The University of Toledo

Transcription

1 The University of Toledo Project Kronos Preliminary Design Review 11/03/2017 University of Toledo UT Rocketry Club 2801 W Bancroft St. MS 105 Toledo, OH 43606

2 Contents 1 Summary of Proposal Team Summary Launch Vehicle Summary Size and Mass Motor Selection Recovery System Milestone Review Flysheet Payload Summary Payload Title Payload Experiment Summary Changes Since Proposal Changes to Vehicle Criteria Changes to Payload Criteria Changes to Project Plan Vehicle Criteria Design Rationale Mission Statement Mission Success Criteria Vehicle Design Alternatives Vehicle Design Alternatives Nose cone Vehicle Design Alternatives Body Design Vehicle Design Alternatives Mass Management Vehicle Design Alternatives Motor Mount Sizing Vehicle Design Alternatives Fin Designs Ballast Weight Preliminary Vehicle Design Nose cone Payload Bay Main Parachute Bay Electronics Bay Drogue Parachute Bay Lower Body Assembly

3 3.2.8 Materials Motor Alternatives Recovery Subsystem Recovery Design Alternatives CO 2 Canisters Black Powder StratoLogger Collects Flight Altimeters Drogue Parachute Sizes Preliminary Recovery Design Preliminary Analysis Redundancy Mission Performance Predictions Preliminary Simulation Stability Margin Kinetic Energy Drift Calculation Safety FMEA Procedure Overview Personnel Hazard Analysis Failure Modes and Effects Analysis Environmental Hazard Analysis Payload Design Rationale Payload Mission Statement Payload Success Criteria Payload Design Alternatives Solenoid Deployed Solar Panels Flexible Solar Panels Clam Shell Solar Panels Sliding Solar Panels Gear Wheels Non-Gear Wheels Payload Bay Orientation System

4 Orientation-Independent Rover Solenoid Payload Latch Rotary Payload Latch Black Powder Payload Hatch CO 2 Payload Hatch Sonic Distance Measurement GPS Distance Measurement Preliminary Payload Design Rover Design Solar Panel Design CO 2 Deployment System Payload Electronics Subsystem Interfaces Project Plan Requirements Verification Team Derived Requirements Education Budget Funding Plan Timeline Appendix

5 Table of Tables Table 1: Mission Success Criteria Table 2: Pre-Flight Success Criteria Table 3: Mass of vehicle sections Table 4: Kinetic Energy under 32 inch drogue Table 5: Kinetic Energy under 28 inch drogue Table 6: Kinetic Energy under 24 inch drogue Table 7: Kinetic Energy under Main Parachute at Ground Contact Table 8: Drift Distances Table 12: Payload Success Criteria Table 13: Requirements Verification Table 14: Team Derived Requirements Table of Figures Figure 1: Fin Design Figure 2: Fin Design Figure 3: Fin Design Figure 4: Fin Design Figure 5: Ballast configuration Figure 6: Vehicle Configuration Figure 7: Fin Assembly Figure 8: Recovery Sled Figure 9: Recovery System Circuit Diagram Figure 10: StratologgerCF Figure 11: Flight Profile Figure 12: K1000T-P Thrust Curve Figure 13: Stability Margin vs Time Figure 14: Center of Gravity and Pressure positions Figure 15: Drift distance with 0 mph wind Figure 16: Drift distance with 5 mph wind Figure 17: Drift distance with 10 mph wind Figure 18: Drift distance with 15 mph wind Figure 19: Drift distance with 20 mph wind Figure 20: Solenoid Deployed Solar Panel Design Figure 21: Flexible Solar Panel Figure 22: Gear Wheel Figure 23: Solenoid Payload Latch Figure 24: Rotary Payload Latch Figure 25: Peregrine CO 2 Kit Figure 26: Isometric view of payload bay Figure 27: Micro Servo Figure 28: Top view of Payload Bay

