University of Florida NASA Student Launch

Transcription

1 1 University of Florida NASA Student Launch Vertical Landing Utilizing Autorotation Preliminary Design Review November 4, 2016

2 2 Table of Contents TABLE OF CONTENTS... 2 TABLE OF FIGURES... 5 TABLE OF TABLES SUMMARY OF REPORT TEAM INFORMATION LAUNCH VEHICLE SUMMARY Size and Mass Motor Choice Recovery System PAYLOAD SUMMARY CHANGES MADE SINCE PROPOSAL CHANGES MADE TO VEHICLE CRITERIA CHANGES MADE TO PAYLOAD CRITERIA CHANGES MADE TO PROJECT PLAN VEHICLE CRITERIA MISSION INFORMATION Mission Statement Mission Requirements Mission Success Criteria SUBSYSTEM LEVEL DESIGN Airframes Composites Payload Subsystem Forward Section Avionics Subsystem Electronics Aft Airframe Motor Selection Motor Integration Flight Dynamics RECOVERY SUBSYSTEM Deployment Process Deployment Method Attachment Scheme Parachute Sizing Component Selection MISSION PERFORMANCE PREDICTIONS Performance Criteria Simulations Vehicle Stability Kinetic Energy during Descent Drift Scenarios... 44

3 DIMENSIONAL DRAWINGS MASS STATEMENT MANUFACTURING PLAN Facilities Components to be Manufactured New Member Introduction and Training Manufacturing Schedule INTERFACES AND INTEGRATION PLAN Integration Strategy Implementation Strategy Impact Assessment Phase 1 Integration Revision History Structural Interfaces Vehicle to Ground Station Vehicle to Launch System SAFETY SAFETY OFFICER AND RESPONSIBILITIES PROJECT RISK ASSESSMENT SAFETY REQUIREMENTS LAUNCH PROCEDURES PERSONNEL HAZARD ANALYSIS ENVIRONMENTAL HAZARD ANALYSIS FAILURE MODES AND EFFECTS ANALYSIS PAYLOAD CRITERIA SELECTION, DESIGN, AND VERIFICATION Payload Objectives System Level Design and Design Matrix Performance characteristics Verification plan Payload Integration Instrumentation Drawings and Electrical Schematics Payload Key Components PAYLOAD CONCEPT FEATURES AND DEFINITION Creativity and Originality Uniqueness and Significance Suitable Level of Challenge ALTERNATIVE DESIGN Alternative Recovery Subsystem SCIENCE VALUE Payload Objectives Payload Success Criteria Logic, Approach, and Method of Investigation Measurement, Variables, and Controls Relevance and Accuracy of Data Experiment Procedures

4 4 6. PROJECT PLAN PERFORMANCE CHARACTERISTICS REQUIREMENTS EVALUATION REQUIREMENTS COMPLIANCE AND VERIFICATION PLAN TESTING PLAN BUDGET PLAN FUNDING PLAN TIMELINE EDUCATIONAL ENGAGEMENT CONCLUSION APPENDIX

5 5 Table of Figures Figure 1. Solid Model of Mission Vehicle with Major Components Labeled... 9 Figure 2. OpenRocket Model of Mission Vehicle showing CG and CP... 9 Figure 3. Solid Model Depicting Different Airframes Figure 4. Extended Blade Figure 5. Blade Curling Figure 6. Autorotation System of Payloads Figure 7. Electronics of Payload Figure 8. Avionics Bay Figure 9. Thrust Bulkhead Figure 10. Aft Assembly with Payload left out Figure 11. Forward Face of Boat Tail Figure 12. Transparent View of Motor Retention Boat Tail, and Centering Ring Interface Figure 13. Clipped Delta Shaped Fins with Landing Legs Figure 14. Trend Established Between Impulse and Simulated Apogee Based on Various Rocket Motors Figure 15. Thrust Curve of the Class K Cesaroni K260-CL Rocket Motor Figure 16. Simulated Altitude versus Time Plot Using the Class K Cesaroni K260-CL Rocket Motor Figure 17. Simulated Stability Margin versus Time Plot Using the Class K Cesaroni K260-CL Rocket Motor Figure 18. Separation Events and shock cord lengths Figure 19. Open Rocket simulation depicting drag forces imparted onto the rocket Figure 20. Altitude vs. Time Figure 21. Velocity vs. Time Figure 22. Acceleration vs. Time Figure 23. Drag Force vs. Time Figure 24. Relative CG and CP locations... 43

6 6 Figure 25. Stability Margin vs. Time Figure 26. Drift radius for forward section Figure 27. Drift radius for autorotation payload Figure 28. Avionics Subsystem Figure 29. Fin Figure 30. Forward Section Figure 31. Payload Subsystem Figure 32. Aft Airframe without Payload Figure 33. Mission Vehicle Profile Figure 34. V-model Integration Strategy Figure 35. System Configuration Figure 36. Avionics Subsystem Integration Figure 37. Recovery Subsystem Integration Figure 38. Payload Subsystem Integration Figure 39. GPS and Communications Block Diagram Figure 40. Control Logic for Autorotation Figure 41. Number of Pixels vs Altitude Figure 42. Arduino for controls Figure 43. The electrical schematic for the autorotation payload Figure 44. Assembled Payload Figure 45. Payload Components Figure 46. Autorotation Blades Figure 47. Swash plate height and spacing Figure 48. Servo, Bulkhead and Ring Positions Figure 49. Payload Dimensions Figure 50. Rocket Structure

7 7 Figure 51. Alternative rocket design Figure 52. UF Rocket Team Overall Budget Distribution Figure 53. UF Rocket Team Total Expenditures Figure 54. UF Rocket Team Launch Costs Figure 55. UF Rocket Team Pre-Launch Costs Figure 56. Breakdown of the Structures Bill of Materials Figure 57. Breakdown of the Recovery Bill of Materials Figure 58. Breakdown of the Avionics Bill of Materials Figure 59. Breakdown of the Propulsions Bill of Materials Figure 60. Breakdown of the Testing Bill of Materials Figure 61. Breakdown of the Payloads Bill of Materials Figure 62. Breakdown of the Educational Engagements Bill of Materials Figure 63. Breakdown of Travel Cost Estimates Figure 64. Gantt Chart Schedule

8 8 Table of Tables Table 1. Bulkhead Material Selection Justification Matrices Table 2. Separation System Comparison Table 3 Parachute Comparisons Table 4. Parachute Dimensions Table 5 Recovery Components Factor of Safety Table 6: Launch Site Simulation Conditions Table 7. Kinetic Energy of Descent Table 8 Total Mass with Margin Table 9 Severity and Likelihood Table 10: Risk Matrix Table 11: Criticality Descriptions Table 12: Project Risk Assessment Table 13. Safety Requirements Table 14. Personnel Hazard Analysis Table 15. Design Matrix Table 16: Launch Vehicle Requirements Established by NASA Table 17 Recovery System Requirements Table 18: Competition and Payload Requirements Table 19: Safety Requirements Table 20: General Requirements Table 21: Launch Vehicle and Payload Requirements Established by the Team Table 22. Testing Plan Table 23: Summary of Funding Plan

9 9 1. Summary of Report 1.1. Team Information University of Florida Rocket Team Department of Mechanical and Aerospace Engineering MAE-A Building: 231 MAE-A Building P.O. Box , Gainesville, FL Mentor: James Yawn, Level 2 Certified, NAR #: Launch Vehicle Summary Size and Mass The rocket will have an outer diameter of 4.03 inches and has a projected mass of lb m. Plus a 20% mass margin, the mission vehicle will have an upper mass of lb m. The total length of the rocket will be inches from the tip of the nose cone to the ends of the landing legs. The finalized length will be determined with Test 6 from Table 22 in order to determine the amount of spacing required for the recovery subsystem. The general vehicle layout is shown in Figure 1. Figure 1. Solid Model of Mission Vehicle with Major Components Labeled The center of gravity of the rocket is located inches from the tip of the nose cone. The center of pressure is located inches from the tip of the nose cone. These center of pressure and gravity locations produce a stability margin of approximately 4.11 throughout flight. Center of gravity and center of pressure locations are illustrated in Figure 2. Figure 2. OpenRocket Model of Mission Vehicle showing CG and CP Motor Choice The propulsion system used in the rocket will be a single commercially available, rearward-firing, solid ammonium perchlorate composite propellant rocket motor without titanium sponge additives, and not

10 10 exceeding 5,120 N-s of impulse. This motor will also satisfy the performance requirement of achieving an apogee of one mile. Simulations performed using OpenRocket were the primary means of selecting the motor to satisfy the performance requirements of the rocket to reach a one mile apogee. Another consideration for motor selection was the rocket geometry, and how the motor would fit within the rocket based on its dimensions. Based on these simulations and considerations, it was determined that a Class K Cesaroni K260-CL will best satisfy the requirements Recovery System The recovery system utilizes a dual deployment system triggered by black powder charges for returning the rocket safely back to earth. During the first separation event at apogee, the rocket will separate into two disconnected sections. The first section is the forward section which will return to earth using dual deployment, releasing the drogue parachute at apogee and subsequently a main parachute at 1000ft. The second section is the experimental payload, which will return to earth using an auto-rotation system with a back-up main parachute system Payload Summary The vertical landing payload consists two systems. The first of these systems is a hazard detection system capable of identifying both hazards and a specific colored landing area. The second is an autorotation system with cyclic control of the aft section on descent as well as the ability to have near zero velocity upon completion of a rotor flare maneuver through cyclic control. The hazard detection system uses a ground facing camera that feeds data in real time to an onboard microprocessor in order to monitor the ground for identification of the specific colored landing tarps. It also contains an altimeter that will be used to determine the altitude of the aircraft in real time. The auto-rotation system consists of a single rotor system with full cyclic and collective pitch control. Four servo motors wired to the microprocessor control the pitch of the blades. This allows the aircraft to be maneuvered into a position above the landing zone. The servo motors also allow for the control of descent speed, and the arresting of velocity by modulating the lift generated by the blades. A backup parachute will also be included to ensure the safe landing of the payload.

11 11 2. Changes Made Since Proposal 2.1. Changes Made to Vehicle Criteria For the avionics bay, A PAM-7Q u-blox GPS Module will be used to provide GPS data to the Ground Station through the XBee-PRO 900HP. This module is a good candidate because of small size, as well as a sizeable database of information available from u-blox, specific to the product. Stratologger altimeters will be implemented instead of Raven3 altimeters. Raven3 altimeters need to be oriented vertically, and perpendicular to the ground to maintain calibration. The vertical orientation of the avionics subsystem does not allow this. Remote radio frequency power switches may be implemented through the existing Xbee RF module in the avionics subsystem. Three 9V batteries will power the components of the avionics subsystem. The battery for the payload subsystem will be separate, as it is now in the aft section instead of the forward section, and the payload subsystem and aft section of the launch vehicle will be coming down separately from the avionics subsystem and the nose cone. Electronics for the payload and the avionics subsystem have been separated distinctly, and servo motors will not be in the avionics subsystem. The components in the avionics subsystem will be two altimeters, one for redundancy, three 9V batteries, and an Xbee RF/antenna/GPS module. The altimeter section of the avionics subsystem will be coated with copper tape to prevent RF interference and the design of the avionics subsystem will ensure separation of compartments, which is detailed under A few changes were made to the payload. Instead of landing the nosecone only nose side down, the payload will land the entire aft section. The camera will now be mounted to a fin in a fairing. There has been the addition of permanent landing legs to the fins. A parachute has been added to the top of the payloads as a backup safety mechanism for all flights. Instead of a custom script the payload will use an altimeter that uses the Arduino to run the evaluation script. Based on updated mass estimates of the rocket components, the performance of the motor initially evaluated as a favorable candidate was determined to be excessive to achieve an apogee of one mile. By considering these updated mass estimations, simulations were performed using OpenRocket to determine a new candidate motor. According to the evaluation of the OpenRocket simulation data, it was determined that a Class K Cesaroni K260-CL motor would satisfy the performance requirements of the rocket system. The average thrust provided by this motor is 60.2 lbf, while the peak thrust is 96.8 lbf. With 513 lbf-s of total impulse, this motor is projected to carry the rocket to an apogee of 5152 ft, with a velocity off the launch rod of 55 ft/s and a maximum flight velocity predicted to be 507 ft/s. For the recovery system, originally proposed was that the rocket would experience two separation events and return to earth in three tethered sections using a drogue and main parachute. This has since been changed to two separation events that results in two completely disconnected sections returning to earth under different systems. The aft section will return to earth with an auto-rotation system, while the forward sections will use a dual deployment system of drogue and main to return to earth. Changes to the recovery systems are largely based around the changes made to the location of the payload. Whereas the original placement of the drogue and main parachutes were more centrally located in the rocket, the changes to the payload to be moved to the aft section resulted in the main and drogue parachute being pushed closer to the nosecone. The positioning of the main and drogue parachutes is also switched, with the main chute being closer to the nosecone, rather than the drogue being closer.

12 12 Lastly for the overall structural design, the payload components moved to the aft section, and avionics bay moved to the forward section of the rocket. The nosecone changed back to regular, and landing legs were added to the fins. The parachute locations changed with the main chute being towards nosecone and drogue chute being in aft end of forward section. Also, a bigger motor has been selected, so a longer motor tube will be needed Changes Made to Payload Criteria Following the proposal a few changes were made to the payload. Instead of landing the nosecone only nose side down, the payload will land the entire aft section. The camera will now be mounted to a fin in a fairing. There has been the addition of permanent landing legs to the fins. A parachute has been added to the top of the payloads as a backup safety mechanism for all flights. Instead of a custom script the payload will use an altimeter that uses the Arduino to run the evaluation script Changes Made to Project Plan Every element of the project plan has evolved since the proposal. As the design has matured, the budget has changed as well, shrinking to better represent the needs of this year s vehicle and payloads. The budget has decreased from $9,075 to $6,620. Furthermore, the details on the schedule have been expanded considerably. Detailed dates for design, analysis, simulation, manufacturing, and educational engagements have been added to the schedule. Possible launch days as well as back up launch days are now included in the schedule as well. The full schedule can be seen in Figure 64. Finally, the educational engagement plan has been updated to include the schools the team plans to attend and the material to be presented.

13 13 3. Vehicle Criteria 3.1. Mission Information Mission Statement Two mission statements reflect the general objectives of the UF Rocket Team and the specific objectives associated with NASA Student Launch. UF Rocket Team Mission Statement The UF Rocket team exists to serve as a platform for students at the University of Florida to design, build, test, and fly high powered rockets. We will develop essential skills that can be applied across all fields of engineering and science for members internally, as well as student youth in the local area. The team will pursue meaningful projects that positively impact the rest of the world. NASA Student Launch Mission Statement The UF Rocket Team will develop two payloads that will be flown to an altitude of 5,280 feet. All components will be safely recovered after each flight. By doing so the team hopes to gain experience in a research-based, competitive, and experiential exploration project while also providing relevant and cost effective research and development. During the process, the UF Rocket Team will focus on developing an interest in STEM fields and STEM education in the local area Mission Requirements In addition to the requirements set in the NASA Student Launch Statement of Work, the team has developed requirements specific to the payload system the team is developing. Each requirement has a verification plan in progress: Refer to Table 16 in Section 6.3. for Launch Vehicle Requirements Established by NASA. Refer to Table 17 in Section 6.3 for Recovery System requirements as laid out by the Student Launch Handbook. Refer to Table 13 in Section 4.3 for Safety requirements as laid out by the Student Launch Handbook. Refer to Table 21 in Section 6.3 for Competition and Payload requirements as laid out by the team. Refer to Table 20 in Section 6.3 for General requirements as laid out by the Student Launch Handbook. Refer to Table 18 in Section 6.3 for Team Launch Vehicle and Payload requirements as laid out by the team Mission Success Criteria a) The launch vehicle will leave the launch rail with sufficient velocity and stability to continue stable flight to apogee. b) The launch vehicle will maintain stable flight after motor burn out and until apogee is reached c) Apogee will occur at 5,280 feet Above Ground Level (AGL). d) The drogue parachute will be deployed from the airframe at apogee. e) The drogue parachute will prevent the launch vehicle from continuing a ballistic trajectory

14 14 f) The launch vehicle will descend on the drogue quickly enough that it will not leave the site from which it was launched g) At 1000 feet AGL, the launch vehicle will separate once more to release the main parachute. The launch vehicle will land attached to this parachute. h) The autorotation payload will vary its descent speed throughout the flight and perform a flare to reduce airspeed to 0 ft/s at 225 AGL. i) The autorotation payload will release a parachute at 200 feet AGL and will make an upright vertical landing j) The landing zone detection camera shall observe the area beneath the launch vehicle and analyze it for landing location by determining the color of the tarps. k) The launch vehicle will land in a manner that does not damage it so that it can be prepared for launch once more within the same day. l) The launch vehicle will transmit its location to the ground station via GPS Subsystem Level Design Airframes The mission vehicle is divided up into the forward and aft airframe. The forward section will house the recovery and avionics subsystems while the aft airframe will house the propulsion and payload subsystems. Figure 3 depicts the division of the subsystems. Later sections will go into further detail about the components that make up each airframe and subsystem. Refer to Integration Strategy as well for an integration plan illustrating the relationship between the subsystems in full assembly. Figure 3. Solid Model Depicting Different Airframes Material Selection Bulkheads are structural discs that are epoxied to the body tubes or coupler tubes within the mission vehicle to which the rest of the subsystems are then attached to the bulkheads through epoxy or fasteners to secure itself in the mission vehicle. Bulkheads will be made of plastic, specifically type II PVC, because of their machinability and homogeneous properties as well as not affected as much by weather and temperature conditions, unlike wood. Both plastics and metals would be candidate materials to be used for bulkheads, however, plastics are lighter than metal. The strength of the plastic suffices for the design of the rocket

15 15 specifically the tensile strength to withstand the max motor thrust of 177 lbf. The previously mentioned qualitative comparisons are quantitatively justified is found through the use of design matrices in Table 1. Table 1. Bulkhead Material Selection Justification Matrices HIGH Density PE PEEK Parameter Weight Value Score Weighted Score Value Score Weighted Score Tensile Strength (psi) Machinability HSS Carbide Impact Strength (ft-lb/in) Density (lb/in^3) Heat Conductivity (BTU/hr-ft-F) Cost $ ABS Nylon Parameter Weight Value Score Weighted Score Value Score Weighted Score Tensile Strength (psi) Machinability 0.20 HSS HSS Impact Strength (ft-lb/in) Density (lb/in^3) Heat Conductivity (BTU/hr-ft-F) Cost $ Polycarbonate Parameter Weight Value Score Weighted Score Value Score Weighted Score Tensile Strength (psi) Machinability 0.20 Carbide Carbide Impact Strength (ft-lb/in) Density (lb/in^3) Heat Conductivity (BTU/hr-ft-F) Cost $ PVC 1 Document 8574KAC, 8747KAC, 8539KAC, 8657KAC, 8545KAC;

16 16 Strengthened PVC Parameter Weight Value Score Weighted Score Tensile Strength (psi) Machinability 0.20 Carbide Impact Strength (ft-lb/in) Density (lb/in^3) Heat Conductivity (BTU/hr-ft-F) Cost $ In the aft airframe, a motor retention ring, boat tail, and centering ring will be used to capture the thrust of the motor and transfer that thrust force to the rest of the mission vehicle. The motor retention and centering rings will be made out of aluminum. Aluminum was chosen because it will be able to withstand heat exposure from the motor, has a large tensile strength of psi 6, and is manufactural allowing for a tighter tolerance on dimensions than previously used rings that were made of plastic or wood. The boat tail will be made of polycarbonate due to the lower thermal conductivity and tensile strength in comparison to ABS. In the forward section, blast caps are containers that hold black powder charges that will be used to create a pressure between the payload subsystem and avionics subsystem to deploy drogue parachute as well as between avionics and the nose cone to deploy main parachute at a lower altitude further along the forward section descent after apogee. The blast caps for the recovery subsystem of the forward section will be made out of aluminum. Two will be on each exterior surface of the avionics bulkheads. One will be for usage while the other for redundancy should the primary black powder charge not ignites. All the blast caps will have countersunk holes for 6-32 threaded screws to attach them to the avionics bulkheads. They are also reusable and easier to clean than alternatives such as PVC tubing and less complicated and expensive than CO 2 canisters. The structural components in the autorotation system for controlled descent will use aluminum for its strength and machinability, specifically the swash plate, blade attachments, and main plate much for the reason that aluminum has tensile strength that will be able to hold the entire payload together and be able to deliver the force to control the blades. In the forward section, the avionics subsystem uses G10 fiberglass sheets as the basis for its discs to which the electronics will attach to. In the aft airframe, the payload will also contain its own avionics section of it subsystem and G10 fiberglass sheets will be used in the manufacturing of fins to be attached to the mission vehicle. G10 fiberglass sheets will be used for the base of the fins, payload electronics, and the avionics

17 17 subsystem. The fiberglass is strong, rigid, and has a high impact strength 7, but will specifically be used for its ability to allow electrical signals to pass through Composites Airframe Material Selection To choose the optimal airframe material several were considered with the top choices, phenolic, Blue Tube, and carbon fiber. The following properties were compared: individual tensile strengths, weight, price, and ease of manufacturability as well as the educational and manufacturing experience opportunities they offered to new members. Another consideration was the team s manufacturing experience with each material and the facilities and equipment available. The tensile strength of phenolic tube was found to be 7,000 psi, with that of Blue Tube being slightly less, at 6000 psi. Phenolic is very brittle with cracking as its leading cause of failure. One of the advantages of Blue Tube is that it does not suffer the same brittleness of phenolic tubing. Carbon fiber on the other hand has a tensile strength of 500,000 psi making it by far the strongest material we considered. It is more brittle than its common composite counterpart, fiberglass, but not near as brittle as phenolic tube. Carbon fiber s main disadvantage is its price, at approximately $100 per pound. Blue Tube cost approximately $23 per pound making it an attractive option. Phenolic is nearly as cheap at $26 per pound. The early parachute deployment event at last year s competition, is a reason why the added expense of carbon fiber would be worthwhile. The shock cords tore through two inches of the carbon fiber airframe indicating that if a weaker material had been used, the results would have been catastrophic. Phenolic tube and Blue Tube require only being cut to length and surface finished making them attractive materials in the realm of manufacturability earning them an excellent rating in the analysis. However, the team already has the equipment and experience needed for composite manufacturing readily available. In addition, all stages of the wet layup process offers new members many opportunities to become involved in hands-on manufacturing. As one of the team s main goals is to educate others, carbon fiber is ranked excellent in educational value. Durability was the last comparison factor that was used in the material analysis. As previously mentioned, the brittleness of phenolic tube is its main weakness and a common complaint among amateur rocket builders. It is a strong material however, giving it a fair rating in comparison. Blue Tube does not suffer from brittleness, but is easily damaged with water. As we cannot predict our launch conditions and water is common in the environment, and with high humidity levels in Florida, it also has a fair rating. The high strength of carbon fiber indicates a high durability and last year s deployment malfunction further demonstrates the team s ability to effectively manufacture a strong and durable composite. Carbon Fiber Blue Tube Phenolic Tube 7 Document 8549KAC

18 18 Price per Pound [$/lb] Tensile Strength [psi] 500,000 6,000 7,000 Specific Strength [σ/ρ] Manufacturability Fair Excellent Excellent Educational Value Excellent Poor Poor Durability Excellent Fair Fair Carbon fiber has been selected as the primary fuselage material. Testing has shown that a single layer of uni-axial fiber weave layered with two layers of bi-axial fiber weave then followed by a layer of bi-axial fiberglass weave (to act as a sacrificial layer to protect the carbon fiber during the wet sanding process) to sufficiently meet our strength requirements as well as prove less expensive than pre-made composite counterparts on the market by a factor of four. Carbon fiber is electrically conductive and would therefore act as a Faraday cage around the electronics and avionics subsystem preventing radio communication. Therefore, the avionics subsystem will be composed out of pure fiberglass in the same fiber orientation as the carbon fiber sections (1 uni-axial sleeve followed by 3 bi-axial sleeves). This section will also meet the strength requirements as fiberglass is almost as strong as carbon fiber with a strength of 257,000 psi. Research and Development Last year s composite team developed manufacturing methods that yielded very strong laminates using the wet layup process. The goal for the team this year is to improve the devices used to form the layups. The first being the tube that supports the body tube layups. Previously this has been a phenolic tube mounted on an axle that can be turned by an electric drill. This has worked well for wrapping and compressing the wet laminate, but the team is currently looking for ways to improve laminate removal from the tube after the epoxy has cured as well as improving the bulkheads which integrate the axle with the larger diameter phenolic tube as these have proven difficult to use. The second has to do with the fins. Last year during subscale manufacturing the team had three aluminum plates, which had been CNC milled to produce the rounded fillets at the fin base and the attachment point on the body tube. These worked extremely well, but unfortunately, were not large enough for full scale manufacturing and thus the team improvised with plates of G-10 fiberglass. This year the team will manufacture the plates with the full-scale fin size in mind so the plates can be used for both sub-scale and full-scale manufacturing.

