunction with each block to secure the payload to the outer body tube, which means each screw will need to be removed when unloading the payload. Figure 3.17 displays how the blocks and screws will be used to secure the payload within the body of the rocket. Figure CAD Model of the FOP and its Integration Figure View of the Upper Half of the Vehicle with the FOP 32

34 Recovery Subsystem Integration The vehicle s recovery subsystem is designed to guarantee a smooth, safe descent of the vehicle from apogee to landing. It ensures that the vehicle will descend with a low kinetic energy to ensure safety and reusability of the rocket. The recovery subsystem will be located in the center of the center section of the rocket, as seen in Figure This placement allows the rocket components to avoid colliding with each other during descent after both parachutes have been deployed. The even spaces on either side of the CRAM within the section allow the ejection charges to be the same size, as each side separation will require the same amount of pressure to shear the shear pins. This allows for easier preparation of the rocket for launch. The recovery section of the rocket as a whole is the center section of the rocket s three main sections. In this way, it makes up a large portion of the overall structure of the vehicle. The nose and tail sections of the rocket are attached to the recovery section via shear pins that shear when charges are detonated in the CRAM to deploy the vehicle s parachutes. The rocket s recovery system and avionics module will be located in the middle section of the rocket. A schematic of the system can be found in Figure Details of each labeled component can be found in Table 3.5. Figure Schematic of Recovery System Table 3.5. Description of Recovery System Components Component Location in Figure 3.19 Section Main Parachute A Drogue Parachute B Nomex Cloth C Nomex Shock Cord Protectors D Shock Cords E Quick Links F CRAM G Eye-Bolts H Shear Pins I

35 Parachutes Misc. Hardware and Adhesives N/A The rocket will use a two-stage deployment recovery system, with a drogue parachute deploying at apogee, and a main parachute deploying at 600 feet AGL. Both are made of nylon. When packed, the drogue parachute will take up approximately 10 cubic inches and the main parachute will take up approximately 130 cubic inches. They will be packed on opposite sides of the CRAM Harnesses The parachutes are attached to the rocket using 1 inch tubular nylon shock cords. These allow the independent rocket sections to remain tethered to each other after separation, and also absorb the large forces generated by separation and parachute deployment. The shock cords are each 5 times the length of the rocket to ensure that there is enough for this purpose. 12 inch long tubular pieces of Nomex cloth will be wrapped around the portion of each shock cord that is closest to the separation charges to protect the shock cord when the charges detonate. Additionally, a sheet of 18 x 18 Nomex cloth will be wrapped around the end of each parachute that is facing the separation charges to protect these critical components from the explosion Bulkheads The bulkheads are attached using epoxy to Section I and Section II of the launch vehicle and are made of fiberglass. These bulkheads keep the rocket attached even at the recovery events. Extra care will be taken epoxying the bulkheads, as they will be subjected to a high impulse when each parachute is deployed Attachment Hardware The shock cords will be attached to the rocket sections using Quick Links and 6, 5/8 diameter eye bolts. The Quick Links at the ends of the shock cords will be attached to closed-loop eyebolts in each of the sections of the rocket. The parachute cords will be threaded through the quick links which will in turn be attached to the eye bolts. Shear pins will be utilized to connect the three main sections of the rocket during ascent, thus ensuring structural integrity. At each parachute deployment, charges will be detonated from the CRAM in the corresponding section of the rocket. The shear pins will break, thus allowing the desired sections of the rocket to separate CRAM Mount The Compact Removable Avionics Module (CRAM) is the principal component of the recovery payload. It is located in the middle of the center section of the rocket and contains the altimeter, charges, and connects the parachutes to the vehicle. It attaches to the vehicle by screwing into a 3D-printed mount which is in turn epoxied to the rocket body. The mount s outer diameter is the same as the inner diameter to ensure a close fit and is 2.3 inches tall. To ensure that the CRAM does not become unscrewed from its mount during launch, a screw holds the mount and CRAM together. To allow the altimeters to properly measure altitude, small holes will be drilled in the body of the rocket to allow atmospheric pressure readings Propulsion The launch vehicle's ability to reach the target altitude of 5,280 feet AGL will be dependent primarily on motor selection. For the Preliminary Design Review, three possible motors were selected as possible candidates. These were the Loki L930, the AeroTech L1150, and the Cesaroni Technology Inc. (CTI) 3300-L3200-VM. All motors provided satisfactory (greater than ~55 ft/s) off rail velocities and the vehicle never exceeded Mach 1 with either choice. All 34

36 motors projected to achieve an altitude higher than the target altitude. Two other options were considered previously, such as the Loki L1482 and the Loki L840CT. Both of these featured an apogee that was significantly greater than the target altitude, and these options were quickly disregarded. These simulations were performed with a slightly different version of the rocket. Since the PDR, the rocket and simulation have both changed. Table 3.6 shows the specific characteristics of these motors. Table 3.6. Motor Characteristics for Loki L930, AeroTech L1150, and CTI 3300-L3200-VM Motor Classification Loki L930 AeroTech L1150 CTI 3300-L3200-VM Diameter (in) Length (in) Total Weight (lb) Propellant Weight (lb) Average Thrust (lbf) Maximum Thrust (lbf) Total Impulse (lbf*s) Burn Time (s) Thrust to Weight Ratio The Thrust Curves for all motors are shown below in Figure 3.20, Figure 3.21, and Figure 3.22 Figure Thrust Curve for Loki L930 35

37 Figure Thrust Curve for AeroTech L1150 Figure Thrust Curve for CTI 3300-L3200-VM The results of the simulations for these motors in identical conditions are shown in Table 3.7. All simulations were performed using 10 mph winds with a standard deviation of 1 mph. The rail had a length of 10 ft. Table 3.7. Launch Simulations Motor Classification Loki L930 AeroTech L1150 CTI 3300-L3200-VM Apogee (ft) Off Rail Velocity (ft/s) Maximum Velocity (ft/s) Flight Time (s) Time to Apogee (s)

38 Motor Mount and Retention The propulsion system for the launch vehicle will be integrated into the fin can section of the rocket through the motor mount. The motor mount will consist of a smaller motor mount tube that will contain the motor casing for the solid rocket fuel motors. This will then be held securely along the axis of the launch vehicle by three centering rings positioned along the tube. The top of the motor mount will be capped by a bulkhead as well in order to seal the motor casing inside the motor mount tube. The dimensions for these components are given in Table 3.8 below. Table 3.8. Motor Mount Dimensions Component Outer Diameter (in.) Inner Diameter (in.) Length (in.) Motor Mount Tube Centering Ring Bulkhead 5.38 NA 0.25 The motor mount tube will be made out of the same carbon fiber tubing as the body tube of the rocket, while the centering rings and bulkhead will be made out of quarter inch thick fiberglass. This is to reduce cost and weight of the rocket while still providing sufficient structural stability. The bulkhead will be located at the foremost end of the motor mount tube to act as a cap and prevent the motor from traveling up through the rocket during burnout. The centering rings will be positioned along the motor mount tube at 0.75, 13.5, and 20.0 inches aft of the base of the tube. This allows sufficient space to integrate the fins and motor retention system for the rocket. The centering rings provide structural stability to the fin can and also prevent the motor from coming out of alignment with the launch vehicle and creating a gimbaled thrust system during burnout. Gimbaled thrust is dangerous because it creates a torque about the rocket s center of gravity and will greatly affect the rocket s stability during flight. Both the centering rings and bulkheads will be attached to the motor mount and fin can body tube using JB Weld epoxy in order to keep all the components in place during flight. JB Weld is used rather than regular epoxy because it is better equipped to handle the intense heat generated by the motor during burnout. The components were assembled in CAD and are displayed in Figure 3.23 below. Figure Motor Mount Component and Integration The motor will be held securely in the motor mount tube by a positive motor retention system consisting of both a burnout retention system and a descending retention system. Since the motor is generating thrust during burnout there must be a way of keeping it from shooting up through the rocket. However, after burnout there will be no longer be a 37

39 force keeping the motor in the vehicle and there needs to be a system in place to ensure that the motor will not fall out of during its descent or when there is not sufficient thrust to keep it in the launch vehicle. This is much more dangerous because the motor may still be burning as it falls out of the rocket and this is a high risk failure mode. The burnout retention system is composed of the centering rings and a bulkhead that ensure that the motor remains fixed in alignment with the axis of the launch vehicle. The descending retention system will be located at the aft end of the motor mount tube and will ensure that the motor remains in the launch vehicle. The team has been considering four types of motor retention systems to use for the launch vehicle. The first design is what the team has used in past years. It is a set of washers offset by 180 degrees and attached to two screws protruding from the aft most centering ring. A variation of this design was to use two stainless steel clamps instead of washers to secure the motor casing. The other two alternatives consisted of using commercially available quick change motor retainers for 75 mm rocket motors. One version included an aluminum cap that would be bolted into the aft most centering ring of the rocket. The other version of this system consisted of an adapter that could be epoxied to a piece of the motor mount tube protruding from the aft most centering ring. A cap would then be screwed onto the adaptor to hold the motor casing in the rocket. These designs are shown respectively in Figure 3.24 below. Figure Previous Motor Retention Designs Since PDR, the team has considered each of these designs and has determined that the two piece quick change motor retainer assembly is the best option. The motor retention system the team has used in the past was determined to be insufficient for properly securing the motor in the launch vehicle. The leading design at PDR had been to use the bolt on clamps to secure the motor. However, after further analysis the quick change retainers were chosen as the better option. The two-piece quick change motor retainer assembly is manufactured by Aero Pack and is available to order online for approximately $44. They are manufactured out of precision machined 6061-T6 aluminum and are compatible for all of the 75 mm diameter solid rocket motors that the team has been considering for the propulsion system of the rocket. They are integrated into the launch vehicle by epoxying the adaptor piece to a section of the motor mount tube protruding at least 0.75 inches from the aft centering ring. The threads on the adaptor align with the threads on the retainer cap to seal the motor casing in the tube. An image of this installation process was taken from Aero Pack s website and included in Figure 3.25 below. 38

40 Figure Motor Retainer Installation This system much simpler than the alternative quick change retainer, which would involve multiple bolts being driven into the aft centering ring. There is far more uncertainty and room for error in this system since the team has not worked with carbon fiber or fiberglass. It would also be advantageous to avoid complicating the construction of the vehicle unnecessarily. Staying with the two piece retainer and using JB Weld epoxy to secure it to the motor mount tube is, therefore, the most feasible option. A similar system was used in the subscale rocket and the team is confident that it is a viable option for the full scale as well Fin Integration and Placement The four fins have been placed at the rear of the rocket, 6.0 in. forward of the bottom of the fin can, spaced at 90 o intervals. They will be offset 45 o from the roll control fins, so that the disturbance in the airflow from the roll control fins has minimal effect on the airflow over the fins. The purpose of the fins is to move the center of pressure, which is where the lift and drag forces act, aft of the center of gravity, providing stability whenever the rocket wobbles during flight. This restoring force will compensate for any wobble in the rocket due to wind gusts, realigning the orientation of the rocket with the desired vertical flight path. The fins will be constructed by using pre-fabricated carbon fiber plates. First they will be cut into the desired parallelogram shape seen in Figure 3.27, with a base of 7.0 in., a height of 7.2 in., and a sweep angle of 28.3 o. They will be cut using a diamond-edged saw. The thickness of the fins will be in. After they are cut into the proper shape, they will be sanded, to create a symmetrical airfoil. The leading edge will be sanded at 90 o and the trailing edge will be sanded at 45 o. To attach the fins to the rocket, there will be four in. slots cut into the fin can. These will be part of the fin can when it is ordered. The fins will be inserted into these slots and then adhesive will be applied to completely attach them to the body of the rocket. Rocketpoxy will be the adhesive used. To ensure that the fins are indeed perfectly aligned, a fin alignment mechanism has been developed. Figure 3.26., which is a reproduction of a photo from Section , shows part of the alignment mechanism that was used in the construction of the subscale. A similar mechanism, albeit bigger, will be used for the full scale launch vehicle. 39

41 Figure Fin Alignment Mechanism Figure Fin Placement on Fin Can 40

42 Ballast Integration If the stability of the rocket is not within its limits and needs to be changed, it is not usually ideal to change the design of the rocket in order to fulfil this requirement. Therefore, ballast is used in order to change the center of gravity (CG) of the rocket, and therefore the stability. Ballast is usually placed near the nose of the rocket, moving the CG towards it. Since the center of pressure (CP) will not change (the mass is added inside the rocket and does not impact the aerodynamic properties), this increases the distance between the CG and CP, and therefore the stability of the rocket. Ballast will be incorporated into the two inch area between the shoulder of the nose cone and the coupler housing the Fragile Material Protection Payload. Here there will be a phenolic coupler with two plywood bulkheads attached, the one closest to the Payload coupler will be epoxied and immobile, while the one closest to the nose cone will be free to move. Ballast in the form of sand (96 lb/ft 3 ) will be measured to the correct weight and placed in durable plastic bags, these bags will be put in coupler between the two bulkheads. In order to keep the ballast and top bulkhead from moving during flight, three bolts will be evenly spaced and put through drilled holes in both bulkheads, and they will be tightened together using nuts, securing the ballast. To verify that the retention system will not fail during flight, a shake test will be performed once the ballast is in place. A model of the ballast retention coupler can be found in Figure Figure CAD Model of Ballast Container 41

43 3.1.7 Integrity of Design Fin Shape and Style The fins style chosen was described in Section as excellent for maximizing stability and apogee. This style was also chosen for the flexibility is offers the team. Given the results of the full scale test launch, the team will be able to make last amends to perfect the launch vehicle. The integration of the fins is described in Section This integration system was used on the subscale, although the material was ply-wood for the fins and phenolic for the body tubes. The fins performed well. The launch vehicle reached apogees higher than predicted in some instances. The team has high confidence in the fins. Another visual representation of the fin shape and style is shown in Figure This CAD drawing includes the part that extends below into the body tube to the motor mount. Figure Drawing of Final Fin Design Materials Carbon fiber was selected for the body tube and for the fins. This is due to its desirable strength-to-weight ratio. Its first advantage is its strength. Carbon fiber has a high compressive strength and a high shear strength, so it is much less likely to fail than other materials, such as kraft phenolic. Its second advantage is its weight. Because of its high strength, less material is required to maintain structural integrity, meaning that while it is denser than many materials, the small volume of material required ultimately results in less weight. This allows for more weight in the payloads which in turn allows for a more robust design for the fragile material protection system and for the roll control system. 42

44 Kraft phenolic was selected for the couplers because of its easy workability. It isn t necessary for the couplers to be as strong as the body tube itself, so they do not need to be carbon fiber. Phenolic is easier to work with and is still strong making it a good choice for the couplers. Because the couplers will be housing the payloads, they need to be easily adaptable to the needs of the payload and phenolic allows for this. Polypropylene was chosen for the nose cone. This is due to its very high compressive strength and because it is easy to shape. The high compressive strength is an advantage because of the position of the nose cone on the rocket. Because it is the foremost point on the rocket, a large portion of the compressive forces due to the airflow will act on it, so to maintain the integrity of the rocket it must be able to withstand those forces. The ease of use is important because the shape of the nose cone is important for the aerodynamic shape of the rocket. For the bulkheads, G10 fiberglass was chosen. This is because it is strong, but can also be machined using a CNC router. The use of the CNC router is important because it makes the manufacturing of the bulkheads simpler. Strength is important because the bulkheads have been a point of failure in the past. Therefore, a stronger material was chosen to prevent this failure and sustain the structural integrity of the rocket. Table 3.9. Summary of Material Composition Part Nosecone Body Tube Motor Mount Fins Bulkheads Couplers Material Polypropylene Carbon Fiber Carbon Fiber Carbon Fiber G10 Fiberglass Kraft Phenolic Centering Rings G10 Fiberglass Motor Mounting The chosen motor and its casing will be held in the motor mount component of the rocket located in the fin can section of the launch vehicle. As mentioned in Section , this will be accomplished through the use of centering rings and bulkheads to secure the motor mount tube along the axis of the launch vehicle. This prevents the motor from traveling up through the rocket during burnout and prevents the development of a gimbaled thrust system that would result in the loss of stability in the rocket. After burnout, to ensure that all components of the propulsion system remain in the launch vehicle, the team has decided to use an Aero Pack quick change motor retainer assembly as a form of positive motor retention. This assembly consists of two pieces. The first component is an adaptor that will be attached using JB Weld epoxy to 0.75 inches of the carbon fiber motor mount tube protruding from the aft fiberglass centering ring. The second component is a threaded cap that is then placed over the motor and its casing and screwed onto the adaptor to secure the motor in the launch vehicle. 43

45 The use of bulkheads and centering rings to secure the motor mount tube has been used by the team in the past and it has proven to be trustworthy, the only difference this year will be the materials chosen to construct it. So long as the motor mount tube can withstand the forces exerted on it during burnout, it can be trusted to be structurally sufficient throughout the entire launch. To verify this, a simple load analysis was performed on the motor mount tube using the Loki L1482 configuration. This motor was chosen as the test motor because out of all the motors under consideration it created the largest force on the rocket. This force was found by dividing the total impulse by the burn time to get lbf applied to the motor mount tube. The rest of the analysis was done using the cross sectional area of the motor mount tube to calculate stress exerted on the component by the rocket motor. The cross sectional area for the dimensions given in Section was found to be inches squared. The stress was then calculated by dividing the force exerted by the motor by this cross sectional area to obtain a value of psi. Since the tensile strength of the carbon fiber tubing used to construct the tube is given as 500,000 psi a margin of safety for our design could be calculated by dividing the tensile strength by the maximum stress any of the motor could exert. This value came out to be 719.5, and with such a high factor of safety, the team is confident that the chosen material and dimensions of the motor mount tube will be more than sufficient to properly secure the motor in the launch vehicle. The Aero Pack motor retainer, however, is new this year and the team has no experience with this system. A similar system was used for the subscale this year and the team was satisfied with its ease of installation and its ability to prevent the motor from falling out of the rocket. Information online from Aero Pack and other model rocketry companies suggests that this assembly is more than capable of securing the rocket motor in the launch vehicle. However, to further validate the system, another load analysis was conducted on the adaptor piece of the retainer to ensure that the JB Weld epoxy would hold throughout the flight. For this analysis, the Loki L1482 was also used because it was had the heaviest fully loaded motor weight of lbs. The retainer requires there to be at least 0.75 inches of motor mount tubing protruding from the aft centering ring. This meant that the minimum surface area for the epoxy to be applied is inches squared. The stress on the cured epoxy would be the weight of the motor assembly divided by the surface area covered by epoxy, yielding psi. Since the tensile strength of JB Weld epoxy is listed as 3960 psi, a factor of safety for this cure was found to be 3155 by dividing the tensile strength by the maximum predicted stress on the epoxy. These analyses show that the dimensions the team has chosen for the motor mount tube are sufficient to withstand the stresses exerted by the most powerful motor under consideration, and that the location where the motor retention system is connected to the launch vehicle is capable of withstanding the maximum loads that can be applied to it. In the future, more analyses such as these will be conducted on the retainer assembly in order to be certain that the part can withstand all the loads the current design places on the component. Source of Motor Impulse: Mass of Launch Vehicle Because mass is a driving factor for altitude and rocket performance, it is important to have an appropriate mass estimate. Table Weight of Various Rocket Components shows a detailed breakdown of the weight of each component of the vehicle. Table Weight of Various Rocket Components Component Weight (oz) Nose Cone