6 Figure 29: Isometric view of the rover Figure 30: Front view of the rover Figure 31: 1/8 Inch Rack Figure 32: 5/16 Inch Rack Figure 33: Panel Carrier drawing Figure 34: Payload Tube Base drawing Figure 35: Rover Base Drawing Figure 36: Solar panel mechanism detail Figure 37: Rover with deployed solar panels Figure 38: Peregrine CO2 Canister Figure 39: Rover Deployment Coupler Figure 40: CO 2 Canister drawing Figure 41: ATMega32U2 Development Board Front and Back Figure 42: ATMega32U2 Development Board Circuit Figure 43: SAM-MBQ-O Development Board Front and Back Figure 44: AX5043 RF Transceiver Circuit Figure 45: AX5043 RF Transceiver Development Board Front and Back Figure 46: Electrical Block Diagram Figure 47: Payload Integration Figure 48: Education Progress Figure 49: Solar Panel drawing Figure 50: Servo Figure 51: Gear Wheel drawing Figure 52: Solar Panel Gear drawing Figure 53: Wheel Ball drawing Figure 54: Rover Deployment Bulkhead drawing Figure 55: Motor Selection Form

7 1 Summary of Proposal 1.1 Team Summary Team Leader: Michael Blackwood (440) NAR # L2 Certified Safety Officer: Victoria Raber (419) Team Mentor: Art Upton (419) NAR #26255 L3 Certified 6 6

8 1.2 Launch Vehicle Summary Size and Mass The current design of the rocket is built using a 5-inch G-12 fiberglass airframe body. The nose cone is a filament wound fiberglass 4:1 Ogive style nose cone. The combination of the nose cone and the airframe puts the total length of the rocket at 87 inches or 7.25 feet. The mass of the rocket, without the motor, has been estimated as close as possible to what the team expects in real life. These estimations come from obtaining mass measurements for each component from their external sources. Any extra weights, such as epoxy resin and paint, are estimated by using standard design practices. Once all components have been accounted for, the rocket, without a motor, will weigh 325 oz. or 20.49lbs. The final projected weight of the rocket, including the motor, is 416 oz. or lbs Motor Selection The rocket will be powered by an Aerotech K1000T-P. This motor is commercially available for purchase, and meets all requirements and standards set by NASA and UT Rocketry. Several simulations using OpenRocket ensured that the rocket will have a safe and successful flight. During simulations in OpenRocket, the following data was recorded for the Aerotech K1000T-P motor. The Aerotech K1000T-P is a solid-state class K motor approved by NAR, TRA, and CAR that does not expel metal shards. The Aerotech K1000T-P has an impulse of 2,497 N*s and an average thrust of 1,012 N. The physical dimensions of this motor are 396 millimeters in length and 75 millimeters in diameter. With a weight of 90.8 ounces and a burn time of 2.47 seconds, the rocket will achieve a maximum altitude of 5,355 feet. The rocket will be below the maximum altitude of 5,600 feet and come within 75 feet of the 5,280 feet target. Peak velocity of the rocket will be 660 feet per second or 0.59 Mach. This velocity is below the speed of sound, and transonic speeds and will therefore ensure that the rocket will have a safe flight. Similarly, the exit rail velocity, using an 8 eight-foot rail, is projected to be 69.5 feet per second. Using a 12-foot rail the exit velocity is projected to be 84.5 feet per second. The team will use an 8-foot launch rail as the rocket exceeds the exit rail velocity requirement of 52 feet per second. The thrust-to-weight (TTW) ratio is 9:1, which is well above the minimum 5:1 TTW ratio Recovery System The recovery system will consist of two main sections. These sections are as follows: drogue parachute and main parachute. The drogue parachute will be connected in the secondary body tube via a forged eye bolt and a quick link on the electronics bay section with a braided nylon shock cord holding the electronics bay and the secondary body tube together on the back. The shock cord will be tied with a non-slip knot on both ends and will be epoxied to prevent slippage. The main parachute will be connected to the electronics bay with a forged eye bolt and a quick link. There will also be a braided nylon shock cord that connects the electronics bay to the aft section of the nose cone. The nose cone will be secured to the body using shear pins. 7 7

9 This shock cord will also be tied on both ends using a non-slip knot and will be epoxied to prevent slippage. The two sections will be separated via an ejection charge Milestone Review Flysheet 8 8

10 1.3 Payload Summary Payload Title The team's payload is the Deployable Rover Payload. The Rover's name is Zeus. Section 4.5 of the NASA Student Launch Handbook will be described below Payload Experiment Summary The payload will be a four-wheeled rover that will deploy itself from the rocket after landing and a bulkhead is released. To ensure that the rover deploys, and to allow for analysis post launch, a camera will be installed at the base of the Payload tube. After the rover has driven out of the rocket, the rover will execute a small turn and drive for an amount of time that will ensure the rover has traveled five feet. Once the rover has driven five feet it will deploy both of its solar panels, creating a total exposed surface area of square inches of solar panel surface area. 9 9