19 Payload Subsystem In order to ensure the aft bay s descent is slow enough to make vertical landing possible, an autorotation system will be utilized. The autorotation system located at the forward end of the aft bay will be composed of a rod with blades at one end whose rotation and pitch is controlled by a swashplate several inches below as can be observed in Figure 6. The other end of this rod will be fixed to a bulkhead using epoxy and this bulkhead will then be attached to a centering ring using removable machine screws. The centering ring will be epoxied to the inside of the coupler located at the front end of the aft bay. This arrangement will allow us to remove the payload if necessary to access the electronics in the aft bay while assuring the autorotation system stays attached to the aft bay during descent. Figure 7 shows the model of the electronics needed for controlled descent. The blades, being carbon fiber, effectively curl up, as seen in Figure 4 and Figure 5, along the circumference of the body tube of the forward bay. Figure 4. Extended Blade

20 20 Figure 5. Blade Curling To control the swashplate and direct the aft bay s decent, four identical Kevlar string will run from equidistant points on the circumference of the swashplate through holes in the bulkhead, each attaching to a separate servo motor. This system will pull on the swashplate, changing the pitch of our blades and allowing change in direction. Throughout the flight, there will be 1/8 in. threaded rods approximately one foot in length running parallel to the body tube extending from each of the fins past the end of the aft bay. These will be the legs that stabilize the rocket during the vertical landing. Figure 6. Autorotation System of Payloads

21 21 Figure 7. Electronics of Payload Recovery Separation of Payload The separation of the aft and the forward sections will be initiated by a black powder charge of 2 grams that will force a piston in the back end of the forward section with two pushrods extending towards the aft bay. The blades of the autorotation system, being curled around the inner rim of the forward bay body tube along their length, take up little space within the tube itself. As such, upon firing of the piston, the push rods can go through this space, soon coming into contact with and pushing off the payload bulkhead, causing separation. By pushing off the bulkhead, we prevent damaging the swashplate or other components of the autorotation system. The piston will be connected to a drogue parachute located in the forward bay using wire. The firing of the piston pulls on the wire, deploying the drogue parachute. Later, at a predetermined altitude the nose cone will be separated from the forward bay by a black powder charge and the main parachute will be deployed Forward Section The forward section consists of the avionics subsystem, the nose cone, a drogue chute, and main parachute. The forward section does affect the payload subsystem and is completely detached from the aft airframe during the first separation event. The body tube of the forward section is restricted to a fiberglass body frame, rather than carbon fiber. Carbon fiber disrupts electrical signals, creating a Faraday cage for everything in the avionics subsystem. Installed on the forward and aft ends of the avionics subsystem are 1/2" type II PVC bulkheads. These bulkheads serve as the mounting points for the recovery subsystem for the forward section. Installed at the aft base of the nose cone is another 1/2" type II PVC bulkhead, which serves as a tether point for the shock cord connecting the nose cone and the avionics subsystem after the second separation event. Test 6 and 7 of Table 22 will assist in determining the distance between the nose cone bulkhead and avionics bulkhead in packing the recovery subsystem between the two. The drogue chute is stored at the aft end of the avionics subsystem, between the forward and aft airframes of the rocket. The drogue is connected to the rear avionics bulkhead through a simple system of shock cord and a U-bolt installed on the bulkhead. The main parachute is stored between the nose cone and the forward avionics bulkhead. The main parachute is connected to both the forward avionics bulkhead and the nose cone bulkhead via shock cord and U-bolts, same as the drogue chute. Test 30 and 31 of Table 22will help determine the method of packing the drogue and main chute in their respective sections of the forward section to be implemented on final launch. The length of shock cord attaching the nose cone and avionics subsystem is set at a different lengths relating to the attachment point of the main parachute. From nose cone to main parachute, the shock cord length is set at 18", while from avionics subsystem to main parachute shock cord length is set at 45", resulting in a length ratio of 2.5:1. These lengths reduce the chance of mid-

22 22 air collisions between the nose cone and avionics subsystem during descent. Test 11 of Table 22 will be of great importance of testing the theory of no mid-air collisions given a 2.5:1 ratio of shock chords Avionics Subsystem As seen in Figure 8, the avionics subsystem is located aft of the forward section of the rocket. It is comprised of two type II PVC bulkheads, one circular G10 fiberglass disk, one circular carbon-fiber disk (referred to as Electronic disks), two threaded nylon spacers, an airframe, and a U-bolt. The two type II PVC bulkheads are on either end, with two threaded nylon rods that run the length of the bay. Steel hex nuts and washers attach to the rods to provide structural support for the bay itself. Situated vertically, parallel to the bulkhead and to each other are 2 composite disks where the avionics equipment and power supply will be located. The boards will have holes machined out for the threaded rods to run through. The threaded rods will compress the bulkheads around an inner coupler tube permanently mounted to the inside of the body tube, to allow for the avionics subsystem to be easily removable, and still provide the inertia needed for parachute deployment. The vertical disk design of the avionics subsystem provides several benefits that a horizontal orientated avionics subsystem through ease of assembly of the bay and into the rocket, the capability to modify length, and structural support. The bay can be assembled and made launch ready quickly by preparing the Electronic disks in advance with the appropriate electronics in a subassembly and then sliding threaded rods through the holes on the Electronic disks, followed by bulkheads, and securing with appropriate hardware. The length of the avionics subsystem can be modified within the bounds of the threaded rod lengths through preparation of more Electronic disks. Simply removing the hex nuts and either bulkhead, an Electronic disk can be added to the existing Electronic disks. The reverse is also true that unnecessary Electronic disks can be removed to shorten the length of avionics subsystem. Each disk disperses vertical forces throughout their area, whereas a horizontal configuration depended upon the tensile strength of solely the threaded rods. The bulkheads are secured with compression from hex nuts through the bulkheads on either end of the avionics subsystem. The internal Electronic disk subassembly is secured through compression with hex nuts to the threaded rods, but independently of the bulkheads, as the internal subassembly can traverse the length of the avionics subsystem within given space, while the bulkheads remain fixed. The compression of the subassembly disks secure the batteries and the switches mounted between disks. The assembly does not require permanent bonding which allows for modifications. A coupler tube equal to the dimension between the inner face of each bulkhead (5.46 ) permanently secured in the body tube will provide a lip for the bulkheads to catch and compress around to anchor the avionics subsystem into the rocket through launch and deployment. The main flaw of the vertical design of the avionics subsystems is the vertical design itself. As a result, Raven3 Altimeters cannot be used, as these altimeters require the altimeter to be oriented vertically in the avionics subsystem, perpendicular to the ground, prior to launch and during flight to maintain accurate calibration of the altimeter. Orientation does not affect Stratologger altimeters, which will be used. This design forces a decision in altimeters, and limits the altimeters compatible with the avionics subsystem.

23 23 To minimize interference of radio frequency exciting the recovery systems the first disk is carbon fiber to encourage radio wave interference. The forward bulkhead and face of the carbon fiber disk, as well as the walls of the inner coupler tube surrounding the altimeters will be coated with copper tape to further promote radio wave interference. Figure 8. Avionics Bay Electronics The electronics on the vehicle will interface several subcomponents to complete various tasks, including: (1) GPS to locate the vehicle and its sections (2) communication between the vehicle s sections and ground, and (3) recovery electronics. Ground Station The purpose of the ground station is to receive live serial data used for locating the rocket through GPS in the case it proves necessary. The ground station will consist of a laptop as the interface, an Arduino UNO board microcontroller board, one XBee Module to receive data, an XBee Adaptor Kits to interface with the Arduino UNO, and appropriate circuitry. The Arduino UNO microcontroller transfers the raw GPS serial data into recognizable numbers and units that can be viewed through the serial monitor available as a live feed. The XBee modules use transmit serial data through radio frequency in a network consisting of one coordinator and two router modules. Communication Communication between the vehicle s sections and the ground station will be accomplished through three XBee-PRO 900HP modules to transceive GPS data. A Radio Frequency (RF) XBee module is excellent in that its base operation is to act as a wireless serial cable. The XBee-PRO 900HP modules will be connected to Adafruit XBee Adapter Kits that allow easy access to interface with the pins of the modules to connect with GPS modules and batteries. Alternatives for wireless communication include Bluetooth and WiFi. Bluetooth has a limited range of approximately 328 feet (1) while Wi-Fi has a similar range, although the

24 24 main issue with Wi-Fi is the necessary ground equipment that can be burdensome and time consuming as well as less affordable. The XBee-PRO 900HP has an outdoor range of 10 Kbps up to 9 miles at a RF data rate of 10 Kbps or up to 4 miles at a RF data rate of 200 Kbps(with 2.1dB dipole antennas)(2). The modules can be configured to work together through the XCTU XBee module program. The XBee RF module s small size, yet powerful capabilities, provide an ideal candidate for this avionics subsystem. GPS A PAM-7Q u-blox GPS Module will be used to provide GPS data to the Ground Station through the XBee- PRO 900HP. This module is a good candidate because of small size, familiarity with this product, as well as a sizeable database of information available from u-blox specific to the product. The data will be continuously transmitted to an XBee RF module via UART, a simple wire connection, and from one XBee to the ground station. Recovery Electronics The Stratologger altimeter will be used for both the primary and redundant systems, as aforementioned in , as a result of the orientation of the avionics subsystem the Raven3 altimeter cannot be used. An e- match connects the altimeter to blast caps on either end of the avionics subsystem for black powder recovery charges. Power Each altimeter will be powered by a single 9V battery. Both XBee RF and GPS modules require 3.6V [8].These systems will have power supplied by a single 9V battery with appropriate resistors while the ground station XBee and Arduino will be powered through USB to a laptop. Power will be supplied to the payload through a separate power source detailed in 5.1. External access is provided in the form of manual switches mounted directly above the batteries in the middle of the subassembly, snug in between two Electronic disks. A hole in the coupler and body tube of the rocket will provide direct access with a screwdriver to the switch. The Xbee RF Module has the potential to provide a remote power switch option to switch the avionics subsystem components on and off Aft Airframe Motor Centering System The motor centering system will fix and coaxially align the motor tube to the aft airframe with a centering ring and thrust bulkhead, each attached to the motor tube and aft airframe. It will also be designed to harness the thrust of the motor in order to propel the launch vehicle. The centering ring transfer thrust to the airframe from the epoxy bonds. The epoxy permanently binds the thrust bulkhead and centering ring to the airframe and motor tube. The thrust bulkhead is designed to manage the thrust of the motor in the event that the centering rings fail to properly transfer thrust. The motor tube extends beyond the aft edge of the aft airframe and into the boat tail. The assembly of the system is depicted in Figure 10. The thrust bulkhead has a coaxially aligned extruded circular impression on the aft side and approximately one-third of the thickness of the bulkhead deep as seen in Figure 7 below. This extruded circular cut accepts 8

25 25 the forward end of the motor tube. Epoxy mates the motor tube in the aft end of the thrust bulkhead. Test 10 and 18 of Table 2Table 22 will be important in gathering necessary data of motor thrust and what thickness of the bulkhead is required to withstand such force. Test 21 of Table 22 will also be of importance as it determines how much contact area between the inner diameter of the aft body tube and epoxied covered outer diameter of the thrust bulkhead is needed should the thrust bulkhead stand up to the thrust force of the motor. The extruded circular cut serves two main purposes: 1. To provide proper coaxial and lateral alignment of the motor tube with the thrust bulkhead. 2. To increase the thickness of the thrust bulkhead at the airframe interface to distribute the shear loading from the thrust over a larger surface area, which will decrease the shear loading per unit length on the airframe to bulkhead epoxy joint? (This second purpose is based on the bulkhead material having a larger shear strength than the airframe to bulkhead epoxy joint.) Figure 9. Thrust Bulkhead The centering ring is designed to provide the second point of coaxial alignment of the motor tube to the aft airframe. It also carries a portion of the thrust load via the centering ring to the motor tube epoxy joint. The centering ring has a coaxially aligned hole for the motor tube to slide through. Figure 10. Aft Assembly with Payload left out

26 26 Boat Tail The mission vehicle is designed with a boat tail to increase aerodynamic performance and propulsion speed. It will be designed such that it satisfies an optimal shape, given the flight parameters of the launch vehicle. Figure 11 illustrates the fastening system; the boat tail is attached to the rocket with 6-32 threaded screws that extend from the aft end of the boat tail through the centering ring at the aft end of the launch vehicle fuselage. In order to decrease drag by eliminating external fasteners, a series of cylindrical holes will run through the length of the boat tail, concentric with the threaded holes on the back face of the centering ring for the fasteners to be inserted. Especially important for the boat tail design is the reduction of drag as airflows over the airfoil shape of the mission vehicle. Test 16 of Table 22 to determine the curvature of the boat tail will greatly assist in optimizing the reduction of drag on the aft end of the vehicle. The other is the heat resistance of polycarbonate material in relation to the temperature of 1600K that the chosen motor gives off during lift off and continued ascent. Test 19 of Table 22 will be able to test polycarbonate s ability to withstand and hold up to such a temperature. Figure 11. Forward Face of Boat Tail Motor Retention On the aft end of the aft airframe, a motor retention ring will be attached with 6-32 threaded screws in three holes, aligned with the threaded holes in the aft face of the boat tail. To constrain the motor from sliding out past the boat tail inner diameter hole, the inner diameter of the motor retention ring will be less than that of the outer diameter of the motor casing, and thus serving as the motor retainer for the launch vehicle. The assembly of the motor retention mating is illustrated in Figure 12.

27 27 Figure 12. Transparent View of Motor Retention Boat Tail, and Centering Ring Interface at Aft Airframe Fins The fins for the missionvehicle will have a clipped delta shape. The trailing edge will be positioned two inches from the bottom of the body tube of the aft airframe, as shown in Figure 14. They will be parameterized by dimensions of 6 inches for the root chord, 2.75 inches for the tip chord, 4 inches for the fin height, and 39.1 degrees for the sweep angle. They will have a 0.19 inch thickness, constructed from G10 fiberglass and biaxial carbon fiber. Layers of fiberglass and carbon fiber will be alternated in the construction of the fins. The justification for using both materials entails each of their benefits and shortcomings: carbon fiber has a better strength to weight and volume ratio while fiberglass is cheaper while still maintaining sufficient structural integrity. Merging these two materials will allow the fins to attain the desired thickness. The method of merging materials will especially be important as the payload s landing legs will be merged into the body of the fins themselves in order to secure the landing legs to the fins of which the fins are secured to the mission vehicle. The merging of the landing legs will help to distribute stress during landing throughout the fin rather than at a single point on the fin exterior.

28 28 Figure 13. Clipped Delta Shaped Fins with Landing Legs Motor Selection Motor Options After reevaluation of the rocket system, including an updated estimate of total rocket mass and mass distribution within the rocket, new motors were reviewed as candidates for use in the full-scale rocket. OpenRocket was used as the primary tool in determining the optimal motor selection for the rocket. To systematically select a motor to satisfy the performance requirements, various launch simulations were performed in OpenRocket using a variety of motors yielding apogees within an interval of approximately ±400 ft of one mile. The apogee data was plotted versus the respective total impulse values of the motors used in the simulations using Microsoft Office Excel. From this data, a trend was established, and thus the total impulse required to achieve a one mile apogee was determined to be approximately 516 lbf-s. This estimate was used as a benchmark for selecting an optimal motor.

29 Apogee (ft) University of Florida Impulse (lbf-s) Figure 14. Trend Established Between Impulse and Simulated Apogee Based on Various Rocket Motors Based on OpenRocket launch simulations, the motor found to result in optimal performance is the Class K Cesaroni K260-CL. This motor has a burn time of 8.5 seconds and 43.8 oz of propellant. The average thrust is 60.2 lbf, while the peak thrust reaches 96.8 lbf. With 513 lbf-s of total impulse, this motor is projected to carry the rocket to an apogee of 5152 ft. Equipped with this motor, the rocket velocity off the launch rod is projected to be 55 ft/s, while maximum velocity of the rocket during flight is projected to be 507 ft/s. This motor provides a high initial thrust, which gradually decreases throughout the duration of the burn. This thrust behavior ensures that the motor has adequate initial thrust to provide the necessary acceleration for launch when the rocket it is at rest on the launch pad at its fully-loaded mass. As propellant is then combusted and the rocket s overall mass is decreased, the gradually decreasing thrust provided by the motor will ensure a stable rocket flight by minimizing sharp variations in thrust. Figure 15. Thrust Curve of the Class K Cesaroni K260-CL Rocket Motor

30 30 Other motors evaluated using OpenRocket simulations included the Class K Cesaroni K600-17A, and the Class K Cesaroni K660-7, which had initially been the leading motor candidate based on preliminary rocket mass estimates; however, these motors provided unfavorable performance based on simulations. The Cesaroni K600-17A provided insufficient total impulse, and the rocket only achieved an apogee of 4855 ft, while the Cesaroni K660-7 was overly powerful, and the rocket achieved and apogee of 5668 ft. Although these apogee values are unfavorable, these motors were nonetheless noted as possible alternative motor choices for future evaluation should the mass of the rocket change appreciably based on further design. Implementing the Cesaroni K-260-CL motor, the results of the OpenRocket simulations, including the altitude versus time and the stability margin versus time of the rocket s flight, can also be plotted to visually assess the performance. From the altitude versus time plot, it is evident that the motor sufficiently satisfies the desired performance requirements by achieving an apogee of nearly one mile. From the stability margin versus time plot, it is evident that the stability margin increases gradually from engine ignition up to the point of burnout, at which point it remains relatively constant until stage separation. This behavior is expected, given that the propellant mass is being ejected from the rocket, thus shifting the rocket s center of gravity towards the nosecone, and widening the distance between it and the center of pressure. Figure 16. Simulated Altitude versus Time Plot Using the Class K Cesaroni K260-CL Rocket Motor

31 31 Figure 17. Simulated Stability Margin versus Time Plot Using the Class K Cesaroni K260-CL Rocket Motor Current Motor Selection Based on the excellent simulated results produced by the Cesaroni K260-CL motor in terms of the mission performance requirements, as well as the fact that it satisfies all motor requirements discussed in the Student Launch Handbook, it has been selected as the motor for use in the full-scale rocket Motor Integration The motor and motor casing are inserted into a motor tube which is attached to a PVC thrust bulkhead and two aluminum centering rings. The thrust bulkhead and centering rings are adhered with epoxy to the inside of the airframe and they take the full thrust load of the motor during flight by transferring thrust from the motor to the motor casing and then to the centering ring and thrust bulkhead itself. The top of the motor rests in a cutout of the aft face of the thrust bulkhead to aid in axial alignment. The motor also rests in the aft end of the boattail. After the motor is inserted into the motor tube, the boattail is attached. It is attached by fasteners, which pass through the boattail to the centering ring, holding the boattail to the rocket. Helicoils are used to strengthen the threads in the plastic boattail to prevent failure from loading or fatigue due to many fastening cycles. The aft end of the motor is shaped closely to a cylindrical cutout of the boattail in which it rests. This provides alignment and prevents movement of the motor during flight. The motor is held in place with a motor retention ring. The motor retention ring is fastened to the boattail. The middle section of the retention ring is hollow so exhaust from the nozzle may exit the motor and vehicle Flight Dynamics The two major objectives in fin design included a target apogee of 5,280 ft above ground level and a stability margin off the launch rod of 2.0. The altitude at apogee was a parameter given by the requirements for the competition. The static stability margin, which is required to be a minimum of 2.0 off the launch rod, can also drastically affect the flight of the rocket. If the rocket is under stable, it will not be able to correct itself properly after experiencing wind perturbations. If it is over stable, it will correct its flight path too much

32 32 and fly horizontally into the wind. Typically, stability margins below 1.0 are considered to be under stable, and stability margins above 4 are over stable. To increase flight predictability, stability margins should lean toward the over stable range. This provides further justification for the stability margin objective. The selection of a suitable stability margin was a priority to minimize the possibility of either of these circumstances. In previous competition flights, static stability margin was considered, but test flights showed that rockets flew on unpredictable paths due to understability at launch rod clearance. Therefore, it was necessary to consider the dynamic stability margin at this flight event. Instabilities arise when a wind or a disturbance guides a rocket off its initial flight trajectory. A disturbance may cause the rocket to rotate or translate in some direction, meaning that the axis of the rocket is no longer aligned with its velocity. The rocket pitches around its center of gravity. Its stability will affect its ultimate flight path. An unstable rocket is likely to be pushed off its fight path by a wind such that the rocket is likely to tumble over its aft end. Stable rockets, however, tend to counteract disturbances by altering the angle between the rocket and the oncoming wind. This angle, the angle of attack, tends to zero for stable rockets. Thus, stability can be achieved without any external control. Stability was improved by optimizing fin design. Through numerous iterative attempts of running simulations and changing fin height and root chord incrementally, the optimal design point was found to be a fin height of 6 inches and a root chord of 6 inches. This returned a dynamic stability margin at launch rod clearance of 4.11 with an expected apogee of feet. In the future, surrogate modeling and Kriging methods may be used in combination with computational fluid dynamics (CFD) simulations in STAR-CCM+ to optimize the curve of the boat tail for the least drag and least mass added. The surrogate modeling and Kriging techniques would be executed using MATLAB to minimize a cost function of two fin design parameters, most likely fin root chord and fin height, since these most affect stability margin and apogee. The fin and boat tail CFD simulations could be further verified by wind tunnel test data. Kriging methods could be used to optimize performance for various design parameters, and will be investigated in the future for these uses Recovery Subsystem The recovery subsystem of the rocket must safely return the vehicle back to Earth within the guidelines stated in the NASA Student Launch handbook. The rocket will employ two systems to achieve this task. The first is a standard drogue and main parachute deployment at specified altitudes, while the second is the use of an auto-rotation system. The recovery systems will utilize a black powder ignition system to trigger the separation events at the designated altitudes. The forward section of the airframe houses the drogue and main parachutes which will have a dual deployment parachute system. The aft section houses the autorotation system and a redundant main parachute system Deployment Process The rocket will experience its first separation event at apogee where a black powder charge situated on the avionics subsystem bulkhead between the avionics subsystem and the aft section is ignited. A Statologger altimeter will send an electrical signal to an e-match which will ignite the black powder, separating the rocket into two different sections and releasing a drogue parachute connected to forward section. This separation event will separate the experimental payload situated in the aft section which will return to earth through the use of an auto-rotation system, rather than a parachute.

33 33 A second separation event will occur in the forward section at 1000 feet, where the Stratologger altimeter will send an electrical signal through an e-match to the black powder charge situated between the avionics bay and the nose cone, releasing the nosecone. Figure 18 below details the separation events and their locations on the airframe. Figure 18. Separation Events and shock cord lengths The nose cone, avionics subsystem, and aft section will be held together with #2 Nylon shear pins. Four shear pins will be placed where the nosecone and avionics subsystem connect and where the avionics subsystem and the payload section connects. #2 Nylon shear pins are rated at a maximum of 63 lb f, which is above the kinetic forces the rocket will experience during flight. The shear pins will break during the separation events. When the black power charges are ignited, the gas forces the airframe sections apart, breaking the shear pins and releasing the parachutes/payload. The black powder charges are stored in aluminum cylinders installed on the outside of the avionics subsystem bulkheads. Electronic matches will be inserted into the aluminum cylinders and connected to their respective altimeter. To protect the parachutes and airframe from the hot gases, a flame retardant material will be packed in with the black powder charges. The recovery system will make use of four black powder charges, a set of two on each of the avionics subsystem bulkheads. One acts as the primary system of separation, while the second is a redundant back-up system. One Stratologger altimeter will be the primary altimeter and be set to ignite its black powder charges at apogee and 1000 ft. respectively, while the second Stratologger altimeter is the back-up one set to ignite its black powder charges at apogee + 2 seconds and at 500 ft. The aft section has a back-up main parachute system that will deploy based on designated flight parameters not being met Deployment Method Determining which separation method to use required taking a closer look at the method chosen for the NASA Student Launch rocket. A CO2 ejection system was used to trigger the separation events. The CO2 ejection system has been reassessed this year in comparison to a simple black powder ejection system. Table 2. Separation System Comparison details the comparison.

34 34 Table 2. Separation System Comparison System Cost Reliability Manufacturability Weight Safety Black Low High Simple Medium Low Powder CO 2 Ejection High Medium 1 Moderate Medium Medium 1. Suffered critical failure during competition launch. The first factor to consider is the cost between the two systems. More-so, the repeated costs of testing and flight of the two systems, rather than the initial costs of developing the system. The black powder charges are very cheap, with a 16 oz (454g) container of costing around $30 [9] lasting for hundreds of tests and launches, while a 25 cartridge pack of CO 2 canisters costs $20 [10] and good for about a dozen launches, using a dual deployment system. Even though the CO 2 ejection system was chosen as the primary separation system last year, the redundant back-up system was black powder charges because of the reliability of the system from previous experiences. The ground based tests of the CO 2 ejection system show ample reliability to continue with the program and the CO 2 system proved reliable during two test launches in early spring However, the CO 2 system saw a critical failure during the NASA competition launch in 2016, nearly resulting in a catastrophic failure. The manufacturability of the two systems are roughly the same. They require aluminum tubing housing to be manufactured and installed onto bulkheads, although the CO 2 system required some additional components, such as several small 3-D printed parts necessary for housing the CO 2 system in the aluminum casing. The weight of the two systems were very similar, both requiring aluminum housing for operation. The primary difference between the two systems is that the CO 2 system requires use of a CO 2 cartridge, which adds a bit more weight to the system in comparison to the black powder system. Safety of the two systems is based around two factors. The first is the safe installation and preparation before launch. Both systems required black powder of some amount in order to function. The CO 2 system used a very small black powder charge of 0.75 grams to push the CO 2 canister into the puncturing device. The black powder system uses a larger amount, about 2 grams, in order to trigger the separation event. Use of black powder in general is a dangerous prospect, thus reducing the safety rating of each system. The second factor for safety involves what occurs during the separation event. The CO 2 system is an exhaustless system upon initiation that has minimal chance of damaging parachute or shock cords during separation. Black powder on the other hand releases hot particles during the separation event. If the parachute is packed too closely to the blast cap, there is a chance that the black powder change may cause damage to the parachute during the separation event

35 35 Lastly, and the largest factor in which system to use, is the necessity of the system. One of the primary factors to switching away from a black powder system to an alternate system is the target altitude of the rocket. The NASA Student Launch target altitude is 1 mile, which is well within a safe operating range for black powder, thus not requiring an alternative method. Because of this, the CO 2 ejection system does not have enough positive aspects in comparison to the black powder system to justify continuing the CO 2 ejection system for the NASA Student Launch rocket Attachment Scheme The attachment mechanisms for the recovery system will involve shock cord, U-bolts, and swivel links. U- bolts will be installed on the bulkheads that will connected to a parachute. These bulkheads are the two bulkheads on the avionics bay and a third bulkhead installed in the nosecone. Shock cord will tether the nosecone and the avionics bay together after the second separation event occurs. The other avionics bulkhead with have a length of shock cord attached to the drogue chute. The parachute used as the aft section back-up will be connected to a U-bolt installed on the auto-rotation system. Swivel links will be attached to the end of the shock cords via carabiners. The swivel links will attach the parachute cords to the shock cords. The swivel links are to provide a greater range of movement for the parachute during the separation event and during descent, to reduce the chance of the parachute cords becoming tangled. The length of shock cord will be chosen to reduce the chance of tethered objects colliding with each other during descent. Figure 18 details the length of shock cord used for each section during recovery. The ratio of shock cord length between the nosecone and avionics bay is 2.5:1 to reduce the chances of collision Parachute Sizing NASA Student Launch guidelines require that the maximum amount of kinetic energy any section of the rocket land with is 75 ft-lb f. This competes with the desire to return to earth quickly in order to minimize drift and the land within the designated launch area. For the sake of having an additional factor of safety, the aim is to have no section impact the earth with a kinetic energy greater than 65 ft-lb f. The parachute sizes are determined from the heaviest section that the parachute is attached to. In the forward section, the avionics had the heaviest weight of 2.9 lbs, which results in a maximum velocity of 37.7 ft/sec. The aft section has a weight of 7.6 lbs which results in a maximum velocity of 23.5 ft/s. Equation D = 8mg πv 2 C D ρ maximum velocity. (1) details how the diameter of the parachutes can be determined with respect to the D = 8mg πv 2 C D ρ (1) The other determining factor in what the parachute diameter will be is based on the coefficient of drag (C D) of the parachute. The coefficient of drag of the parachute is a function of the shape of the parachute Table 3 detail comparisons between various parachute shapes commonly used for main parachutes.