46 Ballast 0 Fragile Materials Payload Bay Payload Bay Body Tube 12.7 Tube coupler 10.8 Fragile Materials Payload 60 Bulkhead 8.84 Bulkhead 8.85 Parachute Bay Parachute Bay Body Tube 42.2 Main Parachute 27.5 Drogue Parachute 35.2 CRAM 60 Roll Control Mount Roll Control Body Tube 3.17 Coupler 12.9 Bulkhead Top 3.13 Bulkhead Bottom 3.13 Roll Control Equipment 80 Roll Control Fins 13.6 Ballast 0 Fin Can Body tube 28.2 Fins 22 Engine Mount

47 Center Rings (3) 4.03 Motor 124 Bulkhead 6.08 Miscellaneous Weight Total 602 The above estimates are based on manufacturer provided information, as well as informed guesses (miscellaneous weight). In all likelihood, the rocket s weight will not change by more than 20% over the remainder of the design phase. This is in contrast to the normally projected 25% to 33% increase expected because the mass estimate is drawn from prior knowledge and experience where that kind of increase would be highly unlikely. It must be noted that between the two simulations of the rocket (OpenRocket and RockSim) there is a mass difference of 7.76oz (3.83oz in the nose cone and 3.93oz in the main fins), making the RockSim simulation heavier. However, this is being neglected since the calculated Center of Gravity and Stability Margin in RockSim only differ from the OpenRocket simulation by.026 inches and.02, respectively If needed to change the center of gravity and increase stability, ballast will be used in areas of the rocket where it can be accommodated. See section for details. Table Mass Verification Plan Requirements Plan of Verification Method of Verification Progress The center of mass and individual masses are accurate. -Software will be used to estimate center of mass and individual masses prior to construction. -During and after construction, each component will be individually weighed, and the overall rocket mass and center of mass will be measured -OpenRocket and RockSim -Inspection with scale Complete. The final weights include epoxy and paint Not started. Material in transit. Ballast will not move throughout flight. -A shake test will be performed before the ballast is loaded to ensure the durability of the containers. -Another shake test will be -Shake Tests before loading and after loading Not started. Material in transit. 46

48 performed after the ballast is loaded to ensure that the containers are secured properly. -The ballast containers and retention will be inspected after launch to ensure that they did not break or move during flight. -Full Scale Test Not started. Full Scale scheduled for Feb 2017 KEY: Not started; In Progress; Finished Construction and Assembly All necessary components to construct and assemble the launch vehicle will be ordered and delivered to Notre Dame in order to commence the construction phase. On resuming the spring semester on January 17 th, 2017, construction will start immediately. A general construction plan was drafted and is detailed for each of the three sections which form part of the rocket. Further, the assembly and integration of the three sections is also detailed in the construction plan. Section I is comprised of the Nose Cone and Fragile Materials Payload Bay. Section II is comprised of the Parachute Bay Finally, Section III is comprised of the Roll Control Payload Bay and the Fin Can and Motor Mount. The plan is divided into the three pertinent sections and is further detailed for the major components of each section. This plan may be amended as construction continues and should only serve as a general guide. Also a brief outline on the construction of Carbon Fiber Tubes is outlined before the construction Plan. Though custom carbon fiber tubes will be ordered, extra material may be ordered. This plan will apply in case there is not enough time for custom orders to be prepared and shipped. Most of the necessary materials have been procured and there are no anticipated delays in construction. The workshop will be available during open hours for all team members to contribute to the construction. Carbon Fiber: In order to create a new carbon fiber body tube, use the steps outlined below from the following website: The creation of carbon fiber tubes can be divided into four stages: 1) Preparation 2) Layup 3) Mandrel Removal 4) Finishing What you will need (Figure 1): A full-length coupler Mylar film (.005 thick) Carbon fiber cloth Peel ply Chip brushes Epoxy resin Scotch tape 47

49 The necessary materials and the stages of carbon fiber creation are previously listed. In order to view a fully detailed plan one must consult the guide to fabrication carbon fiber airframes attached. For the custom body tubes that have been ordered, they will be cut using a Dremel tool and a reinforced cutter. This is the easiest way to cut carbon-fiber parts. Construction Plan: Section I Fragile Materials Section 1. The Fragile Materials payload is located entirely in the coupler connecting the top body tube with the recovery body tube. The coupler will be made out of phenolic tubing and be 12 in length and 5.38 inches in diameter and will be ordered with these dimensions and the necessary length will be cut in the workshop. As previously explained the coupler will be connected by shear pins to the recovery body tube where the sections separate during parachute ejection and descent. The top smaller body tube that connects with the FOP will be made of Carbon Fiber and be 12 in length and 5.5 in diameter and will be made with the outlined carbon fiber process previously explained. The connection to the coupler will stay throughout flight, and will be screwed in by bolts. The FOP will be two concentric cylinders in which the outer cylinder will be secured with wooden blocks and screws going through the blocks and coupler. The specific construction of the FOP is described in the relevant section. 2. The nosecone will be 13 in height and 5.5 in diameter made out of polypropylene. It will be secured to the top body tube by screws with a nut and washer on either end of the body tube. The nosecone will remain in place throughout the entire flight. Section II Recovery Section 3. The recovery body tube is 40 long. It will be a carbon fiber body tube. First the body tube will be cut to that length in the workshop. The aft end of the tube will be attached by shear pins to the Fin Can. When the black powder charges go off from the CRAM, will shear the pins the are attaching Sections I and II as well as Sections II and III. The CRAM is inserted in the payload as described in Section 3.3. For the aft end of this section, just drill in four holes for shear pins through the body tube and through the RCP coupler. Consult with RCP squad to determine where it is most appropriate to drill the holes. Section III Roll Control Section 4. The Roll Control Section will be comprised of two payload bays which are stacked on top of each other and placed in the 12 phenolic coupler. There will be a bulkhead with an eye bolt directly in front of the payload and a bulkhead aft of the payload in front of the Fin Can. Four threaded steel rods will be used to connect a bulkhead in the fin can to the RCP. These rods will run the entire length of the payload until the bulkhead with the eyebolt thus integrating the payload coupler to the Fin Can. The rods will be secured with nuts to the bulkheads and epoxied in since this section of the rocket will remain attached. Fins 5. A CAD drawing will be used of the fins according to the dimensions on OpenRocket to construct the parallelogram shaped fins. Be sure to include the fin tabs that extend into the body tube. Save files that are used to manufacture with the mill (refer to Creo textbook). Carbon fiber plates will be ordered with the exact thickness required and will be cut using a Dremel with a reinforced cutter. Once the fins are cut, double check their dimensions to ensure that they are 48

50 consistent with our actual dimensions on OpenRocket and that they fit in the body tube slots. Next, mark each side of each fin trailing edge and leading edge. The fins need to be constructed with the appropriate shape and must have a more rounded leading edge with a sharper trailing edge in order to keep the flow attached over the fins. 6. Take out the fin sanders, and place sand paper in each of them. The fin sander with the larger angle is for the leading edge, while the one with the smaller angle is for the trailing edge (like an airfoil). Place sandpaper in the wedge and the fin on the platform. Sand the fin diagonals, not the side parallel to the flow. Do not sand anything that will be underneath the body tube. The expectation is that the sanding is symmetrical and thorough. Ensure that that is the case. Body Tube: 7. Take the carbon fiber body tube with the fin slots that are pre-cut. Cut the body tube to a length of in the workshop. This will be the Fin Can and will house the motor mount and RCP Rail Buttons: 8. Use a straight edge to make a line in between two of the fin slots, parallel to the rocket. Make one point on that line one inch from the aft end of the body tube, and another 17 inches from the aft end. To be continued. Motor Mount: 9. The motor mount will be made out of carbon fiber following the process listed in the Carbon Fiber Section and will have a 3.13 diameter and a length of in the workshop. Ensure the three centering rings and fit around the motor mount and while still around the motor mount, put it through the main body tube to make sure it can fit in without trouble. If there is anything that doesn t fit well, sand accordingly. 10. The centering rings will be made out of fiberglass and will have an outer diameter of 5.38 and an inner diameter on Epoxy the aft 2 centering rings onto the motor mount. They are located at 13.5 and 0.75 from the aft end of the motor mount. Mark these two locations. Both the centering rings and bulkheads will be attached to the motor mount and fin can body tube using JB Weld epoxy in order to keep all the components in place during flight. Let cure. Use wood glue to fillet around the motor mount and centering ring. Let dry. 11. Motor Retention: For Motor Retention use the two-piece quick change motor retainer assembly that is manufactured by Aero Pack. They are manufactured out of precision machined 6061-T6 aluminum and are compatible for all of the 75 mm diameter solid rocket motors that the team has been considering for the propulsion system of the rocket. They are integrated into the launch vehicle by epoxying the adaptor piece to a section of the motor mount tube protruding at least 0.75 inches from the aft centering ring. The threads on the adaptor align with the threads on the retainer cap to seal the motor casing in the tube. 12. Place the motor mount with 2 centering rings inside the body tube. Slide in through the aft side of the rocket until about half an inch off from target location (so the aft centering ring is hanging off the end). Continued (rail buttons): Place the aft rail button on the point made an inch from the aft section of the body tube. Use a nut and washer and a small wooden block to secure it from the inside. (This step is located here so that there is no interference with the centering rings with the inner part of the rail buttons). Place epoxy on the outside of the most aft centering ring, and at the location on the inside of the body tube where the other centering ring goes. Slide the motor mount to its final location. Add epoxy around the contact points as needed. Secure as best as possible and let cure. 13. Place the last centering ring about 1 from the top of the motor mount and try as best as possible to epoxy the contact points with the body tube and motor mount (this will really only be possible to do from the forward side, since the aft is blocked by the other two centering rings). Let cure. 49

51 13.a Take a bulkhead, drill two holes for the threaded rods for RCP payload (consult RCP on where the holes go, spacing, size, etc.). epoxy a nut and washer on one side of the holes. That side will face down towards the motor mount. Place the bulkhead just forward of the top of the motor mount, and epoxy it in. Let cure. 13.b Take the forward rail button, and screw it into the bulkhead from the outside. Fin alignment and fin placement guide: When placing the fins, it is important that they stay aligned when epoxied onto the rocket. If they aren t parallel to the rocket, the flight path and stability will be adversely affected. 14. Use the fin alignment guide without any adhesives to ensure that it will fit and align the fins properly. Place epoxy (30 min cure) on the bottom side of the fins and place them inside the slots and connect them to the motor mount. Place the fin alignment guide around the body tube and fins and let the epoxy cure. Once they are set, remove the guide and place wood glue on the contact points of the fins with the external main body tube (make a fillet around the fin). Place the fin alignment guide around the fins again. Ensure the fins are place in the correct direction (i.e. trailing edge and leading edge directions are consistent) Verification of Vehicle Design Table Vehicle Design Verification Requirement Requirement will be verified by: Method of Verification Status Nose Cone and Ballast Stability - Applying an impulse to a ballasted nose cone as by a baseball bat and - Analyzing damage - Testing a ballasted nose cone in a practical sense - Ground Test in a Solids laboratory No Longer Necessary: Ballast Design changed. - Full Scale Test To occur in Feb 2017 Airframe Strength and Stability - Creating a finite element model and analyzing potential loads - Analyzing airframe prior and after each launch for damage - FEM Analysis in ADINA - Post Full Scale Launch Test Inspections To be done in Jan 2017 To occur in Feb 2017 Fin Strength and Alignment - Creating finite element models and analyzing potential loads - Analyzing fins prior and after each launch for damage - Ensuring proper alignment during construction with fin alignment mechanism - FEM Analysis in ADINA - Post Full Scale Launch Test Inspections - Construction Quality Started, October 31 st To occur in Feb 2017 Construction to start Jan /Feb

52 Roll Control Payload Integration Fragile Object Protection Payload Integration Recovery Integration - Ensuring the payload fins do not endanger the structural integrity of the launch vehicle - Ensuring Recovery bulkhead attached to the coupler is robust -Ensuring the coupler holding the Payload is robust in the launch vehicle - Ensuring that the shear pins shear as envisioned - Ensuring that the bulkheads and eye-bolts supporting the system are robust - Full Scale Test To occur in Feb Shake Test Prior to Launch in Feb Full Scale Test To occur in Feb Full Scale Test To occur in Feb Shake Test Post construction Jan 2017 Motor Integration and Retention - Verifying the sizes of purchased material prior to construction - Performing load analysis on chosen system - Launching a full scale test with the chosen motor - Inspection Jan Load Analysis Finished. Section Full Scale Test To occur in Feb 2017 Motor Performance - Ensuring that the chosen motor can achieve the required altitude - RockSim and OpenRocket Simulations The chosen motor achieves needed altitude -Full Scale Test To occur in Feb2017 Accuracy of Individual Masses and Center of Mass - Using Software to estimate center of mass and individual masses prior to construction. - Weighing individual components during and after construction, and measuring the overall rocket mass and center of mass -OpenRocket and RockSim - Inspection with scale Ongoing through design phase To occur during construction in Jan 2017 Ballast will not move throughout flight. - Performing shake tests before the ballast is loaded to ensure - Shake Tests before loading and Post construction in Jan

53 the durability of the containers and after the ballast is loaded to ensure that the containers are secured properly. - Inspecting the ballast containers and retention after launch to ensure that they did not break or move during flight. after loading - Full Scale Test To occur in Feb2017 KEY: Not started; In Progress; Finished Risks and Mitigations Concerning the risks involving the launch, three main categories of risks to consider are of the utmost importance. Any of these three risks can have detrimental effects if not properly analyzed and prevented given the necessary criteria. These three risks are structural failures, propulsion, and stability. Proper mitigation of these potential hazards is instrumental in ensuring the safety and performance of the launch vehicle. If any of these forms of failure are not adequately analyzed and prevented, the rocket and personnel can suffer. A major proponent of each risk is presented in Table Table Launch Risks and Mitigations Type Risks Cause Effects Controls/Mitigations Structural Failure Bod Tube (airframe) Failure Load Capacity of Body Tube exceeded during Launch and or deployment Airframe Cracks or Breaks, compromising structural integrity and rendering the vehicle non-reusable Critically Analyze design to identify weak points. Carry out appropriate calculations to determine expected stresses and use a reasonable safety margin. These have been done. Full Scale launches will verify these. Propulsion Motor Casing Explosion Nozzle is clogged by a detached chunk of propellant Motor casing explodes under pressure and partially or totally destroys the fin can Make sure that the certified personnel properly checks the motor casing and correctly packs the motor. Stability Vehicle is unstable The Center of gravity is behind the center of pressure Flight path will be unpredictable and erratic. Use OpenRocket and RockSim to ensure proper placement of CG and CP. Physically verify the location of CG prior to all launches 3.2 Subscale Flight Results In order to ensure a stable full-scale launch vehicle design, two subscale rockets were built. The primary goal of these subscales was to verify the accuracy of the simulations the team ran in OpenRocket and RockSim. The accuracy of the simulations determine how reliably the team can use those programs to predict the launch of the full-scale. One 52

54 was built with four additional fins to represent those that would be used for roll induction, and the other was built without these fins. The reason for having two models was to compare the flight results and see the effect the additional roll induction fins have on the rocket s flight. The subscale rockets were built at approximately 30% scale in comparison to the full-scale rocket. The scaling was not exactly 30% on most components, as certain parts such as the body tubes had to be purchased from vendors, and they only come in certain dimensions. The overall design of the subscale was similar to the full-scale, except for the integration of payloads. There were minor structural differences regarding the mounting of roll induction fins and the number of sections the launch vehicle splits into for recovery purposes. The main fin span was also changed in order to more accurately reflect the stability of the full-scale vehicle. The subscale s recovery system was to use an ejection charge attached to the motor that ignites about 8 seconds after burnout. This would split the rocket into two sections and deploy a parachute. A major difference between the subscale and full-scale is the choice of building materials. While the full-scale will be built using mostly carbon fiber and phenolic, the subscale was built from phenolic and birch plywood. However, both nosecones will be made of polypropylene plastic. The difference in material is negligible. In order to validate the software simulations, the material is noted so that error is minimized. For wind tunnel data comparison, it will be essential that both carry smooth surfaces so results from the subscale can be scaled accurately to the full scale. Calculated stability margins were above 2 for both versions of the subscale, which is the recommended value. The fin size was changed to more accurately represent the stability of the full-scale model. Both subscale models are shown below in Figure The subscale was launched twice. One configuration was the launch vehicle without roll induction fins. The other featured roll induction fins offset 45 degrees from the main fins. Figure Diagram Showing Both Subscale Designs The outer diameter of the launch vehicle was inches with an inner diameter of 1.5 inches. The overall length of the subscale was inches. The main fins had a height of 1.4 inches, with a root chord and tip chord length of inches. The sweep angle was 28.3 degrees. The weight of the launch vehicle with the roll induction fins, parachute, and altimeters was g. The weight of the launch vehicle without the roll induction fins was g. Dimensions and components are shown below in Figure 3.31 and Figure

55 Figure Dimensions and Components of Launch Vehicle with Roll Induction Fins Figure Dimensions and Components of Launch Vehicle without Roll Induction Fins 54

56 Figure Dimensions of Main Subscale Fins The motor used for both launches was an AeroTech F44-8. The characteristics of this motor is given in Table Table Motor Characteristics for AeroTech F44-8 Motor Classification AeroTech F44-8 Diameter (in) Length (in) 2.76 Average Thrust (lbf) Maximum Thrust (lbf) Total Impulse (lbf*s) Burn Time (s) 0.9 Total Weight (lb) Propellant Weight (lb)

57 Table 3.16 and Table 3.17 shows characteristics of the launch vehicle using the F44-8 with Roll Induction Fins offset 45º from the main fins. Table 3.18 and Table 3.19 shows characteristics of the launch vehicle using the F44-8 with Roll Induction Fins offset 45º from the main fins. Simulations were performed in both OpenRocket and RockSim for both configurations. Launch simulations were done using the same conditions as those present on launch day to best gauge the simulation accuracy. The launch day conditions are shown in Table 3.15 below. Table Launch Day Conditions in Three Oaks, MI (12/10/16) Temperature (ºF) 32 Wind (mph) 4 Pressure (IN) Latitude (º) 28.6 Longitude (º) Table Characteristics of Subscale Vehicle with Roll Induction Fins Offset at 45 (OpenRocket) Motor Classification F44-8 Length (in) Diameter (in) Total Weight (oz) 15 Stability Margin 2.43 Projected Apogee (ft) 914 Table Characteristics of Subscale Vehicle with Roll Induction Fins Offset at 45 (RockSim) Motor Classification F44-8 Length (in) Diameter (in) Total Weight (oz) 15 Stability Margin

58 Projected Apogee (ft) Table Characteristics of Subscale Vehicle without Roll Induction Fins (OpenRocket) Motor Classification F44-8 Length (in) Diameter (in) Total Weight (oz) 14.8 Stability Margin 2.69 Projected Apogee (ft) 969 Table Characteristics of Subscale Vehicle without Roll Induction Fins (RockSim) Motor Classification F44-8 Length (in) Diameter (in) Total Weight (oz) Stability Margin 2.83 Projected Apogee (ft) 1040 For the launch vehicle configuration without roll induction fins, OpenRocket predicted an apogee of 953 ft. This configuration had a stability of OpenRocket predicted a flight time of 100 s, an off the rail velocity of 86.9 ft/s, and deployment velocity of 44.3 ft/s. The actual launch resulted in an apogee of 1019 ft and a flight time of 71.8 s. The off the rail velocity was about 79 ft/s and the velocity at deployment was 50 ft/s. The actual apogee was noticeably higher than that predicted by OpenRocket, which means that in this configuration, OpenRocket overestimated the coefficient of drag. The flight time was also shorter than projected. This is clearly because of the failed deployment, in which the fin can was separated from the body tube and parachute due to a failure of the shock cord. The velocities both off the rail and at deployment were fairly accurate. For the launch vehicle configuration without RCP fins, RockSim was a better predictor. It predicted an apogee of 1040 ft, which is almost within 20 feet. OpenRocket was not so good, as it predicted an apogee of 954 ft. This difference could be explained by the differences in off-rail velocity and time of flight. RockSim overestimated the former, but underestimated the latter. It seemed they cancelled each other out. OpenRocket overestimated both, but 57