11 2 Changes Since Proposal 2.1 Changes to Vehicle Criteria The vehicle has undergone multiple changes that have been critical to the overall design. These changes include; changes to the airframe design, fin design, payload exiting procedure, and method of keeping a constant altitude. The changes in the airframe design have taken place in three principal areas. These changes include: the motor mount, the vehicle airframe, and the nose cone. Within the motor mount, the diameter of the motor mount tube was increased from 54mm to 75mm. This change was made in order to accommodate for a larger motor to reach the desired height. This change affected the sizing of the centering rings and increased the mass of the vehicle. Changes to the vehicle airframe are based around the concept of removing the transition. By removing the transition, a weight and cost savings, an increase in stability, and a decrease in drag can all be achieved. These changes are detailed below in the Alternative Designs section. The fin design was developed based off of four designs and each was analyzed to determine the most efficient design for the rocket. The design analysis is reviewed below in the alternative designs. The payload deployment procedure has been an area of heated debate. Specifically, the payload deployment procedure has been developed and redeveloped multiple times in order to formulate the most efficient and safe design possible. This design is reviewed in more detail below and is compared to the previous designs that were proposed. Finally, a method of keeping a constant altitude was developed in order to assist in reaching the one-mile goal. This is achieved through a simple system of washers and is described in detail below. 2.2 Changes to Payload Criteria The Payload team decided to drop the goal of sonic distance testing in favor of a simpler, and likely more accurate GPS method. The GPS locations will be saved when the deployment signal is received. During rover deployment, the GPS position will be run through the Haversine formula to evaluate the distance the rover has moved from the rocket. A sufficient safety distance margin (to be evaluated) will be used to ensure the GPS coordinates will position the rover at least five feet from the rocket. Additional sensor inputs can be provided by an IMU and utilized within a Kalman Filter, or similar processing technique, to improve the positional accuracy of the system. Changes have also been made to the design of the rover s solar panel deployment mechanism. Instead of a simple spring-loaded retaining mechanism, a servo driven rack-pinion system will be utilized to simultaneously slide out two solar panels to the sides of the rover. This will increase the surface area of exposed solar panels (only two panels would have been exposed before, now 10 10

12 there will be three panels exposed to the sun with the use of a centrally located solar panel), while creating a more robust, lower power design for the panel deployment. 2.3 Changes to Project Plan There have not been many changes to the project plan since the proposal. The main change that the club has seen is there is a tentative launch date set for December 2. The club will take this opportunity to launch the subscale for Kronos. This time frame gives the club approximately one month to prepare for the launch. Other substantial changes were outlined above in their appropriate sections

13 3 Vehicle Criteria 3.1 Design Rationale Mission Statement It is the goal of UT Rocketry to provide a stable and safe vehicle for the payload from launch to recovery, while ensuring all parts of the rocket are as reliable and safe before launch as they are after recovery. UT Rocketry seeks to deliver a payload as close to the altitude of 5280 feet as possible, while guaranteeing the safety of team members, students and spectators Mission Success Criteria Table 1: Mission Success Criteria Requirement How the requirement will be fulfilled How to verify that the requirement is fulfilled All subsystems must be accurately represented in simulations. The rocket must assemble in a way that allows all subsystems to work as planned. Minimize chances of system failure wherever possible. Maintain a straight flight path. Properly attach the rocket to the launch pad. Properly attach the igniter. Simulations and tests in OpenRocket will use all known variables. The variables tested will be the same variables interacting with the rocket. The distances from each component will be measured so that the components may be secured in the correct location. Will have a minimum of two qualified team members attach each component after being briefed on the correct procedure beforehand. Properly attach the fins to the rocket body using a jig. Will follow all the steps on the setup launcher's checklist. Will follow all the steps on the ignition installation check list. All masses and dimensions of components will be measured once the components are obtained, then they must match the OpenRocket model. After each component is installed and secured, it will be remeasured. Have the vehicle team leader confirm each part has been properly attached. Measure the angle between each fin and use a level to make sure the fins are straight. All the items on the setup launcher's checklist will be met and checked off. All the items on the ignition installation check list will be met and checked off