36 36 Table 3 Parachute Comparisons 11 Shape Vertical Stability 1 Coefficient of Drag Cost Cruciform Good at all speeds 0.4 Medium Flat Fair at low speeds, poor elsewhere 0.7 Low Panel Good at all speeds 1.1 Medium Elliptical Best at low speeds, good elsewhere 1.6 Medium Toroidal Best at low speeds 2.2 High 1. Vertical stability is referencing the parachute's tendency to sway during descent at a designated altitude. Referring back to equation D = 8mg πv 2 C D ρ (1), the coefficient of drag will be a significant component in choosing the diameter of the parachute. A high C D will result in a smaller diameter, but the costs tend to be higher. The toroidal shaped parachute offers the best C D while being the most expensive. The next cheapest alternatives are the cruciform, panel, and elliptical shapes. The cruciform's C D is considerably low that the parachute diameter would be enormous in comparison to the other two sizes. Between the panel and elliptical shapes, the elliptical parachute's higher C D would result in a smaller parachute size than the panel. Comparing the increased cost for smaller parachute size of the toroidal and the elliptical parachutes, it would come down to the overall necessity of a smaller parachute. For the current designs, packing efficiency is a high priority, thus the elliptical parachute is the chosen parachute shape. With the shape chosen, the diameters of the parachutes needed for the sections can be determined. Table 4 below shows the parachute diameters for both the forward and aft sections to comply with a kinetic energy limit of 65 ft-lb f. Table 4. Parachute Dimensions Forward Section Aft Section Parachute Size 14 inches 36.5 inches The data above lists the bare minimum parachute diameters to comply with the designated 65 ft-lb f chosen. Since parachute dimensions are sold in nominal sizes, the parachute diameter purchased will be the next nominal size up, further increasing the factor of safety of the descent. 11 Engelgau,Gene.Rocketry_Recovery_Technology._NARCON

37 Component Selection The components for each aspect of the recovery system must be able to withstand the forces imparted to the rocket during each separation event. Each component has a max rating that it can handle before failure and these ratings must exceed the maximum forces experienced by the rocket. The maximum forces are applied during each separation event when either the drogue or main parachute are deployed, resulting in a sudden deceleration of the rocket. Figure 19 below shows an Open Rocket simulation for the deployment of the drogue and main parachutes in the forward section. Figure 19. Open Rocket simulation depicting drag forces imparted onto the rocket Drogue deployment sees a drag force of roughly 7.5 lb f and the main parachute deployment experiences a drag force just under 37.5 lb f. These values will be compared to the components selected for the attachment scheme. Table 5 details these comparisons and lists a factor of safety for each component based on the particular separation event experienced. Table 5 Recovery Components Factor of Safety Component Max Rating (lb f) Factor of Safety: Drogue Factor of Safety: Main Swivel Link U-Bolt Avionics/Nosecone Bulkheads 1/2" Kevlar Shock Cord Paracord

38 38 #2 Nylon Shear Pins Caribiners Mission Performance Predictions Performance Criteria The rocket should have a stability margin value between one and three so that the rocket will remain stable during flight, while also not becoming over-stable. Other mission performance characteristics include: The rocket must fly safely to an apogee of 5,280 ft. AGL. On descent, the rocket must deploy a parachute that will permit safe landing within 1 square mile of the launch pad. There needs to me minimal to no damage to the payloads or the launch vehicle itself, such that the rocket and payloads can be reused the same day with few to no repairs. The kinetic energy of all independent or tethered sections of the launch vehicle must not exceed 75 ft-lbf at landing. All independent or tethered sections of the vehicle must not travel more than 2,500 ft. of the launch pad, assuming 20 mph winds. The motor used must not have a total impulse exceeding 5,120 N-s (L-class). The autorotative descent payload must be able to flare at 200 ft. AGL to reduce descent speed. The landing hazard detection camera must be able to analyze images of the launch site for possible landing hazards. The launch vehicle and autorotative descent payload must broadcast their locations to a ground station via GPS Simulations Flight Simulations were conducted using OpenRocket software, which utilizes a 6-DoF Runge-Kutta simulation to predict rocket flight with dynamic wind perturbations and accurate geometry to give a true approximation of the behavior of the rocket during flight. The software calculates center of pressure and center of gravity using similar methods to Barrowman. These simulations were run with conditions similar to what those at the launch site in Huntsville, Alabama will be: approximate global position, elevation of launch site above sea level, average temperature, pressure and wind speed for what is predicted in the month of April, and a launch rod of 144 inches. These conditions are shown in Table 6.

39 Altitude (ft) University of Florida 39 Table 6: Launch Site Simulation Conditions Huntsville, AL Launch Site Conditions Avg. Wind Speed (mph) 20 Avg. Temperature ( F) 63 Pressure (atm) 1 Latitude ( ) N Longitude ( ) W Altitude (ft) 600 The complete flight of the rocket, including the deployment of the autorotation payload, was modeled using OpenRocket. Because the payload separates from the fore section of the rocket, simulations determined the behavior of the fore and aft sections after separation. The data for the aft section assumes it is neither falling under parachute nor being slowed down by autorotation. The descent of the autorotation payload cannot be simulated in OpenRocket. Numerous simulations were done at varying wind speeds to determine how the rocket will behave. In order to determine the best wind speed to use for the final plots, weather data from the last 10 years at the launch site was analyzed, and it was determined that 20 mph was a good baseline to analyze the rocket against. The plot of altitude versus time can be found in Figure Altitude Time (s) Motor burnout Apogee Main parachute deployment Figure 20. Altitude vs. Time The increasing section of the graph represents the ascent of the rocket to apogee, which is broken into both the powered and unpowered flight regimes. Powered flight starts at 0 seconds and ends at motor burnout 8.70 seconds later. The latter is represented by the rest of the ascent, which begins at burnout and ends when the rocket reaches its apogee of 5,176 feet at 19.5 seconds into flight. At this point, the fore and aft sections separate, and the drogue parachute will deploy from the fore section which will begin its return to Earth, represented by the steady decline from apogee to 225 feet. Once reaching 225 feet, the fore section will deploy its main parachute, which will slow it down to its landing speed. Since the main parachute is

40 Velocity (ft) University of Florida 40 significantly larger than the drogue, it has more capability to slow the rocket down, hence the change in inclination of the altitude curve from 225 feet to the ground Velocity Time (s) Motor burnout Apogee Main parachute deployment Figure 21. Velocity vs. Time Figure 21. depicts the rocket velocity throughout the entire flight. Initially, there is a sharp increase in the velocity, which is caused by the motor accelerating the rocket and causing a momentum change. Maximum velocity of 507 feet per second is reached right at the motor burnout time of 8.70 seconds. After this, the rocket switches to the unpowered flight regime, and the velocity will continue to decrease until it reaches apogee, where it will reach zero. Right after apogee is reached, the velocity moves into the negative range, meaning direction is now reversed, and the rocket is returning to earth. At 23.3 seconds, the velocity levels out at roughly 49 feet per second, which corresponds to the drogue parachute deploying and restricting the free fall speed. At seconds into flight, the main parachute will deploy, which is represented by the quick decrease to 20.4 feet per second.

41 Acceleration (ft/s^2) University of Florida Acceleration 50 Motor burnout Apogee Main parachute deployment -100 Time (s) Figure 22. Acceleration vs. Time Figure 22 shows the acceleration of the rocket for the duration of the flight. During the powered flight, the rocket experiences its maximum acceleration of 128 feet per second squared during the motor burn. From 0 seconds, the graph shows a rapid increase of acceleration to its maximum value. After motor burnout, a sharp drop signifies that the only acceleration acting on the rocket is that of gravity and the drag force. The drag is a function of velocity squared; as the rocket slows down, the drag force decreases, and as the rocket reaches apogee, the only acceleration of the rocket is the acceleration of gravity at 32.2 feet per second squared since the rocket velocity is zero at that point. After drogue deployment at apogee, which appears as a slight change in slope of the curve, the rocket will experience an impulse in acceleration as the rocket slows down to terminal velocity. During this phase of constant velocity, the net acceleration will be 0 until the main deploys. Since the main deploys rather violently, the rocket will experience a quick increase in the magnitude of acceleration as it shifts to a negative velocity of lesser magnitude, and this can be seen as a spike in the graph. The net acceleration will then return to 0 for the rest of the descent until landing.

42 Drag Force (lbf) University of Florida Drag force Motor burnout Apogee Main parachute deployment Time (s) Figure 23. Drag Force vs. Time Shown in Figure 23 is the drag force experienced by the rocket over the duration of the flight. Initially the rocket has very high velocity as it moves through its powered flight regime. During this time, it is heavily dominated by drag caused by the velocity of the rocket and the overall outer diameter of the area. The rocket reaches a zenith of 20.9 pounds of drag at max velocity. After burnout, the velocity decreases until it reaches apogee, which as previously stated has zero velocity at that point which implies that the drag is 0. However, the graph shows a step jump which is representative of the drogue deploying and causing drag to increase until it reaches the maximum drag force the parachute can give. The large spike at 126 seconds corresponds to main parachute deployment, which is measured to be 68.5 pounds at the peak. It causes a large jump because of the rapid change in acceleration experienced briefly as the parachute deploys. One of the main design parameters is to ensure the rocket will survive this shock, which is why extensive testing will be done on the airframe to ensure it can handle the load Vehicle Stability For a rocket to be stable during flight, the center of pressure must be located aft of the center of gravity. An ideal static stability margin for safe stable flight should be between 1 and 3.5 body tube diameters, but it is preferable for the value of the margin to be above at least 2. The center of gravity (CG) for the rocket is located at 72.5 inches from the nose cone tip, and the center of pressure (CP) is 56.1 inches from the nose cone tip. The static stability margin can be calculated as in SM= x CP x CG (2), d where SM represents the static stability margin, x CP is the position of the center of pressure relative to the nosecone, x CG is the position of the center of gravity relative to the nosecone, and d is the diameter of the rocket. SM = x CP x CG d (2) The static stability margin of the rocket was determined to be Though this number would normally lie in the overstable range, some added stability was introduced because the landing legs are likely to

43 Stability Margin (cal) University of Florida 43 produce a significantly high amount of drag to reduce stability. Figure 24 is a visual representation of the rocket with CG and CP marked in their respective locations. CP is shown in red, and CG is shown in blue in in Figure 24. Relative CG and CP locations The stability margin off the rod will be As the propellant in the motor burns, the overall rocket mass will drop by about 43.8 ounces; as this occurs, the CG will move forward toward the nose cone until the burn ends. The variation of CP will be very small compared to the change in CG, although it will tend to oscillate slightly as the rocket passes through its coasting phase. The stability margin over time is plotted in Error! Reference source not found Stability margin Motor burnout Apogee Time (s) Figure 25. Stability Margin vs. Time The rocket starts off stable, and stability margin increases at a fairly steady rate for the first few seconds of flight. Initially the stability margin sees an overall increase. This stems from the decrease in mass from the motor burn, which causes the CG to shift forward in the rocket. After the powered flight regime ends at 8.70 seconds, the CG stops moving. Also encountered is the introduction of wind gusts. These gusts lead to rapid oscillations in stability as the rocket goes through the ascent phase where it attempts to correct itself

44 44 from encountered disturbances. The overall trend of the graph descends slightly downward as the rocket moves through the air until the parachute deploys at apogee, after which stability goes to 0 because it cannot be measured anymore by the software. Overall from the graph, the rocket starts off less stable but quickly corrects itself to become very stable during flight and maintains it for the duration. For the majority of the flight, it falls above the established stability margin range of 1.5 to 3.5, but since OpenRocket cannot adequately model the drag of the landing legs, stability was added to the design to increase stability with landing legs Kinetic Energy during Descent Kinetic energies calculated for each section based on attached parachute. Table 7. Kinetic Energy of Descent Parachute Nosecone Forward Section Aft Section Forward Main (15 in. diameter) Aft Parachute (42 in. diameter) N/A N/A N/A Drift Scenarios Figure 26 and Figure 27 show drift calculations for varying wind speeds. The solid lines represent the maximum areas within which the aft section of the rocket can land. The dotted lines represent the same values for the autorotation payload. This will help identify safety concerns for spectators on the day of the launch. Also, in the event of GPS failure, this will allow for a good approximation for where to search for the rocket so that it can be recovered. Figure 26. Drift radius for forward section

45 45 Figure 27. Drift radius for autorotation payload

46 Dimensional Drawings Figure 28. Avionics Subsystem

47 47 Figure 29. Fin

48 48 Figure 30. Forward Section

49 49 Figure 31. Payload Subsystem

50 50 Figure 32. Aft Airframe without Payload

51 51 Figure 33. Mission Vehicle Profile

52 Mass Statement Table 8 Total Mass with Margin Component Mass (lbm) QYT Total Mass (lbm) Aft Body Tube Motor Retention Ring Boat Tail Centering Ring Thrust Bulkhead Motor Tube Fin Aft Coupler Tube Ebay Forward Bulkhead Ebay Aft Bulkhead Ebay/Forward Body Tube Electronics Protection Tube Blast Caps Nosecone Nosecone Bulkhead Threaded Rods U-Bolts Ebay Electronics Push Rods Rotation Shafts Blade Attachments Rotation Plate Payload G10 Cap Payload Electronics Bulkhead Payloads Electronics Blade Swash Plate Payload Parachute Container Fasteners Motor Landing Legs TOTAL MASS (lbm) % Margin TOTAL MASS with 20% Margin 22.24

53 Manufacturing Plan Facilities For the completion of the subscale rocket and full-scale rocket in coordination with testing and competition, an efficient and effective manufacturing plan must be developed. One paramount item to consider in this plan is available facilities. The team must be aware of available machining processes and equipment. Furthermore, all components must be designed in a way so that they can be manufactured with only the available machines and processes. The main facilities the team will utilize for manufacturing is the Student Shop, the Student Design Center at the Solar Park, and the UF 3D Rapid Prototyping Lab. The Student Shop is a machine shop on campus open during specific hours. Students under the Mechanical and Aerospace Engineering College who have completed the Design and Manufacturing Laboratory course, EML2322L, are permitted to handle the machines and tools in the shop for personal, academic, and extracurricular projects. Designated teaching assistants (TA s) operate the Student Shop and a University of Florida faculty instructor oversees them and the Shop. Within the Shop, there stand three milling machines (mills), an engine lathe, a vertical and horizontal bandsaw, and an assortment of tools and attachments. The Student Shop is where the majority of the manufacturing for both scaled rockets will take place especially in the manufacturing of metal components. This is the only facility that contains the highprecision mills and lathe that can turn down metal for student use. There will be many components that require the high-precision features that can be quickly and easily achieved on these machines. The Student Design Center at the Solar Park is a newly built and open warehouse center located approximately two miles from campus. The team is allocated a portion of the warehouse floor to store materials and tools as well as a workbench area to perform maintenance, assembly, and stationary tests on the rocket. In addition, there are a couple drill presses, sanding machines, vertical and horizontal bandsaws, and a large number of hand and power tools such as hand drills, screwdrivers, handsaws, power drills, and impact drills. The 3D Rapid Prototyping Lab, also located on campus, is the primary source of producing any non-critical plastic components as needed per the design. The benefit of 3D Printing is more complicated parts can be produced without the need for an experienced manufacturer to machine Components to be Manufactured The components to be manufactured have main component categories: structures, composites, payloads, and recovery. Structures Structures is the umbrella category that all components that are included in the manufacturing and assembly of the rocket, but not including the electronics or airframe pieces. The majority of structures manufacturing is internal rocket components in addition to components to be implemented in the payload bay.

54 54 Bulkheads divide the rocket into separate bays, add structural support to the airframe, and serve as mounting surfaces for other internal components such as the electronics boards. The bulkheads will be made out of plastic. They have a diameter that will allow a close fit inside the coupler tube or body tube, depending on their location in the rocket. If the bulkheads are being placed in inside the coupler tube, they will be manufactured to have a coupler diameter, and if they are placed inside the body tube, but not the coupler tube, they will be manufactured to have a body tub diameter. The manufacturing processes to produce the proper diameter will be approximately the same for the subscale and full-scale rocket. Components with a more consistent diameter and fit more effectively inside the rocket provide more support for the airframe and handle loading better. The subscale rocket will have very few components attached to the bulkheads and there will be much less loading on the bulkheads because of this. Additionally, the motor is smaller than full scale, meaning the subscale rocket does not go as high or as fast. The loading due to aerodynamic forces transferred from the airframe to the bulkheads is significantly less on the subscale when compared to the full scale. On the full scale however, there are more components attached to the bulkheads and the rocket will go higher and faster resulting in greater loading on the bulkheads. To successfully counteract these forces and prevent failure, the bulkheads must fit more precisely inside the rocket and must be of a higher quality than the subscale bulkheads. The centering rings are inside the coupler tube of the rocket and are considered coupler diameter. To achieve this diameter, the same processes used for the bulkheads are used. The new design of the centering rings has a larger thickness than previous designs in order to slide in without cocking and not require the use of dowel rods to align with the body tube and each other. The motor retention system will be composed of two aluminum parts. The first will be an aluminum centering ring that will lie inside the boat tail and rest on the bottom lip. The second will be a thin aluminum ring that will lie on the outside, bottom surface of the boat tail that has a diameter bigger than the motor nozzle but less than the motor casing. These two components are made in a very similar fashion. Initially, aluminum round stock larger than the diameter of both components will be turned down on a lathe to the proper diameter for each component. The diameter of the aluminum centering ring that lies inside the boat tail will be verified by sliding the boat tail over the round stock and checking for a snug fit. Then, the round stock will be cut into disks. Following this, both aluminum disks will be faced on the lathe to create smooth, parallel surfaces. The last process common to both will be boring the inner hole out on the lathe. The inner hole will be larger for the ring inside the boat tail because the motor tube must slide through it, while the diameter of the motor retention ring will be smaller than the motor tube. Although these rings will have a different diameter, the process will be the same for each. Drill bits are first used to create a hole that is large enough to allow a boring bar to be used. Then, a boring bar is used to increase the diameter of hole to the specified diameter. On the subscale, the hole may be too small for a boring bar. In this case, drill bits will be used to make a hole of the proper size. The ring that fits inside the boat tail is taken to the mill where threaded holes are created for the fasteners that attach both rings together. The motor retention ring is also taken to the mill where through holes are drilled in its face for the fasteners that connect it to the ring inside the boat tail. Lastly, the motor retention ring is faced on the mill until it reaches the proper thickness.

55 55 The payload bay bulkheads will be very similar to the bulkheads on the full scale rocket. They will be manufactured in the same way. The only difference in manufacturing is the addition of separate holes. These will be drilled the same way as the holes in the other bulkheads. The payload bay bracket is a custommade ninety degree angle bracket. Aluminum bar stock will be sawed and milled to create the ninety degree profile. The part will be clamped in a way that allows the end mill to create a fillet instead of a completely ninety degree turn. In addition, holes will be drilled in the bracket for fasteners. Composites The body tubes are made by wrapping carbon fiber and fiberglass sleeves over a PVC pipe which is wrapped in parchment paper. This allows for easy removal of the body tubes after the epoxy has cured. Epoxy will be spread over each sleeve layer with paint brushes before the next sleeve is put on and a tool to remove air bubbles will be rolled over the wet laminate. This will eliminate air bubble voids in the finished laminate maximizing our material strength. Last year our team made a spindle which will conveniently hold the pipe and axially spin it with an electric drill giving us an easy way to wrap the layup in wide electrical tape. When slowly and evenly wrapped in tape from one end to the other, the excess epoxy is squeezed out and the composite laminate is put under the pressure it needs for strong curing. The fins will be made by laying the carbon fiber and fiberglass weaves on two parchment paper-wrapped aluminum plates (with dimensions slightly larger than the finished fin size), half of the weaves on one plate and half on the other. The plates will have previously been machined with a large fillet along one edge using a CNC milling machine and the weaves placed on the sides with fillets. This machined fillet will be the radius of the fillets we desire on our finished fin base. A third plate will be used to form the fin base and fuselage attachment point, having one side CNC milled to the radius of our rocket. This third plate will be wrapped in parchment paper and have carbon fiber weave placed on the curved side. After applying epoxy to all weaves and having air bubbles removed with a roller-like tool, the two side plates will be placed together (with wet laminate between them and fillet edges coincident) and the third plate placed against the first two plates, its curved face resting in the filleted edges of the first plates. The entire layup will be placed vertically (third plate on the bottom) and C-clamps applied to the sides of the first two plates squeezing them together and weights placed on top of the entire setup to ensure the all carbon fiber is in tight compression. This method of fin manufacturing worked very well during our team s subscale rocket last year and provided us with strong fillets on our fin bases as well as a fin base that was very easily attached to the rocket fuselage. The finished fins will be attached to the outside of the rocket fuselage by spreading epoxy over fin base which will then be placed against the outside of the fuselage. Payload The swash plate and rotor hub will only be made for full-scale, as the complexity of the design does not allow for linear scaling as compared to the rest of the components for other subsystems for the subscale rocket. Both components will be manufactured through the 3D Rapid Prototyping Lab. The holes on the exterior perimeter would be drilled using hand tools or the drill press at the correct measurements. The blades of the autorotation system will undergo composites treatment in order to create the streamline shape conventional manufacturing machines will have difficulty producing. The blades will start as a foam core in the approximate shape of its design and will be wrapped in layers of carbon fiber and coated with epoxy. A final layer is composed of a material that does not adhere to the epoxy in order to squeeze

56 56 excess epoxy from the carbon fiber. The blades are then rotated while the epoxy dries over a period of 24 hours to which a Dremel tool will do any necessary alterations and sizing of the blades. The landing legs will be epoxied to each fin and layered over with carbon fiber. After each successive layer dries, the structure of the legs will be successively strengthened for impact with the ground upon landing. A small camera will be fastened to the base of one of the fins, for the purpose of collecting data necessary for controlling the system. Recovery Manufacturing processes for recovery systems involve manufacturing 5 aluminum casings for the black powder charges. The aluminum casings will be aluminum cylinders with a platform base. The platform base will have four threaded holes drilled for attaching the cylinder to the bulkheads New Member Introduction and Training The involvement of members throughout the design and manufacturing process is a paramount concern within the team. One of the main objectives of the team each year is to recruit new members and get them involved in what they are interested in. This is especially important for the manufacturing team which has many upperclassmen and few younger members. However, manufacturing becomes less meaningful if the younger members do not know what they are manufacturing as the upperclassmen already have an idea from previous experience with the team. Therefore, new and younger members were encouraged to join the other sub-teams in order to gain understanding of the design process from start to finish and to become familiar with what is needed to be done on their parts to be manufactured. A major requirement to be on the manufacturing team is to have prior manufacturing experience or have taken Design and Manufacturing Lab course, EML2322L, at UF. It is expected that interested members will already have some prior exposure if not experience along with basic techniques on the machines, specifically the milling machines, lathes, bandsaws, and drill presses. Once these interested members are identified, a manufacturing meeting is scheduled. In this meeting, members are given a PowerPoint presentation that includes: A rough schedule of when subscale, full scale, and payload manufacturing would take place. A preliminary list of parts to be manufactured Description of common manufacturing processes used during manufacturing such as finding zeros on a cylindrical part and proper use of a dividing head to drill holes in a circular pattern among others Following the introductory manufacturing meeting, a training and safety session is scheduled. Over a two week period, each manufacturing team member, new or experienced, is given a training and safety session in the Student Shop. In this session, the manufacturing and safety leads will show each member proper safety techniques for each of the machines that will be used during manufacturing. Additionally, common manufacturing processes are demonstrated on the machines and the member is allowed to get hands-on experience before manufacturing starts. The members are also shown how to sign up and reserve machines. At this point, members are prepared to begin manufacturing. In early November subscale manufacturing should begin in order to meet the subscale completion deadline of the end of November. If there is no time

57 57 during the fall semester due to circumstantial reasons, when the spring semester starts in early January, full scale manufacturing is set to commence as well. Through the early composite R&D process, numerous team members are given the opportunity to gain hands-on experience and be exposed to a research work pattern; this prior familiarity will allow for more fluid workflow during the official manufacturing periods. Dr. Ifju, a University of Florida professor, is the local expert in experimental composites. For the past several months, he has collaborated with the team to initiate the experimental phase of the composite R&D pursuit. With Dr. Ifju s on-campus composites manufacturing laboratory, the team s composite lead schedules a series of experimental workshops, in which various manufacturing proposals are tested and evaluated on their effectiveness. At these workshop sessions, a maximum of ten interested team members can register to participate in these hands on opportunities. There will be a relatively large manufacturing workload compared to previous years with the addition of another payload and an autorotation system and its complimenting recovery system to manufacture parts for. However, with the previous training completed, members are capable of completing manufacturing assignments in the safe and efficient manner needed to handle this workload. Composites Composites manufacturing gives new members many opportunities for hands on experience as we have many separate components to manufacture during both subscale and full-scale manufacturing times. With the number of individual components that need to be manufactured and the multiple steps that each component must go through from the layup process to post-processing, there will be many meeting times throughout the week so members with tight schedules will still have plenty of opportunities to participate in the manufacturing process Manufacturing Schedule Manufacturing for the subscale rocket is set to begin in late October or early November. The coupler tube is needed before manufacturing of the coupler tube diameter parts commences because the diameter of these parts must be checked to see if they fit properly inside the tube. When the coupler arrives with other shipped materials, it is immediately ready while the fiberglass used to make the body tubes comes in sheets. This means it will be approximately one week after the materials arrive until a body tube is made, and the body tube diameter parts cannot be manufactured until a body tube is available to check the diameter of these parts. Because the coupler tube is shipped completely finished while the body tubes still have to be made, the coupler diameter components are made before the body tube components. After the body tube components are complete, the motor retention system is completed. The last step is to ensure all parts are functional, and modify parts if needed. Only components that are finalized for the full scale rocket may be machined and in the fall most designs are still being altered. Certain components should be finalized earlier than others and these are scheduled earlier than components that are expected to undergo multiple changes. Should there be any advanced progress made on designs for any of the sub systems, manufacturing for the full scale rocket will commence in the late fall semester and roll into the spring semester. The payload bay components are bulkheads, bracket, swash plate, blades, and rotor hubs that will go on the full scale rocket. Accordingly, these components will be completed with the rest of full scale manufacturing.

58 58 Scheduling A schedule will be made for the manufacturing team. Each member will be sent a link to a Google Spreadsheet to fill out their weekly availability. After all members have completed this, the manufacturing lead will design a schedule that meets the team members availabilities. The manufacturing members are expected to manufacture during their assigned time slots each week. Each day, the lead will visit the members that are manufacturing to provide advice and check on their progress. Daily processes will be assigned and modified based on the progress throughout the week and these processes will be based on a predetermined weekly schedule of tasks. The use of a consistent, weekly schedule allows members to plan around other conflicts and to contribute to the manufacturing of the rocket. Inevitably, some members may run into scheduling conflicts from other events that come up throughout the semester such as family events, illnesses, job interviews, etc. In this case, the team members are asked to give the manufacturing lead notice as far ahead of time as possible. It is the responsibility of the manufacturing lead to schedule another member during the vacated time block or manufacture during that time himself. A situation may arise in which nobody can manufacture during an available time block. In this case, the team will have to skip the manufacturing session. As long as this contingency is minimized the team should have adequate time to manufacture all components for the team. Additionally, the manufacturing lead will set up a GroupMe text messaging service for the team members to use. This service allows team members to text the entire group at one time. If a question, problem, or scheduling conflict arises this will give not only the manufacturing lead the ability to respond, but other team members as well. Again, this increases efficiency of the team and helps mitigate possible mistakes. One problem the team may encounter is unavailability of machines. The Student Shop only has one lathe. If the lathe is taken by another person, obviously the team cannot use it. In this case, the manufacturing lead will assign another manufacturing process for that day. There is plenty of manufacturing work to complete, so another manufacturing process can be easily identified. This problem can be mitigated by reserving machines. This ensures the team will have a certain amount of access to machines each week. However, there is a cap to how many hours the team can reserve one machine per week. The team will prioritize manufacturing processes and reserve machines when the most critical manufacturing processes need to be completed. With proper prioritization and reservation of machines, critical manufacturing processes are guaranteed to be completed. Any kind of machine availability conflict will occur while completing less critical processes. In this case, another process can be assigned while the unfinished process will be completed on another day. Scheduling for composite meetings and manufacturing gatherings will mostly be done over so that all rocket team members are aware of our proceedings and can participate. As the manufacturing process advances, the composites manufacturing process will only require one or two members to remove the mold from a finished laminate so scheduling in that regard will simply be done be word of mouth (since many members have classes together) or by text message.

59 59 Preliminary Weekly Schedule for Component Manufacturing October 31 November 5 Subscale Rocket Begin Body Tubes Composite Work Coupler Diameter Bulkheads Begin Electronics Bay Body Tube Bulkheads Recovery System November 6 November 12 Subscale Rocket Finish Body Tubes Composite Work Electronics Bay Fin Construction Motor Retention System Overall Assembly January 9 January 14 Full-Scale Rocket Preparation, Logistics, & Scheduling Coordination for Full Scale Manufacturing January 15 January 21 Full-Scale Rocket Begin Autorotation / Vertical Landing Payload Begin Vertical Landing Avionics System Begin Electronics Bay Begin Body Tube Composite Work Coupler Diameter Bulkheads January 22 January 28 Full-Scale Rocket Begin Body Tube Composite Work Finish Autorotation / Vertical Landing Payload Finish Vertical Landing Avionics System Body Tube Bulkheads January 29 February 4 Full-Scale Rocket Fin Construction Complete Body Tube Composite Work

60 60 Motor Retention System Recovery System February 5 February 10 Full-Scale Rocket Overall Assembly

61 Interfaces and Integration Plan Integration Strategy The integration strategy for the launch vehicle shall follow the V-model development process for verification and validation of the rocket system design. This model was chosen as it is the most commonly used development process used in industry for requirements analysis, verification and validation for systems engineering. This model also provides guidance to minimize project risks, improve quality, and improve the communication between all subsystems. Figure 34. V-model Integration Strategy The first part of the verification process will be the requirements analysis phase. The launch vehicle system must comply and be verified with all requirements in Table 16, and

62 62 Table 21. This shall be followed by low and high level design, and the implementation of hardware & software, which can be found in Section 3.8.2, implementation strategy. As each component are designed and created, the assemblies shall be verified by testing methods outlined in Testing Plan Table 22. Once all assemblies and subsystems have been verified, the final product will be validated by officials at approved launch sites to determine if it is ready for operations. Any obstacles encountered in the development process shall be traced back to the requirements definition and design process, to be decomposed again for implementation and validation before review and operations Implementation Strategy The system shall be implemented by dividing the product and analyzing each subsystem, and how each component is connected/interfaced with other components and subsystems. Figure 35. System Configuration The launch vehicle has four main components which are mechanically assembled. The airframe consists of two sections, the forward section, which will be formed by a mechanical assembly of the avionics and recovery subsystem, and the aft section, which will be formed by a mechanical assembly of the payload and propulsion subsystem. Other components which will be mechanically assembled to the airframe are the nosecone, fins, and a boat tail which will be mechanically assembled to a motor.