59 came out lower, which led the team to believe OpenRocket overestimated the drag coefficient when the fins were absent. Table Launch Results Compared to Simulation Predictions (No RC Fin Configurations) Source of Data Subscale Launch OpenRocket Simulation RockSim Simulation Apogee (ft) Flight Time (s) Velocity off Rail (ft/s) Deployment Velocity (ft/s) For the vehicle configuration with roll induction fins, OpenRocket predicted an apogee of 902 ft. The stability margin of the vehicle was OpenRocket predicted a flight time of 47.1 s, an off the rail velocity of 85.9 ft/s, and a deployment velocity of 49 ft/s. The actual launch of this configuration resulted in an apogee of 917 ft. The flight time from the launch is irrelevant. The recovery device failed to deploy at apogee due to extremely tight packing of the parachute. The off the rail velocity was about 91 ft/s. There was no velocity at deployment because the recovery system failed. The predicted apogee was slightly lower than the actual apogee, meaning a small overestimation in the coefficient of drag. The rail velocity was also fairly close to the actual value. RockSim however predicted an apogee of 987, compared to the 902 of OpenRocket for the configuration containing fins. OpenRocket was much closer, within 15 feet of the correct apogee. RockSim once again overestimated the off-rail velocity. Its much higher apogee suggests that its drag forces were estimated too low. The team was not able to compare flight times because the parachute was too tightly packed for the vehicle flight to be all the way successful. RockSim did indeed prove to significantly underestimate OpenRocket in terms of coefficient of drag. See Table 3.22 and Table Table 3.21 shows the launch results alongside the predicted values for both OpenRocket and RockSim for the configuration with roll induction fins. Table Launch Results Compared to Simulation Predictions (With RC Fin Configuration) Source of Data Subscale Launch OpenRocket Simulation RockSim Simulation Apogee (ft) Flight Time (s) Velocity off Rail (ft/s) Deployment Velocity (ft/s)

60 Table CD for Launch with no RC Fins Simulation Software OpenRocket RockSim CD Pre-burnout CD Post-burnout Table CD for Launch with RC Fins Simulation Software OpenRocket RockSim CD Pre-burnout CD Post-burnout The subscale launch was successful from the standpoint that the team now knows that the design is stable and structurally sound. The failure of the recovery system on the second launch are not relevant as the recovery system for the full scale design is much different than that of the subscale. At this stage, the full scale design has a stability margin higher than either subscale model has, even without any ballast near the nose cone. If anything, the full scale should be more stable than the subscale was. However, the results of the subscale launch show us that we cannot blindly use OpenRocket or Rocksim to predict performance. Based on the results of the subscale launch and the simulations, it seems that each software predicted better for one of the configurations, with OpenRocket working better for the with fin configuration, and RockSim working better for the without fin configuration. For this reason, we will be using OpenRocket as the primary source of simulation data as it appears to work better with the roll induction fins that will be present on the full scale. RockSim will be used if the roll induction payload does not fly as it better predicted the apogee when the launch vehicle did not have the second set of fins. Overall, the subscale gave the team confidence in the design and helped decide which simulation software will be prioritized moving forwards. 3.3 Recovery Subsystem The recovery system is necessary to ensure the vehicle is recoverable and reusable by reducing the vertical velocity during descent. It accomplishes this using a two-stage deployment scheme with a drogue parachute deployed at apogee and a main parachute deployed at 600 feet AGL. The components of the recovery system are shown in the diagram in Figure 3.34 and in Table 3.24, and detailed in the following subsections. 59

61 3.3.1 Hardware Components Figure Recovery System Layout Parachute Table Recovery System Components Component Location on Figure 3.34 Main Parachute A Drogue Parachute B Nomex Cloth C Nomex Shock Cord Protector D Shock Cords E Quick Links F CRAM v3 G Eyebolts H Shear Pins I The rocket will use a two-stage deployment recovery system, with a drogue parachute deploying at apogee, and a main parachute deploying at 600 feet AGL. The drogue will be a 24-inch diameter Compact Elliptical Parachute with a bleed hole. The main parachute will be a 120-inch Iris Ultra Standard Parachute. Both are made entirely of nylon Harnesses The parachutes are attached to the rocket using 1 inch tubular nylon shock cords. These allow the independent rocket sections to remain tethered to each other after separation, and also absorb the large forces generated by separation and parachute deployment. The shock cords are 4-5 times the length of the rocket to ensure that the rocket sections do not collide with each other during descent and to dissipate some of the ejection charge forces. 12 inch long tubular pieces of Nomex cloth will be wrapped around the portion of each shock cord that is closest to the separation charges to protect the shock cord when the charges detonate. Additionally, a sheet of 18 x 18 Nomex cloth will be wrapped around the end of each parachute that is facing the separation charges to protect these critical components from the explosion. 60

62 Bulkheads The bulkheads are attached to the CRAM and are made of aircraft grade plywood. These bulkheads protect the avionics module from the detonation of the separation charges. They were designed using Creo Parametric 3.0 and manufactured using a CNC mill, and then firmly attached to the CRAM with epoxy. The top bulkhead is made with a 5-ply birch plywood, while the bottom is slightly thicker and made out of 7-ply birch plywood. This is because the bottom bulkhead will bear a slightly higher load than the top Attachment Hardware The shock cords will be attached to the rocket sections using Quick Links and eye bolts. The Quick Links selected are zinc-plated steel, 3/16 diameter, with a load limit of 660 lbs. These will be looped through the ends of the shock cords and through the shroud lines of each parachute. The Quick Links at the ends of the shock cords will then be attached to closed-loop eyebolts in each of the sections of the rocket. These will be zinc-plated steel, with 3/8 thread size and 1 eye diameter. They are rated at a lifting capacity of 1300 lbs. lifting force. The eyebolt going through the bottom of the avionics module will be 6 long, to hold in the core of the CRAM. All other eye bolts will be 1 long. To maintain structural integrity during launch and ascent, the rocket will use shear pins to hold the sections. These plastic pins are designed to break upon detonation of the ejection charges. This allows for the deployment of the parachutes when the desired conditions are met, and not before Electrical Components Altimeters The Featherweight Raven3 altimeter is the critical component used to deploy the parachutes. Two Raven3 altimeters are used as the primary and redundant controller for the ejection charges. This altimeter allows for up to four deployment events; however, only two will be used for deployment of the drogue parachute at apogee and the main parachute. It utilizes a barometric pressure sensor operating at 20 Hz to determine altitude and an accelerometer operating at 400Hz to measure acceleration and calculate velocity. Specifications of the Raven3 altimeter are shown below in Table Table Featherweight Raven 3 Altimeter Specifications Power Source Maximum Altitude Altitude Resolution Barometric Sample Rate Axial Acceleration Sample Rate Lateral Acceleration Sample Rate 9V Battery 100,000 ft atm 20 Hz 400 Hz 200 Hz Dimensions 0.8" x 1.8" x 0.5" Weight 6.6 grams Both the primary and backup altimeters were programmed using the interface software provided by the manufacturer, which utilizes Boolean statements to determine under which conditions the deployment events should take place. Based on manufacturer recommendations and specific mission requirements, the settings are summarized in Table 3.26 below for both altimeters. 61

63 Table Deployment Settings for Primary and Backup Altimeters Primary Altimeter Backup Altimeter Drogue Deployment Pressure is increasing Velocity is negative AND Time delay of 0.7 seconds Main Deployment Pressure is increasing AND Altitude AGL is less than 600 ft. Pressure is increasing AND Altitude AGL is less than 550 ft CRAM Electrical Components To streamline the wiring for the CRAM, a PCB was designed to connect the Raven altimeter, the switches, electric matches and 9V battery power source. This PCB will replace a majority of the wiring in the avionics bay, reducing the probability of tangled wires and increasing the reliability of the connections. The PCB will be mounted on top of the core of the CRAM and will be connected to the electrical components. A schematic of the connections for these components is shown in Figure 3.35 below. Figure Recovery System PCB Schematic Looking at the left half of the plane, wires coming from the terminal connector of the raven altimeter will be connected to the PCB. Two of the pins will be connected to a main electric match igniter and apogee electric match igniter. The other ends of both igniters are connected to a switch to the 9V battery power source. Pin 1 of the Raven will also be connected to the switch and pin 2 will be connected to the negative terminal of the battery. Before flight, the screw switch will be closed, presenting current to one side of the electric matches. When the Raven signals the appropriate time for deployment, it will close the circuit internally on the other side of the electric match. With the circuit completely closed, the respective electric match will ignite and deploy the relevant parachute. These connections are copied on the right half of the schematic for the backup altimeter. The schematic was translated into a board layout shown in Figure 3.36 below. 62

64 Figure CRAM PCB Layout This board layout was designed using Eagle Light software. It was then transferred to PCB g-code to be used in a CNC mill. Using the g-code file, the PCB traces were milled on a copper plated board. The bit of the mill removed copper from the board, etching out the black outlines seen in Figure Mission Performance Predictions Validity of Analysis Both OpenRocket and RockSim had significant error when comparing simulation results with the subscale launch results. The team analyzed the data from the subscale launch and determined that the likely cause of these discrepancies was either the overestimation or underestimation of the coefficient of drag. Each performance prediction software performed better than the other in one of the launch configurations. Both software featured vehicles with equal weights and components, meaning that the most likely reason for these difference was smoothness of the vehicle. For the first launch in which the vehicle did not have roll induction fins, OpenRocket vastly underestimated the apogee. RockSim predicted the apogee much better. Based on the results, it seems OpenRocket overestimates the coefficient of drag when the vehicle does not have the 4 additional fins, while RockSim provides a more accurate estimation. In the case that the roll induction payload does not fly for one of the test launches, the prediction from RockSim will be weighted heavier than the prediction of OpenRocket. For the second launch in which the vehicle did have 4 additional fins, RockSim vastly overestimated the apogee while OpenRocket provided a very accurate estimate. Based on these results, OpenRocket provided an accurate estimation of coefficient of drag while RockSim underestimated the coefficient of drag. When predicting launches in which the roll control payload does fly, the prediction from OpenRocket will be more heavily weighted than that of RockSim because it estimates drag better in the roll control configuration Performance Prediction Program In previous years, the team has depended on commercially available predictive software to estimate the performance of the launch vehicle. This year, the team is developing several ways to verify this software. The one described here will be developed further later years. 63

65 Last year, the team predicted an altitude of more than 5300 feet. Instead, the launch vehicle reached an altitude of 4477 feet. While there were many oversights that will be avoided this time around, the team will use many methods to ensure the launch vehicle is within 100 feet of 5280 feet. This program accepts inputs such as the mass of the rocket, mass of engine, mass of propellant, coefficients of drag, thrust, burnout motor time etc. The masses are important because the altitudes will partially depend on this. The mass of the rocket changes during flight because the propellant disappears. The coefficients of drag are important. This is because the launch vehicle will go through several flight phases. While the motor is burning, the velocity will be increasing and thus the drag will increase; therefore, the team needs a reliable coefficient of drag at different points of velocity. After burnout, the launch vehicle coasts for a little bit, before being subjected to a rolling moment and then coasts to apogee once more. A constant multiplier of 1.72 was carried over from OpenRocket calculations to make the performance prediction program more accurate. This was from analysis of subscale data and the code that existed. After careful analysis of the physical equations, the results from the code, OpenRocket and RockSim results and the actual apogee, it was determined that some multipliers would be necessary, one for fin-included model and one for the fin-less model. There are three distinct rocket phases: the thrust phase, the roll phase, and the coast phase. The performance prediction program does not take into account exactly when the rocket begins rolling; it rather uses the percentage time the rocket will be rolling and thereby uses the rolling drag coefficient for a percentage of the flight and the coasting drag coefficient for the rest of the flight. This simplifies the performance prediction program because one less coasting phase needs to be calculated the coasting time after motor burnout and before the rocket begins to roll and also the time does not need to be taken into account during the calculations vastly simplifying the program. We know from simulations that the rocket will be rolling for approximately 6 seconds which is about 37.5% of the rocket s time after motor burnout. We used this information and applied the rolling drag coefficient which was calculated in the prediction program to 37.5% of the coasting flight while we applied the coasting drag coefficient to the other 62.5% of the flight. There OpenRocket and RockSim cannot simulate the roll portion of the flight. This is the most important reasoning for the necessity of an independent program. However, the team can still use OpenRocket and RockSim to guide the design of the program, as will be described in Section below. The program gives high apogee numbers, more similar to RockSim than OpenRocket. Which of these simulations is proven to be superior can only be found out after the full scale launch. The program breaks those flight phases apart in its predictions because they all affect the altitude differently. The program s raw code is included in Appendix A. The stability was also calculated in a second Python program in addition to it being calculated in OpenRocket and RockSim. This program first calculates the center of pressure through the use of Barrowman equations by breaking the rocket into three separate terms: nose cone terms, conical transition terms, and fin terms. The program then calculates the stability using an imported center of gravity. The stability was calculated to be This is slightly lower than the stability calculated by OpenRocket and RockSim but a stability of 2.22 will still yield a stable rocket. This program can still be further developed to yield a more accurate stability and the program is included in Appendix B. The source for the Barrowman equations can also be seen in Appendix B Wind Tunnel Tests On December 16 th, 2016, the Vehicle Sub-team was able to perform wind tunnel tests in the Hessert Laboratory for Aerospace Research. The model used for the wind tunnel tests was the subscale model. For the tests, a model with roll control fins was tested as well as a model without. Section 3.2 elaborates the subscale s dimensions and structure. 64

66 The purpose of the wind tunnel tests was to verify the drag co-efficient values that were attained through OpenRocket and RockSim. Secondly, the team aspired to find out how different the effect of the presence of an extra set of fins on the drag co-efficient. The MATLAB codes used to generate the graphs in this section can be found in Appendix D. To perform the test, a force balance was set up in a wind-tunnel big enough to fit the rocket. Data could be collect with lift and drag transducers. In our case, the drag transducer data was analyzed. Lift data shall be analyzed at a later time. The setup is shown in Figure Subscale Model Wind Tunnel Setup. Care was taken to account for the vertical metallic pole. Figure Subscale Model Wind Tunnel Setup To calibrate the transducers, masses were hung on strings attached to the transducers. The transducers emit values in Voltage and these must be converted to force values. The masses ranged from 50 g to 700 g. The drag force graph as a function of the voltages is shown in Figure Drag Transducer Calibration. The slope is Figure Drag Transducer Calibration 65

67 The drag was calculated for the rocket without extra fins, the rocket with extra fins aligning with the main fins and for the rocket at 45 degrees with the main fins. The main comparison of interest is the one with the former two. The latter was done in case complications from rail size forced the team to go with aligning the two sets of fins; that is, if there is not enough space for the rail within the 45 degrees that is allowed by a main fin and a Roll Control Payload fin, then the RCP fins would have to align with the main fins. This could affect flight. The effects had to be measured. Figure CD as a function of Reynolds Number (Without RC Fins) Figure CD as a function of Reynolds Number (RC Fins at 45 ) Comparing the coefficient of the rocket without fins and one with fins at 45 degrees, it appears there is a 7.72% increase in drag coefficient. The average coefficient with fins at 45 degrees is The average coefficient without fins is An increase was expected. The values themselves proved too high. OpenRocket approximates a 66

68 coefficient of about 0.84 without fins and RockSim With fins, OpenRocket gives an approximate coefficient of 0.95 while RockSim gives With these numbers, OpenRocket gave a difference of ~13% between the rocket without fins and the one with. RockSim gave a difference of ~11%. The wind tunnel suggests a change of about 7.7%. This suggested that both software could not simulate both instances correctly. Indeed, the subscale results described in Section 3.2 show that one was great at predicting one instance while the other predicted the other closer. While the results were inconclusive in validating the coefficient of drag themselves, they helped in determining exactly which of the software estimated which instance closely. OpenRocket for when the team is using fins. RockSim for when the team is not using fins. However, the team will continue to use both for both instances until the full scale results are acquired. The coefficient of drag as a function of Reynolds No. for the rocket with the fins aligning is shown in Figure 3.41 below. Figure CD as a function of Reynolds Number (RC Fins at 0 ) The average coefficient for the rocket with fins aligning was found to be This is lower than the model with RCP fins at 45 degrees. The final design, described in the Critical Design Report has specified that the RCP fins will be at 45 degrees. This is a sort of insurance. If the full scale launch is below the altitude required, the fins alignment will be switched Apogee Approximations Apogee Approximations using launch simulations are one of the primary methods the team uses to ensure a launch that meets the mission success criteria. Reaching the target apogee is primary objective, so accurate estimates are required. Two different softwares are used to obtain the most accurate estimate. Simulations are performed at a variety of conditions in order to find the configuration that best meets the requirements. The changes in condition include changing wind speeds, temperature and pressure. Simulations are performed before any construction or test launches are performed to give a general idea of the launch results, and those results are then verified later by test launches. If the results from the simulation seem reasonable, the design is likely worth pursuing. Simulations are needed as one of the many factors to determine the feasibility of the design. Apogee approximations are also compared with results from test flights to determine how accurately the simulations can be used to predict future launches if changes are made. 67

69 The launch condition that was changed between simulations was the wind speed because that is the most unpredictable and most influential condition on the launch results. Simulations were performed at ranges from 5 to 20 mph winds. The most common conditions are 5 to 10 mph winds. Pressure, temperature, and latitude have an impact, but it is much less than a change in wind speed. A ballast of 1 lb was used in these simulations. Current analysis (described in Section ) shows that the Roll Control Payload affects flight by only 100 ft. These simulations would put us in the correct area. The team awaits the full scale launch to verify this analysis. It is however better in this instance to estimate with 1 lb, which we can take off than with less weight which we may not be able to meet. The rail used will be 10 ft long and 1.5 inches wide. The rail length will affect the velocity when the launch vehicle leaves the rail. In order to be considered stable, the off the rail velocity should be at least 52 ft/s. The longer the rail, the faster the launch vehicle will be traveling when it exits. Table 3.27 through Table 3.33 show simulation data from OpenRocket and RockSim for various wind speeds while other conditions remain the same. Figure Flight Profile of Launch Vehicle with 10 mph Winds shows the flight profile of the launch with 10 mph winds. Table OpenRocket Simulation at 5 mph Winds Apogee (ft) 5452 Off Rail Velocity (ft/s) 59.9 Maximum Velocity (ft/s) 659 Maximum Acceleration (ft/s^2) 280 Time to Apogee (s) 18.7 Flight Time (s) 212 Table OpenRocket Simulation at 10 mph Winds Apogee (ft) 5410 Off Rail Velocity (ft/s) 59.9 Maximum Velocity (ft/s) 657 Maximum Acceleration (ft/s^2) 287 Time to Apogee (s) 18.6 Flight Time (s)