14 Motor must ignite safely. The recovery system must properly release. Parachute must decrease the rocket to a safe landing velocity. The rocket must be able to fly after refueling. Fins must stay attached and undamaged after landing. Payload must remain stationary during the flight. Payload must properly deploy. Will measure and remeasure all of the mounts of the motor. Will ensure switches are on the proper setting, has new batteries, and wires are attached to the proper terminals. Will ensure the parachute and shock cords are properly attached to the rocket. Will ensure the parachute properly deploys and be of the proper size to slow down the rocket to a safe velocity. Will add more epoxy than the minimum required quantity and use proper fin design. Will interlock the payload wheels to the inside of the body tube. CO2 separates the rocket nose cone to enable the payload to deploy. All the items on the motor preparation check list will be met and checked off. Will construct and complete the check list for the preflight recovery system. Will construct and complete the check list for recovery system assembly. Will perform the proper calculations, double check the assembly of the recovery system, and a create a working sub scale model. Calculate the force experienced at landing and compare to the strength of the fin design. Construct and complete a check list for payload and body tube assembly. Create and complete a check list for nose cone assembly

15 Table 2: Pre-Flight Success Criteria Requirement How the requirement will be fulfilled How to verify that the requirement is fulfilled All systems will be in good, operational order Team leads will review their respective systems for damage. A pre-flight checklist will be observed and followed. The vehicle will show no signs of structural damage Ensuring proper care of the rocket during transportation and after launch. Proper padding will be used during transport and post flight damage inspections will be carried All systems will be properly prepared for launch All team leads ensuring that their respective systems are prepared for launch. out. Each team lead will sign off after verifying their system is prepared for launch. All team leads will ensure their sub-system is properly prepared and installed into the rocket The recovery system will be properly prepared and armed Any means of body tube retention (shear pins, screws, etc.) will be installed before flight The rocket will be properly positioned on the launch stand The rocket will be prepped and armed for the minimum duration specified in the NASA SL Handbook Proper motor preparation and installation will be followed All team leads will verify that their respective system is prepared for launch Team confirms all black power charges are prepared and parachutes are properly secured. Team double checking all shear pins and screws before the launch begins The team will ensure the stand is in the desired position for launch, considering wind and terrain. The payload and recovery system will be designed in accordance to the NASA SL guidelines Instructions that are provided with the motor will be followed and the motor will be securely installed with the proper motor retention method. Each team lead will sign off after verifying their system is prepared for launch. Recovery team double checks all charges and parachute for holes and tares that would compromise the safety of the launch. A preflight check of sheer pins and tightening of screws. The teams calculate for wind and check the stand to make sure it is not pointed at an angle. The required sub systems will be tested for compliance with the guidelines prior to launch. The propulsion team will ensure that all steps are followed, and the team lead will sign off on all procedures

16 Table 3: During-Flight Success Criteria Requirement How the requirement will be fulfilled How to verify that the requirement is fulfilled Rocket successfully leave launch stand Altimeters record altitude data Payload systems deploy at specified times The rocket safely and successfully achieves lift. Proper design considerations will be followed including off-the-rail stability margins and rail exit velocity. The electronics bay will be properly armed and prepared. The rover leaves the rocket once it reaches the ground and receives the deployment signal. Careful observation of the rocket prior to, and just after motor ignition. Motor selection and vehicle design must consider safety margins. Altimeters are in the proper standby mode before the launch. After landing the rover moves 5 feet from the rocket and deploys solar panels. Drogue parachute deploys at apogee Main parachute deploys at 700ft The rocket successfully lands without damage The electronics bay will be properly prepared, and the blast charges will be properly sized and prepped. The electronics bay will be properly prepared, and the blast charges will be properly sized and prepped. During launch and recovery, the rocket stays in one piece and lands without external or internal damage. The recovery team will prepare the blast charges and electronics bay in accordance to pre-tested configurations. The recovery team will prepare the blast charges and electronics bay in accordance to pre-tested configurations. There is no damage to the body tubes, payload, fins, or electronics bay