63 63 Figure 36. Avionics Subsystem Integration The Avionics subsystem will consist of three 9V batteries, two altimeters, and the GPS/Antenna/Xbee. Each battery will be mechanically interfaced with both altimeters and the GPS/Antenna/Xbee. The Altimeters will have an electrical interface with the recovery subsystem, for signaling the deployment of parachutes at desired altitudes. Figure 37. Recovery Subsystem Integration The recovery subsystem consists of the mechanical assembly of black powder ejection charge, the drogue parachute is mechanically interfaced with the avionics subsystem, and the main chute is mechanically assembled with the nosecone and the avionics subsystem.

64 64 Figure 38. Payload Subsystem Integration The payload subsystem consists of a parachute and Arduino, which is electrically interfaced with a landing detection camera, a GPS, a battery, and a barometric altimeter. The Arduino is also interfaced to a servo, which is electrically interfaced with the autorotation system, and an e-match, which will trigger the black powder charge to deploy the parachute Impact Assessment Should the primary payload design not meet the mission requirements and control decent, specifically by not providing enough lift to compensate for the weight of the payload subsystem and the aft section of the airframe, this would require a redesign of the entire launch vehicle. To provide more lift would mean larger and/or more blades, or a powered system to combat the free-falling of the aft section of the airframe. This could also mean decreasing the weight of the system being controlled in its descent, meaning that the payload would not interfere with the rest of the aft section in controlling descent. This would shift the new payload design to the forward section of the airframe, behind the nose cone, and coupling the avionics behind the payload subsystem. A drogue parachute will be interfaced between the avionics and payload subsystems, and there will be a separation event between the propulsion subsystem from the thrust bulkhead and the avionics aft bulkhead where the main parachute will be deployed. This scenario leaves the nose cone, recovery, and the new payload subsystem to be further discussed in how they will interface during the descent. Although structurally and aerodynamically unfavorable, one of the mission objective favorable suggestions would be to run a shock chord from the avionics subsystem along the exterior of the forward section of the airframe, to the nose cone U-bolt in the forward section. This would decrease the number of separated bodies coming down to two, instead of three. The nose cone will be mechanically assembled to a bulkhead using epoxy, with a U-bolt fastened to the bulkhead. The new payload design will likely need to have a piston system much like the current design, to separate the forward section and the aft section of the airframe, with the nose cone interfaced to the forward section using shear pins, to deploy the autorotation blades and landing legs at apogee. Excess chord will allow for larger degree of distance separation between the forward and aft section for the new payload design to begin its mission

65 65 in controlled descent, while the remainder of the launch vehicle implements the recovery subsystem to make a safe landing Phase 1 Integration This section will refer to section 6.5.2, Integration strategy, to identify each step in the integration process. Payload choice was made prior to proposal by payloads sub-team, and approved by faculty adviser. Aerodynamics, Avionics, propulsion and recovery sub-systems went through changes such that the payloads will be delivered as stated in the payloads verification and requirements Avionics sub-team will work closely with payloads sub-team, and shall be responsible for all necessary implementation of hardware & software. Testing shall be conducted to verify proper functionality of implemented hardware The Composites sub-team shall be responsible for the fabrication of vehicle airframe, and fins. Manufacturing sub-team shall be responsible for all other supporting components, such as centering rings and bulkheads. Sub-systems will be tested as components are made and assembled using necessary equipment, to assure quality of final system configuration. Sub-systems and final configuration verification will be conducted with safety officer and certified safety stewards before launches, to ensure safety procedures are implemented, as listed in safety verification plan. The team shall validate performance using analysis and simulations to assure compliance of NASA Handbook requirements Competition week operations and validation review shall be held in Huntsville, Alabama. All other sub-scale and full-scale test launches shall be conducted in approved NAR and TRA launch sites.

66 Revision History Significant changes and revisions have made to launch vehicle design since proposal Version Sub-System Revision Status Description Aerodynamics In Progress Avionics Completed Preliminary Bay Design Payloads In Progress Propulsion System Selected Motor Recovery System In Progress Structural Design Completed Preliminary General Design Refer to section 1.1 Changes made to Vehicle Design Refer to section 1.1 Changes made to Vehicle Design Refer to section 1.1 Changes made to Vehicle Design Refer to section 1.1 Changes made to Vehicle Design Refer to section 1.1 Changes made to Vehicle Design Refer to section 1.1 Changes made to Vehicle Design

67 Structural Interfaces The aft airframe (consisting of the motor mount system, fin assembly, and boat tail, and payload subsystem) interfaces with the avionics subsystem (consisting of both the main and drogue parachutes, and the recovery subsystem) through a coupler tube epoxied to the aft airframe. The avionics subsystem interfaces with the ogive plastic nose cone by being connected from the shock chord leading to the main parachute attached to the U- bolts on the forward avionics bulkhead and nose cone bulkhead. The distance between the avionics subsystem and the nose cone bulkhead will be determined in Test 6 and 7 of Table 22. The boat tail interfaces with the aft centering ring by being fastened with 6-32 threaded screws, and the motor retention ring is then fastened to the end of the boat tail by 6-32 threaded screws as well. The payload is fastened by 6-32 threaded screws to a permanently epoxied centering ring attached to the coupler tube connecting the aft and forward airframes. In order to push the aft airframe from the avionics subsystem at apogee, a bulkhead acting like a piston is used to separate the two airframes of the rocket and release the drogue parachute for safe descent of the avionics subsystem. The avionics subsystem is fastened together with 1/4 inch threaded rods in a stacked disc format in order to save space between the two subsystem bulkheads. The bulkheads then act as clamps holding the entire avionics subsystem together by enclosing around a coupler tube that will house the electronic components of the rocket Vehicle to Ground Station An Arduino UNO board with a coordinator Xbee-PRO 900HP module will allow the GPS analog data to be read directly on a laptop from router Xbee-PRO 900HP modules in the Electronics Bay and Payloads bay. The Arduino UNO will be directly connected to the laptop through USB. Figure 39. GPS and Communications Block Diagram

68 Vehicle to Launch System A set of two rail buttons along the side of the aft airframe will be inserted onto the rocket and will be used to attach the rocket to the launch rail. Rail buttons have a lower area of contact than launch lugs and the rail button material of Nylon has a low drag coefficient that most conventional launch lugs. The rail buttons also have a round profile meaning a tangential point of contact and less area of contact that would induce drag in comparison to launch lug geometry 12. Rail buttons reduce movement of the rocket on the launch pad allowing for a more predictable flight path of the rocket. The reduction in flight path on the launch rails thus produce a low amount of friction between the rocket and the rail in order to gain relatively more velocity as the rocket launches off the rail. A higher velocity off the rail also corresponds to more favorable stability margin which is important during the early period of the flight. 12

69 69 4. Safety 4.1. Safety Officer and Responsibilities The team safety officer is Aaron Dagen, a third year aerospace engineering student and third year member of the UF Rocket team. When Aaron is not available the project lead or the Systems Lead, Nahien Chowdhury, will act as Safety Officer. In addition to the safety officer, the team mentor, Jimmy Yawn, will be responsible for team safety as well during any processes that require his involvement. The following are the Safety Officer s responsibilities: 1. Enforcing all protocols outlined in the MAE Student Design Center: Rules for Facility Use document. This includes policies for personal safety; equipment use; facility cleanliness, organization, and respect; proper language; use of the Material & Tool List and Broken / Lost Tooling List; and all other miscellaneous policies. The MAE Student Design Center: Rules for Facility Use can be seen in the Appendix. 2. Having a strong understanding of each machine at the UF Student Design Center. The Safety Officer can t effectively train students in proper equipment use unless he/she possess a solid understanding of each machine and process. The Safety Officer will also be ineffective at proactively identifying and preventing mistakes that cause injury or damage. 3. Training students on machines, administer knowledge quizzes, and sign authorization sheets. If a student requests machine training, it is the Safety Officer s responsibility to train him/her to the standard expected and outlined in the safety protocols. After training, he/she will administer the knowledge quiz to assess their understanding of the safety protocols for the specific machine. If the student passes the quiz, the Safety Officer will add their name to the approved list of users for that machine so (s)he can use the machine with steward supervision in the future. 4. Verifying students are trained and authorized on each machine they use. The Safety Officer s primary responsibility is to ensure students are trained and authorized on each machine they use by referencing the lists of authorized users located by each machine. Students using machines on which they are not authorized lose facility use privileges, effective immediately. 5. Keeping watch over powered machinery as it is being used. Accidents can happen to the best trained users. Therefore even though all users of powered machinery must be trained to be allowed to use them, the Safety Officer should still keep a watchful eye to make sure that machinery is being used safely and correctly. 6. Ensuring students clean machines after each use and accept responsibility for stations not up to the UF Student Design Center cleaning standards. 7. Ensuring students keep the UF Rocket Team bay neat and clean in the UF MAE Student Design Center. 8. Understand and abiding by Federal Aviation Administration rules and regulations concerning highpowered rocketry, as listed in Title 14: Aeronautics and Space, Part 101, Subpart C Amateur Rockets.

70 70 9. Understanding, abiding by, and enforcing the NAR Safety Code for high powered rockets 10. Looking up the Material Safety Data Sheets for materials used by the team and making them easily available for the team to reference 11. Reading, understanding, and enforcing rules, regulations, and legislation concerning propellant and motors from: a. The National Fire Protection Agency b. The Bureau of Alcohol, Tobacco, Firearms, and Explosives 12. Attending manufacturing sessions not supervised by proper authorities, and enforcing proper safety protocol is observed. Manufacturing safety guidelines can be seen in Appendix C. 13. Overseeing all relevant vehicle and other team activity testing including: a. Design of vehicle and launcher b. Construction of vehicle and launcher c. Assembly of vehicle and launcher d. Ground testing of vehicle and launcher e. Sub-scale launch test f. Full-scale launch test g. Competition launch h. Recovery activities i. Educational Engagement activities 2. The Safety Officer is responsible for making sure that all members that qualify are to become Safety Stewards of the UF Student Design Center. Refer to the Safety Steward Responsibilities document that is listed in Appendix.

71 Project Risk Assessment The criticality of a risk or failure is identified as the product of the severity of failure and the likelihood of occurrence. The table below identifies how the level of severity or likelihood associated with an issue is defined. Table 9 Severity and Likelihood Scale Severity of Failure Likelihood of Occurrence 1 Minimal or no impact. Remote likelihood. 2 Some reduction in margin and some impact Unlikely to occur. 3 Significant reduction in margin and moderate impact. Likely to occur. 4 No margin and major impact. Highly likely to occur. 5 Unacceptable. Near certainty. The risk matrix in Table 10: Risk Matrix and criticality description in Error! Reference source not found. identifies when an issue is considered a low, moderate, or high risk. The risk plot used here will be utilized for the remainder of the report. Table 10: Risk Matrix Likelihood Severity Table 11: Criticality Descriptions Criticality Low Criticality Description Minimum impact on mission. Criticality is less than or equal to 6, except when likelihood is 1 and severity is 5.

72 72 Moderate Some impact to the mission. Criticality is greater than or equal to 8 and less than 15, except when likelihood is 1 and severity is 5. High Significant impact to the mission. Criticality is greater than or equal to 15. Table 12 identifies the risks associated with project. The effects of these risks, their criticality, and efforts to mitigate the risks are also included. Table 12: Project Risk Assessment Risk Effect Severity (1-5) Likelihood (1-5) Criticality Mitigation Vehicle surpasses available budget Insufficient funds to complete project More efficiently use money available. Utilize past years parts and knowledge. Develop a budget with additional funds set aside for overdraft Team members too busy to finish project Late submission of reviews, unfinished project Allow time for setbacks in the schedule. Divide work Injury to team members Injury to team members, team members unable to work, decrease in morale Strictly follow in-place safety procedures, educate members of possible associated safety risks. Difficulty of manufacturing Part unable to be manufactured accurately/true to design, parts require more time to be manufactured Have designs approved by manufacturing leads and Dr. Braddock. Ensure manufacturers have experience and manufacturing is done under supervision. Key personnel no longer available/not available for the competition in April Launch operations at competition are rushed, incomplete, or incorrect Ensure that all members know the project in detail. Keep all activities and all meetings documented and posted online. Prepare a guaranteed launch team significantly before competition. Identify alternate leads. Materials delivered late Team delayed in manufacturing, incomplete rocket Identify multiple vendors, order materials early, order materials that are in stock, track orders

73 73 Team members unfamiliar with available software* Gap of knowledge exists between team members; particular responsibilities can only be delegated to certain team members Hold sessions for tutoring new members in available software. Materials unavailable for purchase Team unable to work on project Choose common materials, identify multiple vendors and sources for less reliable materials Test results indicate faulty plan Redesign portion of the vehicle Simulate test with software Insufficient personnel* Portions of project incomplete/improperly done Incentivize member attendance, recruit help of colleagues and members of other teams. Unrealistic timeline* Portions of project incomplete/improperly done Adjust timeline to have available time overdraft/leeway. Hold members accountable for keeping deadlines. Low supervision Stray from original project design and objectives, safety procedures overlooked Coordinate schedule with supervisors. Keep close and open communication with mentors. Manufacturing equipment unavailable Compromise on rocket design, incomplete rocket or failure Find equipment elsewhere or find a separate method. Set up multiple options for manufacturing sites Damage to property Potential fines incurred, privileges to property potentially revoked Strictly follow in-place safety procedures, educate members of possible associated safety risks. Communication difficulties Portions of project incorrectly done, incomplete reports, Ensure all information is easily accessible and that all members know how to contact one another. Have frequent meetings where information is shared As demonstrated above, the highest risks in the project risk assessment are that the vehicle goes over budget, team injury, or the team members are too busy to complete the project.

74 74 The current budget has been thoroughly examined to include as many facets of the project as possible. In addition, a contingency of roughly 15% has been added to characterize any spending that is not yet expected. Both measures should counteract a surprising increase in cost. This can be seen in the funding plan in Section 5.2. Hopefully, more funds will be raised than required. Several steps have been and will continue to be taken to address the possibility that manufacturing equipment becomes unavailable. A plan that estimates the time necessary to manufacture each of the components has been developed. Time will be reserved in the student shop according to this plan. Additionally, if the shop is still unavailable, back-up manufacturing sites like the Solar Park and the Capps Sutton woodworking shop are available as less-desirable, but nonetheless functional facilities. Additional precautions can counteract a lack of time for team members. One of which is that a buffer zone has been added to the schedule for many milestones. Furthermore, each subsystem lead will work with a team that is well-informed and can take over responsibility should the leader be unavailable to work. Ultimately, even the greatest potential project risks can be mitigated Safety Requirements Table 13. Safety Requirements Req. Number Requirement Statement Design Feature Verification Method 4.1 Each team shall use a launch and The team s safety officer is Inspection safety checklist. currently creating a launch and safety checklist for the FRR report, LRR, and launch day operations. 4.2 For all academic institution teams, a student safety officer shall be identified, and shall be responsible for all items in section 4.3. Aaron Dagen is the team s student safety officer. Inspection The safety officer shall monitor team activities with an emphasis on Safety The safety officer shall Implement procedures developed by the team for construction, assembly, launch, and recovery activities The safety officer shall manage and maintain current revisions of the team s hazard analyses, failure modes analyses, procedures, and MSDS/chemical inventory data The safety officer shall assist in the writing and development of the team s hazard analyses, failure modes analyses, and procedures. The safety officer will attend team activities to monitor and emphasize safety. The safety officer will implement the team s safety procedures during team activities. The safety officer will maintain a current log of team s hazard analyses, failure modes analyses, procedures, and MSDS/chemical inventory data. The safety office will help the team with developing hazard analyses, failure modes analyses, and procedures. Inspection Inspection Inspection Inspection 4.4 Each team shall identify a mentor. Jimmy Yawn is the Team Mentor. Inspection

75 During test flights, teams shall abide by the rules and guidance of the local rocketry club s RSO. The team will properly communicate with the local rocketry club before attending any launches and abide by any rules at the launch. Inspection The team has reviewed and acknowledged regulations regarding unmanned rocket launches and motor handling. Federal Aviation Regulations 14 CFR, Subchapter F, Part 101, Subpart C, Code of Federal Regulation 27 Part 55: Commerce in Explosives; and fire prevention, and NFPA 1127 Code for High Power Rocket Motors documentation is available to all members of the team in the team safety manual. The University of Florida Rocket Team team understands and will abide by the following safety regulations declared by NASA. The following rules will be included in the team safety contract that all team members are required to sign in order to participate in any builds or launches with the team. 1. Range safety inspections of each rocket before it is flown. Each team shall comply with the determination of the safety inspection or may be removed from the program. 2. The Range Safety Officer has the final say on all rocket safety issues. Therefore, the Range Safety Officer has the right to deny the launch of any rocket for safety reasons. 3. Any team that does not comply with the safety requirements will not be allowed to launch their rocket Launch Procedures Pre-Launch Procedures The majority of the vehicle preparation will occur before launch day activities in the field. Major portions of the project can be assembled, transported, and prepared beforehand. By launch day, the vehicle should consist of four discrete groups: 1. The nosecone and the payload bay airframe containing the payload bay, the associated electronics, and the payload parachute located inside to the payload bay. 2. The main electronics bay airframe containing the main electronics bay, the main parachute tethered to the aft end of the electronics bay, and the drogue parachute tethered to the forward end of the electronics bay. 3. The aft airframe containing the motor centering mechanism with thrust bulkhead, centering rings, and boattail. 4. The motor casing, motor (unassembled/unloaded), and motor retention. Before attending the launch, each of these systems will be analyzed and checked. If any of the following are out of place, the appropriate response will be taken to remedy the situation. In this checklist and all the following lists, red text indicates a step that requires caution. Following a list item, the responsible lead is in parenthesis. The Systems Lead will ensure all checklists are completed properly and in order. On launch day, the responsible engineer will be required to sign off on a sheet to indicate the task has been completed properly, and the project manager will sign off to prove oversight of the task.

76 76 Group 1 1. Ensure payload electronics are properly functioning (Electronics Lead). 2. Examine nosecone for imperfections and faults (Structures Lead). 3. Examine payload parachute for rips, burns, or charring (Recovery Lead). 4. Fold payload parachute (Recovery Lead). 5. Examine payload bay airframe for imperfections or faults (Structures Lead). 6. Ensure payload bay functions properly (Payloads Lead). a. Rotors spin freely b. Linear actuator raises and lowers swashplate 7. Ensure all shock cord joints are securely fastened (Recovery Lead).. Payload bay attachment a. Payload parachute attachment 8. Ensure hardware is properly tightened in payload bay airframe (Structures Lead). 9. Ensure coupler tube and payload bay airframe are securely fixed by the epoxy (Structures Lead). Group 2 1. Examine main parachute for rips, burns, or charring (Recovery Lead). 2. Fold and pack main parachute into aft airframe (Recovery Lead). 3. Examine drogue parachute for rips, burns, or charring (Recovery Lead). 4. Fold and pack drogue parachute into forward parachute and payload holding area (Recovery Lead). 5. Examine main electronics bay airframe for imperfections or faults (Electronics Lead). 6. Ensure forward and aft airframes slide smoothly away from main electronics bay airframe (Recovery Lead). 7. Ensure all shock cord joints are securely fastened (Recovery Lead). a. Main electronics bay attachments b. Main and drogue parachute attachments 8. Ensure all hardware properly tightened (Structures Lead). 9. Ensure launch lug on main electronics bay airframe is securely in place (Structures Lead).

77 Ensure electronics bay coupler tube and main electronics bay airframe are securely fixed by the epoxy (Structures Lead). 11. Ensure all main electronics bay components are functioning properly (Electronics Lead). 12. Test continuity on all e-match terminals (Recovery Lead). 13. Verify proper wiring of all altimeters (Recovery Lead). 14. Charge batteries for at least four hours before flight (Electronics Lead). Group 3 1. Examine aft airframe for imperfections or faults (Structures Lead). 2. Examine fins for imperfections or faults (Structures Lead). 3. Ensure fins are attached securely to aft airframe by attempting to manipulate them (Structures Lead). 4. Ensure launch lug on aft airframe is securely in place (Structures Lead). 5. Ensure coupler tube and motor centering mechanism are securely epoxied to the aft airframe (Structures Lead). 6. Ensure all shock cord joints are securely fastened (Recovery Lead). a. Attachment at ballast mass. b. Attachment at drogue parachute. 7. Ensure all motor retention components are accounted for (Propulsion Lead). 8. Examine boattail for imperfections or faults (Structures Lead). 9. Examine motor retention mount and ring for imperfections or faults (Structures Lead). Group 4 1. Ensure all motor components are accounted for (Propulsion Lead). 2. Examine motor casing for faults and imperfections (Propulsion Lead). 3. Reread motor installation instructions (Propulsion Lead). In addition to the checklists above, the Systems Lead will ensure that the following items are packed and prepared for the launch: A. All necessary hardware and fasteners B. Hand tool box (screwdrivers, pliers, wrenches, etc.) C. Hand drill with 2 charged battery packs D. Fresh batteries E. Shear pins F. Pyrodex powder

78 78 G. CO2 cartridges H. Flame retardant recovery wadding I. Aluminum tape J. E-matches K. 5-minute epoxy and hardener L. Zip ties M. Scissors N. Spare wires and connectors O. Wire cutters P. Additional ballast mass Q. Computer prepared to simulate flight conditions R. Power inverter S. Ground station computer to communicate with the vehicle T. Igniters U. Small weighing scale V. Spoon W. Tape X. Multimeter Y. Tables Z. Small ladder AA) Launch day checklists Recovery Preparation On launch day, there are three major components to recovery actions in the field: tying and attaching shock cord, preparing charges, and folding the parachutes. Each has their own checklist listed below. The Recovery Lead is responsible for the Recovery Preparation checklists. The Systems Lead will verify the Recovery Lead has completed each list properly. Shock Cord Preparation 1. Verify proper lengths of shock cord are placed correctly in each airframe. 2. Tie and verify shock cord attachment between payload bay and payload parachute is secure. 3. Tie and verify shock cord attachment between main electronics bay forward bulkhead to forward piston is secure. 4. Tie and verify shock cord attachment between forward piston and main parachute is secure. 5. Tie and verify shock cord attachment between main electronics bay aft bulkhead to aft piston is secure. 6. Tie and verify shock cord attachment between aft piston and drogue parachute is secure. 7. Tie and verify shock cord attachment between ballast mass and drogue parachute is secure. Charge Preparation 1. Trim e-matches to appropriate length.

79 79 2. Attach e-matches to the proper terminals on the bulkhead, and place the head of each e-match into the blast cap. 3. Walk away to a minimum distance of 50 feet from all other individuals. Caution: Proper distance is essential for the safety of others. 4. Turn on the altimeters. Verify continuity on the payload parachute terminals for both the primary and secondary altimeter. Caution: Keep e-matches away from the body or face, on the chance they may be accidentally ignited. 5. Turn off altimeters and return to workstation. 6. Place small scale on level surface, shielded from wind. 7. Place parchment paper onto the scale, and zero the scale. 8. Carefully use the spoon to gradually place more Pyrodex powder onto the parchment paper. Caution: Take care not to accidentally spark Pyrodex 9. Continue to add Pyrodex until the scale indicates the proper amount of Pyrodex for the main parachute separation has been added. 10. Carefully transfer the Pyrodex to a blast cap on the parachute housing mounting plate of the payload bay. Caution: Take care not to accidentally spark Pyrodex powder. Load the blast cap with flame retardant wadding to compress the Pyrodex. 11. Secure Pyrodex and flame retardant wadding in blast cap by taping over the blast cap with aluminum tape. 12. Repeat steps 6-12 for each of the other remaining blast cap 13. Check that the aluminum tape is still firmly in place. Parachute Preparation On launch day, all three parachutes will be refolded for proper deployment. The procedure for packing the parachute was developed to ensure a quick and reliable deployment of the parachute after it has been ejected. First, the parachutes are laid out flat with all the shroud lines neatly pulled straight. The gores of each parachute are then neatly folded in half from the edges into the center, and repeated until the parachutes sizes fit within their allotted spaces. The remaining shroud lines are wrapped around their respective parachutes, taking care to not twist the lines. The parachute is then placed into their respective spaces and attached to the swivel links on the shock cords via a quick link. Motor Preparation Motor preparation will follow the instructions set forth by Aerotek for this particular motor. Two people will assemble the motor. A third person will supervise to ensure that all steps are carried out properly. A clear table far from any hazards or sparks reserved solely for motor preparation will be used. The Propulsion Lead will verify proper preparation of the motor and the Systems Lead will verify that the motor has been prepared.

80 80 Caution: This entire procedure should be treated with caution. The materials dealt with in this process are explosive and hazardous. Final Assembly 1. Once the recovery teams and propulsion teams have finished their preparations, the final assembly can begin. This process will be as follows: 2. Install new batteries, and secure into place (Electronics Lead). 3. Activate payload electronics (Electronics Lead). 4. Attach nosecone to payload bay airframe (Structures Lead). 5. Pack payload parachute into forward piston (Recover Lead). 6. Slide payload bay airframe onto forward end of main electronics bay airframe. Secure in place with shear pins (Structures Lead). 7. Slide aft airframe onto aft end of main electronics bay airframe. Secure in place with shear pins (Structures Lead). 8. Attach boattail to aft end of aft airframe with fasteners (Structures Lead). 9. Load motor into aft airframe (Recovery Lead). 10. Attach motor retention ring using fasteners (Structures Lead). Setup on Launcher When out at the launch site, the following steps must be taken. 1. Carry the vehicle out to the launch rail with at least 3 people (Structures Lead). 2. Bring a screwdriver, a small ladder, tape and the appropriate fasteners (Structures Lead). 3. Carefully slide the lower rail button into the slot in the rail (Structures Lead). 4. Move the rocket down until the second rail button is in place on the rail (Structures Lead). 5. Slide the rocket to the stopper on the rail (Structures Lead). Launch Procedure 1. All individuals except for one retreat to a safe distance 200 feet away from the vehicle (Safety Officer). 2. Activate electronics and recovery switches (Structures Lead). 3. Listen to ensure that all altimeters verify continuity (Structures Lead). 4. Confirm wireless communication with the rocket is functional (Electronics Lead). 5. Re-place the electronics bay within the vehicle (Structures Lead).