70 Table OpenRocket Simulation at 15 mph Winds Apogee (ft) 5334 Off Rail Velocity (ft/s) 59.9 Maximum Velocity (ft/s) 656 Maximum Acceleration (ft/s^2) 293 Time to Apogee (s) 18.4 Flight Time (s) 208 Table OpenRocket Simulation at 20 mph Winds Apogee (ft) 5275 Off Rail Velocity (ft/s) 59.9 Maximum Velocity (ft/s) 655 Maximum Acceleration (ft/s^2) 295 Time to Apogee (s) 18.3 Flight Time (s) 208 Table RockSim Simulation at Low Winds (0-2 mph) Apogee (ft) Off Rail Velocity (ft/s) Maximum Velocity (ft/s) Maximum Acceleration (ft/s^2) Time to Apogee (s) Flight Time (s)

71 Table RockSim Simulation at Medium Winds (3-7 mph) Apogee (ft) Off Rail Velocity (ft/s) Maximum Velocity (ft/s) Maximum Acceleration (ft/s^2) Time to Apogee (s) Flight Time (s) Table RockSim Simulations at High Winds (8-14 mph) Apogee (ft) Off Rail Velocity (ft/s) Maximum Velocity (ft/s) Maximum Acceleration (ft/s^2) Time to Apogee (s) Flight Time (s)

72 Figure Flight Profile of Launch Vehicle with 10 mph Winds As shown in Section , the apogee is not heavily influenced by the roll induction payload. The simulation shows that the apogee is going to decrease slightly due to the presence of the 4 extra fins (by about 100 feet). The team will keep the current estimates and await for the full scale launch to make any changes that may be required after examining the real-life effects of the roll control. These estimates were done with 1 lb ballast in anticipation of unforeseen weight increases and as a sort of double-insurance in case the effect of the roll control is significantly larger than analyzed. The maximum apogee using these simulations was found to be 5602 ft using RockSim and having winds ranging from 0-2 mph. This is very slightly above the maximum allowed apogee of 5600 ft. If test launches indicate that the launch vehicle exceeds 5600 ft, additional ballast will be added in order to decrease the apogee within the competition limits. Based on results of subscale launches discussed in Section 3.2, the team is using OpenRocket as the primary simulation software for any configuration that includes roll induction fins. The apogee predicted by OpenRocket in 10 mph winds is roughly 120 ft over the target apogee of 5280 ft. This was done to account for the roughly 100 ft decrease in apogee due to the roll induction fins. The 1 lb of ballast will be changed based on test launch results and the conditions at the launch site. OpenRocket gives an average coefficient of drag of pre-burnout and post-burnout. The applied correction for the rolling provides a post-burnout coefficient of drag of The rolling is expected to last for about 6 seconds. This analysis is outlined in Section Results from both simulation software give confidence in the design. These are the apogee values that the team designed the launch vehicle for to counteract the effect of the roll induction fins. Now that simulations show that our initial prediction of the extra fins slightly decreasing the apogee was correct, the team must wait until the result of test launches to make any additional changes to the launch vehicle. The test launch results will also determine how reliably the team can depend on OpenRocket and RockSim estimates in the future 71

73 3.4.3 Stability The diagrams of the rocket in Figure 3.43 and Figure 3.43 show the relative locations of the center of gravity (CG) and center of pressure (CP) for both the OpenRocket and Rocksim predictions, respectively. As seen, the CG is forward of the CP, which is necessary for the rocket to be stable. The locations of the CG, CP, and stability margin for both simulations can be found in Table Locations of CG and CP, and Stability Margins for Both Simulations. The stability margin was measured to be 2.71 in OpenRocket and 2.70 in Rocksim, both of which are comfortably above the minimum required 2.00 margin. If the stability margin comes too close or lower than 2.00 due to an increase in weight in the aft section of the rocket or a decrease in weight towards the nose, ballast will be used to move the CG and therefore change the stability. The ballast will be placed towards the nose of the rocket, which will move the CG forward. Since this will not impact the aerodynamics of the rocket, and therefore the CP, the distance between the two points will increase, and thus the stability margin will also increase. Table Locations of CG and CP, and Stability Margins for Both Simulations OpenRocket Rocksim CG (in inches from nose) CP (in inches from nose) Stability Margin (without motor) Stability Margin (with motor) Figure OpenRocket Stability 72

74 3.4.4 Kinetic Energy Figure RockSim Stability Kinetic energy management is one of the most important roles of the recovery system. Minimization of kinetic energy upon landing is essential to ensuring recovery and reusability of the rocket, as well as satisfaction of the competition rules. Toward this end, the rocket uses a large main parachute near the ground to slow the rocket down to safe levels before landing. To reduce wind drift, the rocket descends at a faster rate under a smaller drogue parachute until the main chute deploys. The kinetic energy of the rocket and the heaviest rocket segment during each significant phase of flight was calculated using the velocity data obtained in the Runge-Kutta Matlab code. The results are shown in Table 3.35 and the velocity plot after apogee is shown in Figure Table Kinetic Energy at Key Launch Phases Kinetic Energy (ft-lb) Phase Velocity (ft./s) Total Rocket oz. Front Section (Heaviest Stage) oz. Motor Burnout , ,500 Apogee/Drogue Deployment Main Deployment Landing

75 3.4.5 Wind Drift Figure Vertical Descent Speed Calciulations Figure Wind Drift Calculations below shows the predicted horizontal drift of the rocket under various wind conditions. Under the maximum operable speed of 20 mph, the Runge-Kutta Matlab calculations predict a drift of 1870 ft, which is below the 2500 ft limit. 74

76 Figure Wind Drift Calculations 3.5 Launch Procedures The team has developed procedures to follow for every launch, including the test launches and the launch in Huntsville. The launch procedures are going to be followed in order to ensure a flawless flight and to help meet the mission success criteria. Each sub-team has its launch procedures to follow for assembling its payload/sub-systems. Designated members of the sub-team sign off on the launch procedures followed by the leaders of the sub-team to ensure that the procedures were followed correctly. These procedures are listed in Appendix G. 3.6 Vehicle Test Plan Table Vehicle Test Plan Time Period of Test Test Purpose of Test Status November 2016 FEM Analysis To analyze load paths and Started; To be stresses in order to reinforce continued in January effectively. Body 2017 Tube and Fins to be prioritized 75

77 4 Safety November 20, 2016 December 3, 2016 February 2017 Subscale Test To verify simulations performed in OpenRocket and RockSim and utilize correction factors where applicable Material Stress Test To verify the strength and stress properties of chosen material November 2016 CFD analysis To calculate the coefficient of drag To finalize fin design To attain a more refined calculation of altitude December 2016 Wind Tunnel Analysis January 2017 February 2017 January 2017 February 2017 Shake Tests Ballast Test To verify drag estimates during flight. To verify the effect of fin placement on flight path - To verify the required robustness of integration systems before flight - To verify the stability of ballast systems in case they are ever needed January 21, 2017 Full Scale Test To verify simulations performed in OpenRocket and RockSim To verify the requirements in Section February 18, nd /Back-up Full Scale Test KEY: Not started; In Progress; Finished To verify the requirements in Section To test features added for safety and efficiency after first launch Complete. Vehicle design tweaked to account for results. To occur after material deliveries Unable to schedule; working to fit into schedule Complete. Results described in Section To occure after material deliveries To occur after material deliveries To occur on January 21, 2017 To occur on February 18, Checklist of Final Assembly and Launch Procedures A detailed pre-launch checklist will guide the final assembly process for the rocket with step-by- step instructions. A repeatable launch procedure will also be developed to mitigate risk of failure at the launch site, and a post-launch procedure will ensure the all personnel retrieve the rocket in a manner that is safe for both the personnel and the rocket. These steps must be followed precisely to ensure successful execution of the project. These procedures can be found in Appendix G. 76

78 4.2 Personnel Hazard Analysis The Notre Dame Rocketry Team understands that the construction, testing, and launch of the rocket pose several potential hazards to team members. The table below explores the personnel hazards that may occur during different phases of constructing or testing the launch vehicle and its subsystems. Similar to the FMEA table, a severity, likelihood, and overall risk level was assigned to each hazard to better understand what mitigations are necessary. The risks and likelihoods were assessed assuming that all team members have been properly trained, are following the correct procedures, and are wearing the proper personal protective equipment (PPE). By recognizing these hazards now, the team can be better prepared to mitigate them and to take the proper actions in the event that an accident occurs. This table can be found in Appendix H. 4.3 Failure Modes and Effects Analysis (FMEA) A Failure Modes and Effects Analysis (FMEA) table was developed to identify the potential technical failures of the vehicle. For each failure, the effects and causes were identified, as well as their likelihood of happening, and the severity of their occurrence. The last two parameters were used to assess the risk of each failure through the Risk Assessment Codes (RACs) suggested in the handbook for the competition. The risk matrix used, based on the one shown in the handbook s appendix, is shown in Figure 4.1. Severity Probability Catastrophic Critical Marginal Negligible Frequent Probable Occasional High Risk Moderate Risk Figure 4.1. Risk Assesment Matrix Low Risk Minimal Risk After classifying the risk of each failure mode, mitigations and controls to prevent said failures were developed. It is important to note the importance of first determining the level of risk of each failure as to implement appropriate mitigation levels. Figure 4.2 depicts how the failure modes were divided into six categories: Structural, Recovers, Propulsion, Stability, and relating to the specific payloads. The FMEA table for all possible failure modes the launch vehicle and its subsystems may experience can be found in Appendix I. 77

79 Figure 4.2. Failure Mode Classification 4.4 Environmental Concerns The environment in which the rocket will be operated also poses a certain amount of risk. Specific problems related to inclement weather at the launch and landing sites have been identified and solutions have been devised to decrease the negative effects of the environment on the rocket. Additionally, many of the materials used in the construction of the rocket pose a significant hazard to the environment if they are improperly handled. By considering these things, one can ensure that the rocket is able to adequately perform and not be negatively affected by the environment. Failure modes tables have been constructed for both the environmental effects on the rocket and the rocket s effect on the environment. These tables can be found in Appendix J and Appendix K, respectively. 4.5 Project Risks There is the possibility of encountering a number of roadblocks throughout the rocket design and launch process. Each of these risks has been identified and categorized in terms of their potential impact on the project and the likelihood of that specific problem occurring. Risk is minimized with specific mitigation plans for each scenario. Failure to mitigate these risks will result in significant time delays for the project, which in turn lowers the chance of success on launch day. A table has been constructed outlining potential risks associated with the project, their likelihood, their impact on the project, and how they will be mitigated. This table can be found in Appendix L. 78

80 5 Payloads 5.1 Roll Control Payload Design and Testing of Payload Equipment System Level Design Review Roll Fins The roll-control system designed for this rocket requires a set of external, movable control surfaces to control rotation about the roll axis. To accomplish this, the roll control payload will feature four external fins, equally spaced around the circumference of the fuselage. The fins will be 3 tall, 1 deep, and thick. They will be made of 3D Printed carbon fiber, and the 3/16 shaft and fin will be printed as one piece to remove possible uncertainty and weakness in the joint. This 3/16 shaft will mate to a drive shaft from the bevel gearbox using an aluminum clamping coupler. The fins will initially be positioned vertically, such that the long edge is parallel to the roll axis of the rocket, and they will remain in this position until burnout so as to avoid inducing a static roll to the rocket or cause unexpected drag which might prevent the rocket from reaching the target apogee. After burnout, the fins will be driven by a servomotor connected to the bevel gearbox, which will allow the payload to induce the required roll and counter-roll motions Servomotor Figure 5.1. CAD Rendering of Roll Fin In order to provide the necessary physical motion to control the bevel gear system, which will control the angle of the roll fins, a ClearPath CPM-MCPV-RLN model servomotor will be used. This servomotor has a Shaft Diameter 79

81 of and can provide a RMS torque of 0.41 N-m. The servomotor will be set up in 2-Postion mode which allows for 2 user defined positions and a home position. In this context the home position will be set so the fins are perpendicular to the rocket. The 2 user defined positions will be set so that the fins are 4 offset from the home position. The servo can be controlled in this 2-Position mode using three control signals. When the motor receives a +5V signal across the enable control, the motor coils are energized. When the motor receives a +5V signal on Input B the motor is set to the homing position which will be used throughout the Rocket motor burn. Once burnout has happened Input B will switch to +0 V and Input A will toggle between +5 V and +0 V to control which position (+4 degrees) or (-4 degrees) relative to the home position (0 degrees). All of the control signals will be from the digital output pins of an Arduino Uno Gearbox A bevel gearbox will be used to drive the four exterior control surfaces of the rocket. This design was chosen because it allows all four fins to be controlled at the same time by one servo, which will also improve reliability and make the payload easier to control. The gearbox has a 20-tooth steel drive gear that will be driven by the servo, and four 10-tooth, steel, driven pinion gears. Each of the four pinions is connected to one of the fins via a 3/16 aluminum shaft, which interfaces with the shaft on the fin through an aluminum clamping coupler. The gears all interface with their respective shafts through a 6-32 set screw and a machined flat on the shaft to prevent slip. The gearbox is machined from a solid piece of aluminum in order to maintain precision and strength. The shafts are all supported by flanged bronze bushings, which allow precise alignment of the servo shaft and all the interfacing gears, and also reduce friction from the lateral forces that any right-angle drive produces. After aligning the servo shaft with its bushing, the gearbox is attached to the servo directly via the four bolt holes in the corners of the box. A recess is cut into the bottom of the box as well, to remove extra mass and to make sure that the mounting boss on the servo doesn t interfere with alignment. 80

82 Electronic Control System Figure 5.2. CAD Rendering of Bevel Gearbox An Adafruit Feather microcontroller is the central feature of the Electronic Control System (ECS). The system will be powered by a 9V battery. This source of power was chosen because 9V is one of the more common battery types; they are inexpensive and readily available. The Adafruit 10DOF sensor package will be interfaced with the Arduino providing a 3-axis gyroscope, 3-axis accelerometer, and an altimeter for real-time data on acceleration, altitude, and velocity. This single package offers streamlined communication in comparison to the alternative of monitoring each sensor individually. An algorithm fed with readings from the gyroscope will control the roll fins that are responsible for rotating the rocket. The other on-board electronic component will be a ClearPath CPM-MCPV- 2310D-RLN servomotor powered by two 22.2V rechargeable batteries. The motor was chosen because it can be programmed to turn the roll induction fins precisely to their three preset positions. It is small enough to fit inside the body of the rocket, but its max torque of 0.45 N*m is adequate for turning the gears that will in turn rotate the fins. The servo was manufactured in the USA and backed by a three-year warranty so reliability should not be an issue. As a sustainability measure, the batteries powering the motor will be rechargeable. This also saves money over purchasing many single use batteries. The specific voltage level of the batteries was chosen due to their high current output to power the motor. The servo will physically turn the fins that roll the rocket and will either be controlled by external switches on the rocket body or by the Arduino, based on input from a wireless ground station. A gearbox with a 2:1 gear ratio will mediate between the fins and the motor. A single drive gear is attached to the shaft of the servo with one smaller gear attached to each of the four fins. 81

83 Figure 5.3. Servomotor Position Control Diagram 82

84 Figure 5.4. Servomotor Input Diagram Figure 5.5. Servomotor Example Inputs On the outside of the rocket body are three physical switches used to control varies aspects of the roll control system. The first of three on-board switches connects the 9V power to the Arduino, providing control of the payload and the opportunity for power conservation when not on the launch pad. The second switch overrides the Arduino by sending a signal voltage directly to the servomotor. In this way, Arduino control of the motor is bypassed completely. This signal will cause the servo to turn the fins to the vertical home position. In this position, the secondary fins will be locked in a vertical orientation as they should be during launch and the rocket will not spin unnecessarily. This is useful because if there were an issue with the Arduino, the rocket would at least be able to launch without rolling uncontrollably. The third and final switch controls the connection between the servomotor and its 44.4V power source. With so much current flowing to the motor, it is important to make sure it is not unnecessarily connected to the 83

85 batteries when it is not being used. This prevents a large discharge, and turning the third switch off outside of the launch window will conserve energy. Control of the servo will be maintained by the Arduino, connected sensors, and the printed transistor circuit. At the time of launch, the motor will hold the roll induction fins in the vertical home position because a +5V signal will be sent to the Enable control and to Input B. After burnout, the signal on Input B will switch to 0V, taking the motor out of the locked in home position. The signal on Input A will toggle between 0V and +5V to rotate the fins +/-45 degrees, causing the rocket to roll clockwise or counterclockwise depending on input from the gyroscope. The override button on the ground station or switch on the rocket will restore the +5V signal to Input B, causing the motor to rotate the fins back to their vertical home position. Example code for the servomotor can be seen below: void Roll_Control(void) { /* Get a new sensor event */ sensors_event_t event; gyro.getevent(&event); accel.getevent(&event); // Assume Z-axis is vertical for accel/gyro? bmp.getevent(&event); if(!startrollflag) { //For roll initialization, cant fins in the direction of current roll digitalwrite(homepin, LOW); startrollflag = true; if(event.gyro.z > 0) { digitalwrite(statepin, HIGH); finstate = true; } else { digitalwrite(statepin, LOW); finstate = false; } } else if(!endrollflag) { //Wait until two revolutions have ben completed to begin counter-roll if(rotationcounter > 2) { finstate =!finstate; digitalwrite(statepin, finstate); endrollflag = true; } } else { //At this point, continue to make adjustments to prevent roll if(abs(event.gyro.z) < MIN_ROLL_THRESHOLD) { digitalwrite(homepin, HIGH); } else if(event.gyro.z > 0) { digitalwrite(homepin, LOW); digitalwrite(statepin, LOW); 84

86 } } } else if(event.gyro.z < 0) { digitalwrite(homepin, LOW); digitalwrite(statepin, HIGH); } All flight data gathered by the sensors will be written to an SD card for further analysis. In addition to recording the readings from the 3-axis accelerometer and 3-axis gyroscope, the rocket will also keep track of pressure, temperature, altitude, and time during flight. Example code for data logging can be seen below: void Record_Data(void){ //Subroutine for saving sensor data to the SD card sensors_event_t event; gyro.getevent(&event); accel.getevent(&event); bmp.getevent(&event); float temperature; bmp.gettemperature(&temperature); //Fill data buffers with sensor readings acceldata[0] = event.acceleration.x; acceldata[1] = event.acceleration.y; acceldata[2] = event.acceleration.z; gyrodata[0] = event.gyro.x; gyrodata[1] = event.gyro.y; gyrodata[2] = event.gyro.z; barodata[0] = event.pressure; barodata[1] = temperature; barodata[2] = bmp.pressuretoaltitude(sealevelpressure, event.pressure); timedata[0] = millis(); timedata[1] = millis() - starttime; datalog = SD.open("flight_data.txt", FILE_WRITE); //Open the file flight_data.txt in write mode 85

87 if (datalog) { //log data only if the file opened properly datalog.println(); //Start a new line for(int c = 0; c > 1; c++){ datalog.print(timedata[c]); datalog.print(", "); //Separate data entries by a comma and a space } for(int c = 0; c > 2; c++){ datalog.print(acceldata[c]); datalog.print(", "); } for(int c = 0; c > 2; c++){ datalog.print(gyrodata[c]); datalog.print(", "); } for(int c = 0; c > 2; c++){ datalog.print(barodata[c]); datalog.print(", "); } } } datalog.close(); //Close the file Verification tests will be run to ensure proper function of the three external switches on the rocket, including the override feature, as well as the two switches, ten LEDs, and override button on the ground station. These tests will include checking the correct voltages are sent through the circuit board and that these signals produce the appropriate response actions in the servo motor and roll induction fins when the corresponding switches and buttons are selected. Further testing will be done to ensure the Arduino and its code is properly interfaced the servomotor, and to verify the radio communication between the ground station and rocket, including storage and relay of sensor data and enabling 86