17 Table 3: Post-Flight Success Criteria Requirement How the requirement will be fulfilled How to verify that the requirement is fulfilled The rover successfully deployed and data is sent back to the ground receiver The panels on the rover are deployed and a signal is sent Back. The data is showing up at the ground station and Being recorded. The rocket is recovered with no critical damage incurred. All material is recovered from the launch field Altimeters and energetic systems are Properly Disarmed after obtaining relevant data 3.2 Vehicle Design Alternatives All prior flight preparation procedures are followed The rocket will land as intended and will not separate in an excessive manner Qualified members who fully understand these energetic systems will disarm and confirm with the team lead The team leader will Verify all flight preparation lists are completed satisfactory All parts are returned to The launch preparation site The team lead will confirm that the systems have been properly disarmed Design alternatives have been a key driving point in developing a suitable and efficient rocket. Many of the design alternatives have been in crucial areas of the rocket that will ultimately affect the output and performance. These alternatives vary from modifying the overall shape of the body and fins, to changing the internal layout of the payload and motor mount. All of these changes, again, have been made to develop a better overall rocket that will have superior performance, provide a safer and more efficient flight, and successfully launch and deploy a rover upon landing. Below are alternative design criteria that were evaluated and ultimately decided upon to make the preliminary design Vehicle Design Alternatives Nose cone The initial design of the nose cone was based off of the current stock of the team's main supplier. The nose cone that was available was a 5-inch filament wound fiberglass 5:1 Von Karman style nose cone. This nose cone would successfully fill many of the requirements set forth for the design, but proved to be costly, and excessive in length and weight upon further discussion and review. As a replacement, a 5-inch filament wound fiberglass 4:1 Ogive style nose cone was chosen. The justification for this change came in the form of a weight savings. The change from a 5:1 to 4:1 style produced a weight savings of 5 ounces and reduced the overall length of the body by 5 inches. This savings will help save on propellant, allowing for a slightly smaller motor to be used

18 3.2.2 Vehicle Design Alternatives Body Design One of the major concerns regarding the overall design of the rocket from the initial proposal was the usage of a transition. This transition would have taken the larger fore section of the vehicle, at 5 inches in diameter, and transitioned it down to 4 inches in diameter towards the aft section. The stability of the rocket was one of the main concerns. One key item in the discussion was how the rocket would "cut" through the air throughout the flight. In the original design, the transition would have created a zone where air would become trapped and create drag, much like how air can become trapped behind a car. With the new design, a straight 5- inch diameter airframe throughout the entire design, this problem is eliminated as the fluid flow across the body is not being obstructed or diverted in any way. Another design that was considered was also with the usage of a transition. A 4-inch diameter body tube would be used in the fore section of the rocket with a transition leading to a 5-inch diameter body tube. From a fluid flow standpoint, this design was still able to "cut" through the air and did not create as much drag as the fluid flow over the body is being pushed away at the transition and not swirling together as in the previous transition set-up. Though this design is more feasible from a fluid flow standpoint, it is must less efficient in achieving the overall goal: to successfully fly a rover and return to the ground safely. Having a 4-inch diameter fore airframe limits the total useable area for integrating the rover and its associated systems and forces the overall design to be overly compact and more difficult to successfully produce. With the 5-inch diameter fore section, the overall useable area is increased and the overall design of the rover and its associated systems can larger in scale. One item that must be considered with this larger area is the increase in weight and the distribution of the weight throughout the payload section of the rocket. The overarching design goal is to have the payload section be as rotationally balanced as possible in order to avoid any shifting or outbalance that could affect the flight performance of the rocket. One of the means used in overcoming this obstacle is the usage of ballast weight to maintain a constant weight throughout the design process Vehicle Design Alternatives Mass Management A ballast weight will be used in the rocket to assist the propulsion team when deciding which motor will produce the apogee closest to the targeted one mile altitude. The usage of a ballast will compensate for any lost or gained mass that may occur throughout the building and design stages as exact, real-life part masses have not been measured and entered and minor changes to the design may occur. A few designs were considered for the ballast including a permanent, fixed ballast that would be made of epoxy resin or similar material, removeable ballast such as metal shot, or a fixed removable ballast such as washers. After discussion, the usage of washers was decided upon. Washers ended up being the design of choice due to a number of advantages it had over the other designs. Primarily, washers allow simple customization of the weights allowing them to be easily adapted to each unique situation. This customizability was not an option in the fixed epoxy resin method, and concerns over the accurate measurement of epoxy during application kept the fixed epoxy resin method out of reach. Another key advantage of the washer method is the ability to securely retain the washers during flight, 17 17