81 81 6. All non-essential personal return to the ground station (Safety Officer). 7. All members will remain a minimum of 200 feet from the launch pad, in accordance with NAR regulation (Safety Officer). Caution: any closer could place the individual in harm s way from the motor. 8. Have a small segment of the recovery team set up downwind of the vehicle to better monitor the descent. This will help locate the vehicle (Recovery Lead). 9. Send a current through the igniter to launch the vehicle (LCO). 10. All members will be actively paying attention and standing ready (Safety Officer). Caution: if the vehicle becomes unstable and flies in a trajectory towards bystanders, they must be prepared to move out of the way. 11. During the flight, the electronics team will monitor the incoming data (Electronics Lead). Troubleshooting Several unexpected scenarios may occur. This list of procedures aims to establish guidelines for several of these possibilities. Misfire 1. Turn off power to the launch area, preventing accidental ignition (LCO). 2. Do not approach the vehicle for five minutes, or until the all-clear is given (RSO). 3. No more than four individuals approach the vehicle (Safety Officer). 4. Remove the igniter (Propulsion Lead). 5. Inspect the igniter for discharge (Propulsion Lead). 6. If the igniter has not discharged, check for continuity (Propulsion Lead). 7. If the igniter is faulty, replace the igniter (Propulsion Lead). 8. If the igniter is not faulty, bring the vehicle back to the workstation to inspect the motor, and for other potential causes of failure (Systems Lead). Discontinuity 1. Follow this checklist in the event that one of the altimeters does not indicate that there is continuity between both main and drogue parachute ejection charges. The likeliest cause is a faulty connection. 2. One individual will turn off the altimeters to conserve battery and prevent premature detonation (Structures Lead). 3. Visually examine the interior of the main electronics bay airframe to identify the cause (Recover Lead). 4. If this cannot be readily seen, remove the vehicle to the workstation (Structures Lead).

82 82 5. Open the main electronics bay airframe to access the altimeter (Structures Lead). 6. Use a multimeter to measure continuity between smaller, discrete sections connecting the e-matches to the altimeters (Electronics Lead). 7. If the cause has yet to be found, replace additional wires and e-matches (Recovery Lead). 8. Check for continuity using steps 3-6. Post-Flight Inspection 1. Locate the vehicle with the GPS information, if visual location cannot be determined (Recovery Lead). 2. A team of four will approach the vehicle (Recovery Lead). 3. Upon approach, proceed carefully (Safety Officer). Caution: be aware of any debris or hazards that may be harmful. 4. Determine if any explosive charges are still armed. If so, disarm them immediately by cutting the e- matches (Recovery Lead). 5. Examine the rocket for structural damage (Structures Lead). 6. Take pictures of the vehicle and the landing site (Recovery Lead). 7. Record altitude and separation data by listening to the beeps from the altimeters (Recovery Lead). 8. Reassemble the vehicle for ease of transportation back to the workstation (Structures Lead). 9. Remove the used ejection charges and motor. Safely store them to discard properly (Propulsion Lead). Caution: let these cool temporarily. They may still be hot. 10. Clean the motor casing to remove residue from the expended motor (Propulsion Lead). 11. Disassemble the vehicle for easy transportation from the launch site (Structures Lead). 12. Record any anomalies (Systems Lead) Personnel Hazard Analysis Table 14. Personnel Hazard Analysis Source of Hazard Hazard Cause Result Severity Likelihood Criticality Mitigation Belt Sander Dust Particles in eyes and breathing Contusion Incorrect use of equipment -Irritated eyes, throat, or nose -Damage to person with cuts or burns Utilize protective eyewear and ventilation masks. DO NOT wear gloves or long sleeves. Sand in a

83 83 wellventilated area. Hand Drill Particles in eyes Contusions Lacerations Incorrect use of equipment -Damage to person with cuts or burns -Damage to rocket parts -Damage to hand drill and/or drill bit Utilize protective eyewear, close toed shoes, tie hair up, and remove all jewelry. Always assume the hand drill is plugged in and on. Never look away when working on a part. Be aware the hand drill could slip if part is smooth, NEVER do more work than the hand drill, be patient and let the hand drill cut. Dremel Particles in eyes and breathing Contusions Lacerations Incorrect use of equipment -Irritated eyes, throat, or nose -Damage to person with cuts or burns Utilize ventilation masks, protective eyewear, close toed shoes, tie hair up, and remove all jewelry. Always assume the Dremel is plugged in and on. Never look away when working on a part. Be aware the

84 84 Dremel could slip if part is smooth, NEVER do more work than the Dremel, be patient and let the Dremel do the work. CNC Mill Contusions Hearing Damage Lacerations Burns Incorrect setup of machine -Damage to CNC Mill -Damage to rocket parts Utilize protective eyewear, close toed shoes, hearing protection, tie hair up, remove all jewelry, and roll up sleeves. Be sure to use proper etiquette when using the CNC Mill. Always keep one hand on the stop button to cancel any operation that is potentially harmful. Never put hand within a six inch radius of the cutting zone. If part is hot then use a rag over your hand to remove part. Drill Press Particles in eyes Contusions Lacerations Incorrect use of equipment -Damage to person with cuts or burns Utilize protective eyewear, close toed shoes, tie

85 85 -Damage to rocket parts -Damage to drill press drill and/or drill bit hair up, and remove all jewelry. Grinder Particles into eyes Contusions Hearing Damage Burns Incorrect use of equipment -Irritated eyes, throat, or nose -Damage to person with cuts or burns Utilize protective eyewear, close toed shoes, and hearing protection. Tie hair up, remove all jewelry, and roll up sleeves. Be sure to use tongs if part is too small, six inches or less, to hold. If grinding for an elongated be sure to cool off part in water before touching, and use a rag. Lathe Particles into eyes Contusions Lacerations Burns Incorrect use of equipment -Damage to person with cuts or burns -Damage to rocket parts -Damage to hand drill and/or drill bit Utilize protective eyewear, close toed shoes, tie hair up, remove all jewelry, and roll up sleeves. Be sure to use proper etiquette when using the lathe. Always ease into parts and then pick

86 86 up speed when cutting. Never put hand within a six inch radius of the cutting zone and use the brake to make sure the part is not spinning when being touched. If part is hot then use a rag over your hand. Mill Particles into eyes Contusions Lacerations Burns Incorrect use of equipment -Damage to person with cuts or burns -Damage to rocket parts -Damage to mill and drill bits Utilize protective eyewear, close toed shoes, tie hair up, remove all jewelry, and roll up sleeves. Be sure to use proper etiquette when using the Mill. Always ease into parts and then pick up speed when cutting or milling. Never put hand within a six inch radius of the cutting zone. If milling a large amount, over inches, then use the screen to guard from hot pieces of

87 87 metal being cut from the part. If part is hot during removal then use a rag over your hand. Table Saw Particles into eyes and breathing Contusions Lacerations Hearing Damage Incorrect use of equipment -Damage to person with cuts or burns -Damage to rocket parts -Damage to saw blade Utilize ventilation masks, protective eyewear, close toed shoes, and hearing protection. Tie hair up and remove all jewelry. DO NOT use gloves or have anything distal to the elbows. Soldering Severe burns and fumes irritate eyes, nose, and throat -Overheating -Incorrect use of equipment -Burns - Inoperable electronic equipment Operate correctly with proper safety equipment and a clear understandin g of safety procedures. Heat Gun Skin damage -Overheating -Incorrect use of equipment -Burns Utilize protective gloves Bandsaw Contusions Lacerations Incorrect use of equipment -Damage to person with cuts or burns -Damage to rocket parts -Damage to saw blade Utilize protective eyewear, close toed shoes, tie hair up, remove all jewelry, and roll up sleeves. DO NOT wear gloves. Be

88 88 sure to use proper etiquette when using the Bandsaw. If part is too small, less than six inches, then use a pushblock to cut. Air Gun High velocity air irritates eyes, hearing damage - Overpressuri zed gun -Incorrect use of equipment -Irritated eyes and ears Do not point directly at self, gradually buildup airflow, be cautious of high velocity air. Hammer Contusion Pinching -Dropping the hammer -Improper use of equipment -Damage to person with bruises and possibility of breaking body parts Utilize protective gloves and use a solid workbench. Hatches Contusion Pinching -Improper use of equipment -Human error -Damage to person with cuts -Damage to rocket parts Utilize protective gloves and use a solid workbench. Files Contusion Skin Irritant -Improper use of equipment -File breaking -Skin abrasion or burns -Wearing away part Utilize protective gloves and always file away from the body. Lighters Burns -Lighter explosion -Improper use -Damage to person with burns Always direct away from body when in use. Vise Pinching -Human error -Clamping to body part and causing injury NEVER look away when clamping a part to the vise.

89 89 Extensio n Cord Electrocutio n -Malfunction of equipment -Injury to person in form of burns DO NOT use when there water on or near the sockets of the extension cord. Always assume extension cord is actively moving electricity. Shears Contusions -Improper use of equipment -Human error -Injury to person in form of cuts Always be attentive when using and NEVER leave shears open or on the ground. Wire Stripper Contusion -Improper use of equipment -Equipment malfunction -Injury to person in form of cuts Always be attentive when cutting wires and never look away while working. Source of Hazard Hazard Cause Result Severity Likelihood Criticality Mitigation Epoxy Toxic fumes Skin irritant -Chemical splash or fumes -Burns on skin or eyes -Irritation to lungs Utilize ventilation masks and latex gloves. Use in a well-ventilated area. Fiberglass When sanding: eye and skin irritant and inhalation hazard -Chemical splash or fumes -Irritating skin and eyes Utilize ventilation masks, long sleeves, and latex gloves while sanding. Sand in a well-ventilated area. Pyrodex Explosive if contained improperly -Oxidation of propellant -Damage to person and equipment Insure proper storage. Keep away from sparks, heat, and open flame. DO NOT arm altimeters until ready.

90 90 APCP Propellant Unexpected Explosion -Oxidation of propellant -Damage to person and equipment Insure proper storage. Keep away from sparks, heat, and open flame. DO NOT install igniter until on launch pad. Steel Dust inhalation, harmful to eyes, hot when machined -Cutting steel -Injury to person from cuts or irritation Wear safety glasses; be aware and control inhalation of dust; be aware of hot surfaces. Aluminum Metal chips lacerations, harmful to eyes, hot when machined -Cutting aluminum -Injury to person from cuts or irritation Wear safety glasses; be aware of flying chips and risk of lacerations; be aware of hot surfaces Copper Wires Dust and fumes irritate eyes, nose, and throat -Cutting wires -Irritation to person body(ie. Nose, skin, eyes) Wear safety glasses; be aware and control exposure to dust and fumes Carbon Fiber Skin Irritant Eye Irritant -Cutting carbon fiber -Irritation to person body(ie. Nose, skin, eyes) Utilize safety glasses; be aware and control inhalation of dust. Lead Skin Irritant Inhalation Hazard -Cutting lead -Irritation to person body(ie. Nose, skin, eyes) After touching lead, use soap and water. If inhaled, go to an openly vented area to obtain fresh air. Electrical Shrink Wrap Eye Irritant Inhalation Hazard -Cutting shrink wrap -Irritation to person body(ie. Nose, skin, eyes) Be aware of vapors created while heat shrinking, if an overwhelming amount of vapor consists, skin gets irritated, or breathing causes lightheadedness, stop the heat shrinking process and move to a well ventilated area. Spray Paint Skin Irritant Eye Irritant Inhalation Hazard -Spraying paint improperly -Irritation to person body(ie. Nose, skin, eyes) Utilize protective ventilation masks, protective gloves, use long sleeve shirts, and use in a well ventilated area.

91 91 Compressed Air High velocity air irritates eyes, hearing damage -Improper use of equipment -Irritation to person body(ie. Nose, skin, eyes) Do not point directly at self, be cautious of high velocity air Petroleum Jelly Skin Reaction Fire Hazard -Human error -Irritation to person body(ie. Nose, skin, eyes) Follow directions for use, DO NOT use if allergic, and keep away from open flames. Paint Skin Irritant -Painting improperly -Irritation to person body(ie. Nose, skin, eyes) Wash hands with soap and water after use. If skin irritation consists consult a doctor. Foam Skin Irritant Flammable -Accidental fire -Irritation to person body(ie. Nose, skin, eyes) Wash hands with soap and water after use. If skin irritation consists consult a doctor. Keep away from open flames. Simple Green Skin Irritant Inhalation Hazard Ingestion Hazard -Getting in contact with person -Irritation to person body(ie. Nose, skin, eyes) If adverse effect occurs on skin wash hands. If adverse effect occurs with breathing, move to open ventilation. Batteries Release acid if not stored correctly -Dropping batteries -Irritation to person body(ie. Nose, skin, eyes) Insure proper storage. Fast Orange Skin Irritant Eye Irritant -Getting in contact with person -Irritation to person body(ie. Nose, skin, eyes) Use as directed and check ingredients if allergic to anything. If caught in eye rinse with water, if condition consists get to a doctor immediately. If swallowed get to a doctor immediately Environmental Hazard Analysis Source of Hazard Trees Hazard Cause Result Damage to rocket or parachutes N/A Rocket getting stuck in tree Severity (1-5) Likelihood (1-5) Criticality Mitigation Drogue and main parachute will be

92 92 Dryness High Wind High Temperatur es Humidity Coupler tube cracking More brittle epoxy joints Drastically increased drift on descent Rocket at risk of being dragged along ground. Overheatin g electrical component s Warping of body tubes Coupler tube swelling N/A N/A Exposu re to Sun N/A -Internal failure -Have to launch at an angle -Unable to launch - Structur al integrity jeopardi zed - Saturatio n of sized to prevent excessive drift in high winds Coupler tube will always be stored in a humidity controlled environme nt Longer curing time epoxy will be used at crucial points Drogue and main parachute will be sized to prevent excessive drift in high winds Rocket will be launched during favorable wind conditions All assembly will be done outside of direct sunlight Vehicle will be recovered as soon as possible Coupler tube will always be stored in a

93 93 Fog Bodies of Water Freezing Temps & Snow Adverse effect on altimeters Weakened epoxy joints Increased risk of losing rocket Damage to electrical component s Damage to rocket Damage to electronics Irretrievabl e rocket component s Faster battery discharge Warping of body tubes N/A N/A N/A propella nt -No ignition -Can t recover rocket - Damage d electroni cs -Rocket wil not separate humidity controlled environme nt Longer curing time epoxy will be used at crucial points A GPS tracking device will be put in use to find rocket if not visible All electronics bay bulkheads will be tightly sealed Drogue and main parachute will be sized to prevent excessive drift in high winds All electronics bay bulkheads will be tightly sealed All batteries will require lifetimes far exceeding the time needed to assemble and launch

94 94 the vehicle. Rain Lightning Dust & Dirt Damage to electrical component s Severe Damage to entirety of Rocket Damage to finish on body tubes N/A N/A N/A - Damage d electroni cs -No altimete r data - Damage d rocket and electroni cs -Rocket becomes dirty All electronics bay bulkheads will be tightly sealed Launches will be suspended in case of inclement weather Monitor weather reports before the launch date, if sever weather arises then postpone launch date. Rocket will be put in clean storage. Source of Hazard Recovery failure Hazard/Result Rocket lands with high kinetic energy causing harm to anything in its path. Breakup of rocket on impact will litter surrounding area Harm to local vegetation and wildlife Severity (1-5) Likelihood (1-5) Criticality Mitigation Inspect entire recovery system prior to every launch Ensure all explosive charges are properly packed Ensure all parachutes are properly folded Fire Harm to team members and range personnel The team will only use launch rails with blast deflectors Damage to rocket

95 95 Rocket drifting on descent Explosion of rocket Littering Leaching of toxic chemicals Rocket dragging may harm range personnel, local vegetation, and local wildlife Hot components can harm local vegetation and wildlife after landing Rocket materials litter surrounding area Harm to local vegetation and wildlife Batteries leak battery acid into ground, harming local vegetation and wildlife Drogue and main parachute will be sized to prevent excessive drift in high winds Sirens from altimeters will prevent wildlife from approaching rocket Inspect motor casing prior to every launch All members will be instructed to place all trash in designated trash bins during vehicle preparation Only properly sealed batteries will be used. All batteries will be disposed of in designated trash bins 4.7. Failure Modes and Effects Analysis Potential Failure Mode Potential Failure Cause Consequence Severity (1-5) Likelihood (1-5) Criticality (1-25) Mitigation Rail button breaks during launch Rocket trajectory is improper Proper installation, alignment, and location of rail buttons. External Structural Failure Fins fail from aerodynamic forces Instability during flight Use proper materials and construction techniques for fins. Body tube fails due to stress during flight Total failure of vehicle and safety hazard to bystanders Ensure that the material selected exceeds expected compressive forces. Internal Structural Failure Thrust capture system fails to transfer thrust Motor accelerates through body of rocket completely destroying vehicle Use proper materials for centering rings and motor tube.

96 96 from motor to body tube Coupler tube fails from inadequate length or strength Rocket breaks apart Make couplers at least two diameters in length to create unwanted separation. E-bay threaded rods fail E-bay becomes insecure, forward section may separate from rocket Use the proper material for the threaded rods. Motor retention fails Motor falls out, safety hazard to people below Ensure that the motor retention can withstand the forces seen throughout flight and on launch. Bulkheads fail during parachute deployment Part of vehicle will come down ballistic or on smaller parachute than intended Ensure that bulkheads are capable of withstanding applied loads. Altimeter signals separation before proper point is reached A separation event occurs prematurely, the mission might be considered a failure due to less time collecting data Do proper testing of all altimeters used during launch. Make sure that they work consistently well. Altimeter Failure Altimeter runs out of battery power Altimeter is not wired correctly Altimeter does not signal separation, and the rocket goes ballistic. Altimeter acquires too much noise or unusable data Make sure that each battery is fully charged before each launch. Check electronics before launch. Loose wires should be contained, and all connections should be accounted for. Altimeter is not calibrated Parachute deployment at improper altitude Do proper testing of all altimeters used during launch. Make sure they are properly calibrated. Ejection Event Failure Explosive charge fails to ignite Parachutes are not deployed and rocket becomes ballistic Inspect e-matches while preparing ejection charges.

97 97 Comes untied from U-bolts A portion of the rocket is separated from the parachute and comes down ballistic Use self-tightening knots. Shock Cord Failure Snaps due to excessive loading A portion of the rocket is separated from the parachute and comes down ballistic Use simulations to have estimates of the loads the shock cord will see. Buy the appropriate rated shock cord, and then test it with the sub-scale and full-scale launch vehicles. Cut by sharp object inside rocket A portion of the rocket is separated from the parachute and comes down ballistic Keep sharp objects away from the shock cord before launch. Parachute burns from ejection charges. Damage to parachute causes faster descent rate Shield the parachute from the ejection charges with fireretardant material. Parachute Failure Parachute detaches from shock cord. Rocket falls at a high descent rate without parachute Use self-tightening knots and check the integrity of the parachute before launch. Parachute is tangled and does not open completely. Rocket falls at a faster descent rate. Possible damage to rocket Use a systematic method of wrapping the parachute such that it will not tangle. Igniter fails to ignite. Rocket does not launch Inspect igniter for continuity. Make additional igniters. Launch Failure Propellant fails to ignite. Launch rail breaks. Rocket does not launch. Rocket is sent in random direction or out of range Use proper igniter. Store propellant properly. Use launch rail appropriate to handle the thrust. Make sure rocket path is clear. Igniter does not ignite the entire propellant. Rocket does not reach the intended altitude Properly inspect motor and igniter installation.

98 98 Motor becomes overpressurized. Motor casing fails. Permanent damage to rocket Ensure proper motor hardware and preparation. Motor Failure Centering rings aligning the motor fail. The thrust is no longer along the central axis of the rocket. Rocket is unstable Manufacture centering rings that maintain true center and can sustain loads. Motor explodes while on the launch pad. Catastrophic damage to the rocket Ensure proper motor preparation and that personnel are at a safe distance during launch. Blades become lodged in body tube during separation Payload does not separate, but is also unsecured to the airframe. Runs the risk of becoming a projectile Fold blades properly before launch. Test this function during fullscale launch. Rotor Recovery Failure Actuator fails Blades break during flight or upon impact with ground Payload becomes a projectile. Payload becomes a projectile Ensure that the actuator has been treated well, check it for defects before launch. Design the blades for a shock load that simulates impact with the ground. Unscheduled deployment of backup parachute Lines may become tangled in rotors rendering the system inoperable, and making it a projectile Ensure that the altimeter is working properly before launch. Ensure that the parachute cannot easily tangle with the rotor recovery.

99 99 5. Payload Criteria 5.1. Selection, Design, and Verification Payload Objectives As posted in the NASA Student Launch handbook the payload objectives are as follows: Teams shall design an onboard camera system capable of identifying and differentiating between 3 randomly placed targets Each target shall be represented by a different colored ground tarp located on the field Target samples shall be provided to teams upon acceptance and prior to All targets shall be approximately 40 X40 in size The three targets will be adjacent to each other, and that group shall be within 300 ft. of the launch pads After identifying and differentiating between the three targets, the launch vehicle section housing the cameras shall land upright, and provide proof of a successful controlled landing Data from the camera system shall be analyzed in real time by a custom designed on-board software package that shall identify and differentiate between the three targets System Level Design and Design Matrix The autorotation payload will be stored in the rocket for the ascent to apogee. Once apogee is reached, the payload section will be ejected from the rocket. The payload will use a camera in a fairing on the fin. The payload will control the vertical descent of the forward section until approximately 200 ft above ground level. At this point, the autorotation system will flare, arresting the vertical descent, and a small parachute will be deployed. The flare maneuver consists of increasing the angle of attack of the blades through a swash plate. This increases the collective lift, slowing the descent of the payload to near zero. The payload section will then continue descent under both the auto rotation system, and the parachute. This payload was selected using a design matrix. Table 15 shows the selection of the decent method for the payload. Table 15. Design Matrix

100 100 The objective definition and justification is as follows: Complexity is defined by the amount subsystems that operate to comprise the entire payload landing system. For the purposes of choosing a particular design, complexity is quantified by the number of potential points of failure during rocket flight that are affected by the choice of vertical landing system. These include control surfaces and separation events, as the amount of separation events depends on the desired manner in which the payload is to be executed. Relative decision matrix scores are calculated by the fraction of the least possible number of potential failure points over the number of potential failure points in a given design, then multiplied by ten. As calculated, the design with the fewest potential points of failure is assigned the highest score, as the fewest number of system errors can occur during flight and landing. The weighting factor of payload complexity is significantly high (25%) because with a rocket on such a small scale, a single error in payload operation could compromise the mission. Ascent stability is the overall stability of the rocket as it travels towards apogee. It is measured by the stability margin that is outputted when the rocket is modeled in OpenRocket software. The magnitude will be determined by subtracting the stability margin from 2 since 2 is the optimal stability margin. The weighing factor of ascent stability is 20% because one of the main objectives is to have the rocket ascend 1 mile into the air. If the rocket is not ascending stably, it will not successfully complete that goal. Similarly, if the rocket crash lands out of the air due to unstability, other objectives will be rendered obsolete. The design with the stability margin closest to 2 (lowest magnitude according to the design matrix) will be given a score of 10 out of 10 possible points using a linear score assignment. Velocity controllability is the ability of the payload to decrease the velocity of the rocket after apogee, in order to land the rocket on the designated tarp in a controlled manner. It is measured by the maximum lifting force of the control surfaces. Control surfaces consist of the mechanisms used to provide lift to the payload after apogee (such as rotors or air brakes). Maximum lifting force of the control surfaces gives an idea of the magnitude to which that control surface can provide a force to counteract the force of gravity. The weighting factor of descent controllability is 10% because being able to slow the payload down efficiently is very important. Though, success of other objectives (i.e. reaching one mile) is not dependent upon this objective, and therefore the weighting factor is not higher. Score is determined by calculating combined lifting force of the rotors and the air brakes. Rotor lifting force is calculated via induced airflow equation, and air brake lifting force is calculated via drag equation. Terminal velocity for all calculations are taken from the theoretical terminal velocity from last year s Critical Design Report, at a value of approximately 7.65 m/s^2 for standard of comparison. With regards to air brakes, design matrix magnitudes are calculated assuming the drag coefficient is 1.00 for a standard of comparison. Scores are calculated by dividing each magnitude by the largest possible magnitude and then multiplying by ten. Flight Path Controllability is defined as the maneuverability of the control surfaces in the desired direction. It is measured qualitatively depending upon how easily it is judged to move toward the desired tarp. Prior research of the control mechanisms efficiency and degree of maneuverability will be the main factor behind score assignment. The flight path controllability weighting factor is 15% because it directly affects the objective of landing on the tarp. The design with the most maneuverability will be given a score of 10 out of 10 possible points using a linear score assignment.

101 101 Descent Weight is the weight of the payload after apogee. It is measured in newtons. The weighting factor of descent weight is 5% because descent weight determines the ease with which the rocket can descend. With a higher weight, the difficulty associated with slowing the rocket down after apogee. The design with the lowest descent weight will be given a score of 10 out of 10 possible points using a linear score assignment. Manufacturability is how easily the payload can be manufactured. It is measured by the number of components that have to be manufactured. The more components that need to be manufactured, the longer it ll take to create the rocket. The weighing factor of manufacturability is 25% because this is a time sensitive project. In order to have sufficient testing as well as put out the rocket without rushing, manufacturability must be weighed heavy. The design with the highest manufacturability will be given a score of 10 out of 10 possible points using a linear score assignment. The score justifications for the design matrix is as follows: Complexity: Design 1: Received a maximum score of because this design contained the least number of potential points of failure related to the payload system. Points of failure include a nose-cone separation event, and autorotation system. Design 2: Received a score of 4.44 because this design contained nine potential points of failure, including eight air brakes, and separation event. Design 3: Received a score of 4.00 because this design included an objective-high ten potential points of failure, including the eight air brakes, separation event, and autorotation system. Design 4: Received a score of 8.00 because this design included five potential points of failure, including separation event and autorotation system. Ascent Stability: Ascent stability was measured using Open rocket simulation. A stability margin was calculated and the difference between the stability margin and a stability margin of 2 was used to calculate the magnitude. Design 1: Received a score of 7.27 because this design yields a stability margin with second greatest distance from 2.00 at 2.05 difference. Design 2: Received a high score of because yields design had a stability margin closest to 2.00 at a low distance 1.49 away from Design 3: Received a low score of 6.26 because this design yields a high stability margin with greatest distance from 2.00 at 2.38 difference. Design 4: Received a score of 8.61 because this design yields a stability margin with second lowest distance from 2.00 at 1.73 distance.

102 102 Velocity Controllability: Velocity Controllability was determined by calculating the lift force at terminal velocity. The following equation was used: L = C L 2 ρv2 A 1 Where ρ is the density of air, v is the velocity and A 1 is the area of the disk. For airbrakes the following equation was used to calculate force: F = K 2 ρv2 A 2 K is a constant, and A 2 is the surface area of the rotary blades. Design 1: Received a score of 8.15 because this design produces tie for second highest relative lifting force of N from the autorotation system, as with Design 4. Design 2: Received a maximum score of because this design produces a high relative lifting force of N from both the autorotation system and air brakes. Design 3: Received a low score of 3.69 because this design produces a low relative lifting force of N from only air brakes. Design 4: Received a score of 8.15 because this design produces a tie for second highest relative lifting force of N from the autorotation system, as with Design 1. Descent Weight: Design 1: Received a maximum score of and a rank of lightest design because this design employs only an autorotation rotor and the nose cone. Design 2: Received a low score of 2.50 and a rank of heaviest design because this design lands a large portion of the airframe plus both the autorotation and airbrake systems. Design 3: Received a score of 5.00 and second heaviest design because this design lands a large portion of the airframe plus the airbrake system. Design 4: Received a score of 7.50 and second lightest design because this design lands a large portion of the airframe plus the autorotation system. Manufacturability: Design 1: Received a score of 4.44 because at nine components, this design is tied for second for most components directly related to the vertical landing system that require manufacture by the team. Design 2: Received a low score of 3.08 because at thirteen components, this design requires the most components directly related to the vertical landing system that require manufacture by the team.