88 the payload remotely. Much of this functionality can be easily verified with quick and inexpensive ground testing, but a greater level of confidence will be instilled through full scale test launches. Two circuit boards were designed; one for use onboard the rocket and the other for the ground station. Board size was a major factor for the circuit inside of the rocket. The final board dimensions of 63.50mm x 54.60mm not only had to be small enough to fit in the rocket cylinder, but also avoid the four support rods that secure the payloads. On the ground station, size was less of a concern. The outer casing of the station was designed to accommodate the 90.17mm x 81.28mm board. All boards were printed at OSH Park. They produced low priced boards that arrived quickly because they were not made overseas. Waste was also avoided because OSH had a minimum order of three boards. Many other companies required a purchase of additional boards we had no intention of using. A printed circuit board is superior to point-to-point wiring all of the connections in the rocket and ground station because there are fewer parts that can come loose or short circuit. Once the electrical components are securely soldered to the PCB, it is a very robust design even under the stresses of launch and landing. To further increase reliability and make use of some of the extra PCBs that were produced, duplicate boards for both the rocket and ground station will be soldered and tested in case a faulty electronic component such as a diode or transistor causes a failure in one of the boards. Schematics for the boards can be seen below. 87

89 Figure 5.6. Payload Schematic Figure 5.7. Payload Board 88

90 Figure 5.8. Ground Station Schematic Figure 5.9. Ground Station Board A ground station will take advantage of the Feather board s radio capabilities to remotely communicate with the rocket. Powered by a simple 9V battery, the ground station uses two physical switches and a button to electronically control the rocket. The first switch is an on/off toggle to conserve power in the station itself. The other switch remotely gets the payload on the rocket ready for launch by taking it out of low power mode. The manual override button 89

91 bypasses the programming in the controller and sends a signal to the servo to rotate the roll induction fins to a neutral, vertical position. The ground station will also output live altitude and GPS data from the rocket Roll Control Algorithm The primary goal of this Roll Control Payload is to induce two full revolutions during post-burnout ascent and then prevent further rotation on the roll axis. With the variable nature of rocket launches and possible aerodynamic conditions, this payload will not be effective unless it employs an adaptive algorithm that can actively assess the situation and react accordingly. With this active control in mind, a control algorithm must be used to track the rocket's flight staging and rotational velocity to determine how and when to cant the fins. The control algorithm that will be used in this design will first detect liftoff and burnout using onboard sensors, along with a time-based system as backup. Once burnout has occurred, the algorithm will cant the fins and begin tracking roll using data from the onboard gyroscope. When two full revolutions have been completed, the algorithm will cant the fins to the opposite preset position, reducing the rocket's roll. For the rest of the flight, the algorithm will continue to minimize the rocket's roll by canting the fins to counteract any rotation it detects Power Control System In order to control the adjustable secondary fins, batteries must be connected to the servomotor in order to provide power. In order to achieve this, a switch will be used that will connect and disconnect the batteries to the servo. A total of two 22.2 V batteries will be used to power the servo, and a diode will be connected across the servo to prevent any back Electromotive Force (EMF). A separate battery will be used to independently power the Arduino, this battery will also be able to be connected and disconnected by a switch. Verification of the batteries will be through testing the given voltage of both batteries. The system will be evaluated through ground inspection to observe if each subsystem runs smoothly. Power to the servomotor that controls the turning of the secondary fins will be provided by two 22.2V rechargeable batteries. These batteries are rechargeable to reduce waste and increase the reusability of the rocket. It is more desirable to recharge the same batteries than to order new ones and reinstall them between each launch. A switch will be used for connection and disconnection between the power source and the motor, as mentioned in the Electronic Control System section. Risk of back Electromotive Force (EMF) will be mitigated by use of a diode across the servo. The diode used is rated for a max reverse voltage of 50V so it will be able to handle the 44.4V total of the two batteries connected to the motor. A second switch and 9V battery will power the Arduino Feather independently of the motor. Tests will be performed to verify all three batteries perform at their specified voltage levels. Additional ground inspection will be performed to evaluate each subsystem. Control over the power supply is crucial to the competition because the batteries need to have enough charge to reliably operate during flight even if the rocket has spent a significant amount of time waiting on the launch pad. Use of the ground station allows the payload to be powered on remotely, closer to the time of launch. This reduces energy consumption over having to power up the payload with the switch directly on the rocket Structural Support Measures To ensure stability of the payload and the rocket, four steel rods, acting as stringers, will run throughout the payload and will anchor this system to the other sections of the rocket. At each end of the payload there will be wooden bulkheads to anchor onto. Each subsection of the payload will be separated into 3D printed interlocking bays. The structural support will be evaluated through visual inspection Ground Station For communication with the Adafruit Feather, a wireless ground station will be constructed. In addition to processing power and control capabilities comparable to an Arduino Microprocessor, the Feather board offers radio 90

92 communication to relay live flight data from the rocket to the ground station. It also is capable of sending signals from the ground station to control the roll induction fins. This station will be powered by a 9V battery. The ubiquity of this type of battery saves money and lead-time in purchasing them, while providing ample power to the ground station. It will feature a LCD display to provide access to real-time flight position data. The most important outputs will be GPS and altitude. Position readings aid in recovery of the rocket, while altitude is helpful for verification that the roll occurs before the target apogee. A total of ten LED indicators will clearly display the connection status between the ground station and the rocket in addition to the position of the secondary, roll-inducing fins. The first of two switches will connect the entire ground station to its power source. The second switch will take the roll control payload onboard the rocket out of low power conservation mode so that it is ready for launch. This sends a +5V signal to the Enable Control of the servomotor so it is ready for input and a +5V signal to Input B to lock the fins in their vertical home position. Finally, a manual override button will send a signal to the rocket to turn the fins back to the vertical home position and stop the rocket from spinning in event of a failure of the Arduino algorithm. This supplies a +5V signal to Input B if it had been turned off during flight to allow to motor to turn to one of its preset positions. The LED indicators will also illuminate to show whether the motor is in transit between preset positions or if it has turned the fins to the desired cant of ±4 degrees. This rotation is accomplished by turning the signal on Input B to 0V and the signal on Input A to +5V or 0V depending on the desired motor position and fin angle. The circuit for the ground station was designed in-house using the CadSoft EAGLE PCB software, and a PCB was fabricated at OSH Park. Printing a circuit board was a much more robust option over wiring all of the connections in the ground station. The PCB is particularly advantageous for a ground station because it is small and portable, and no wires will come loose or short circuit during transport and use. OSH Park was a smart choice for a supplier because they were able to produce a smaller quantity of boards than most companies. Shipping was also relatively quick because all of the PCBs were made in the USA. In addition to the benefits of providing live position data and offering the option of emergency override to put the fins in their upright position, the ground station will also aid in the conservation of power prior to launch. By remotely signaling the payload on the rocket to go from the low power mode to standby, lots of energy can be conserved in the time leading up to the launch. The rocket can sit on the pad for an extended period of time using a minimal amount battery power, and only be switched to full power immediately prior to launch. Radio communication between the ground station and occurs in the license-free ISM band of MHz. This eliminates the need for any sort of HAM radio certification to communicate remotely with the rocket. With the proper antenna configuration, the ground station radio should have a reliable range of communication of at least 1.2 miles. The ground station control panel consists of an LCD, several LEDs, switches, and a button. A model of the ground station structure is shown in Figure

93 Testing Figure CAD Rendering of Ground Station To ensure reliability of the payload, extensive testing methods will be carried out on every part of the payload. Ground testing will be used to ensure reliability of the batteries and electronic control system. Ground tests will also be used to pass data into the sensors in order to test the reliability of the algorithm to ensure correct activation of the servomotor. Similar tests will be conducted to make sure that the sensor package and communications systems are working properly, along with tests to ensure that the payload can be enabled remotely. Further, to test the functionality of the payload, a post-burnout simulation has been created in Matlab to verify that the system will work quickly enough to complete the requirements before apogee. Below is a graph of the rocket s altitude and rotational position as functions of time simulated after burnout up to apogee. 92

94 Manufacturing and Assembly Figure Graph of Altitude and Rotation Angle as Functions of Time The Roll-Control Payload will be manufactured using a variety of machines and techniques. The unique bays and some internal parts will be printed on an engineering-grade Fortus 3D printer owned by Notre Dame s College of Engineering, which gives a much higher precision versus consumer-level machines. The control fins will be 3D printed as well, using a Markforged Mark Two, a new machine designed to print carbon fiber at production quality, through a collaboration with local engineering firm Springboard Engineering. The custom circuit board will be outsourced due to complexity and time constraints, but then all the electronics and wiring will be installed on campus. Finally, all the aluminum components will be machined from raw stock using a manual mill and lathe, as well as various other power tools. The bays interlock for easy assembly and disassembly, and the wiring that runs through the payload will have connectors at each bay so they can be separated as well. Permanent assembly items will be glued in place with high-strength epoxy or other adhesives, and the removable components will be assembled using nut-and-bolt fasteners, as well as thread locker compounds where needed Integration Plan External Structure The payload is secured to the rest of the rocket via four threaded metal rods that run through the entirety of the payload. These rods are epoxied perpendicular to the top of the fin can, and the payload will slide onto the rods. The rods are secured to the payload with washers and nuts on the top bulkhead of the payload. The payload is held to the parachute bay by shear pins. The payload will be utilizing the rocket s common ground for the grounding of all electronics in the payload Internal Structure The internal structural integration of the Roll Control Payload consists of 3D printed bays and their respective connections. The four bays are: the Battery Bay, the Electronics Bay, the Arduino Bay and the Servomotor and Gearbox Bay. Each bay is held in place through the metal stringers and compression from the nuts at either end. The 93

95 fins will connect to the payload in the Servomotor and Gearbox Bay. Holes cut into the coupler and the side of the rocket will accommodate the fin rods Instrumentation Precision and Repeatability The breakout board being used for measurements in this payload is the Adifruit 10DOF breakout board. This board combines a 3-axis gyroscope, 3-axis compass, a 3-axis accelerometer, and a barometric pressure/temperature sensor. Each senor can be set to the desired sensitivity, the gyroscope can be set to a ±250, ±500, or ±2000 degree-persecond scale, the compass can be set to a ±1.3 to ±8.1 gauss magnetic field scale, and the accelerometer can be set to a ±2g/±4g/±8g/±16g selectable scale. The pressure sensor has a range of hPa, and a resolution of 0.17m. Since this payload is focusing on completing two complete rotations the scale will be selected as to ensure that each time at least two complete rotations can be accurately recorded. This should make repeating the same measurements fairly easy for this task Payload Electronics Electronic Components A transistor circuit is used to control the turning of the secondary, roll-inducing fins. During launch, a +5V signal will be sent across the enable control of the motor and a +5V signal will be sent across Input B, keeping the servo in the home position and the secondary fins at a vertical 0 degrees. When the Arduino detects it has reached burnout, or receives the appropriate signal from the ground station, its resulting output will bias one of the two NPN transistors and switch Input B to +0V. Additionally, Input A will then toggle between +5V and +0V, causing the fins to rotate to +4 degrees and then -4 degrees relative to vertical based on gyroscope input. The ClearPath CMP-MCPV- RLN servo motor is rated for an input voltage range of 24-75VDC so the 44.4V provided by a pair of batteries will be adequate Arduino Roll Calculations To determine when and how to cant the fins, the Arduino calculates the number of completed revolutions by applying Simpson's Rule for numerical integration to rotational velocity data from the onboard gyroscope. This data is recorded on an SD Card mounted on the Arduino for post-flight analysis, and any actions the algorithm takes are communicated to a ground station using a long-range packet radio transceiver. Once the Arduino detects burnout, either through accelerometer data or using a backup timer system, it sends a digital signal to the servo to cant the fins in a direction favorable to any roll the rocket is experiencing at the time. Then, once the Arduino determines that two full revolutions have occurred, it cants the fins to one of two preset locations to counteract any rocket roll above a minimum detectable threshold Power Management Methods Power consumption will be minimized and safety maximized by connecting the roll control payload to a series of switches on the body of the rocket in addition to a ground station. The physical switches will control the connections between the Arduino, servomotor, and their respective batteries. One of the switches will also send a digital signal directly to the motor to turn the fins back to their home position of 0 degrees from vertical. The ground station will implement this same override capability, preventing the rocket from rolling prior to burnout. This is a safeguard against the failure of the Arduino control system Payload Safety and Failure Analysis Roll Controller Failure Modes and Mitigation 94

96 Table 5.1. Roll Controller Failure Safety Table Hazard Cause Effect(s) Likelihood Severity Risk Controller Malfunction: Fins canted throughout flight or not in response to postburnout roll Control fins breaking off rocket Power failure Ground station communicati on failure Unstable drag profile Arduino/cont rol algorithm error Excessive force applied to fins during ascent Battery depletion during/befor e flight Transceiver error Improper initial fin Increased/ continuous rocket roll throughout flight Uneven drag distribution on rocket No control system function, no experimental data collected Payload unable to be turned on remotely; potential to cause complete payload nonfunctionalit y if not addressed Rocket could lose D 2 Moderate D 2 Moderate D 2 Moderate D 2 Moderate E 1 Low Controls/Mitigati on Repeatedly test/simulate control algorithm and electronics system to ensure functionality Structural testing/calculatio ns; cant angle will be sufficiently small to prevent significant shear force on fins Insure batteries have sufficient charge to exceed duration of launch preparation and flight during ground testing Test communication between payload and ground station to ensure functionality; ground station will have indicator light to determine if communication with the payload is occurring. Repeat testing to ensure 95

97 from fin deployment alignment; broken teeth in fin-canting bevel gear system stability and veer out of control, potentially impeding proper parachute deployment redundancy; fins controlled by bevel gear system such that all have the same cant angle Control structural failure Design lacks necessary robustness to handle takeoff forces Damage to payload, loss of data, failure of entire superstructure D 2 Moderate Controller will be structurally contained, failure will not affect data, isolation from other payloads Payload Concept Features and Definition Creativity and Originality The creativity of the Roll Control Payload lies in its ability to utilize passive aerodynamic forces and therefore eliminate the need for a complex mechanical system. The roll is induced by a single servomotor rotating fins located aft of the post burnout center of gravity of the rocket. This increases the computations necessary for determining the angle that the flaps need to rotate to induce the roll, but also limits the need for a large power source and additional motors Significance The Roll Control Payload is designed to induce a roll in the rocket and to collect data on the aerodynamic forces present on a rocket that is rotating post burnout. Without this payload, there would be no way to induce roll in the rocket, or prevent the rocket from rolling if needed. The Roll Control Payload allows for roll induction by measuring the forces on and the position of the rocket. Using this information, it uses the roll fins to increase or decrease angular momentum as necessary. The accelerometer and altimeter data will also be useful in the study of the rotational forces present on the rocket Challenge One of the challenges the team faced when designing the ground station was in trying to create a custom PCB board. The schematic of the board was created using Eagle Light Software and then converted into g-code instructions for a CNC mill. However, as the board was being milled, the team realized that some of the traces on the PCB were spaced too closely together to cleanly etch each of the traces without any overlap. Unfortunately, because of the configuration of the board, the traces could not be arranged in a way that avoided overlap without making them too narrow, and therefore compromising them. In addition, the size of the copper pads where the components of the board would be soldered were also too small, and could not be expanded due to the configuration of the board. Due to these hindrances, the team decided to order a custom PCB online. This decision was made due to the relatively inexpensive price of custom PCBs, and to save the extensive time and labor that would have been required to make our milled PCB useable. 96

98 Additionally, due to weight concerns, the team decided to print the roll fins and connection rods as one piece using a carbon fiber 3D printer. This innovative tool allows for the reduction of weight in using carbon fiber instead of steel, without compromising strength, which would have occurred if using a standard plastic 3D printer. Finally, with the powerful capabilities of the Adafruit Feather microcontroller, it was decided that a ground station could be implemented with the payload as well. While not a requirement for the payload, a ground station allows for increased team safety, and also provides the team with real-time information regarding altitude and GPS location for recovery, as well as status indicators to ensure that the payload is functioning correctly on the pad Science Value Objectives The objective of the Roll Control Payload is to perform an aerodynamic analysis on the post-burnout rotating rocket in order to perform a sequence of controlled rolls and then minimize rotation before the rocket reaches apogee Success Criteria The Roll Control Payload will be considered a success if the following conditions are met: Rocket completes two full revolutions in a fin cant-induced direction during post-burnout ascent, and then does not experience any significant roll until the recovery system is deployed. The payload performs its tasks without causing unstable flight or failure of the rocket. The payload does not experience any failures during flight preventing success of its desired mission, or preventing reuse in subsequent flight. 5.2 Fragile Object Protection Payload Payload Objective The objective of the Fragile Object Payload is to safely carry and protect the provided unknown object throughout the rocket s flight. The payload must be able to contain the object no matter its size or shape, due to the fact that the object is unknown until competition day, and ensure that it is not damaged during the rocket s liftoff, flight, parachute deployment, and landing. The role of the viscous fluid will be to suspend the fragile object without imparting additional stress to the object from inertial acceleration. Additionally, a fluid, rather than solid padding, can morph its shape to support an object of any size within the constraints. A successful experiment, thus, would include the intact,undamaged survival of the object, i.e., the object is still functional or can serve its intended purpose. In order to be classified as reusable, there must be a lack of leakage from the viscous fluid to the exterior cavity or out of this payload. The primary vibrational control will be accomplished by the springs and a foam tube. The springs above and below the inner cylinder are designed to dissipate the axial forces imparted by the rocket during flight as intended, and the foam tube is designed to dissipate the radial forces. If both perform as anticipated, the payload has a high chance of surviving unharmed. For a successful implementation of the payload, the inner cylinder and springs effectiveness should not be dependent on the stability of the rocket. Finally, a successful launch would also render the payload able to be reused without the need for any modifications or repairs. This payload offers great working knowledge of vibrational dissipation, a common obstacle in many engineering systems, especially those related to Aerospace Engineering. The Fragile Objects Protection payload is not unlike safety measures that need to be taken by space-faring rockets carrying extremely delicate equipment during much more ferocious flight vibrations Payload Design The Fragile Object Payload design will utilize a combination of a spring and high density foam system and high-viscosity fluid protection. The payload will consist of two cylindrical shells, separated by a foam tube insert. The 97

99 inner shell will be filled with a fluid of relatively high viscosity to protect the unknown fragile object through suspension and some additional cushioning. The springs separating the two cylinders will dissipate the axial forces of takeoff and landing while the foam dissipates the radial forces. This design is the leading choice because it covers the two most important aspects of this payload: impulse mitigation and the ability to hold an object of unknown size and material. A CAD model of the design can be found in Figure 3.1. Figure Fragile Object Protection Payload Design Unlike other payloads, the precision of instrumentation of the FOP payload is not directly applicable due to a general lack of instrumentation. Nonetheless, the data output from the payload can be analyzed as a binary outcome system - the fragile object will either return intact or damaged. Due to the fact that it is a binary set of potential outputs, no extra precision needed to measure its performance. As for analyzing the repeatability of measurement, it is once again difficult to specifically qualify for the FOP payload. This is because, technically speaking, the payload will be at 98