19 keeping the stability of the rocket in check. This was a major safety concern that was found with the metal shot as the shot would have to be secured in place to prevent shifting of the material. If the material did shift during flight, the overall stability of the rocket could be compromised. Therefore, for reasons of safety, ease of customizability, and ease of use, the washer method was chosen to be the ballast method Vehicle Design Alternatives Motor Mount Sizing In order to accommodate a large enough motor to propel the rocket to the desired altitude, two motor mount tube diameters were researched and simulated. A 54-mm motor mount and a 75- millimeter motor mount were the sizes tested. These two motor mounts are both known to be able to hold a motor of appropriate size to propel the rocket to the desired altitude, but further investigation was needed to determine which would be the most appropriate. Initially, the 54- mm motor mount tube was chosen and simulations in OpenRocket were designed and run in order to get a baseline height that could be achieved by that motor mount. After a few simulations were run, the average altitude was in the low to mid 2000 feet range. This distance is nowhere near the 5280 feet that the rocket is required to reach. This prompted the change to go to a 75-mm motor mount tube. From this change and the subsequent simulations that were run, the 75-mm tube ended up being the most viable option to hold a motor that could propel the rocket to the required altitude

20 3.2.5 Vehicle Design Alternatives Fin Designs Numerous fin designs were considered for this year's rocket. Considerations included the resulting stability margins, the cost of material, difficulty to obtain or manufacture, and the rigidity of construction. The designs were all tested within OpenRocket for identical rocket models (not including the main fins) under identical conditions. All fin designs utilized a inch sheet of G12 fiberglass cut into 4 identical trapezoidal fins with a 0.25-inch fillet radius. The first fin design is illustrated in figure 1. The dimensions of the design are as follows: 11-inch root chord, 3 inch tip chord, 6 inch sweep length, and a height of 4.5 inches. The fins would be attached 1 inch above the base of the rocket in relation to the fin tabs. According to the simulations, the rocket would have a stability margin of 2.53 calibers. The velocity at which these fins would break due to fin flutter is 1424 feet per second, significantly higher than the expected maximum flight speed. Figure 1: Fin Design 1 The second fin is illustrated in figure 2. The dimensions of the design are as follows: 8-inch root chord, 6 inch tip chord, 5 inch sweep length, and a height of 4.5 inches. The fins would be attached 0.25 inches above the base of the rocket. According to the simulations, the rocket would have a stability margin of 2.85 calibers. The velocity at which the fins would break due to fin flutter is 1958 feet per second. Figure 2: Fin Design

21 The third fin design is shown in figure 3. The dimensions of the design are as follows: 5-inch root chord, 5 inch tip chord, 4 inch sweep length, and a height of 6 inches. The fins would be attach 3.2 inches above the base of the rocket. According to the simulations, the rocket would have a stability margin of 3.99 calibers. The velocity at which the fins would break due to fin flutter is 1600 feet per second. Figure 3: Fin Design 3 The fourth fin design is shown in figure 4. This design differs from the others in that it uses two sets of 4 identical fins rather than a single set of symmetric fins. The dimensions of the upper four fins are as follows: 5-inch root chord, 2-inch tip chord, 2-inch sweep length, and a height of 4 inches. The upper fins would be attached 6.5 inches above the base of the rocket. The dimensions of the lower four fins are as follows: 4-inch root chord, 2-inch tip chord, 2-inch sweep length, and a height of 4 inches. The lower fins would be attached 0.5 inches above the base of the rocket. According to the simulations the rocket would have a stability margin of 3.44 calibers. The velocity at which the upper fins would break due to fin flutter is 2038 feet per second. The velocity at which the lower fins would break due to fin flutter is 2049 feet per second. Figure 4: Fin Design