103 103 Design 3: Received a maximum score of 10.0 because at four components, this design requires the fewest components (by a relatively large margin) directly related to the vertical landing system that require manufacture by the team. Design 4: Received a score of 4.44 because at nine components, this design is tied for second for most components directly related to the vertical landing system that require manufacture by the team. When each parameter score is multiplied by its parameter weight and these values are added together, a comparison can be made. The design matrix shows that the autorotation design is the better design for the application in question. Final Score: Nosecone Autorotation (7.68) / Body Autorotation & Airbrakes (6.51) / Body Airbrakes (6.50) / Body Autorotation (7.52). Even though Nosecone Autorotation was the winning design in the design matrix, the Q&A defined upright to be at the same orientation as the component was on the pad, therefore the body autorotation was selected as it was the next highest score from the design matrix. Payload Description Before launch, the autorotation blades are stored in their rolled configuration. The autorotation payload is attached to the aft airframe at the top of the coupler tubing. The blades of the autorotation system are inserted into the aft end of the electronics bay. The electronics bay is attached to the aft airframe through shear pin. The blades are kept in their rolled orientation by the airframe. At apogee, the payload will be ejected from the rocket. The blades are no longer held in the rolled configuration by the airframe, and spring to their open position through their geometry. When opened, the blade chord will be oriented in the direction of the payload shaft axis. This will allow the blades to act like fins, producing minimal lift for the majority of the descent. This minimizes the wind drift of the payload, while still having directional control over the payload. Gradually over the course of the descent, the blades will be rotated to begin the autorotation procedure. Initial angles will be small, with the purpose of reducing the descent speed below terminal velocity. Additionally, the blades will begin to spin about the bearing that attaches them to the shaft. Blade angles are controlled by a swash plate, which is moved by a servo motors housed in the aft airframe. The servo is controlled with an Arduino, which reads data from the altimeter, and camera. The Arduino will analyze the photos taken. After analysis the servo angle will change to change the cyclic pitch of the swash plate. By changing the cyclic pitch of the swash plate, the pitch of the blades relative to each other changes resulting in greater lift on one side of the payload and pushing the payload to one side.

104 104 Figure 40. Control Logic for Autorotation Figure 40 shows the control loop for the autorotation payload. Altitude data is read, and a picture is taken by the microcontroller, which controls the position of the servos. The changing position of the servos moves the swash plate, which changes the pitch angle of the blades. As the blade angle changes, the descent rate and position of the payload changes. The altitude is then read by the altimeter, a picture is taken by the camera, and a new signal is sent from the microcontroller to the servos. At a designated height above ground level, the payload will begin the flare maneuver. The pitch angle of the blades will be significantly increased, and rotation energy that has been stored in the blades will be converted into a higher lift force. This lift will slow the descent velocity to near zero, completing the operation of the payload, Test 26 in Table Performance characteristics The performance of the autorotation payload is characterized by four criteria the landing flare, the closeness of the actual descent speed to the desired descent speed at all points during the flight, if the payload lands vertically, and the closeness to the tarp. The flare will be conducted at an altitude of approximately 200 ft above ground level. If the payload operates successfully, it will be deployed from the airframe and the vertical descent rate will be controlled through changing pitch angle. Coupled with the completion of the landing flare maneuver, the payload will also deploy a parachute to ensure a safe recovery of the entire system, vertically on the designated tarp.

105 Verification plan Requirement Design Feature Verification # Team chose one design experiment option from the given list in NASA Handbook Target detection and upright landing Payload selected utilizing autorotation Proposal accepted (req. 3.2) Team will design an onboard camera system, successfully capable of identifying and differentiating between three randomly placed objects There will be a small camera in a fairing on a fin. The microprocessor will determine the tarps from their color, through a custom code Ground testing, and fight testing, starting with drop tests Launch vehicle housing the camera shall land upright, and will provide of successful controlled landing after identifying and differentiating the three targets The camera will be on the aft airframe on a fin, and the autorotation system will guide the aft airframe to the ground Visualize inspection that the aft can separate and can be guided to an upright landing Data from camera system shall be analyzed in real time by a custom designed on-board software, capable of identifying and differentiating between three targets There will be a small camera in a fairing on a fin. The microprocessor will determine the tarps from their color, through a custom code utilizing the concepts of gravity wells Ground testing, and fight testing, starting with drop tests

106 106 The following are team defined requirements for the payload: Requirement Design Feature Verification Status Requirement # The autorotation system shall fold properly within the rocket airframe. The rotors will have a geometry that allows them to roll around the center axis The rolled diameter of the rotors will be measured and compared to the inner diameter of the airframe. unverified The autorotation system shall not catch on separation. The rolled diameter of the blades will fit within the diameter of the airframe The payload will undergo ejection testing to ensure a smooth deployment. unverified The autorotation system shall extend its rotors immediately following separation. The geometry of the rotors allow them to extend to a rigid position Deployment testing will be conducted to ensure the rotors extend as intended unverified The autorotation system shall collect energy from the air upon descent. The rotors will rotate around a bearing, storing energy in a rotational form Descent testing will be conducted to ensure rotors rotate when exposed to airflow. unverified The autorotation system shall slow the descent of the payload. The rotors will follow an airfoil shape to generate lift Testing of the swash plate mechanism will be conducted to ensure blade angles change when commanded by the microcontroller. unverified The backup main parachute on the autorotation system shall not be The parachute will be located above the swash plate of the Visual inspection of the parachute packing will ensure the unverified

107 107 rendered inoperative by any means before, during or after separation. autorotation system. deployment of the backup parachute is possible at all stages during flight. Shock cord and recovery harness shall not become untied. Shock cord knots will be epoxied after being tied. Visual inspection will confirm the epoxy around the shock cord and harness has cured. unverified Shear pins for payload deployment shall break at apogee. The minimum strength of shear pins required will be used. Ejection testing of the payload will be conducted. unverified The payload airframe shall be able to withstand the jolt seen during deployment. The airframe will be designed to withstand forces of deployment. Impact testing of the payload will be conducted to simulate deployment of the payload. unverified The payload swash plate shall be able to withstand the forces generated by the rotor blades during flight. The swashplate will have significant thickness to withstand the forces of flight. The swash plate will be analyzed using FEA tools to ensure it will withstand flight forces. unverified The autorotation system shall be attached to the payload airframe in such a way that it will not break or detach during flight. Shafts will attach the autorotation system to the airframe and have sufficient strength to withstand flight forces. Visual inspection of the shaft will ensure the system is attached as designed. unverified The body tube encasing the payload shall be made in-house. The body tube encasing will be made from carbon fiber. Visual inspection of the body tube casing will confirm manufacturing occurred in-house. unverified

108 108 The rotor blades shall be able to collect enough energy to achieve autorotation. The rotor blades will be attached to bearings allowing rotational energy to be stored. Descent testing will confirm the lift generated by the blades can achieve autorotation. unverified The payload shall have its own altimeter that is accessible from the external airframe. An altimeter separate from the rocket will be included in the design. Visual inspection will confirm the altimeter is located in the aft airframe, and accessible from the external airframe. unverified The payload shall have its own tracking device that relays its position to a ground station. A GPS receiver will be included in the design. A GPS receiver will be confirmed by visual inspection. unverified The payload shall have an onboard flight controller for the purpose of autorotation. A micro controller will be included in the design to control the linear actuator. A microcontroller will be confirmed by visual inspection. unverified The payload electronics shall be completely independent of the rest of the electronics on the rocket. The electronics will be powered by and run on a separate circuit from the rest of the rocket. Ground testing will confirm that the electronics of the payload do not affect the electronics of the rocket. unverified The autorotation system shall have a back up main parachute in the event of autorotation failure. A reserve parachute will be included in the design of the system. Ensure a parachute is included by visual inspection. unverified The back up parachute shall be contained such that The parachute will be enclosed in a separate case, Ensure the parachute is packed in its unverified

109 109 it does not deploy upon separation from the main rocket airframe. and ejected with a separate system not related to the main rocket airframe. casing by visual inspection. The autorotation system shall not cause the main parachute to get tangled. The autorotation system will separate from the rest of the rocket, keeping sufficient distance from the main parachute. Ejection testing will confirm the autorotation payload does not interfere with the main parachute. unverified The autorotation shall slow the velocity when the payload is close to the ground. The autorotation system will flare at approximately 200 ft above ground level. Ground testing will confirm the blades are rotated at the designated altitude. unverified Payload Integration The payload will be integrated in the aft airframe. The mechanical systems composing the autorotation system will be attached using a centering ring and bulkheads. The electronics governing the system will be stored above the thrust bulkhead and attached to the centering ring at the top of the coupler tubing. At apogee the aft section will become a free system from the remaining airframe. For the camera the, CMOS MT9M112 SOC camera was chosen. This camera was chosen out of all of the other cameras for its compatibility with an Arduino, and its resolution. At 1.3 Megapixels and a 62-degree field of view the CMOS MT9M112 SOC camera was the best compromise between size and image quality. The CMOS MT9M112 SOC is small enough to have a small fairing on the fins to minimize drag. The CMOS MT9M112 SOC also allows for one tarp to be about 70 pixels, see Figure 41. With 70 pixels the algorithm has enough pixels to find the center of the tarp and guide toward it.

110 110. Figure 41. Number of Pixels vs Altitude Instrumentation To reach its objective, the payload must accurately measure its altitude. As such we have chosen to include the BMP085 Barometric Pressure/Temperature/Altitude Sensor in the payload bay to measure altitude. The BMP085 Barometric Pressure/Temperature/Altitude Sensor relays barometric data to the microcontroller where a library from the manufacturer converts it into altimeter data in real time with a precision of +/-1ft. When deciding what instrumentation to use for measuring altitude the StratologgerCF, and the raven 3 altimeter was also considered. Their inability to relay data to the computer in real time is ultimately what makes it a less viable option.

111 111 Figure 42. Arduino for controls In order to control the swash plate, the payload will use Servo - Generic High Torque Continuous Rotation model ROB ROHS. The servo weighs 0.1 lbs, can provide 4 in-lbs with 3.5 in-lbs required. The servos also have 360-degree motion. These servos were compared to a lot of servos. The ROB ROHS servo was chosen as it can provide the torque required at max lift on the swash plate. For the camera the, CMOS MT9M112 SOC camera was chosen. This camera was chosen out of all of the other cameras for its compatibility with an Arduino, and its resolution. At 1.3 Megapixels and a 62-degree field of view the CMOS MT9M112 SOC camera was the best compromise between size and image quality. The CMOS MT9M112 SOC is small enough to have a small fairing on the fins to minimize drag Drawings and Electrical Schematics Models of the Payload System are shown below. Figure 43. The electrical schematic for the autorotation payload

112 112 Figure 44. Assembled Payload

113 113 Figure 45. Payload Components Figure 46. Autorotation Blades

114 114 Figure 47. Swash plate height and spacing

115 115 Bulkhead Centering Ring Figure 48. Servo, Bulkhead and Ring Positions Figure 49. Payload Dimensions

116 Payload Key Components Three key components exist in this payload. The autorotation system is the mechanical system used to control the vertical descent rate of the payload. The electronic component is used to control the mechanical system, track location, and broadcast data during the descent. The parachute system is used to guarantee successful recovery of the payload, and reduce the risks associated with uncontrolled landings. To successfully meet the payload objectives, all three of these components are required to interact together. The electrical system interacts with the mechanical system via an Arduinoand servos. Pitch angle is controlled with the servos of the mechanical system. The servo position is determined by the PWM pulse sent to it by the Arduino. The Arduino will read altitude data from the altimeter, and analyze the image from the camera to determine pulse sent to the servos. The parachute component interacts with the electrical system by reading altimeter data. At the desired altitude, the parachute will deploy. The remainder of the descent from 200 ft will be under the interaction of both the parachute and the mechanical autorotation system Payload Concept Features and Definition Creativity and Originality In most rockets, the recovery system used consists of a system of parachutes that land the rocket in a random orientation. The payload design offers a way to control the descent of the payload in variable environments and control the orientation of how it lands. Instead of relying on favorable weather conditions to determine the site of landing, the payload allows for a directed landing to a designated site. Despite similar sized masses being recovered with rotors, this payload is original in controlling both vertical descent, orientation, and direction of landing Uniqueness and Significance While utilizing design elements from last year, the payload recovery system offers a unique solution not utilized by current system designs. Through adding a dynamically controlled swash plate to the carbon fiber rolled fins used last year, the payload offers a gentle touchdown at a near-zero velocity at a specific orientation to the ground in a designated area. With these new features in mind, the landing method described will allow for a more delicate payload to be delivered to specific landing sites. All of these aforementioned features would not be possible if utilizing a system of parachutes, which fall based on the designed descent rate and wind drift and arrest velocity upon impact with the ground. Autorotation allows for more control capabilities and guaranteed safety and reusability of the payload, which offers a large potential for future objectives Suitable Level of Challenge The concept of vertical landing and landing zone detection offer challenges both in technical and mechanical aspects. The technological challenges we must assess are the problems with controlling the pitch angle of our blades as the payload descends. Mechanically, we must address the problems in designing and manufacturing a rotor system similar to that of a helicopter.

117 117 In order to control the rate of descent and position of the payload, systems must be put in place that measure the altitude in real time. This includes using an altimeter, camera and microcontroller to send voltages to servos. In order to monitor its relative position, a GPS and antenna must be used. The payload also provides mechanical challenges in design and manufacturing. These include converting linear motion to a change in pitch of the carbon fiber blades and storing the blades during flight. To convert linear motion in to a change in pitch, pitch horns and rods, and a swash plate must be designed. These elements must be able to withstand the aerodynamic forces produced by the blades and the change in forces due to adjusting the pitch of the rotors. Before the apogee and during flight, the rotors must be stored in the rocket itself. This requires a design that allows for the rotors to be folded into the aft airframe spring open to their full diameter at apogee Alternative Design Using multiple sources 13 the lift from the rotors can be calculated by the following equation: 2 N L = c lα 8 ( ρπd2 v 3 tc pα D) 8 ( 4 ( v + ρπd4 v 3 tc pα 16 ρπd2 v 3 tc pα 8 D ) Where, c lα is the coefficient of lift relative to angle of attack, is the air density, D is the rotor disk diameter, v is downward velocity, t is time, c pα is the power coefficient relative to angle of attack, and n is the number of blades. Using the formula, the maximum lift occurs under flare and is approximately 9.5 lbf. Since the max lift produced is very close to the weight that the payload must lift, testing will be done to confirm that the payload can handle the weight of the aft section. As a precaution that the payload cannot produce the required lift a second design was developed. The alternative design is to use powered rotors on a new forward section. The aft end of the section will hold landing legs and the camera. The legs will have springs on them to open to the proper orientation after separation. The rotors will be coaxial and spin in opposite directions to balance any torque from the driving motor. The coaxial rotor system will be made from the EskyHeli-E515A-BigLama-24G RC helicopter. An Arduino will communicate with the transceiver of the EskyHeli-E515A-BigLama-24G RC helicopter to control the cyclic, collective, and rpm controls of the swash plate and rotors. An altimeter and camera will still be connected to the Arduino to control the vertical speed and direct the payload to the landing tarp v D ) 13 Flight Performance of Fixed and Rotary Wing Aircraft by Antonio Filippone

118 118 Figure 50. Rocket Structure Figure 51. Alternative rocket design As a safety precaution, there will be a backup parachute that will be ejected sideways if the rotary system is not working properly. At 250ft the Arduino will signal the rotary system to create high amounts of lift. If the payload velocity slows the lift from the rotors will be decreased in preparation for landing. If the payload velocity is not slowed by the high lift generated the payload will eject the backup parachute to ensure that the payload lands within the energy requirements.

119 119 The payload section will now be in the forward end of the rocket. Only the payload bay will descend under the powered rotors. The legs will be housed in an empty space of the avionics subsystem airframe. The rotors will be housed by a small body tube section attached to the nosecone. Shock chord will connect the forward end of the avionics subsystem and the nosecone. The shock chord will run through a slit in the forward end of the avionics subsystem airframe, on the outside if the forward section housing the payload electronics, and in through a slit on the aft end of the small body tube section attached to the nosecone. At apogee the nosecone will be ejected, immediately followed by the forward section, allowing for a non-tethered payload, while having the nosecone still attached to the avionics subsystem Alternative Recovery Subsystem The alternative design has the payload section relocated from being part of the aft section to its own separate section in the middle of the rocket. To accommodate this, the parachutes and the separation events must change accordingly. The rocket will separate into four sections, three of which will be tethered together, the nosecone, avionics section, and aft section, while the payload section return to earth separate from the others, under the a powered helicopter system. The nosecone, avionics, and aft sections will utilize a dual deployment system for returning to earth. There are no changes made to the use of the black powder charges used to trigger the separation events. There will be two separation events during the vehicle's flight, one occurring at apogee and the other occurring at 1000ft. The first separation event will deploy the drogue chute and separate the payload from the rest of the rocket. The separation event occurs between the payload and the avionics section. The drogue chute is stored between the avionics and payload sections of the rocket. At this time, the payload will initiate its powered

120 120 helicopter system, using a piston to push the nosecone off of the forward end of the payload section. The nosecone is tethered to the avionics bay via 1/2" tubular Kevlar which is attached to bulkheads inside the nosecone and the avionics bay. In order for the payload section to separate completely from the rocket, the Kevlar tether will run down the length of the body of the rocket on the outside of the body frame, entering the airframe through slits made in the frame at the nosecone and avionics sections to attach to the bulkheads. The second separation event occurs at 1000ft where the main parachute for tethered sections of the rocket is deployed. The separation event occurs between the avionics and aft sections. The aft and avionics sections will be tethered together with 1/2" tubular Kevlar attaching to bulkheads in each section. The separated payload section will have a back-up parachute installed. The parachute is set to deploy when designated flight parameters have not been met, deploying through a hatch installed on the side of the air frame Science Value Payload Objectives The primary objective of this payload is to control the vertical descent rate of the payload, and land the payload safely with minimal drift. This capability will be demonstrated by changing the pitch of the blades during descent. A second objective of the payload is to safely land the payload on a specific 40 by 40 tarp. The third objective is to land the payload in a reusable state Payload Success Criteria The payload will be considered to be a success if it achieves the defined objectives. If the payload controls the vertical descent speed, and lands in a safe manner, the primary objective has been met. Further, if the payload lands on the designated tarp, it will meet the second objective. If the parachute deploys and the payload can be reused without any repair, it will meet the third objective. In addition to meeting these three objectives, the payload must also deploy from the rest of the airframe at apogee to be successful Logic, Approach, and Method of Investigation The descent rate of the payload will be controlled by a micro controller. A function will be developed to adjust the pitch angle of the blades as the payload descends. The input to the control system is the altitude data read from the altimeter. The output of the control system is a PWM pulse sent to the servos. After the flight of the payload, the data collected will be analyzed to determine the performance of the payload. A plot of altitude against time will visually show the effectiveness of the payload. From this plot, the actual descent speed vs. time can be determined, and compared to the desired descent speed vs. time Measurement, Variables, and Controls The payload will be continuously measuring altitude, and using the altitude measurements to feed into the pitch control system. Tests of the system (test 8 of the testing plan) will occur in a custom made vacuum chamber to simulate changing altitude. Altitude will be an independent variable during these tests, and descent speed will be a dependent variable. A micro controller will use the altitude data from the altimeter to change the pitch angle of the blades Relevance and Accuracy of Data It is expected that the payload will make the majority of the descent from apogee at terminal velocity. The landing flare will be a gradual process starting from an altitude close to the defined ground. The resulting data

121 121 should show a constant descent rate until approximately 1000 ft. At this altitude, the descent rate should begin to gradually decrease, as the lift from the blades begins to increase. Error in the data will be a direct result of inaccuracies in the altimeter readings. With a high precision of +/-1ft associated with the altimeter, it is expected that the error will be low Experiment Procedures The data from the payload will be collected during flight. At apogee, the payload will be ejected from the rocket. The blades of the payload will be deployed when freed from the rocket. The payload will descend at terminal velocity for a majority of the descent. The payload will begin to flare at approximately 30 ft above the desired altitude of 200 ft above ground level. At this point, the backup parachute will be deployed, to guarantee the safe landing of the payload. The payload operation will be complete when it safely lands on the ground. In addition to the data collected in flight on the payload, testing will occur before flight in a custom vacuum chamber (test 8 of the testing plan), simulating varying altitudes.

122 Project Plan 6.1. Performance Characteristics The following criteria must be satisfied for mission success: a. The rocket must return after flight and be completely reusable, or need minimal repairs such that it can be launched again on the same day. b. The rocket must go up to an apogee altitude of 5,280 ft.; Points will be deducted for each foot that the vehicle deviates. c. The rocket must not have a horizontal drift of more than half of a mile during descent. d. The rocket must not exceed a vertical velocity of Mach 0.8. e. The rocket velocity of the launch rail must not fall below 52 feet per second. f. The kinetic energy of the rocket and payload upon landing must not exceed 75 ft-lbs. g. The rocket must have a stability margin of 1.5 or greater during ascent Requirements Evaluation Evaluation Methods Analysis Inspection Simulation Testing Evaluation Feature Hand calculations will be done using formulas and knowledge from academic background. Will be used to verify simulation results Inspection will be done by Range Safety Officers and by team Safety Officer and Certified Safety Stewards to ensure all compliance requirements are met Simulations shall be conducted using software to approximate stability margin, thrust, stresses and other parameters Vehicle shall be launched prior during sub-scale launch and full-scale launch to determine accuracy of analysis and simulations.

123 Requirements Compliance and Verification Plan Table 16: Launch Vehicle Requirements Established by NASA Requirement Design Feature Verification Status Requirement # The launch vehicle Correct motor This requirement Unverified 1.1 shall deliver the selection using shall be verified payload at an determined weight through analysis, apogee altitude of will allow launch simulations and 5280 ft. vehicle to reach testing during target altitude. subscale and full scale launch. The launch vehicle The vehicle shall This requirement Unverified 1.2 shall carry one have three shall be verified by barometric altimeter for recording the official altitude barometric altimeters to report altitude via a series of beeps. One in the autorotative payload bay, and two in the inspection. The official scoring altimeter shall report competition altitude via a series of beeps Team may have additional altimeter to control vehicle in the electronics bay or payloads NASA official will mark the altimeter that will be used for scoring at LRR In determination of the vehicle s apogee, all audible electronics except for the official electronics bay The vehicle shall have an official scoring barometric altimeters to report altitude via a series of beeps from the electronics bay. A redundant altimeter will be stored in the electronics bay in case of the primary altimeter failing. One altimeter shall be used in payloads bay, and two in the avionics bay The official scoring altimeter will be in the avionics bay of the rocket. A kill switch will be installed on all electronics and actuators. A remote This requirement shall be verified through inspection and testing during subscale and full scale launch. This requirement shall be verified through inspection and testing during subscale and full scale launch. This altimeter will be tested during separate payload deployment and recovery tests The team will indicate the altimeter to the NASA official during inspection. This requirement shall be verified through inspection by NASA official during subscale and full scale Unverified Unverified Unverified Unverified 1.2.5

124 124 altitude-determining altimeter shall be capable of being turned off Score of zero for altitude will be given if the official altimeter is damaged or does not report altitude properly Score of zero for altitude will be given if the team does not report to the NASA official designated for altitude scoring using official marked altimeter Score of zero for altitude will be given if the altimeter reports an apogee over 5,600 ft., above ground level Score of zero for altitude will be given if the rocket is not flown at the competition site kill switch will also be implemented. A second altimeter shall be stored within the avionics bay. This will ensure that if there is a failure during flight, this altimeter will replace the official scoring altimeter. The team will report the altitude to the NASA official for scoring. A redundant altimeter will confirm the official scoring altimeter. Current expectation for maximum altitude, or apogee, is 5176 ft. The rocket will be flown at approved NAR, and TRA launch sites. At competition, the rocket will be launched out of Huntsville, at NASA's Marshall Spaceflight Center. launch. The NASA official will be able to determine altitude by listening to a series of beeps coming from the altimeter. This requirement shall be verified by successfully demonstrating the backup altimeter's ability to replace the official scoring altimeter in the event of failure, through inspection, testing before and during subscale and full scale launches. The team will extract the data from the official altimeter and present it to the NASA official. This requirement shall be verified through subscale and full scale testing. The final official altitude at apogee will be reported to the NASA official for scoring. This requirement shall be verified by launching at approved launch sites. Unverified Unverified Unverified Unverified

125 125 Recovery electronics shall be powered by commercially available batteries The recovery electronics shall be powered by commercially available 9-volt direct current batteries. This requirement will be met by using only commercially available batteries for the recovery electronics. Unverified 1.3 The launch vehicle shall be designed to be recoverable and reusable The launch vehicle shall have a maximum of four (4) independent sections The launch vehicle shall be single stage The launch vehicle can be prepared for launch within 4 hours The launch vehicle can stay in launch ready configuration for one hour The launch vehicle will be capable of launch via standard 12V firing system The launch vehicle will have a dual stage parachute recovery system to safely descend and land the launch vehicle for reuse upon a new motor installation. The launch vehicle will have three independent sections to allow for dual parachute deployment and payload ejection. The launch vehicle will have only one stage motor. It has been designed with one Cesaroni K260- CL motor. The team will have a set list of safe launch day procedures and will rehearse procedures to ensure safety and flight readiness within a 4- hour time window. The electronics and batteries are chosen to meet this requirement two hours at minimum The vehicle is developed to be launched by a standard 12-volt This requirement shall be verified through repeated testing during subscale and fullscale launch. On the day of launch, reusability shall be verified after flight when presented for inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through practice of launch procedures and safety during subscale and fullscale launch, as well as during launch rehearsals. This requirement shall be verified through repeated testing. This requirement shall be verified through inspection and repeated testing, and Unverified 1.4 Unverified 1.5 Unverified 1.6 Unverified 1.7 Unverified 1.8 Unverified 1.9

126 126 direct current firing system. The launch vehicle shall use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA), and/or the Canadian Association of Rocketry (CAR). The motor system final choices will be made by CDR Motor changes after CDR must be approved by RSO Launch vehicle shall have a minimum static stability margin of 2.0 at rail exit Launch vehicle shall accelerate to a minimum velocity of 52 fps at rail exit Team shall successfully launch subscale rocket prior to CDR Team shall successfully launch full scale rocket prior to FRR in its The team shall purchase and use a commercially available solid K- class motor. The current motor is expected to produce 513 lbf-s. The current motor is expected to be Cesaroni K260-CL motor Launch vehicle is expected to have a stability margin of 4.11 Launch vehicle motor is expected to accelerate at 55 fps rail exit velocity. A subscale launch is scheduled for 11/19/2016. A full-scale launch in final flight configuration is scheduled for 2/11/2017. This requirement shall be verified through inspection. This requirement shall be verified through repeated testing and through inspection. This requirement shall be verified and approved by NASA RSO This requirement shall be verified through simulations and analysis. This requirement shall be verified through analysis, simulations and testing of the subscale and full scale rockets This requirement shall be verified upon the launch of the subscale rocket. This requirement shall be verified upon the launch of the fullscale rocket. Unverified 1.11 Unverified Unverified Unverified 1.13 Unverified 1.14 Unverified 1.15 Unverified 1.16

127 127 final configuration flight Vehicle and recovery system shall function as designed Payloads do not have to be flown during full-scale test flight If a payload is not flown mass simulators shall be used to simulate payload mass Mass simulators shall be in the same locations as the missing payload If the payload changes the external surface of the rocket or manages the total energy of the vehicle, those systems shall be active during fullscale test The full-scale motor does not have to be flown during fullscale test launch The rocket, the payload, and the recovery system will function as they have been designed. Otherwise, redundant systems will ensure that the mission is completed, or protect the launch vehicle from going ballistic. The payload may be flown during fullscale test flight. A mass simulator will be used in its place, if a payload is not flown during launches A mass simulator will be used in the same location on the rocket as the missing payload. The autorotation system in the aft section will be present during full scale launch. If the full-scale motor is not flown during the full-scale flight, the motor simulate shall, as closely as possible, predict the maximum velocity and maximum acceleration of the competition flight This requirement shall be verified through testing of recovery systems before, and during the subscale and full scale launches. This requirement is automatically met regardless of action. This requirement shall be met through inspection. This requirement shall be met through inspection. This requirement shall be met through inspection. This requirement shall be met through inspection and testing. Unverified Unverified Unverified Unverified Unverified Unverified