100 an activated state at the moment that the object is inserted and the inner cylinder is sealed. From that point on, the payload will be continuously working to protect its contents, regardless of whether the payload is even loaded into the rocket or not. This makes defining the measurement period difficult because two separate launches can be viewed as a single run of the payload if it is not disassembled between the two launches. Despite this, we intend the payload to be highly reusable seeing as there is nothing that is expended throughout its use and with minimal moving parts and electronics use, the payload should be reusable to the extent of the lifetime of its components Testing Testing Plan Tests were carried out with different high-viscosity fluids in hopes of determining the best-suited fluid to fill the inner cylinder of the payload. The viscous fluid must be able to suspend and cushion the fragile object by holding it in place, but it must also be able to allow for movement to prevent stress build-up. The full test procedures can be seen in Appendix E. To carry out the test, a transparent container was filled with a fluid, and then different objects were placed in the container to observe how the viscosity of the fluid mitigated movement. Objects to be tested Fluids to be tested (approximate viscosity) - Golf Ball - Agave (1000 centipoise) - Clementine - Corn Syrup (2000 centipoise) - Ping Pong Ball - Black Strap Molasses (6000 centipoise) - Steel Ball Testing Analysis The agave offered little resistance to either the buoyancy or weight of the object, with all of the objects either rising to the surface or sinking to the bottom rather quickly based on their relative density. A screenshot of the golf ball in the agave can be seen in Figure 5.13 below. 99

101 Figure Golf Ball Floating in Agave The blackstrap molasses had the most positive results, offering a measurable resistance to the effects of the both forces. The golf ball and the clementine took several seconds to rise to the surface of the fluid after being submerged to the bottom of the container. The corn syrup also offered a better result than the agave, by resisting for several seconds against the buoyant force exerted on the clementine and the golf ball. The ping pong ball was essentially useless as it floated in every liquid and returned to the surface too fast to measure. None of the fluids performed as hoped in order to suspend the objects for a reasonable amount of time. The effects of each fluid on the steel ball were also minimal, with the ball sinking to the bottom of the container taking almost the same amount of time, regardless of the fluid that was filling up the container. However, the steel ball Is a robust and resistant object, unlikely to exhibit the behavior of a fragile object. The most successful combination was the golf ball in the blackstrap molasses as seen in Figure 5.14, taking approximately 9 seconds to travel through 3 inches of the molasses. 100

102 Figure Golf Ball Floating in Blackstrap Molasses After the final test of the planned objects and failing to see a desired result, it was noted that all of the test objects were spheres or nearly spherical. This is important because spheres provide minimal surface area for drag to develop compared to the potential weight of the sphere. Therefore, for an object of an average density, the sphere may be the worst possible shape to achieve suspension within the fluid. To test this hypothesis that the shape had a significant factor on the behavior of the objects, an ID card was subsequently submerged in the corn syrup. The ID card remarkably stayed suspended in the corn syrup and showed no movement visible to the naked eye as can be seen in a screenshot from the video shown in Figure When looking at footage of the ID card that was over a minute long, it could be seen that the ID card was in fact moving, just very slowly. This outcome proved how important the shape of the eventual fragile object is in the success of a viscous fluid at suspending the object during the time on the pad. It also hints that if a viscous fluid has a high enough viscosity to suspend a sphere, it will likely be able to suspend any other shape that the fragile object may take on. 101

103 Figure An ID Card Suspended in Corn Syrup The results of the tests were encouraging with respect to the black strap molasses and the corn syrup; these fluids showed observable resistance to the forces exerted on the objects, which will be necessary to protect the fragile object enclosed by restricting its movement while under the influence of only gravity. The agave was not effective, offering minimal resistance to the forces on the object. Out of the fluids tested so far, the black strap molasses had a slight edge by offering a higher resistance to the mentioned forces. In the coming weeks, further testing will be carried out with objects that do not display the homogeneity seen in these tests with respect to the shapes of the objects. While the tests offered valuable insight of the effects of the viscous fluid on objects of different densities, the tests were limited to objects of mostly spherical shape. The difference that the shape of an object might make was exemplified by the effects of the viscous fluid on the ID card, which were completely different from the effects on the family of spherical objects that were used throughout the testing. Because the shape of the fragile object will be unknown until the day of the launch, this aspect must be accounted for. From these results, higher viscosity fluids will be tested as well. Preliminary candidates include things such as peanut butter or petroleum jelly which have estimated viscosities of up to between 6 and 10 times the dynamic viscosity of even blackstrap molasses. Another possible option that might be pursued are fluids that display shearthinning properties such that under standard conditions of 1 G, the fluid acts more like a solid, but under higher shear forces, it thins out and flows with lower viscosity. These types of fluids are also sometimes known as pseudoplastics with items such as ketchup, paint, and whipped cream being examples. As further testing is completed, it will be 102

104 important to note the density of the final fluid because if whipped cream, for example, is chosen as the viscous fluid, then there the springs may be stiffer than they need to be for the protection of the fragile object Payload Components Foam Insert In reviewing the design put forth in the PDR, the overall vibration mitigation in the radial direction appeared sufficient by using an array of springs. As plans began to develop concerning the testing of the springs, it became evident that the manufacturing of the payload might prove to be extremely difficult. This was most notable for the radial springs because the original design called for the springs to be welded to the inner canister. Due to the limited space between the inner canister and the phenolic tubing, it would become extremely difficult to assemble the payload reliably using radial springs. The solution to this manufacturing hurdle came from a failed suggestion for the object cushioning - foam. In this instance the disadvantages associated with foam when it was previously explored are not as important for the radial cushioning as the shape of the object that it will be used to protect is known. There are several other advantages that the foam offers. In addition to being easier to assemble than springs, it offers a more uniform cushion because the foam can be in contact with the entire length of the inner canister compared to a small number of points for the spring design. Additionally, the opening in the foam area will also be able to act as a guide track to restrict a majority of the movement to the axial direction. Due to the fact that the foam does not have to be directly connected to the inner cylinder, the foam has the design flexibility to run the length of the payload if desired. Though this is possible, it is rather unnecessary and would interfere with the blocks that are securing the Fragile Object Protection payload to the nosecone of the rocket. As such, it has been decided that the top of the foam will be flush with the top of the inner cylinder when all of the springs are uncompressed. The bottom of the foam will be an inch or two below where the bottom of the inner cylinder is expected to be under standard gravity conditions, as it will be experiencing while sitting on the launch pad prior to ignition. Once it was decided to utilize foam to protect against radial displacement for the Fragile Objects Payload, the exact type of foam needed to be determined. Foam is characterized by density and firmness. Unlike most other everyday objects, these two characteristics are not always correlated such that a denser foam will be more firm. This allows for a great variety of options when researching foam parts. Foams typically have a density range between 1 lb/ft 3 and 3 lb/ft 3, although there are some higher density foams as well. In order to manage the weight of the Fragile Objects Protection payload, the least dense foam with the necessary firmness was chosen. The firmness of foam is quantified by a measurement called the Indentation Load Deflection test (ILD). This resulting ILD value comes from the force required to deform the foam to ¾ of its original thickness. The maximum anticipated force in the radial direction is expected to occur at landing as the rest of the flight profile has accelerations almost entirely in the vertical direction. Based off of this peak acceleration in the radial direction it has been estimated that the ILD required should be approximately 15. Table 5.2 shown below is from courtesy of foamonline.com and highlights the different combinations of density and firmness. 103

105 Table 5.2. Available ILD Values for Various Foam Densities Based on this chart, the foam chosen was a high density foam with an ILD of 18. The foam is ordered in a 5.3 inch diameter cylinder with a length of 6.7 inches. The center hole will be cut for the inner cylinder and the 5.3 inch outer diameter will be trimmed as needed to exactly match the inner diameter of the coupler Springs Since the springs are to be the primary source of impulse mitigation in this payload, it is essential that they are chosen to appropriately react to the accelerations and forces that are anticipated. While all three motors will provide a different acceleration and thrust profile, as a whole they are rather similar overall. Each one provides thrust on the order of 10 g s of acceleration for a few seconds. After burnout, the g-force experienced by the rocket as a whole is miniscule by comparison. That is until the Main parachute deployment, which can produce a momentary spike of upwards of 30 g s. An estimate of the full acceleration profile that the rocket will experience while using the AeroTech L1150 can be seen in Figure Although this is not the current engine, it has very similar characteristics and therefore it is sufficient for the approximation of springs that are needed for this payload. This figure shows the magnitude of the total acceleration although most of the data points are almost completely in the vertical direction. For example the spike at main parachute deployment is 99.98% in the vertical direction. With the updated design of two springs on the top and bottom, calculations show that the spring constant needed to dissipate the spike at main parachute deployment is 20 lbs/inch given the available space for deflection. The current spring to be used for this payload is the LC 072GG 13S spring from Lee Spring. It is a stainless steel spring with a free length of 3 inches. The spring constant is lb/in and has an outer diameter of in. As stated previously, the springs formally protecting against vibration in the radial direction have been replaced by a foam tube insert. 104

106 Figure Acceleration vs. Time from OpenRocket Simulation On the above plot, the two red circles signify apogee and contact with the ground. As can be seen from this plot, the motor ignition through burnout and the main parachute deployment posses the largest threats to imparting large impulses on the payload Inner Cylinder As stated previously, an inner canister will be responsible for directly holding the viscous fluid and the fragile object. The important characteristics of the canister is that it is water sealed and that it can withstand the axial forces present during flight. The current canister design is made out of two ground coffee cans that have been cut and connected together as can be seen in Figure In Figure 5.18, the canister can be seen with the top screw top detached. These coffee cans are made out of galvanized tin with walls that are approximately inches thick. Although having a custom-made cylinder would be ideal, these cans had the desired screw top and were the diameter consistent with our design. Therefore it was more economical to create the inner canister this way. Before utilizing this in our final design, tests will be run to ensure that the can can structurally withstand the anticipated forces. Additionally, the cap on the bottom will be epoxied closed to ensure a watertight seal. 105

107 Figure Inner Cylinder 106

108 5.2.5 Requirements Verification Figure Inner Cylinder with Cap Removed The NASA Student Launch Handbook sets forth requirements necessary to achieve complete this payload. The requirements and the mitigations incorporated into the design can be seen below in Table 5.3. Table 5.3. Fragile Object Protection Requirements Verification Requirement Challenge Design mitigation Teams shall design a container capable of protecting an object of an unknown material and of unknown size and shape There may be multiple of the object, but all copies shall be exact Cannot design for a specific shape object Single design must work for all objects Any open cavity design could allow for - Using a viscous fluid can form to any given shape - Robust design that is simple enough to accommodate all objects A viscous fluid that is capable of suspending 107

109 replicas The object(s) shall survive throughout the entirety of the flight Teams shall be given the object(s) at the team check in table on launch day 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 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. multiple objects colliding with each other causing damage Payload must be able to handle all aspects of launch including: ignition, burnout, main parachute deployment, etc. The design cannot be tailored to any specific object prior to launch day All of the year-long design must accommodate all objects imaginable Cannot add more support if design does not adequately hold object on launch day Makes the idea of airbags questionable to whether it would break requirements or not The object can be quite large, but still has the opportunity to be very small as well The object may be rather heavy in comparison to some proposed solutions objects, they could be left far apart to reduce likelihood of collision The springs should diminish the vibrations experienced through all of the flight phases - The viscous fluid gives a great degree of flexibility in the ability to hold any shape. - Most versatile approach for launch day - Viscous fluid will fill entire inner container before object is added - If anything, viscous fluid will be removed from inner cylinder through displacement which will not hinder payload - Make the cylinder not imaginary and support that cylinder. - Ensures fitting max size object The fluid itself is already heavy and the 4 oz will not be large percentage of payload weight 108

110 6 Project Plan 6.1 Testing and Requirement Verification Plans Test plans and Requirement Verfification Plans can be found in the sections devoted to the individual systems and subsystems they pertain to. 6.2 Budget The Budget Plan for the team and each individual squad is attached in Appendix N. The first section of the budget plan details the overall team budget. It includes the subtotal of each squad s budget, the budget allocated to educational outreach events, and the travel budget. This year s budget plan totals to just over $12,500. Subsequent sections go into detail of each squad s budget plan. All materials necessary in designing and building the launch vehicle and its subsystems are outlined in these sections along with their estimated costs. Please refer to Appendix N for more details of NDRT s Budget Plan. 6.3 Funding Plan Funding will be provided by the Aerospace and Mechanical Engineering (AME) Department of Notre Dame. It is drawn from a general account dedicated to aerospace design projects. Funding is secured by submitting a business plan and projected budget to a review board. The board reviews these proposed documents and decides if the design project merits funding from the department. After submitting these documents, NDRT was able to garner funding for their proposed budget plan. In addition to the funding from the AME Department, NDRT has also received a very generous sponsorship from the Boeing Company that will offset a large amount of the cost of the competition. NDRT is also considering other funding sources like the College of Engineering for educational outreach events. This additional funding may allow NDRT a broader reach to the local community and may positively affect the quality of the events hosted by the team. 6.4 Timeline A GANTT Chart showing the team s timeline can be found in Appendix M. So far, the team has kept up with the timeline. The one exception is the subscale launch, which was delayed by one week due to weather on the scheduled launch day 109

111 Appendix A. CAD Drawings Appendix A.1: Full Scale Model Dimensions 110

112 Appendix A.2: Full Scale Model Exploded View 111

113 Appendix A.3: Section I View 112

114 Appendix A.4: Section III View 113

115 Appendix A.5: Nose Cone 114

116 Appendix A.6: FOP Coupler Appendix A.7: RCP Coupler 115

117 Appendix A.8: Roll Control Payload 116

118 Appendix A.9: Motor Mount 117

119 Appendix A.10: Fin Can 118

120 Appendix A.11: Fin Design 119

121 Appendix A.12: Centering Ring Appendix A.13: Bulkhead 120

122 Appendix B. Performance Prediction Programs import math def apogee(m_r, m_e, m_p, p, Cd_t, Cd_c, A, T, g, t): """ m_r : rocket mass [M] m_e : engine mass [M] m_p : propellant mass [M] p : air density [M/L^3] Cd_t : drag coefficient during thrust phase Cd_c1 : drag coefficient during first coasting phase Cd_r : drag coefficient during first coasting phase Cd_c1 : drag coefficient during second coasting phase A : rocket cross sectional area [L^2] T : thrust [F] g : acceleration due to gravity [L/T^2] t : burnout motor time [T] """ Cd_r = Cd_c *math.sin(4*math.pi/180) phase) phase) phase) k_t =.5 * p * Cd_t * A #aerodynamic drag coefficient [M/L] (thrust k_r =.5 * p * Cd_r * A #aerodynamic drag coefficient [M/L] (roll k_c =.5 * p * Cd_c * A #aerodynamic drag coefficient [M/L] (coast 1 # thrust m_a = m_r + m_e - m_p/2 #average mass [M] q1 = math.sqrt((t - m_a*g) / k_t) #burnout velocity coefficient [L/T] x1 = (2 * k_t * q1) / m_a # burnout velocity decay coefficient [1/T] v1 = q1 * ((1-math.exp(-x1*t)/(1+math.exp(-x1*t)))) #burnout velocity [L/T] y1 = (-m_a/(2*k_t)) * math.log((t-m_a*g-k_t*v1**2)/(t-m_a*g)) #altitude burnout [L] # roll m_c = m_r + m_e - m_p #coasting mass [M] qr = math.sqrt((t - m_c*g) / k_r) #burnout velocity coefficient [L/T] xr = (2 * k_r * q1) / m_c # burnout velocity decay coefficient [1/T] vr = q1 * ((1-math.exp(-x1*t)/(1+math.exp(-x1*t)))) #burnout velocity [L/T] yr = (-m_c/(2*k_r)) * math.log((t-m_c*g-k_r*v1**2)/(t-m_c*g)) #altitude burnout [L] print(yr) 121

123 # coast qc = math.sqrt((t - m_c*g) / k_c) #burnout velocity coefficient [L/T] xc = (2 * k_c * q1) / m_c # burnout velocity decay coefficient [1/T] vc = q1 * ((1-math.exp(-x1*t)/(1+math.exp(-x1*t)))) #burnout velocity [L/T] yc = (-m_c/(2*k_c)) * math.log((t-m_c*g-k_c*v1**2)/(t-m_c*g)) #altitude burnout [L] print(yc) peak_altitude = y1 + yr* yc*.625 return peak_altitude*1.72 rocky = apogee(.418,.0446,.0197, 1.225, , , , 40.36, 9.81, 1.03) ferg = apogee(.425,.0446,.0197, 1.225, , , , 40.36, 9.81, 1.03) full_scale = apogee( , 3.538, 1.839, 1.225, , , , , 9.81, 2.6) print(rocky) print(ferg) print(full_scale) 122

124 Appendix C. Stability Prediction Program ## import math in] def CP(Ln, d, df, dr, Lt, Xp, Cr, Ct, S, Lf, R, Xr, Xb, N): """ Ln : length of nose [13 in] d : diameter at base of nose [5.5 in] df : diameter at front of transition [5.5 in] dr : diameter at rear of transition [5.5 in] Lt : length of transition [0 in] Xp : distance from tip of nose to front of transition [0 in] Cr : fin root chord [7 in] Ct : fin tip chord [7 in] S : fin semispan [6.2 in] Lf : length of fin mid-chord line [6 in] R : radius of body at aft end [5.5 in] Xr : distance between fin root leading edge and fin tip leading edge parallel to body in] Xb : distance from nose tip to fin root chord leading edge [95.75 N : number of fins [4] """ # nose cone terms Cn_n = 2 ## Xn =.666*Ln #for cone Xn =.466*Ln #for ogive # conical transition terms Cn_t = 2 * ((dr/2)**2 - (df/2)**2) 123

125 Xt = Xr #Xp + (Lt/3) * (1 + (1-(df/dr)) )#/ (1-(df/dr)**2)) # fin terms Cn_f = (1 + (R/(S+R)) * ((4*N*(S/d)**2) / \ (1 + math.sqrt(1 + ((2*Lf)/(Cr+Ct))**2)))) Xf = Xb + ((Xr*(Cr+2*Ct))/(3*(Cr+Ct))) + \ (1/6)*((Cr+Ct)-((Cr*Ct)/(Cr+Ct))) # center of pressure calculation Cn_r = Cn_n + Cn_t + Cn_f #sum of coefficients CP = (Cn_n*Xn + Cn_t*Xt + Cn_f*Xf) / Cn_r #CoP d from nose tip return CP def stability(d, CG, CP): """ d: diameter of rocket (in) CG: center of gravity (in from nose ogive) CP: center of pressure (in from nose ogive) """ stability = (CP-CG)/d return stability CG1 = CP1 = CP(13,5.5,5.5,5.5,0,0,7,7,6.2,6,5.5,7.312,95.75,4) rocket_1 = stability(5.5, CG1,CP1) print(rocket_1) 124

126 125

127 Appendix D. Wind Tunnel Testing MATLAB Codes Calibration Code % Notre Dame Rocketry Team % 02 Jan 2017 % Wind Tunnel Data Analysis % Author: PATRICK NTWARI % Calibration clc clear all close all % masses used for calibration mass = [(0:50:200),(300:100:700)]; % convert mass into force in Newtons force = mass*9.81*0.001; % recorded drag and lift voltages drag_cal = [ ]; lift_cal = [ ]; % drag and lift fit equations drag_fit = fit(drag_cal(:),force(:),'poly1'); lift_fit = fit(lift_cal(:),force(:),'poly1'); figure(1) plot(drag_cal,force,'ro',... drag_cal, 3.29*drag_cal+25.1,'r-') xlabel('voltage') ylabel('force (N)') title('drag Transducer Calibration'); legend('drag Experimental', 'Line of Best Fit') figure(2) plot(lift_cal,force,'ro',... lift_cal, -45.4*lift_cal+74.7,'r-') xlabel('voltage') ylabel('force (N)') title('lift Transducer Calibration'); legend('lift Experimental', 'Line of Best Fit') C D Without Fins % Notre Dame Rocketry Team % 02 Jan