22 Fin design one was the final design chosen for the rocket. The velocity at which the fins break from fin flutter is over 100 percent higher than the predicted max velocity. This fin design has a stability caliber above 2.5, well within the safe operating region ensuring stable flight. When the stability caliber approaches 4, the rocket becomes in danger of excessive weather cocking. Weather cocking happens because the rocket pivots on the center of gravity as a result of large forces that occur at the center of pressure, the higher the caliber the greater this acting torque will be. To avoid weather cocking, designs 4 and 2 could not safely be considered. During flight, the stability caliber will increase as the motor weight decreases which puts design 3 at risk of weather cocking. The stability of 2.5 is a good balance between stability and the risk of weather cocking. The shape of fin design 1 also helps reduce the risk of the fins breaking when landing, greatly increasing the chances that the rocket could be re-flown without any notable repairs being required Ballast Weight A Ballast Weight will be added in order to account for variations in manufactured parts as well as the potential modifications to the payload designs. The ballast consists of an eyebolt attached to the opposite bulk plate as the CO 2 tank used to remove the nose cone during rover deployment. This eye bot will carry the washers that form the ballast weights. The ballast allows more accurate predictions of the total rocket's mass to be made. Washers can be added and subtracted as needed in order to maintain a constant mass throughout development. The ballast weight also allows for simple adjustments to be made after testing and allows for more accurate launch conditions. This weight will be enough to increase or decrease maximum height by a few feet while not significantly changing the center of mass. Figure 5: Ballast configuration 21 21

23 3.2.7 Preliminary Vehicle Design Nose cone Figure 6: Vehicle Configuration The nose cone, item number 1, will be made of G12 Fiber Wound Fiberglass in a 4:1 Ogive style with a length of 20 inches and a diameter of 5 inches at its base. The nose cone will be comprised of a main outer section and a coupler section. The coupler section will have a bulk plate attached to the aft section of the coupler. In order to secure the nose cone to the body, two means of attachment will be used. The first method is by four (4) shear pins and the second method are four (4) screws. The shear pins will attach the coupler to the payload bay and will allow the nose cone to detach and allow the payload to deploy from the rover. The second means of retention, screws, will attach the coupler to the nose cone. By having this second means of attachment, the team will be able to place the GPS in the upper portion of the nose cone away from the rest of the electronics in the vehicle. The bulkhead will be securely attached to the coupler by epoxy resin and will act as a stopper to allow pressure to build up from the CO 2 canister in the (2) Payload Bay. The pressure that is built up will then allow the nose cone to shear the shear pins and allow the rover to deploy. More information about this deployment system is described in detail in the Payload section. The nose cone will be attached in this method in order to allow for a multi-purpose nose cone to be developed. This design will allow for the nose cone to act as a separation point and a storage area for the GPS. Screws will be an adequate method of retention as this section will not be experiencing separation forces that will put the section apart. The portion that will be secured using shear pins will experience some shear force, but Payload Bay The Payload Bay, item number 2, will be constructed from G12 Filament Wound Fiberglass with an outer diameter of 5 inches and a length of 18 inches and will house the payload, a rover. The payload bay will be connected to the nose cone via four (4) shear pins and will have a coupler attached to the aft portion of the bay that will house the CO 2 deployment system, see section for more detail on this system, and the mass management ballast system. This coupler will be attached to the Payload Bay with four (4) screws. The coupler assembly will be constructed in a comparable manner to the electronics bay. The coupler will have two removable bulkheads that will be 22 22

24 retained and tensioned with steel rods that will run through the length of the coupler and be secured on the ends with a nut on each rod Main Parachute Bay The Main Parachute Bay, item number 3, will be constructed of G12 Filament Wound Fiberglass with an outer diameter of 5 inches. The bay will be connected to the coupler that is attached to the Payload Bay. Through this section, a ½ inch tubular nylon shock cord will be attached to both the coupler and the fore portion of the electronics bay. This will anchor the rocket together when the ejection charge is deployed. It will also bear the brunt of the shock forces when the body separates. The 80" parachute that is stored in this section will be connected to the electronics bay using a ¼" quick link that will be connected to the forged eye bolt built into the electronics bay. The parachute will be made from rip stop nylon, have 4 shroud lines, and be covered using a Nomex Kevlar blast sheet in order to prevent the ejection charge from burning the parachute and its associated systems Electronics Bay The Electronics Bay, item number 4, will be constructed of G12 Filament Wound Fiberglass using a 5" tube coupler and the associated bulkheads. Much like the aforementioned coupler in the Payload bay, the coupler will be retained using two threaded steel rods that will be tensioned using a nut on each threaded rod. An outer ring of width 1 inch will be attached to the outside of the coupler in order to allow access to the arming switches. Forged eye bolts are attached to the bulkheads in order to have the parachute and the associated shock cord attached Drogue Parachute Bay The Drogue Parachute Bay, item number 5, will be built into the Lower Body Assembly which is made from G12 Filament Wound Fiberglass with an outer diameter of 5 inches. The drogue parachute will be attached to the aft section of the electronics bay on the forged eye bolt. The parachute itself will be attached to a ¼" quick link for easy installation and removal of the parachute. There will also be a 1/2" Tubular Kevlar shock cord that will attach the electronics bay to the motor mount where a forged eyebolt will be built into the motor mount. The forged eyebolt will be built into the top most centering ring and will be secured using both fasteners and epoxy resin. The drogue parachute will be a 24" rip stop nylon parachute with 4 shroud lines. For more information about the parachute Lower Body Assembly The lower body assembly, item number 6, is comprised of two sub systems. These systems are the Drogue Parachute Bay, as described above, and the motor mount assembly. The lower body tube is made of G12 Filament Wound Fiberglass with an outer diameter of 5 inches. The motor mount is comprised of 3 Fiberglass centering rings and 23 23