128 128 The vehicle shall be flown in its fully ballasted configuration during full-scale test After successfully completing the fullscale test the vehicle shall not be modified without the concurrence of the NASA RSO Structural protuberance on the rocket shall be located aft of the burnout center of gravity Launch vehicle shall not utilize forward canards The launch vehicle will be flown in its fully ballasted configuration during full scale test launch. The launch vehicle will not be modified after a successful full-scale test launch. Structural protuberance on the rocket shall be located aft of the burnout center of gravity The launch vehicle will not have forward canards This requirement shall be met through inspection. This requirement shall be met through inspection during competition week. This requirement shall be met through analysis, simulation and inspection. This requirement shall be met through inspection. Unverified Unverified Unverified 1.17 Verified Launch vehicle shall not utilize forward firing motors Launch vehicle shall not utilize motors that expel titanium sponges Launch vehicle shall not utilize hybrid motors The launch vehicle will not utilize forward firing motors The launch vehicle will not utilize motors that expel titanium sponges The launch vehicle will utilize a solid motor system This requirement shall be met through inspection. This requirement shall be met through inspection. This requirement shall be met through inspection. Verified Verified Verified Launch vehicle shall not utilize a cluster of motors Launch vehicle shall not utilize friction fitting for motors Launch vehicle shall not exceed Mach 1 at any point during flight Vehicle ballast shall not exceed 10% of the total weight of the rocket The launch vehicle will utilize one Cesaroni K260-CL motor The launch vehicle will not utilize friction fitting for motors The launch vehicle will not exceed Mach 1 at any point during flight Ballast shall not exceed 10% of the total weight of the rocket This requirement shall be met through inspection. This requirement shall be met through inspection. This requirement shall be verified through analysis, simulation and inspection. This requirement shall be met through inspection. Verified Verified Unverified Unverified

129 129 Table 17 Recovery System Requirements Requirement Design Feature Verification Evaluation Requirement # The launch vehicle The recovery system This requirement Unverified 2.1 shall stage the will deploy a drogue shall be verified deployment of its parachute at apogee through inspection, recovery devices and a main parachute at 1000 feet AGL. and testing during subscale and full scale launch. The team shall perform ground ejection test of both drogue and main chute At landing each independent section shall have max kinetic energy of 75 ft-lbf Recovery system electrical circuits shall be independent of any payload electrical circuits Recovery system shall contain redundant altimeters Launch vehicle shall not use motor ejection as a form of deployment A dedicated arming switch shall arm both altimeters The team will ground test the drogue and main parachutes ejection systems of the subscale and fullscale launch vehicles prior to their respective initial launches for proper parachute deployment. Simulations concerning mass and descent velocity of each independent vehicle section will be done to ensure each impact kinetic energy is less than 75 ft-lbf. The recovery system circuitry will be placed independently from the rest of the vehicle s electronics A second altimeter shall be stored within the avionics bay. This will ensure that if there is a failure during flight, this altimeter will replace the official scoring altimeter The recovery system will use black powder ejection system Each altimeter will have a dedicated pin switch accessible This requirement shall be verified upon the successful completion of the ground ejection charge testing. This requirement shall be verified with simulations, testing during subscale and full scale launch, and through observation of a safe and coasting descent. This requirement shall be verified through inspection, and testing during subscale and full scale launch. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement is met through inspection of the design. Unverified 2.2 Unverified 2.3 Unverified Unverified 2.6 Unverified 2.7

130 130 from the exterior of the rocket. Each altimeter shall have a dedicated power supply Each arming switch shall be capable of being locked in the ON position for launch Removable shear pins shall be used for both the main chute and drogue Electronic tracking device shall be installed in the launch vehicle and shall transmit the position of the tethered vehicle or any independent section to a ground receiver Any rocket section that lands untethered from the vehicle shall also carry an electronic tracking device The electronic tracking device shall be fully functional during the official flight at the competition launch site The recovery electronics shall not be adversely affected by any other onboard electronics Each altimeter will have a dedicated, new-per-launch 9V battery. Each pin switch will be locked ON when its pin is removed. Removable shears pins will be used at the separation points for the main and drogue parachutes. The main separation points will be between the payload bay airframe and avionics bay airframe. A GPS chip will be in the avionics bay airframe to track its location. A GPS chip will be in the payload bay airframe to track its location. The tracking device shall be stored in a section of the airframe made of fiberglass for radio translucency. The autorotation system is designed to carry an electronic tracking device and transmit its location to the ground station. The electronic tracking devices will be fully functional during official flights. The recovery system altimeters will be housed in their own compartment. This will be achieved This requirement shall be met through inspection. This requirement shall be met through inspection and testing. This requirement shall be met through inspection and testing during subscale and full scale. This requirement shall be met through inspection. This requirement shall be met through inspection. This requirement shall be met through inspection and testing. This requirement shall be met through inspection and testing. Unverified 2.8 Unverified 2.9 Unverified 2.10 Unverified 2.11 Unverified Unverified Unverified 2.12

131 131 using copper tape to create a shield from other signals. Recovery altimeters will be in a separate compartment within the launch vehicle Recovery electronics shall be shielded from onboard transmitting devices Recovery electronics shall be shielded from onboard devices producing magnetic waves Recovery electronics shall be shielded from any other onboard devices The recovery system electronics will be shielded from other onboard devices. This will also be achieved using copper tape to create a shield from other signals. The recovery system electronics will be shielded from other onboard devices and the altimeters will be housed in their own compartment. This will also be achieved using copper tape to create a shield from other signals. The recovery system electronics will be shielded from other onboard devices and the altimeters will be housed in their own compartment. This will also be achieved using copper tape to create a shield from other signals. The recovery system electronics will be shielded from other onboard devices and the altimeters will be housed in their own compartment. This will also be achieved using copper tape to create a shield from other signals. This requirement shall be met through inspection and testing. This requirement shall be met through inspection and testing. This requirement shall be met through inspection and testing. This requirement shall be met through inspection and testing. Unverified Unverified Unverified Unverified

132 132 Table 18: Competition and Payload Requirements Requirement Design Feature Verification # Team chose one design Target detection and Proposal accepted experiment option from upright landing Payload (req. the given list in NASA selected utilizing 3.2) Handbook autorotation Team will design an onboard camera system, successfully capable of identifying and differentiating between three randomly placed objects Launch vehicle housing the camera shall land upright, and will provide of successful controlled landing after identifying and differentiating the three targets Data from camera system shall be analyzed in real time by a custom designed on-board software, capable of identifying and differentiating between three targets There will be a small camera in a fairing on a fin. The microprocessor will determine the tarps from their color, through a custom code The camera will be on the aft airframe on a fin, and the autorotation system will guide the aft airframe to the ground There will be a small camera in a fairing on a fin. The microprocessor will determine the tarps from their color, through a custom code utilizing the concepts of gravity wells This requirement will be verified through ground testing, fight testing, and drop testing This requirement will be verified through visualized demonstration such that the aft can separate and can be guided to an upright landing This requirement will be verified through ground testing, fight testing, and drop testing

133 133 Table 19: Safety Requirements Requirement Design Feature Verification Status Requirement # The team shall create The team s safety This requirement Verified 4.1 and use a launch and safety checklist officer is currently creating a launch and safety checklist for the FRR report, LRR, shall be verified through inspection. and launch day Student safety officer shall be identified Roles and responsibilities of safety officer shall be outlined Safety officer shall monitor design of vehicle, payloads and launcher Safety officer shall monitor construction of vehicle, payloads and launcher Safety officer shall monitor assembly of vehicle, payloads and launcher Safety officer shall monitor ground testing of vehicle, payloads and launcher Safety officer shall monitor subscale launch tests Safety officer shall monitor full-scale launch tests Safety officer shall monitor competition launch operations. Aaron Dagen is the team s student safety officer. The safety officer will attend team activities to monitor and emphasize safety. The safety officer or certified safety stewards will be present during the design process. The safety officer or certified safety stewards will be present during the construction process. The safety officer or certified safety stewards will be present during the assembly process. The safety officer or certified safety stewards will monitor the testing process. The safety officer or certified safety stewards will monitor subscale launch test. The safety officer or certified safety stewards will monitor full-scale launch test. The safety officer or certified safety stewards will monitor competition launch This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. Verified 4.2 Verified 4.3 Verified Unverified Unverified Unverified Unverified Unverified Unverified

134 134 Safety officer shall monitor recovery activities Safety officer shall monitor educational engagement activities Safety officer shall implement team procedures for construction, assembly, launch and recovery Safety officer shall manage and maintain current revisions of the team's hazard analyses, failure modes analyses, procedures, and MSDS/chemical inventory data Safety officer shall assist in writing and development of the hazard analyses, failure modes analyses, and procedures Team shall identify a mentor Team shall abide by rules of local rocketry club RSO Team shall abide by rules and regulations of FAA The safety officer or certified safety stewards will monitor recovery activities The safety officer or certified safety stewards will monitor educational engagement The safety officer or certified safety stewards will strictly implement team procedures for construction, assembly, launch and recovery activities The safety officer will be responsible for management of maintaining current revisions of team s hazard analysis, failure modes analysis, procedures and chemical inventory data. Safety officer shall assist in writing and developing the hazard, failure mode, and procedure analyses. The team has identified Jimmy Yawn as the team mentor. See section 1.1 The team will properly communicate with the local rocketry club before attending any launches and abide by any rules at the launch. The team will follow and rules and regulations declared by the FAA. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified by the safety officer through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. Unverified Unverified Unverified Verified Verified Verified 4.4 Verified 4.5 Verified 4.6

135 135 Table 20: General Requirements Requirement Design Feature Verification Status Req. # Team members shall do 100% of the project The students of the team will do 100% of This requirement Unverified 5.1 the project except for shall be the handling of verified ejection charges and electric matches will through inspection. be done by the Team Mentor. The team shall provide and maintain a project plan Foreign nationals shall be identified by The team shall identify all members attending launch week by CDR The team shall engage a minimum of 200 participants in educational, hands-on STEM activities by FRR The team shall develop and host a website for project documentation The team shall post, and make available for download, the required deliverables to the team web site by the due dates specified In every report, the team shall include a table of contents including major sections and sub-sections The Project Manager will maintain a project plan. Foreign National team members will be identified for the. Team members attending launch week activities will be identified for the CDR. The team has planned events to engage at least 200 participants (of which at least 100 will be middle school students). See Section 5.4. The team s website is available at m The website will contain the required deliverables available for download. The team will provide a table of contents in each report. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified Verified 5.2 Verified 5.3 Unverified 5.4 Unverified 5.5 Verified 5.6 Verified 5.7 Verified 5.8

136 136 through inspection. In every report, the team shall include the page number at the bottom of the page The team shall provide any computer equipment necessary to perform a video teleconference with the review board The team shall use the launch pads provided by the student launch s launch service provider Teams must implement the Architectural and Transportation Barriers Compliance Board Electronic and Information Technology (EIT) Accessibility Standards (36 CFR Part 1194) Subpart B- Technical Standards ( gov) The team will include the page number at the bottom of each page. The team will provide necessary equipment for video teleconferences with the review board. The team will utilize provided launch at the subscale, full-scale and competition site The team will implement the Architectural and Transportation Barriers Compliance Board Electronic and Information Technology Accessibility Standards Subpart B- Technical Standards. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. This requirement shall be verified through inspection. Verified 5.9 Verified 5.10 Verified 5.11 Verified 5.12

137 137 Table 21: Launch Vehicle and Payload Requirements Established by the Team Requirement Design Feature Verification Status Requirement # The autorotation The rotors will have a The rolled diameter Unverified 6.1 system shall fold geometry that allows of the rotors will be properly within the them to roll around measured and rocket airframe. the center axis compared to the inner diameter of the The autorotation system shall not catch on separation The autorotation system shall extend its rotors immediately following separation. The autorotation system shall collect energy from the air upon descent. The autorotation system shall slow the descent of the payload. The backup main parachute on the autorotation system shall not be rendered inoperative by any means before, during or after separation. Shock cord and recovery harness shall not become untied. Shear pins for payload deployment shall break at apogee. The rolled diameter of the blades will fit within the diameter of the airframe The geometry of the rotors allows them to extend to a rigid position The rotors will rotate around a bearing, storing energy in a rotational form The rotors will follow an airfoil shape to generate lift The parachute will be located above the swash plate of the autorotation system. Shock cord knots will be epoxied after being tied. The minimum strength of shear pins required will be used. airframe. The payload will undergo ejection testing to ensure a smooth deployment. Deployment testing will be conducted to ensure the rotors extend as intended Descent testing will be conducted to ensure rotors rotate when exposed to airflow. Testing of the swash plate mechanism will be conducted to ensure blade angles change when commanded by the microcontroller. Visual inspection of the parachute packing will ensure the deployment of the backup parachute is possible at all stages during flight. Visual inspection will confirm the epoxy around the shock cord and harness has cured. Ejection testing of the payload will be conducted. Unverified 6.2 Unverified 6.3 Unverified 6.4 Unverified 6.5 Unverified 6.6 Unverified 6.7 Unverified 6.8

138 138 The payload airframe shall be able to withstand the jolt seen during deployment. The payload swash plate shall be able to withstand the forces generated by the rotor blades during flight. The autorotation system shall be attached to the payload airframe in such a way that it will not break or detach during flight. The body tube encasing the payload shall be made inhouse. The rotor blades shall be able to collect enough energy to achieve autorotation. The payload shall have its own altimeter that is accessible from the external airframe. The payload shall have its own tracking device that relays its position to a ground station. The payload shall have an onboard flight controller for the purpose of autorotation. The payload electronics shall be completely independent of the rest of the electronics on the rocket. The airframe will be designed to withstand forces of deployment. The swashplate will have significant thickness to withstand the forces of flight. Shafts will attach the autorotation system to the airframe and have sufficient strength to withstand flight forces. The body tube encasing will be made from carbon fiber. The rotor blades will be attached to bearings allowing rotational energy to be stored. An altimeter separate from the rocket will be included in the design. A GPS receiver will be included in the design. A micro controller will be included in the design to control the linear actuator. The electronics will be powered by and run on a separate circuit from the rest of the rocket. Impact testing of the payload will be conducted to simulate deployment of the payload. The swash plate will be analyzed using FEA tools to ensure it will withstand flight forces. Visual inspection of the shaft will ensure the system is attached as designed. Visual inspection of the body tube casing will confirm manufacturing occurred in-house. Descent testing will confirm the lift generated by the blades can achieve autorotation. Visual inspection will confirm the altimeter is located in the aft airframe, and accessible from the external airframe. A GPS receiver will be confirmed by visual inspection. A microcontroller will be confirmed by visual inspection. Ground testing will confirm that the electronics of the payload do not affect the Unverified 6.9 Unverified 6.10 Unverified 6.11 unverified 6.12 unverified 6.13 unverified 6.14 unverified 6.15 unverified 6.16 unverified 6.17

139 139 electronics of the rocket. The autorotation system shall have a back up main parachute in the event of autorotation failure. The back up parachute shall be contained such that it does not deploy upon separation from the main rocket airframe. The autorotation system shall not cause the main parachute to get tangled. The autorotation shall slow the velocity when the payload is close to the ground. A reserve parachute will be included in the design of the system. The parachute will be enclosed in a separate case, and ejected with a separate system not related to the main rocket airframe. The autorotation system will separate from the rest of the rocket, keeping sufficient distance from the main parachute. The autorotation system will flare at approximately 200 ft above ground level. Ensure a parachute is included by visual inspection. Ensure the parachute is packed in its casing by visual inspection. Ejection testing will confirm the autorotation payload does not interfere with the main parachute. Ground testing will confirm the blades are rotated at the designated altitude. unverified 6.18 unverified 6.19 unverified 6.20 unverified 6.21

140 Testing Plan Several validation exams will be conducted to ensure the success of each component. These tests are displayed in Table 21. Table 22. Testing Plan Test Number Test Name Test Description/Purpose Requirement Verification and number Status 1 Airframe Material Test Verify the strength of the composite airframe so that the further use of carbon fiber is empirically justified. This test includes both tensile and compressive stress tests In Progress 2 Avionics Subsystem Material Test Test the transmissions from the electronics bay section to ensure no GPS radio interference Ensure no radio interference from any carbon fiber components #2.12 Future Test 3 Induced Vacuum Chamber Test Determine altimeter functionality by creating a pressure change on the altimeters Verify that the altimeters are functioning properly #2.11 Future Test 4 Subscale Wind Tunnel Test Place a 3D prototype of the rocket in the wind tunnel to determine the drag on the rocket by relating it to the prototype using Buckingham-Pi Theorem of dimensionless coefficients. In Progress 5 Target Detection Test Verify the functionality of the target detection software and camera. Ensure that the camera can successfully detect the targets #3.2.1, In Progress 6 Parachute Packing Test With a length of body tubing, the parachute material will be packed as for the launch, and the packed length will be measured Verify the length of body tubing space needed for parachute space allocation Future Test

141 141 7 Black Powder Charge Ejection Test Simulate recovery separation for both drogue and main parachutes to verify that shear pins will be broken by the ejection charges and that the recovery successfully clears the vehicle Ensure the ejection system provides enough force to break the shear pins during flight #2.2 Future Test 8 Battery Life Check Place all components in the ready configuration of the final assembly and verify lifespan of power units and stability of armed system. Verify the lifespan of critical components after remaining on the launch pad for one hour #1.8 Future Test 9 Shock Resistant Drop Test Test for shock resistance to simulate the effect of launch and landing on the entire assembly Verify components will sustain impact from launch and remain usable #1.4 Future Test 10 Static Motor Test Conduct a static motor test of the full scale motor and capture resulting load cell data for analysis Determine the thrust curve of the selected motor and ensure that motor selection is as expected and appropriate # Future Test 11 Subscale Demonstration Launch Launch a subscale of the rocket that includes all major external features and the recovery scheme Ensure that the aerodynamics of the rocket and the recovery scheme are appropriate. #1.16 Future Test 12 Transmission Test Transmit sensor data from vehicle body tube to ground computer at varying ranges and orientations Ensure that sensor data can be received while the rocket is untethered to ground devices through the fiberglass body tube # Future Test 13 Launch Rehearsal Prepare the rocket as it would be on the official launch day following all launch operation procedures Ensure that the rocket can be prepped in the required four-hour window #1.7 Future Test

142 Safety Inspection Ensure every test conducted complies with safety procedures Complete safety inspection # In Progress 15 Full Scale Launch Test Ensure the rocket meets all timing and component requirements for launch day by launching the fullyloaded rocket. The rocket should be reusable without repairs or modifications Verify all requirements #1-5 Future Test 16 Concave Boattail Wind Tunnel Test Determine viscous drag effects of a concave boattail shape Future Test 17 Bulkhead Shock Test Verify the structural integrity of bulkheads after experiencing sudden loading via a drop test Verify components will sustain impact from launch and remain usable #1.4 Future Test 18 Bulkhead Compressive and Tensile Stress Test Determine the effects of continuous loading on the bulkheads to ensure they will maintain their integrity during powered ascent of the rocket Verify components will sustain impact from launch and remain usable #1.4 Future Test 19 Heat Test of Polycarbonate Boattail Analyze the effects of temperature increases due to motor heat on the polycarbonate boattail, which is susceptible to melting and deformation Verify components will sustain impact from launch and remain usable #1.4 Future Test 20 Epoxy Fin Joint Test Subject the fins to shear loading to determine the stresses that the fins can withstand before the epoxy fails and the fins detach Verify components will sustain impact from launch and remain usable #1.4 Future Test

143 Bulkhead Epoxy Test Subject the epoxy-secured bulkheads to loading along the main axis of the rocket to determine the maximum force the bulkheads can withstand before detaching Verify components will sustain impact from launch and remain usable #1.4 Future Test 23 GPS Range Test Ensure the GPS transmitter is capable of transmitting location signals and that the rocket can be recovered from excessive drift in the case of main deployment at apogee Verify components will sustain impact from launch and remain usable #1.4 Future Test 24 Electronics Bay Shock Test Verify that the electronics and GPS still function after an impact similar to one experienced by the rocket after landing Verify components will sustain impact from launch and remain usable #1.4 Future Test 25 Camera Fairing Wind Tunnel Test Determine the aerodynamic effects of a camera fairing or cowl for the landing detection software Future Test 26 Landing Struts Deployment Test Verify that the landing struts deploy correctly and in sufficient time to land in a stable manner Successfully launch and recover the full-scale rocket in its launch configuration #1.17 Future Test 27 Landing Rotor Lift Test Demonstrate the lifting capacity of the rotor lifting module Successfully launch and recover the full-scale rocket in its launch configuration #1.17 Future Test 28 Landing Rotor Maneuverability Test Determine the maneuverability and glide ratio of the rotor module Future Test

144 Middle Bay Separation Test Verify that the sections of the rocket separate with sufficient clearance to deploy the lifting module Verify components will sustain impact from launch and remain usable #1.4 Future Test 30 Backup Parachute Deployment Test Test the backup parachute deployment to ensure the functionality of the backup system Verify components will sustain impact from launch and remain usable #1.4 Future Test 31 Main Parachute Deployment Test Verify that the main deploys correctly and clears the airframe without tangling Verify components will sustain impact from launch and remain usable #1.4 Future Test 32 Recovery Module Landing Test Demonstrate that the rotor module can land properly and in the desired location Future Test 33 Recovery Module Shock Test Subject the entire loaded recovery module to a shock test to determine that the lifting module can withstand sudden stresses and to verify the stability of the black powder upon landing Verify components will sustain impact from launch and remain usable #1.4 Future Test

145 Budget Plan These estimates for the budget show that the team will spend $6620 on the entire project, including travel. Excluding travel, the total expenditures will be $2206, falling short of the maximum budget of $7500. Figure 1, the pie chart, gives a visual representation of the breakdown of the team s expenditures. The sections of the budget are shown again in Figure 2, the table that lists each section s total cost. The sections are grouped according to which expenses pertain directly to the rocket, shown in Figure 3. Additional, extraneous expenses are listed and totaled in the second group, shown in Figure 4. All expenses enumerated are projections based on the current path of the project, and are subject to change. The tabulated items include required items for general carbon fiber and airframe construction as well as necessary components for testing. In order to remove the cost of a novel testing apparatus for a 400ft drop as described in the testing section, the team will look to other design teams in the Mechanical and Aerospace Engineering Department at UF, such as the High Altitude Balloon Team under the Small Satellite Design Club for minimally-modified assemblies that can be reused without substantial cost while providing a reliable platform for economic testing. Figure 52. UF Rocket Team Overall Budget Distribution

146 146 Subgroup Figure 53. UF Rocket Team Total Expenditures Expenditures for the construction of the rocket itself are listed below. This figure is well below the maximum allowed budget. Figure 54. UF Rocket Team Launch Costs Other costs are outlined below. These costs include testing and travel costs, which make up the majority of the team s total expenditures. Total Structures and Composites $ Recovery $ Avionics $ Propulsion $ Testing $ Subscale Rocket $ Payload $ Manufacturing $ Educational Engagements $ Travel (20 People) $ 3, TOTAL: $ 6, Subgroup Budget Structures and Composites $ Recovery $ Avionics $ Propulsions $ Payload $ TOTAL: $ 2, Subgroup Budget Testing $ Subscale Rocket $ Manufacturing $ Educational Engagements $ Travel (20 People) $ 3, TOTAL: $ 4, Figure 55. UF Rocket Team Pre-Launch Costs

147 147 Item Number Subgroup Assembly Component Source and/or Link (URL) Unit Cost Quantity Material Processing Total Cost Cost Cost PNC-3.0 Composites/Recovery Nose Cone Plastic Nose Cone publicmissiles $ $ - $ - $ Payload Payload Simulant Mass Simulant- Lead Shot (2lb) scuba $ $ - $ - $ 7.99 Payload Payload Simulant Mass Simulant- Lead Shot (1lb) scuba $ $ - $ - $ Recovery ALL Nylon Shear Pins apogeerockets $ $ - $ - $ Structures ALL 75 mm Blue Tube Coupler 48" apogeerockets $ $ - $ - $ A030 Recovery Forward Parachutes 5/16 hex nuts McMaster $ $ - $ - $ K52 Payload Payload Simulant PVC Cap McMaster $ $ - $ - $ A708 Propulsion Aft Motor 6-32 Helicoils (0.345" length) McMaster $ $ - $ - $ 5.63 LOCAL PURCHASES Recovery Parachutes U-bolt $ $ - $ - $ 3.24 Payload Payload Simulant zip ties $ $ - $ - $ 8.57 Structures ALL 6-32 washers $ $ - $ - $ 1.17 Payload payload Simulant PVC Tube $ $ - $ - $ N/A Propulsion Boat tail Boat Tail 3D printed $ - 1 $ - $ - $ - Subgroup Total $ Figure 56. Breakdown of the Structures Bill of Materials Subgroup Assembly Component Description, Source, and/or Link Unit Total Quantity (URL) Cost Cost Recovery Recovery Standard-Wall Aluminum Threaded Pipe Nipple McMaster-Carr $ $ Recovery Recovery Low-Pressure Aluminum Threaded Pipe Fitting McMaster-Carr $ $ Recovery Recovery Single Piece Chute-12" giantleaprocketry $ $ Recovery Recovery Single Piece Chute-24" giantleaprocketry $ $ Recovery Recovery Shock Cords with Pre-sewn Loops(25ft) giantleaprocketry $ $ Subgroup Total $ Figure 57. Breakdown of the Recovery Bill of Materials Item Number Subgroup Assembly Component Description, Source, and/or Link (URL) Unit Cost Quantity Material Processing Total Cost Cost Cost B0131V0L5I Avionics Avionics Bay XBee Module Amazon $ $ - $ - $ PAM-7Q Avionics Avionics Bay GPS Module stop-america $ $ - $ - $ B00NAY3M2Q Avionics Avionics Bay XBee Adapter Kit Amazon $ $ - $ - $ B00DJBNDHE Avionics Avionics Bay FTDI 5V CABLE Amazon $ $ - $ - $ B008GRTSV6 Avionics Avionics Bay Arduino UNO R3 Microcontroller Amazon $ $ - $ - $ B01B272ZUW Avionics Avionics Bay SparkFun Xbee Shield Amazon $ $ - $ - $ Avionics Avionics Bay 0.25-in x 24-in Threaded Rod Lowes $ $ - $ - $ Avionics Avionics Bay 25 count 1/4 Zinc Plated Hex Nuts Lowes $ $ - $ - $ Avionics Avionics Bay 2 pk Insulated Ring Connectors Radioshack $ $ - $ - $ Avionics Avionics Bay 4 pk 9V Duracell Lowes $ $ - $ - $ B00LVI7YA4 Avionics Disk Assembly 10 pk Battery Holders, Clips Amazon $ $ - $ - $ Avionics Avionics Bay 16 pk 1/4" Zinc Washer Lowes $ $ - $ - $ 1.24 B019GU8JC0 Avionics Ground Station Bluetooth XBee Module Amazon $ $ - $ - $ Subgroup Total $ Figure 58. Breakdown of the Avionics Bill of Materials