128 % Wind Tunnel Data Analysis % Author: PATRICK NTWARI % Drag Calculation_Sans Fin [_SF] clc close all %% Dimensions of the Vehicle diam = ; % in meters h = 0.545; % in meters L = 0.748; %area = diam*pi*h+pi*diam*0.5*h; area = diam*diam*0.25*pi; % cross-section % Pressure Cal kp = ; %% Imports from calibration drag_m = drag_fit.p1; drag_b = 0; %drag_fit.p2; %% Drag Data Loading and Calculation load('ndrt_0.mat'); d_off = mean(y_data(:,3)); q_off = mean(y_data(:,1)); load('ndrt_5.mat'); d_sf_5 = mean(y_data(:,3)); D_sf_5 = drag_m*(d_sf_5-d_off)-d_p5; q_sf_5 = mean(y_data(:,1)); Q_sf_5 = kp*(q_sf_5-q_off); cd_sf_5 = D_sf_5/Q_sf_5/area; v_sf_5 = sqrt(q_sf_5/0.5/rho); load('ndrt_10.mat'); d_sf_10 = mean(y_data(:,3)); D_sf_10 = drag_m*(d_sf_10-d_off)-d_p10; q_sf_10 = mean(y_data(:,1)); Q_sf_10 = kp*(q_sf_10-q_off); cd_sf_10 = D_sf_10/Q_sf_10/area; v_sf_10 = sqrt(q_sf_10/0.5/rho); load('ndrt_15.mat'); d_sf_15 = mean(y_data(:,3)); D_sf_15 = drag_m*(d_sf_15-d_off)-d_p15; q_sf_15 = mean(y_data(:,1)); 127

129 Q_sf_15 = kp*(q_sf_15-q_off); cd_sf_15 = D_sf_15/Q_sf_15/area; v_sf_15 = sqrt(q_sf_15/0.5/rho); load('ndrt_20.mat'); d_sf_20 = mean(y_data(:,3)); D_sf_20 = drag_m*(d_sf_20-d_off)-d_p20; q_sf_20 = mean(y_data(:,1)); Q_sf_20 = kp*(q_sf_20-q_off); cd_sf_20 = D_sf_20/Q_sf_20/area; v_sf_20 = sqrt(q_sf_20/0.5/rho); load('ndrt_25.mat'); d_sf_25 = mean(y_data(:,3)); D_sf_25 = drag_m*(d_sf_25-d_off)-d_p25; q_sf_25 = mean(y_data(:,1)); Q_sf_25 = kp*(q_sf_25-q_off); cd_sf_25 = D_sf_25/Q_sf_25/area; v_sf_25 = sqrt(q_sf_25/0.5/rho); load('ndrt_30.mat'); d_sf_30 = mean(y_data(:,3)); D_sf_30 = drag_m*(d_sf_30-d_off)-d_p30; q_sf_30 = mean(y_data(:,1)); Q_sf_30 = kp*(q_sf_30-q_off); cd_sf_30 = D_sf_30/Q_sf_30/area; v_sf_30 = sqrt(q_sf_30/0.5/rho); cd = [cd_sf_10 cd_sf_15 cd_sf_20 cd_sf_25 cd_sf_30]; vel = [v_sf_10 v_sf_15 v_sf_20 v_sf_25 v_sf_30]; Reynolds = vel*rho*l/nu; % From figure it is evident that the first data point is too much an % outlier. figure(1) plot(reynolds, cd,'ro') xlabel('reynolds No.') ylabel('coefficient of drag') title('cd vs. Reynolds No. tested without fins') 128

130 Appendix E. Fragile Object Protection Testing High-Viscosity Fluid Testing Materials: o Corn syrup o Molasses o Honey o Clear container o Golf ball o Ping Pong Ball o Clementine o 1 Stainless balls o Camera o stopwatch/phones Set-Up: o Perform each fluid fully before moving on to next fluid Tape a ruler to the clear container for reference when filling with fluid and measuring distances an object moved over a certain time Fill the clear container to the same marked height of 3 inches with the three fluids o Stabilize and focus the camera pointing towards the container Procedure: o Begin video recording o Hold the object at the surface of the fluid so that it is just touching o Release the object and simultaneously start the stopwatch o Stop both stopwatch and camera filming after object stops moving o If object fails to submerge fully, place at bottom of container record time to resurface o Remove object and clean it off o Repeat steps for all fluid and object types, filling out the results table accordingly Fluid Golf ball; g Clementine 74g Steel ball Ping pong; 2.7 g Agave Floated, but most of ball volume beneath surface; when forced under surface, shot back up to the top rather quickly Floated with about half object in and half out; when forced under surface, shot back up to the top rather quickly Immediately sank to the bottom; NA Black strap molasses Sank till mostly submerged with about ¼ of volume above surface; looked as though it was sinking slowly; when forced to when forced to bottom of container, rose back up and broke surface at 4.46 seconds Immediately sank to the bottom, but probably slower than agave Stayed at surface; when forced to bottom of container, rose back up in about 2 seconds; 129

131 bottom of container, rose back up and broke surface at 9 seconds creepy and gloopy and gross Corn syrup Sank till mostly submerged with about ¼ of volume above surface; looked as though it was sinking slowly; when forced to bottom of container, rose back up and broke surface at 4 seconds when forced to bottom of container, rose back up and broke surface at 3.5 seconds Immediately sank to the bottom, but probably quicker than agave Stayed at surface; when forced to bottom of container, rose back up in less than a second 130

132 Appendix G. Launch Procedure Checklists Launch Procedure Checklists Prior to Departure for Launch Site: Vehicle Sub-team Personnel Safety Concerns Ensure all safety material has been packed Items to bring Safety goggles Gloves Ensure lids to epoxy bottles are appropriately closed Vehicle components to bring: Nose cone Recovery body tube Fin can FOP coupler RCP coupler Shear pins Extra washers Extra nuts Extra screws Extra epoxy Inspect the body tubes and couplers to ensure they have not been damaged during storage. Ensure the items are stored in such manner as to not cause physical damage. Ensure the fin can is stored on the rocket holder so as not to damage the fins during transportation. Subteam Member: Date: Signature: Subteam Lead: Date: Signature: 131

133 Team Lead: Date: Signature: 132

134 Prior to Launch: Vehicles Sub-team Personnel Safety Concerns Ensure everyone operating on the vehicle has proper safety equipment: Safety goggles, gloves, respirator mask if carbon fiber cutting is involved Prepare the vehicle for launch: Section I Insert payloads into the top-most body tube Insert ballast into the top-most body tube if necessary Insert nose cone into the top-most body tube Ensure that the screws on the nose cone are not loose Friction fit the nose cone with masking tape or scotch tape if necessary Prepare the vehicle for launch: Section II Ensure CRAM core is inside the CRAM body Ensure the CRAM can be armed directly from the rocket s rail position Supervise the folding and packing of parachutes Prepare the vehicle for launch: Section III Insert ballast into fin can if necessary Insert payload into fin can Ensure the payload is stable within the fin can Attach rocket sections Check that all interfaces are aligned correctly Insert shear pins to secure each section Ensure the screws are tight Perform Cg test to ensure the center of gravity matches the simulated center of gravity. Ballast as necessary if Cg test fails to keep the stability margin. Prepare and insert the motor (Process performed by Team Mentor Dave Brunsting) Remove motor from packaging Check that motor is properly assembled according to manufacturer s instructions Remove pre-installed ejection charge 133

135 Properly dispose of black powder Insert motor into casing Ensure two spacers precede motor Screw on rear closure Insert motor into rocket Attach motor retainer Check for secure fit Check rocket stability (at least 2 calibers) and final weight Register with LCO and RSO at launch site. Ignite motor right before launch (Process performed by Team Mentor Dave Brunsting) Remove igniter clips from igniter Remove igniter from rocket Ensure igniter has properly exposed ends which are split apart Insert igniter into motor Attach clips to igniter, ensuring good contact Subteam Member: Date: Signature: Subteam Lead: Date: Signature: Team Lead: Date: Signature: 134

136 After Launch: Vehicle Sub-team Personnel Safety Concerns Instruct all personnel to get clearance before starting recovery process Assess there is no harmful physical damage before removal Ensure nothing is on fire Wait for at least 5 minutes before removing due to lingering motor heat Document state of rocket before removing Structural Integrity Check the physical state of the overall body tube Check the physical state of each payload Did the Roll Control Payload suffer damages to the electronic components Are the Roll Control Payload and FOP couplers structurally sound? Is the recovery body tube structurally sound? Ensure parachutes are re-usable Subteam Member: Date: Signature: Subteam Lead: Date: Signature: Team Lead: Date: Signature: 135

137 Prior to Departure for Launch Site: Recovery Personnel Safety Concerns Items to bring Safety goggles Electronics Items to bring CRAM Main parachute Drogue Parachute Shock cords Shear pins Extra batteries Talcum powder Ensure the items are stored in safe boxes at a reasonable temperature. Ensure all applicable electronics are turned off. Structural Integrity Ensure the recovery body tube has not been damaged during storage. Ensure the holes in the recovery body tube are the appropriate size Ensure the recovery body tube is clear of electronics before storage. Squad Member: Date: Signature: Team Lead: Date: Signature: 136

138 Prior to Launch: Recovery Personnel Safety Concerns Ensure everyone operating on the payload has proper safety equipment Safety goggles Prepare CRAM Insert fresh batteries into CRAM core Ensure batteries are connected to altimeters by listening for beeps from altimeters Insert CRAM core into CRAM body Put CRAM core cover on Tighten nuts down onto cover Insert long eyebolt through center of CRAM Place washer against both the bottom bulkhead and the CRAM cover Tighten nut against CRAM cover to hold bolt in place Connect the wires from CRAM core to screw terminals Attach short eyebolt to the long eyebolt with coupling nut Tighten nut on either side of coupling nut Electronics Prepare Avionics Mark the Primary Raven as official contest altimeter Ensure arming switch is safe Properly secure altimeters and batteries Install the CRAM until it locks Ensure the CRAM can be armed directly from the rocket s rail position Structural Integrity Prepare ejection charges Connect e-matches from altimeters to ejection charges Ensure personnel are wearing safety glasses Move all non-essential personnel away from rocket Connect electric matches/ejection charges to altimeter 137

139 Properly load and prepare parachutes Check that shroud lines are not tangled Apply talcum powder to each parachute Ensure that shock cord is not tangled Insert parachutes, chute protector, and shock cord into rocket Attach rocket sections Check that all interfaces are aligned correctly Insert shear pins to secure each section Ensure tight fit of all components Leave hatched door open Check shock cord for brittleness Replace shock cord that appears brittle Squad Member: Date: Signature: Team Lead: Date: Signature: 138

140 After Launch: Recovery Personnel Safety Concerns Instruct all personnel to get clearance before starting recovery process Assess there is no harmful physical damage before removal Ensure nothing is on fire Check that ejection charges have ignited Document state of rocket before removing Electronics Disarm altimeters Disconnect batteries Structural Integrity Check the physical state of the recovery body tube Is it re-usable? Check that components are safely inside the payload Squad Member: Date: Signature: Team Lead: Date: Signature: 139

141 Prior to Departure for Launch Site: Roll Control Payload Personnel Safety Concerns Items to bring: Safety Goggles Electronics Any Arduino connections must be soldered Any batteries must be unplugged to save power. Items to bring (as applicable): Soldering iron, with extra solder Spare batteries Electric tape Extra wire Wire crimpers Wireless Data Receiver GPS Receiver Ground Station Voltage Dividers Microcontroller Sensor Bay Transmitter Ensure the items are stored in safe boxes at a reasonable temperature. Ensure all applicable electronics are turned off. Structural Integrity Perform visual inspection to make sure outer surface has not been damaged during storage. Shake the fin can to ensure the payload components do not wiggle when shaken Squad Member: Date: Signature: Team Lead: Date: Signature: 140

142 141

143 Prior to Launch: Roll Control Payload Personnel Safety Concerns Items to bring: Safety goggles Electronics (as applicable) Before activating electronics, ensure that the time until launch does not exceed battery life. Ensure all connections are correctly soldered. Test all connections to verify there are no short circuits or faulty wiring. Turn on wireless data receiver Turn on GPS receiver Turn on ground station Verify connection between ground station and rocket Ensure all Molex connectors are attached to circuit boards Check that all wire connections are according to design Structural Integrity Perform visual inspection to make sure outer surface has not been damaged during transportation. Ensure all connections are tight and secure. Double-check that the drogue bulkhead is secure. Tighten the nuts on the parachute eye bolt Ensure there are no loose wires or solder Ensure all payload hardware components properly secured to sleds Ensure that the main parachute eyebolt is tight and the screws do not unscrew Perform a shake test to ensure that payload materials do not shift Squad Member: Date: Signature: Team Lead: Date: Signature: 142

144 After Launch: Roll Control Payload Personnel Safety Concerns Instruct all personnel not let the fin can safely land before approaching. Instruct all personnel not begin recovering the payload until given clearance by ground personnel. Ensure the fin can has adequately cooled before handling. Document the state in which the system is before any tampering Electronics Check that all electronics survived the flight intact. Power down ground station Turn off the three control switches on the body of the rocket Structural Integrity Perform visual inspection to make sure outer surface has not been damaged during flight. Assess any damages that may have occurred during operations. Determine if the damages are severe enough to prevent additional launches. Repair any minor damages, where possible. After recovery, re-perform component tests to ensure that operation has been uninhibited. Squad Member: Date: Signature: Team Lead: Date: Signature: 143

145 Prior to Departure for Launch Site: Fragile Object Protection Payload Personnel Safety Concerns Items to bring: Safety Goggles Items to bring: Extra amounts of high-viscosity fluid Ensure the items are stored in safe boxes at a reasonable temperature. Structural Integrity Perform visual inspection to make sure outer surface has not been damaged during storage. Shake the fin can to ensure the springs are still attached and mitigate the inner container s motion appropriately. Squad Member: Date: Signature: Team Lead: Date: Signature: 144

146 Prior to Launch: Fragile Object Protection Payload Personnel Safety Concerns Items to bring: Safety goggles Structural Integrity Perform visual inspection to make sure outer surface has not been damaged during transportation. Check Spring Connections to ensure proper connection Before assembling payload to fuselage: Fill inner canister with fluid Gently place Fragile Object into the inner container Wipe off any displaced high-viscosity fluid from the exterior of the inner container Screw on the top of the inner container Secure the two top springs to the lid of the inner container. Ensure all payload components are properly secured. Ensure all connections are tight and secure. Squad Member: Date: Signature: Team Lead: Date: Signature: 145

147 After Launch: Fragile Object Protection Payload Personnel Safety Concerns Instruct all personnel to let the fin can safely land before approaching. Instruct all personnel not begin recovering the payload until given clearance by ground personnel. Ensure the fin can has adequately cooled before handling. Document the state in which the system is before any tampering Structural Integrity Perform visual inspection to make sure outer surface has not been damaged during flight. Recover Fragile Object Assess any damages to the Fragile Object and record the state that it is in Assess any damages to entire payload that may have occurred during operations. Determine if the damages are severe enough to prevent additional launches. Repair any minor damages, where possible. After recovery, re-perform component tests to ensure that operation has been uninhibited. Squad Member: Date: Signature: Team Lead: Date: Signature: 146

148 Appendix H. Personnel Hazard Analysis Hazard Cause Effect(s) Severity Likelihood Risk Controls/Mitigations Hand and power tools used in assembling systems Improper use of tools due to lack of training; incorrect tool used for a job Minor to severe injury to self and/or others Critical Remote Moderate Train all team members in proper tool use. Ensure the proximity of assistance in case of injury. Flying sawdust chips or solder Loose parts can fly some distance when using tools Eye injury Critical Remote Moderate Ensure all team members working with power tools or soldering have eye protection. Exposure to ammonium perchlorate Handling rocket motor while ignited or hot from ignition Risk of burn or fire Critical Remote Moderate Treat burns quickly and have fire safety equipment ready. Avoid handling motors. Exposure to lead Ingesting solder Lead poisonin g Critical Improbable Low Wash hands after soldering. Avoid eating or drinking while soldering. Exposure to detonated black powder charges or residue Packing, handling, or cleaning black powder charges before or after launch Risk of fire, burns, or irritation of respirato ry system Marginal Probable Moderate Ensure proximity of fire safety equipment; ensure that eye and skin protection is used; wash exposed area thoroughly with water 147

149 Exposure to fumes Joining compone nts with epoxy; soldering Nausea, lightheadedn ess Marginal Occasional Moderate Ensure adequate ventilation when working with solder and epoxy Soldering iron Solderin g irons reach very high temperat ures First or second degree burns Marginal Remote Low Avoid contact with soldering iron. Tie back long hair. Moving heavy objects Transpor ting the ground station Muscle strains; toe injury Marginal Remote Low Ensure the ground station is properly transported Exposure to epoxy hardener or resin Using epoxy during assembl y of rocket Minor skin irritation Negligibl e Frequent Moderate Rinse area immediately with soap and water. Electrical shock Touching exposed wiring Low level shock to person handling payload Negligibl e Remote Minimal Cover up exposed wires. 148

150 Appendix I. Failure Modes and Effects Analysis (FMEA) Table Type Structural Failures Failure Mode /Hazard Failure at Nose Cone and Fragile Materials Payload Bay Interface Cause Effect Probability Severity Risk Improper assembly of components or weak adhesion used to secure components. Separation of Nose Cone and Fragile Materials Payload bay. Can also result in loss of Payload and ballast Remote Catastrophic Moderate Risk Controls/Mitigation s The Nose Cone will be attached with screws to the Fragile Materials Payload Bay. The nuts and washers of these screws will be epoxied into the Nose Cone to ensure that they will remain attached during flight and during assembly. To ensure strength of bond nose cone will first be sanded inside at the point of adhesion. Nose Cone Deformation Loss of Structural Stability in leading component of the rocket. Flight becomes imbalanced altering flight path, potentially leads to further structural failure. Improbable Critical Low Risk The Nose Cone has been chosen to be made of polypropylene, with a tensile and compressive strength of 4,800 psi and 7,000 psi, respectively. These are much greater than the expected load forces during flight. 149

151 Body Tube Deformation Loss of Stability in main structure of the rocket. Flight becomes imbalanced, or fails, leading to structural failure and possible loss of the rocket. Improbable Catastrophic Low Risk The Body of the rocket has been chosen to be made of Carbon Fiber, with a tensile and compressibe strength of 1500 MPa and 1200 MPa, respectively. These are much greater than the expected load forces during flight, Coupler Deformation Loss of internal structural stability. Mass shifts within the rocket, creating an imbalance, and possible alteration of the flight path or stability. Improbable Critical Low Risk Couplers have been chosen to be made of Kraft Phenolic, which has been proven to stand up to expected load forces. Couplers will not be subjected to these forces, since this load will be taken on by the Body Tubes. 150