148 148 Item Number Subgroup Assembly Component Description, Source, and/or Link (URL) Unit Cost Quantity 1596 Propulsion Full Scale Cesaroni K600-17A Motor Reload - csrocketry $ $ $ Propulsion Full Scale Cesaroni 54mm 5-Grain Case Motor Casing - Apogeerockets $ $ $ Propulsion Full Scale Aero Pack 54mm Retainer - L Motor Retainer - Apogeerockets $ $ $ Propulsion Subscale Cesaroni J140-WH Motor Reload - wildmanrocketry $ $ $ Propulsion Subscale Cesaroni 54mm 3-Grain Case Motor Casing - csrocketry $ $ $ Subgroup Total $ Figure 59. Breakdown of the Propulsions Bill of Materials Material Cost Processing Cost Total Cost Subgroup Assembly Component Description, Source, and/or Link Unit Total Quantity (URL) Cost Cost Testing Vacuum Chamber PVC Pipe 6" x 24 " $ $ Testing Vacuum Chamber Acryllic Sheets 1/4 " x 12" x 12" $ $ Testing Vacuum Chamber Neoprene Sheets 1/ 8" x 12 " x 12" $ $ Testing Vacuum Chamber Vacuum Gauge Vacuum Gauge $ $ Testing Vacuum Chamber Vacuum Pump Single Stage Electric $ $ Testing Vacuum Chamber Drain Kit Automatic Compressor Drain Kit $ $ 9.00 Subgroup Total $ Figure 60. Breakdown of the Testing Bill of Materials Subgroup Assembly Component Description, Source, and/or Link (URL) Figure 61. Breakdown of the Payloads Bill of Materials Unit Cost Quantity Total Cost $ $ $ Payloads Detection Microsoft LifeCam Studio Amazon $ Payloads Landing Hazard Raspberry Pi 2 Model B Amazon $ Payloads Detection Landing Hazard Xbee-PRO 900 Sparkfun $ Payloads Detection Cisco Airnet Dipole AntennaAmazon $ $ 7.00 Payloads Landing Hazard Sandisk 64 GB MicroSD Amazon $ $ Payloads Landing Hazard G-10/FR-4 Sheet Stock professionalplastics $ $ Payloads Autorotation 1/2 inch Bearing McMaster-Carr $ $ Payloads Autorotation 3/8 inch Bearing McMaster-Carr $ $ 9.04 Payloads Autorotation Outer Shaft Lowe's $ $ Payloads Autorotation Inner Shaft Lowe's $ $ 1.28 Payloads Autorotation StratoLoggerCF Altimeter perfectflitedirect $ $ Payloads Autorotation Parachute Apogeerockets $ $ Payloads Autorotation GPS reciever Amazon $ $ Payloads Autorotation Linear Actuator McMaster-Carr $ $ Payloads Autorotation ABS Plastic 3D printing, cost per kilogram $ $ 1.33 Payloads Autorotation Antenna Amazon $ $ Subgroup Total $

149 149 Subgroup Assembly Component Description, Source, and/or Link (URL) Figure 62. Breakdown of the Educational Engagements Bill of Materials Unit Cost Quantity Educational Engagements Education 24 Pk Estes A8-3 Bulk Target $ $ Educational Engagements Education 15 Count 2 L Soda Bottles Target $ $ Educational Engagements Education Bicycle Pump Target $ $ Educational Engagements Education 3 Display Board Target $ $ Educational Engagements Education Construction Paper Target $ $ Educational Engagements Education Corks Target $ $ Subgroup Total $ Total Cost Subgroup Assembly Component Description, Source, and/or Total Unit Cost Quantity Link (URL) Cost Travel Travel Group Travel Cost Hotels, Fuel, Food $ 3, $ 3, Subgroup Total $ 3, Figure 63. Breakdown of Travel Cost Estimates

150 Funding Plan In order to complete this year s project, the team is receiving funding from corporate sponsors, the Department of Mechanical and Aerospace Engineering, and internally through UF Student Government. These sponsors will include but will not be limited to The Boeing Company, Aerojet Rocketdyne, Dassault Systemes, and CD Adaptco. Funds already acquired include roughly $8,000 from the MAE Department and $500 from UF Student Government. This will be enough to cover the expense of the NASA Student Launch Competition and educational outreach. To increase this funding for the following year s projects, the team plans to pursue attracting more sponsors and additional support from Student Government. This will allow future teams to pursue more ambitious projects that would require a larger budget. The breakdown of funding is shown in Table 23. In this table, the highlights in green indicate that the funding has already been acquired. In order to offset the travel cost to the competition, a portion of the lodging cost will be distributed amongst attendees, and each member will be in charge of covering any miscellaneous travel expenses themselves. Table 23: Summary of Funding Plan Source Amount UF MAE Department $8,000 UF Student $500 Government Corporate $500 Sponsorships Total $10,750

151 Timeline The Gantt charts shown below highlights both the major milestones for the project, as well as some of the specific tasks for the upcoming CDR. This master schedule includes a manufacturing timeline, a list of educational engagements, deadlines for testing, plans for ordering materials, and many other details. October Start Date: oct oct oct oct sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri Sub-Teams Start Research and Design Work for Start Writing and Continue Research and Design Work Educational Engagement Team 10/1/2016 week 1 week 2 week 3 week 4 Leads Meeting General Meeting NSBE Shadow day Leads Meeting Leads Meeting Submit to Lead

152 152 November Start Date: Team Manufacturing Sub-Team Payloads 10/29/2016 oct nov nov nov nov sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri Website Post General General Subscale Launch Meeting Meeting Launch Subscale Manufacturing December Start Date: week 5 week 6 week 7 week 8 11/26/2016 General Meeting Prototype Design Complete General Meeting Team Sub-Teams week 9 week 10 week 11 week 12 nov dec dec dec dec sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri Work on CDR Meeting CDR Q&A Meeting Subscale Backup January Start Date: 1/4/2017 week 13 week 14 week 15 week 16 jan jan jan jan wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue Avionics Testing Educational Engagement Alachua School Duval School Stephen Foster Manufacturing Payloads Meeting Manufacture Full Scale Meeting Meeting Testing Meeting Testing Recovery Meeting Meeting Testing Sub-Teams Team Submit Revise Submit CDR to CDR CDR to Lead Dr. Lind Leads Meeting General CDR Meeting Due General Meeting

153 153 February Start Date: Educational Engagement Manufacturing Payloads 2/1/2017 week 1 week 2 week 3 feb feb feb feb wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue Shell School Full Scale Manufacturing Testing Irby School T. School Battlebots week 4 Meeting Meeting Meeting Lincoln School Recovery Testing Sub-Teams Team March Leads Meeting General FRR Meeting Q&A Prepare for Launch Full Scale Work on FRR Leads Meeting Backup Launch Leads Meeting General Meeting Submit FRR Revise Submit CDR CDR to Dr. Lind Start Date: 3/1/2017 Team Structures week 17 week 18 week 19 week 20 mar mar mar mar wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue FRR Due FRR Video Teleconfrences Prepare Rocket Components for Competition and Paint the Rocket April Start Date: 3/29/2017 week 21 week 22 week 23 week 24 mar apr apr apr apr wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue wed thu fri sat sun mon tue Team Prepare for competition and travel LRR LRR Rocket Safety Launch Backup Fair Launch Briefing Work on PLAR Figure 64. Gantt Chart Schedule Finish PLAR PLAR Due

154 Educational Engagement The University of Florida s Rocket Team will reach out to as many schools as possible at a variety of grade levels within the Gainesville community. The goal is to provide a quality impact on the students. The primary form of engagement will be through activities and lessons in the classroom. The team has coordinated with administration and teachers to participate and host school events. The most satisfying way of reaching out and engaging local students is personally becoming a part of their day and educational experience in the classroom. Being able to captivate the attention of students on the topic of STEM and rocketry will prove more fulfilling than reaching the requirements for competition. The Educational Engagement Officer organizes these presentations. The presentations will consist of the following: 1. Captivate the Attention and Establish Connection with Students 2. Present information about STEM/Rocketry 3. Opportunity for Q&A session/ Academic Opportunities 4. Individual Hands on Activities 5. Engaging Rocket Launch 1. Captivate the Attention and Establish Connection with Students: Each presentation will begin with the present members of the team introducing themselves providing a brief overview of what will be discussed. Captivating the classroom and connecting with the student s level will prove the major priority and task throughout the presentation. The team will have potential ice-breakers to better hold the attention of students through the class period. Introducing humorous examples of physics in the world around them delivered by an engaging and confident member of the team. 2. Present Information about STEM/Rockets: This is where the bulk of the technical information will be presented. However, slides will be organized dependent on the grade level being presented to. It is important for the team to note the differences required when presenting across a variety of ages and experience levels in the subject. The technical side of the presentation is kept to a minimum and geared toward the current knowledge of the audience to avoid losing the interest of the students. The general layout of the slides should be simple and clean to not overwhelm the viewers. The team should ask questions on different parts of the slideshow to try and keep the audience engaged. Accompanying the presenters will be the previous year s NASA Student Launch Rocket. The 14-foot rocket will allow the students to see the value and potential of the techniques discussed during the presentation able to visually absorb the lesson they just learned. 3. Opportunity for Q&A Session/ Academic Opportunities: This part of the visit should encourage open discussion and interaction with students. Questions do not necessarily need to be about STEM/Rocketry, and may include anything about high school, college, and even other fields. The team should take the time to relate themselves to the students at their age and talk about the steps they took and opportunities they seized to get to where they are today. The discussion includes the benefits of AP testing, dual enrollment, and getting ahead academically, where the team can speak specifically on how advancing one s education has benefitted them already. The team should not alienate students interested in non-

155 155 STEM field and instead talk about the availability of majors in just about any field and the many different clubs available to students on a college campus. 4. Individual Hands on Activities Each event that takes place will incorporate a variety of hands on activities for students. After team discussions, this year students will be able to use and manipulate their own set for activities rather than spending time waiting for one of the few sets to become available. Students will be able to create their own rocket using a pop film canister as a base fuselage and paper stencil guides to create the nose, fins and fuselage length students will then be able to put their rocket to the test, using the reaction of Alka-Seltzer and water to launch their rocket. This will allow students to experiment and customize their design, influencing dependent variables and optimizing their rocket design to reach their desired objective something that is crucial in rocket design as the rest of the team can attest. Students will be also to get their hands on a set of Reaction Rockets student will be able to see visually the concept of energy conservation and energy conversion while dropping the the launcher from a certain height shooting the rocket ending up higher than the original drop height. These activities will give students a greater opportunity to have their own individual experience and create their own questions and with our guidance for an understanding of the laws and concepts on display during the event. 5. Engaging Rocket Launch Students will be given an opportunity to watch a rocket launch at the end of the presentation. The rocket used is a small model is a Grade A or B motor. The team will set up the rocket outside in the school s open field. They will give students the instruction to leave plenty of room between themselves and the rocket as well as to not attempt to catch the rocket on its descent. The spare motors are also kept at a significant distance away from the rocket. A team member will personally recover the rocket after descent. One student will be given the chance to press the ignition button to launch the rocket and this will be decided by either the most involved student during the discussion indoors or by asking the class to answer a question about the presentation. The University of Florida will host the annual E-Week event in February bringing the community of Gainesville together with engaging STEM activities and curriculum. This year a new event will be introduced called Built to Battle, teams will be able to enter a small robot and compete Battlebots style. The Rocket Team will use this opportunity to mentor and sponsor a local middle school as they work to build their robot in preparation for the competition which will take place on February 19 th, 2017 Coordination with Alachua County Schools 21 st CCLC Project Manager, a federally funded after school program with a focus on robotics. The program includes 530 students across 9 local sites across the county. We as a team will be visiting these program sites on these tentative dates: Alachua Elementary January 12 th 2017 Duval Elementary January 19 th 2017 Stephen-Foster Elementary January 26 th 2017 Shell Elementary February 2 nd 2017

156 156 Irby Elementary February 9 th 2017 Terwilliger Elementary February 16 th 2017 Rawlings Elementary February 23 rd 2017 Lincoln Middle School March 2 nd 2017 This would put our student outreach amongst these events alone at 480 student adding UF sponsored events and coordinated events with other organizations we look to have a hands on interactive outreach with close to 500 students, our goal for the competition cycle.

157 Conclusion The UF Rocket Team exists to pursue projects that produce meaningful results. The NASA Student Launch is an opportunity for the team to attempt to overcome new challenges and solve exciting problems while following this goal. The team will fly a rocket to 5,280 feet AGL using a vehicle manufactured with more detail and precision than it has in the past. Constructing a body tube of carbon fiber, developing several electronics bays, and designing and implementing a new recovery system all represent areas of experience and strength for the team. All of this is being done to bolster the success of a vertical landing in support of research and development for NASA. A camera system that scans the ground and detects the three tarps, could potentially be used for finding safe landing zones on unfamiliar and foreign places. An autorotation descent could be used to bring payloads to ground level safely with full descent speed and directional control that a parachute cannot provide. The team feels confident that the preliminary design will perform well and collect useful, exciting information. Many possibilities were considered as options to achieving the goals set forth, and after careful analysis, only the best options were chosen. The design is such that the transition to a final, detailed design will be smooth and successful. The team is looking forward to seeing the project mature into a finalized product.

158 158 Appendix

159 MAE Student Design Center Safety Steward Responsibilities As a safety steward your job is to make sure the SDC remains safe, clean, and professional. Even if you are not the steward on duty for your team, please keep a watchful eye on your surroundings while at the SDC. Your presence is a crucial part to maintaining the standards expected at the SDC. Below is a list of responsibilities expected of all safety stewards. Safety stewards found being remiss in their duties will lose their facility use privileges. 1. Enforce all protocols outlined in the Rules for Facility Use document. This includes policies for personal safety; equipment use; facility cleanliness, organization, and respect; proper language; use of the Material & Tool List and Broken / Lost Tooling List; and all other miscellaneous policies. 2. Have a strong understanding of each machine at the SDC. You can t effectively train students in proper equipment use unless you possess a solid understanding of each machine and process. You will also be ineffective at proactively identifying and preventing mistakes that cause injury or damage. 3. Train students on machines, administer knowledge quizzes, and sign authorization sheets. If a student requests machine training, it is your responsibility to train him/her to the standard expected and outlined in the safety protocols. After training, administer the knowledge quiz to assess their understanding of the safety protocols for the specific machine. If the student passes the quiz, add their name to the approved list of users for that machine so (s)he can use the machine with steward supervision in the future. 4. Verify students are trained and authorized on each machine they use. Your primary responsibility is to ensure students are trained and authorized on each machine they use by referencing the lists of authorized users located by each machine. Students using machines on which they are not authorized lose facility use privileges, effective immediately. 5. Setup each approved equipment user each time (s)he works on a new part. Two trained users should catch more mistakes, so always setup team members each time they use facility equipment. For example, if a team member wishes to use a bandsaw to cut a piece of 4130 alloy steel tubing, check its hardness with a file to ensure it is soft enough to cut with a bandsaw, select the appropriate bandsaw, change the blade so its pitch matches the material thickness, and watch the student make the first cut. If the following day another student desires to cut a piece of steel flat bar, the same checks need to be made with a safety steward. Even if the same student desires to cut more 4130 alloy tubing another day, the same checks will need to be made, which require the presence of safety steward.

160 6. Keep watch over powered machinery as it is being used. Accidents can happen to the best trained users. Therefore even though all users of powered machinery must be trained to be allowed to use them, safety stewards should still keep a watchful eye to make sure that machinery is being used safely and correctly. Do not hesitate to interject if a student is making a mistake on the machines. 7. Manage common use tools access. Common use facility tools like sanders and grinders can be checked out using the Material & Tool Use List and must be returned after use each work session so all users have equal access to them. During checkout, a student must ask a safety steward to retrieve the item from its storage location. Upon return, a safety steward must check that the tool has been respectfully cleaned by the user and that it functions properly prior to returning it to its storage location. Users who fail to clean and return tools each session will lose use privileges. 8. Ensure students clean machines after each use and accept responsibility for stations not up to SDC cleaning standards. Machine stations should always be left cleaner than they are found. Holding students to this expectation helps keep the SDC an efficient facility by preventing premature deterioration of machines, floors, and work surfaces. If a student cannot clean their workstation(s) properly, facility use privileges should be revoked. That said, safety stewards will also be held accountable for dirty stations, so ensure users properly clean each station immediately after use (not at the end of each work session, at which time cleaning will be easily or conveniently forgotten). 9. Ensure students keep bays neat and clean. A clean SDC communicates professionalism and appreciation. Teams should adhere to the cleanliness policies outlined in the Rules for Facility Use document and your job is to ensure they do. This includes spills, general trash, the strict no-food policy, as well as general tidiness. Balance being respectful yet stern. 10. Safety stewards are never required to assist other student groups. This might sound odd at first, but we are never requiring safety stewards to assist other student groups using the facility. The first reason is accountability: if a mistake occurs it s more difficult to assign responsibility. The second reason is that each group using the facility should care enough to put forth responsible members from their team for training who have completed EML2322L instead of burdening other groups safety stewards. That said, please feel free to help other student groups on occasion if they do not have their own steward present, but we ask that you do not make a habit of doing so for the reasons mentioned. 11. Report concerns, problems, or suggestions for improvement to the lab manager (mjb@ufl.edu) in an with SDC in the subject line.

161 MAE Student Design Center SAFETY STEWARD TRAINING OUTLINE Introduction: The following are advanced operations and tasks you should understand as a safety steward. These points are not to replace a thorough understanding of the operating procedures outlined in the SDC Rules document. This material is to provide supplementary, more advanced knowledge intended for those in supervisory positions.

162 MAE Student Design Center DELTA AND MSC BANDSAWS Changing Blades: Note that each blade is stored with a tag denoting its tooth pitch / range and appropriate machine. A higher blade pitch correlates to a lower TPI (tooth per inch) and therefore is meant for thicker materials. More teeth per inch means lower tooth pitch, which is used with thinner materials. Refer to the bandsaw chart above each bandsaw to ensure correct blade pitch. 1. Always wear gloves because the bandsaw blades have sharp teeth. 2. Ensure machine is off and the power is disconnected. 3. Remove current blade from machine. Open blade covers and loosen tension on the drive wheels until the blade can be pulled off and subsequently freed from the bandsaw. 4. Coil blade for storage. Point teeth away from you, step on bottom of blade, and twist around while lowering top half of blade. Blade should coil naturally into 3 loops. 5. Remove tag from hook on machine and attach to coiled blade. Ensure the correct tag stays on the correct blade, as failure to do so will result in ruined blades. 6. Store blade in proper location. Ensure blades are stored coiled in the proper cabinet. 7. Remove and hook new blade tag on machine. This tag denotes the active blade on the machine and informs the next user if a different blade is necessary for the next part / job. 8. Uncoil new blade. Find an open area in the workshop, point teeth away from you, step on one of the three blade loops, and slowly uncoil bandsaw blade. 9. Install new blade on machine. Open covers and place blade on drive wheels and between the blade guides. Ensure teeth are pointing in the right direction (toward the user, and pointed tips down). Tension the blade until snug, and cycle the machine on and off in short bursts for a few rotations until the blade centers itself on the drive wheels. Finally, close the drive wheels and perform a test cut to ensure the new blade functions properly. Changing Speeds: The Delta bandsaw has no variable speed; it is solely for cutting wood/plastics. Never cut metals or other stronger/harder materials on the Delta bandsaw. To change the speed on the MSC bandsaw, rotate the knob until the needle is pointed to the desired speed, while the machine is running. Reference the posted speed chart to know what speed should be used. Note that higher speeds are for softer materials.

163 MAE Student Design Center ROLL-IN VERTICAL GRAVITY FEED BANDSAW Changing Blades: Note that each blade is stored with a tag denoting its tooth pitch / range and appropriate machine. A higher blade pitch correlates to a lower TPI (tooth per inch) and therefore is meant for thicker materials. More teeth per inch means lower tooth pitch, which is used with thinner materials. Refer to the bandsaw chart above each bandsaw to ensure correct blade pitch. 1. Always wear gloves because the bandsaw blades have sharp teeth. 2. Ensure machine is off and the power is disconnected. 3. Remove current blade from machine. Retract the carriage completely using the carriage retract lever. Open blade covers and loosen tension on the drive wheels until the blade can be pulled off and subsequently freed from the bandsaw. 4. Coil blade for storage. Point teeth away from you, step on bottom of blade, and twist around while lowering top half of blade. Blade should coil naturally into 3 loops. 5. Remove tag from hook on machine and attach to coiled blade. Ensure the correct tag stays on the correct blade, as failure to do so will result in ruined blades. 6. Store blade in proper location. Ensure blades are stored coiled in the proper cabinet. 7. Remove and hook new blade tag on machine. This tag denotes the active blade on the machine and informs the next user if a different blade is necessary for the next part / job. 8. Uncoil new blade. Find an open area in the workshop, point teeth away from you, step on one of the three blade loops, and slowly uncoil bandsaw blade. 9. Install new blade on machine. Open covers and place blade on drive wheels and between the blade guides. Ensure teeth are pointing in the right direction (toward the user, and pointed tips down). Tension the blade until snug, and cycle the machine on and off in short bursts for a few rotations until the blade centers itself on the drive wheels. Finally, close the drive wheels and perform a test cut to ensure the new blade functions properly. Changing Speeds: The Roll-In bandsaw has four speeds which are selected via the pulley ratios between the drive motor and drive wheel. The pulley ratios can be set to 1:4, 2:3, 3:2, or 4:1 (slowest to highest). Higher speeds are for softer material. To change speeds: 1. Ensure machine is off by hitting the red button on the front of the machine. 2. Move carriage forward, stopping near the end of its travel by shutting off the hydraulic feed valve. When closing the valve, make sure not to over-tighten the feed knob, as doing so will ruin the valve; a gentle twist is all that s ever required. 3. Uncover pulley set by loosening and removing the black knob holding the orange sheet metal cover. Place cover in a safe place to ensure 4. Using the provided tool, loosen pulley tension by adjusting the threaded rod at the bottom of the machine.

164 Changing Speeds (continued): 5. Change belt location to select desired ratio. 6. Re-tighten belt to proper tension by re-adjusting the threaded rod at the bottom of the machine using the provided tool. 7. Finally, hang appropriate speed indicator on display hook to show others what speed the machine is set to. Put the other indicator on the storage hook on the rear of the machine. Part Orientation: Refer to the bandsaw chart for proper orientation of certain workpieces. In general, try to keep a consistent cross-sectional area in contact with the teeth at all times to ensure consistent cutting. Also, orient the part so it is in contact with the most teeth possible. For example, flat stock has a consistent cross sectional area in both a vertical and horizontal orientation. However, orienting the part so that the cross section is taller than it is wide on a vertical bandsaw ensures the most teeth are in contact with the part.

165 MAE Student Design Center PLASMA CUTTER Changing Cutting Torch Consumables: Consumables are a normal wear item that should last for 5 8 hours of cutting. Worn consumables do not cause any damage to the plasma cutting unit; they just reduce cut quality. Therefore changing consumables prematurely wastes money. The proper procedure for changing consumables is noted below: 1. Turn off power to the machine (it is okay to leave the air supply turned on). 2. Remove the retaining cup. Do not use any tools, only your hands. 3. Remove tip, inspect, and replace if opening is deformed or 50% oversize. 4. Remove electrode, inspect and replace if center has a pit more than a 1/16 deep. 5. Remove swirl ring, inspect and replace if side holes are plugged. 6. Check O-ring for cracks or worn spots, and replace if necessary. 7. Carefully reassemble parts in reverse order. Troubleshooting Tips: 1. If plasma cutter is activating but not cutting, ensure the ground clamp has been connected to the workpiece or the welding table and that the workpiece is conductive.

166 MAE Student Design Center MIG WELDER General Tips for MIG Welding: Besides understanding the safety procedures and basic operation of the MIG welding equipment, the following are a few tips and important things to know about the MIG welder to help students stay safe and prevent them from damaging the machine or their parts. 1. Only weld material which has been cleaned thoroughly (using a sander, grinder, wire brush, wire wheel, etc.) and degreased with acetone or alcohol to remove any residue which will contaminate the weld. 2. Generally, set the feedrate to ten times the voltage (3V/30ipm, 4V/40ipm, etc.). The higher the voltage, the quicker the filler wire melts and the faster you should feed across your part. To ensure enough filler material is in your weld, proportionately increase the feedrate. 3. Do everything you can to position yourself so you are comfortable prior to welding. 4. Keep electrode stick out within ½ of the gun tip for best weld quality. 5. Set welder one heat range higher for tack welds than for normal welding. By definition tack welds start off cold, so the extra heat helps increase penetration. Just remember to turn the heat setting back down prior to final welding. 6. Always run a test pass on a piece of scrap material of similar composition and thickness. If your test passes don t turn out well, your real parts won t either. 7. Make sure the gun nozzle remains clean and free of slag. Periodically remove the brass nozzle on the end of the welding gun and clean the buildup from inside the nozzle with the MIG-specific needle-nose pliers. This prevents the copper contact tip from being damaged prematurely and negatively affecting the wire feed and weld quality. 8. Use nozzle dip to help keep the nozzle clean and free from buildup if welding for prolonged periods of time (>30 min). To apply the nozzle dip, heat up the tip of the welding gun by welding a 1 long bead on a piece of scrap metal for approximately 10 seconds, and then immediately dip the hot nozzle into the jar of nozzle dip approximately ½ deep. This will provide a protective coating that keeps slag from building up inside the nozzle. Troubleshooting Tips: 1. If the MIG welder is on but won t weld, the heat range selector knob may have accidentally been placed between of the power settings. After reselecting the heat range try welding again.

167 MAE Student Design Center TIG WELDER (SEPARATE CERTIFICATION) General Tips for TIG Welding: Besides understanding the safety procedures and basic operation of the TIG welding equipment, the following are a few tips and important things to know about the TIG welder to help students stay safe and prevent them from damaging the machine or their parts. 1. Only weld material which has been cleaned thoroughly (using a sander, grinder, wire brush, wire wheel, etc.) and degreased with acetone or alcohol to remove any residue which will contaminate the weld. 2. Generally, adjust the current setting to 1 amp per of material thickness. 3. Use only ceriated tungsten on the SDC inverter-style TIG welder (never pure tungsten). 4. Match the tungsten diameter (0.040, 1/16, 3/32, or 1/8 ) to the workpiece thickness. 5. Use the smallest filler wire size that won t melt prematurely before dipping into the puddle. 6. Keep electrode stick out within one radius of the TIG torch nozzle s end for best weld quality. 7. Keep the tip of the tungsten within 1/8 of the molten pool at all times when welding. 8. Do everything you can to position yourself so you are comfortable prior to welding. 9. Keep the filler wire in the shielding gas stream when not dipping into the puddle. 10. Upon completing each weld keep the shielding gas aimed at the solidifying weld pool for the entire duration of the post-flow timer setting (typically 7 10 seconds). Pulling the torch away prematurely can cause contamination by surrounding oxygen. 11. Minimize the torch angle to prevent shielding gas vortices from drawing in surrounding oxygen and contaminating the molted weld pool. 12. Be careful when setting the TIG torch down, as the cup on the end of the torch is made of ceramic and is therefore very brittle. 13. Always run a few test passes (welds) on scrap material of similar composition and thickness. If your test passes don t turn out well, your real parts won t either. 14. Spend time on WeldingTipsandTricks.com website for more training tips. Learning to TIG Weld: The following strategy seems to work best for new students desiring to learn how to TIG weld: 15. Purchase 1/16 diameter ceriated tungsten from a vendor like Tig Depot, Cyber Weld, Airgas, or McMaster-Carr. We pay for the shielding gas and electricity, but students must provide their own Tungsten electrodes when using the TIG welder. 16. Sharpen the tungsten on the dedicated tungsten grinder so it has a profile similar to the shape of a pencil tip (see figure 1 below).

168 Figure 1: Properly sharpened tungsten electrodes 17. Begin by making small welds without adding filler across a piece of 1/16 to 1/8 scrap steel paying attention to weld visibility, speed, torch angle (< 20 degrees from vertical), electrode distance (within 1/8 of the molten pool), arc starting and stopping (slowly). Start with your dominant hand and then practice with your other hand. The torch velocity should be very consistent. Every minutes dip the piece in water to prevent from overheating the welds. Keep the torch over the puddle for at least 5 seconds at the end of each weld. Figure 2: Proper torch angle, filler rod angle, and electrode distance