152 Failure of the Fragile Materials Payload Bay's Load Bearing or Eye-Bolt that is attached to the Shock Chord The bulkhead is not properly attached to the Fragile Materials Payload Bay. The stress of the parachute deployment is too great for eyebolt or bulkhead to withstand (Max force on eyebolt: 1,300 lbs. Peak Stress of bulkhead: 300 MPa). Section I of the rocket can become detached from the other sections and can fall to ground with no recovery system and result in more structural damage. Remote Catastrophic Moderate Risk The load-bearing bulkhead will be made of fiberglass (Peak Stress: 300 MPa) in order to ensure it can support the force it is expecting. The strength of the bond between the bulkhead and the body will be maximized. It will be adhered with epoxy and construction/assembl y will be overseen by experienced personnel. Shear Pin Failure at Fragile Materials Payload Bay and Parachute Bay Interface After drogue is deployed there is an excessive force placed on shear pins. The interface between the Parachute Bay and Fragile Materials Payload Bay separates before 600ft. The main parachute deploys at a high altitude. Drift radius is larger that what was modeled. Remote Marginal Low Risk Ground testing will be performed to ensure that the shear pins can sustain the force of the drogue ejection charges. Shear pins have been selected with an average of lbs of force from the ejection charges to separate. Long shock chord will be used that are 5 times the length of the rocket. 151

153 Destabilization of CRAM The CRAM fails under tensile stress The avionics housed in the CRAM can be disturbed or destroyed leading to issues with the recovery system such as proper deployment of the parachutes. Remote Critical Moderate Risk The CRAM features will be produced using additive manufacturing. The material is much stronger in compression that in tension. Great care will be put into the design in order to minimize the tensile stress. Zippering of Parachute Bay Body Tube Parachutes are ejected early or when the vehicle still has significant lateral velocity. Shock chord "rips" through the body tube, rendering the vehicle incapable of immediate reusability. Occasional Marginal Moderate Risk Increase the shock chord area at the point of contact with the body tube. This will distribute the force over a larger area. Use simulations to predict the near vertical path. Rail Button Failure The rail buttons are not properly secured onto the body tube of the rocket. This can cause an unstable flight of the rail and affect the predicted stability margin and flight path of the rocket. Improbable Critical Low Risk The Rail buttons will be secured to the body tube with screws, nuts, and washers in order to prevent any damage to the rocket. 152

154 Failure at Fin and Fin Can Interface The fins were improperly bonded to the fin can. Fins can separate from the fin can, greatly effecting the stability of the rocket and its ability to withstand flight disturbances Improbable Catastrophic Low Risk The fins will extend into the body tube and be adhered to both the fin can and the motor mount with four fillets. The surfaces of the points of adhesion will be sanded prior to adhesion to increase surface area and form a stronger bond between interfaces. Failure of Motor Mount and Centering Ring Interface The centering rings are not properly adhered to the motor mount. The motor mount can shift during flight and cause the motor casing and motor to be unstable. Remote Catastrophic Moderate Risk The strength of the bond between the motor mount and the centering rings will be maximized. They will be adhered with epoxy and the construction will be overseen by experienced personnel. A shake test will be performed to ensure all connections are stable. 153

155 Failure of Centering Rings and Fin Can Interface The centering rings are not properly adhered to the Fin Can. The centering rings and motor mount can shift during flight and can cause the motor casing and motor to be unstable. Remote Catastrophic Moderate Risk The strength of the bond between the fin can and the centering rings will be maximized. They will be adhered with epoxy and the construction will be overseen by experienced personnel. A shake test will be performed to ensure all connections are stable. Failure of Fins Fins are subjected to excessive loads during flight. Fins can fracture or break into pieces, causing destabilization, possible loss of rocket, and endangering team members nearby. Improbable Catastrophic Low Risk The fins have been chosen to be made of Carbon Fiber, with a tensile and compressibe strength of 1500 MPa and 1200 MPa, respectively. These are much greater than the expected load forces during flight. 154

156 Propulsion Failed Ignition Malfunction in the ignition of the solid fuel motor. Large safety concern because a primed motor is on the Launch Pas and could theoretically go off at any moment. Personnel may believe that the motor is inactive when it may still be active. Improbable Critical Low Risk If no activity then a wait of a minimum of 60 seconds before approaching the rocket. Our team will work with trained professionals to insure that there is proper and reliable ignition that will occur on launch day and they will troubleshoot causes of failed ignition. Motor Casing explosion Nozzle can be clogged by a detached chunk of propellant. Motor casing can explode on launch pad under pressure and can partially or totally destroy the fin can. Improbable Catastrophic Low Risk The motor selected will have a high safety rating. It will be inspected for any visual cues of faulty manufacturing prior to installation on launch day. 155

157 Failure of the Motor Retention The screws, nuts or washers could not support the load they were experiencing. Motor casing and spent motor can fall out of the motor mount during descent after main parachute deploys. Improbable Critical Low Risk The motor retention system will be tested during ground testing of the ejection charges. The screws used for the motor retention will be adhered with epoxy and JB weld to insure structural stability of the retention. Inadequate Motor Thrust Motor is too weak to provide the necessary power to keep the rocket on the necessary flight path and reach the desired apogee. Rocket flies off the desired path and underperforms. Remote Critical Moderate Risk Simulations will be run using OpenRocket and RockSim to ensure the motor selected provides the adequate amount of thrust. The motor will be properly packed into the rocket and only motors from reliable manufacturing brands will be used. Stability Vehicle is Unstable Center of Gravity is Behind the Center of Pressure. Flight path will be unpredictable and erratic. Improbable Catastrophic Low Risk Use simulations and computer models to ensure proper placement of the CG and CP. Physically verify location of CG prior to all launches. 156

158 Vehicle is Overstable Center of gravity is too far ahead on the vehicle Vehicle weathercocks into the wind and off path. Will not achieve altitude and drogue ejection may zipper airframe due to lateral speed. Occasional Marginal Moderate Risk Perform multiple center of mass calculations, such as using OpenRocket and RockSim, as well as hand written math. Compare results to ensure the center of mass predictions are accurate. Physically test prior to flights to verify CG is in expected location. Vehicle Unstable off the Rail Vehicle does not achieve a high enough speed to be aerodynamically stable as it leaves the rail. Rocket is launched at an angle and goes off path. May not achieve altitude and drogue ejection may zipper airframe due to lateral speed. Remote Critical Moderate Risk By analyzing the thrust curve of the engine to be used and taking the mass of the vehicle, calculate the speed attained by the vehicle off the rail to ensure it is fast enough, as determined by the appropriate literature (i.e. NAR). 157

159 Rocket does not follow Intended Flight Path Poor construction or significantly overstable Rocket veers from intended flight path Occasional Marginal Moderate Risk Run simulations to ensure approximations are correct. Have experienced personnel oversee construction and utilize resources on fin construction. The Fins are Misaligned The fins were not properly adhered or placed at the correct angle or distance from each other. The rocket becomes unstable and has the potential for flying off angle or spinning. Improbable Critical Low Risk The fin can will be ordered with the correctly designated fin tabs that will be made in the body by the manufacturer. The fins will be examined for proper alignment. The Rail Buttons are Misaligned The Rail buttons were not properly attached or aligned on the same axis. The rocket will have an unstable flight of the rail or not be able to properly sit on the launch pad. Improbable Critical Low Risk The Rail buttons will be secured to the body tube with screws, nuts, and washers in order to prevent any damage to the rocket. The alignment will be ensured using a right angle and the construction will be overseen by experienced personnel. 158

160 Force from main parachute deployment CRAM mount shears and fails Improbable Critical Low Risk Use a robust screwto-lock mounting mechanism with large contact area and large eye-bolts CRAM Torque from main parachute deployment CRAM screws loose from mount Remote Marginal Low Risk Screws lock CRAM into correct position Improper handling of rocket tube or preflight installation of CRAM Mount separates from body tube Improbable Marginal Low Risk Mount designed with large contact area with tube for epoxy, and screws through the body tube hold it in place Recovery Unforeseen failure of altimeters Charges don t deploy Remote Catastrophic Moderate Risk Include dual redundancy to account for unforeseen errors Altimeters Inadequate air flow Inaccurate altimeter readings Improbable Critical Low Risk Ensure air-flow holes are aligned and large enough Incorrect calibration Charges deploy early/late Improbable Critical Low risk Run ground tests to ensure expected performance Parachutes/shock chords Large tear in parachute Frayed or damaged shock cords Parachute fails to slow down rocket Shock cords break Occasional Marginal Moderate Risk Improbable Catastrophic Low Risk Check parachutes before launch to ensure flight readiness and use durable parachutes Ensure shock cords are in good condition (check for tears, frays, and scorching) 159

161 Black powder charges explode Shock cords tangle Parachutes or shock cords are damaged or broken Separated sections collide during descent Remote Marginal Low Risk Occasional Negligible Minimal Risk Use Nomex cloth to protect components Ensure proper folding and installation procedures Broken wire connection Failure to deploy charges Remote Catastrophic Moderate Risk Use minimal wiring of correct lengths to reduce stress on wires Electromagnetic waves Interference with circuitry Improbable Critical Low Risk Use copper shielding to protect the electronics Electronics Batteries out of charge Charges fail to deploy Improbable Catastrophic Low Risk Use fresh batteries for each launch and test for charge Short circuit in PCB Fire or failure to deploy charges Improbable Critical Low Risk Epoxied the top layer of the PCB to provide insulation against external connections Black powder charges Explosion not powerful enough Sections fail to separate Remote Critical Moderate Risk Make sure the necessary amount is measured and installed, and also use extra for backup charge 160

162 Explosion too powerful Shock cords break Improbable Catastrophic Low Risk Consult with mentor regarding on mass of black powder charges, complete ground test of black powder charge, utilize properly rated shock cords in good condition Payload 1: Roll Control System Controller Malfunction: Secondary fins adjusted to wrong angle or adjustment takes place at wrong time Adjustable fins breaking off rocket Arduino/Control code error, calibration failure in servo motor and fin adjustment mechanism Excessive force applied on fins during initial deployment or sustained pressure during flight Increases/decre ased rotation of rocket; fail to meet full rotation requirements Uneven drag distribution on rocket, loss of structural integrity, potential airflow into body of rocket, loss of roll control Remote Marginal Low Risk Remote Critical Moderate Risk Simulate and test control algorithm code, ground test fin adjustment mechanism Fin material (carbon fiber) will be significantly stronger than expected max force, structural testing to be done before flight Power Failure Battery depletion during flight Loss of controller function, no experimental data collected, underpowered servo motor Remote Critical Moderate Risk Insure batteries have sufficient charge to provide necessary power for the motor and sensors exceeding flight duration 161

163 Control structural failure Design lacks necessary robustness to handle takeoff forces or sustained pressure during flight Damage to payload, loss of data, failure of entire superstructure Remote Critical Moderate Risk Controller will be structurally contained, failure will not affect data, appropriate isolation from other payloads to be included in design Servo Motor Failure Torque on servo by fins exceeds maximum value of 0.45 N*m Secondary fins stop turning or do not turn properly Remote Critical Moderate Risk Fins will be designed in a way that they can be turned without putting excessive torque on motor Loss of Connectivity with Ground Station Rocket travels out of range of ground station (1.2 miles/ 2 km) Radio communication stops sending lie flight data from rocket to ground station Remote Marginal Low Risk Radio frequency will be chosen with rang long enough to communicate even at rocket's maximum height. Additionally, multiple antenna will be used so the rocket can transmit while rolling Payload 2: Fragile Material Protection Leakage of viscous fluid Improperly closing inner container Compromise stability of fragile object within inner container Remote Critical Moderate risk Ensure lid is tightly fastened to rest of inner container, forming an airtight seal Disconnection of a spring Larger than anticipated forces applied, resulting in shear stress on spring greater Fragile object jostled around due to imbalanced springs Occasional Critical Moderate risk Ensure spring constants are greater 20 lbf/in 162

164 Structural Failure of inner container Structural Failure of outer container than 1350 ksi Weak interface between spring and outer container Design lacks necessary robustness to handle takeoff and parachute deployment forces; shear force exceeds 30 ksi Design lacks necessary robustness to handle takeoff and parachute deployment forces Inner container ricochets within the rocket Fragile object jostled around due to imbalanced springs Inner container ricochets within the rocket Damage to payload Damage to payload Failure of entire superstructure Occasional Catastrophic High risk Improbable Critical Low risk Improbable Catastrophic Low risk Remote Remote Remote Critical Critical Catastrophic Moderate risk Moderate risk Moderate risk Rigorously test spring interfaces prior to assembly and launch; periodically check interfaces to ensure continued strength Aluminum inner container able to withstand forces both in compression and tension up to 30 ksi Kraft Phenolic coupler is strong in compression (takeoff forces) and tension (parachute deployment forces), payload isolated from other payloads 163

165 Appendix J. Environmental Effects on Rocket Failure Mode/ Hazard Cause Effects Probability Severity Risk Controls/ Mitigations Bodies of Water Launching too close to bodies of water This can damage the electronic components of the rocket if submerged and the rocket can become irretrievable in the body Remote Catastrophic Moderate Risk Ensuring that there are no bodies of water near the drift radius of the rocket Humidity Launching in excessive humidity The charges well become wet do to the humidity and be unable to properly ignite Improbable Critical Low Risk Motors and charges should be stored by certified personnel in a dry place Lightning Launching in a thunderstorm Electrical shock to rocket systems by lightning and can ground the launch Improbable Catastrophic Low Risk This will ground the launch; no rocket should be launched during a thunderstorm Low Hanging Clouds Launching with low cloud cover It is difficult to keep track or rocket and to properly test all the rocket systems Occasional Critical Moderate Risk Low hanging clouds should be avoided during launch days paying careful attention to monitor the forecast Low Temperatures Launching in extremely cold temperatures Batteries can discharge at a faster rate and the fiberglass can shrink; the rocket will not be able to perform at its Occasional Critical Moderate Risk Battery levels will be monitored by the ground station and battery life will be conserved by turning on the systems at designated times and not leaving them on when function is not necessary 164

166 optimum level Rain Launching with risk of rain Can damage electrical components of rockets and ground the launch Remote Catastrophic Moderate Risk This will ground the launch; rockets should not be launched in the rain Swampy Area Launching in swampy area If rocket lands in this area it can permanently damage certain components or become irretrievable in the swamp Remote Catastrophic Moderate Risk Ensuring that there are no swampy areas near the drift radius of the rocket Trees Flying near wooded areas This can damage the rocket and the parachute if caught by a tree and can also cause the rocket to be irretrievable Occasional Critical Moderate Risk Ensuring that there are no trees near the drift radius of the rocket UV Rays Rocket exposed to sun for long periods of time This can weaken material adhesives if exposed for long durations of time Improbable Critical Low Risk The rocket will not be exposed to the Sun for a long period of time and extensive work on the rocket will be performed indoors Wind Launching in over 20 mile an hour winds This can reduce the altitude achieved by the rocket, affect the stability of the flight and increase the drift of the parachute, and will Improbable Catastrophic Low Risk The launch will be grounded if the winds are too severe and there will be no obstructions in the estimated drift radius 165

167 ground the launch in excess winds 166

168 Appendix K. Safety Concerns for Environment Failure Mode/ Hazard Cause Effects Probability Severity Risk Controls/ Mitigations Battery Leakage Improper disposal of damaged or used batteries Contaminate groundwater and in turn contaminate any organic material that is in the water system Remote Critical Moderate Risk Using proper battery disposal techniques and ensuring all batteries are not damaged Carbon Emissions Using cars to travel to launch sites Damage the ozone layer with emissions Occasional Marginal Low Risk Using carpooling as much as possible to minimize the amount of vehicles Epoxy Leakage Improper use and disposal of epoxy resin in an uncontrolled environment Contaminate drinking water, be ingested by wildlife, or pollute as solid waste Improbable Critical Low Risk Using proper techniques in application to ensure the resin is properly dried and disposing of the resin in designated areas Field Fire Igniting rockets near dry grass and shrubs or motor CATO Set the launch site or other nearby objects on fire Remote Critical Moderate Risk Making sure that any field in use is not near any shrubs and using the proper launching pad to ensure the ignition doesn t affect the surrounding area Harmful Gas Emissions Motors emitting gases upon ignition into the environment Pollute the atmosphere with harmful substances Remote Critical Moderate Risk There will not be many launches done by the team so the emissions will not be to a concerning level Harm to Wildlife Launching a vehicle in a non-designated area around an animal's natural habitat Destroy animal habitat and result in loss of food source, water source, or life Improbable Critical Low Risk Ensuring that we only launch in predesignated areas that will have minimal effect on surrounding wildlife Plastic/Wire Waste Stripping a wire on site and not properly disposing of the waste If not properly disposed it can cause solid waste or be ingested by an animal Improbable Critical Low Risk Ensuring that any stripped wires have the waste properly collected and disposed Spray Paint Fumes Spray painting the rocket Can contaminate the water supply or atmosphere Remote Critical Moderate Risk Painting the rocket in a painting booth that properly disposes of waste 167

169 Waste Improper disposal or storage of rocket components Can result in pollution of environment if improperly disposed or stored. Improbable Critical Low Risk Correctly storing any piece of the rocket that is still waste and disposing off the rest in the proper fashion Water/Ground Pollution Leakage of motor chemicals into the ground and water Pollute the water system with improper disposal Improbable Critical Low Risk There will not be many launches done by the team so the pollution will not be to a concerning level 168

170 Appendix L. Project Risks Likelihood - Rare, Unlikely, Even, Probable, Extremely Likely Impact - Negligible, Low, Moderate, High, Critical Risk Likelihood Impact Mitigation Time Possibility of falling behind schedule and/or missing deadlines Probable Low All aspects of the project will be divided up among team members to reduce the chances of falling behind in work. Additionally, multiple team members will coordinate together to ensure that deadlines are met and to keep each other accountable. Budget Failure to have enough funds to purchase rocket materials, cover transportation costs, and pay for other expenses Rare High All material costs will be determined prior to construction. The team will determine how much material must be ordered in order to prevent overspending. Similarly, travel/transportation expenses will be planned out. Overall budget and spending plans will help ensure that this constraint is met. Equipment and Facility Physical injury associated with on- and off-campus facilities and the material/equipment used to build and operate the rocket Unlikely High Dangerous materials and equipment, including power tools, machinery, and rocket engines, will be used. Every team member will have proper knowledge and training before using laboratories, workshops, materials, and/or equipment. In addition, team members will use personal protective equipment when working with the rocket. The team safety officer, and subteam safety liaisons will communicate proper safety practices. Personnel Potential issues involving team members leaving, which may impact time and budget Unlikely Negligible In the case of someone leaving the team, their responsibilities will be spread among other members. Payload Possibility of malfunctioning or inoperative payload(s) Unlikely High The payload subteams will ensure that work is split among members and adequate time is spent on each step of payload design, construction, and testing. Payload functionality will be verified at the full-scale test launch. Launch Launch errors and hazards, including defective launch component(s) Unlikely Critical Prior to launch, the rocket will be thoroughly inspected, and all the launch checklists and procedures will be reviewed. Additionally, the team mentor, David Brunsting, will assist the team at every launch. Recovery Failure of planned rocket recovery, which may result in Unlikely High The recovery subteam will ensure that the recovery system functions properly by thoroughly designing, constructing, and testing the system. On launch day, following the pre-launch procedures and checklists will reduce recovery system issues. 169

171 physical injury or more likely, damage to the rocket and its components Recovery system functionality will be verified at the full-scale test launch. Resources Risk of lacking materials, equipment, and facilities to construct and operate the rocket Rare High Each subteam will outline necessary materials, equipment, and facilities prior to construction. Budget and spending plans will also help ensure that all necessary materials are purchased/obtained. 170

172 Appendix M. Team Timeline 171