25 a G12 Filament Wound Fiberglass motor tube. The centering rings are placed to best support the rocket and develop a suitable space for the fin tabs to mount. Below is a model of the centering ring spacing and an example of how the fin tabs will be supported by the centering rings. The centering rings will be epoxied to the motor tube and a fillet will be made on the connecting face of the centering ring's inner diameter and the motor mount tube for greater strength and retention. The aft most centering ring will be placed 1 inch into the body in order to accommodate the motor retainer. Once the motor mount has been installed, the fins will be tacked in before a proper fillet will be made on the intersecting line between the fin tab and the motor mount tube. This, again, will assist in retaining the fins and will help prevent the fins from breaking free from the motor tube upon landing Materials Figure 7: Fin Assembly The materials used in the rocket vary from plastics to Kevlar to, in certain structural cases, steel. Each of the materials is chosen to best represent the intended function of the part and provide adequate performance. As a general rule of thumb, metal and chemically reactive materials (such as lithium and black powder) are used in limited quantities in order to ensure the safety of the team members and the safety of all involved in case of a failure. These materials cannot be entirely omitted from the vehicle as the roles they play are critical in portions of the design. For an overview of the materials used, reference the bill of materials provided. These materials represent the most pertinent items

26 Separated Bill of Materials, NASA SL , Project Kronos This includes assemblies and their components Item P/N Qty Material Nosecone Assembly 0100-ASM 1 Mixed $ $ Shear Pins A 4 Plastic $3.00 $12.00 TeleGPS A 1 Various $ $ Nosecone A 1 Fiberglass $ $ Payload Bay 0200-ASM 1 Mixed $ $ Payload Bay Body Tube A 1 Fiberglass $0.00 Rover 0201-ASM 1 $0.00 Large 36 T Nylon Gear 2.25" dia A 4 Nylon $15.83 $63.32 Small 16 T Nylon Gear 1/3" dia A 1 Nylon $7.51 $7.51 Continuous Rotation Micro Servo A 4 Plastic $7.50 $30.00 IXYS SLMD121H10L Panel A 6 Silicon $13.13 $78.78 Hitec HS-35HD Ultra Micro Servo A 1 Plastic $24.95 $ /8" Rack Gear A 2 Nylon $5.82 $11.64 Mounting Hardware A 1 Various $20.00 $20.00 Wheel Ball A 4 Plastic NA NA Rover Base A 1 Plastic NA NA Panel Carrier A 2 Plastic NA NA Deployment 0202-ASM $0.00 Peregine Raptor CO2 kit A 1 Various $ $ Payload Base A 1 Plastic NA NA 5/16" Rack Gear A 1 Nylon $8.88 $8.88 Electronics 0203-ASM $0.00 Turnigy 1000mAh 2S A 1 Lithium $8.98 $8.98 Syworks SE4150L-R A 1 Silicon $3.20 $3.20 Single Pull Key Switch A 1 Plastic $9.46 $9.46 STM LSM6DS3USTR A 1 Silicon $3.01 $3.01 ON AX5043 Transceiver A 1 Silicon $3.09 $3.09 Atmel ATmega32U A 1 Silicon $3.05 $3.05 Drogue Chute Bay 0300-ASM 1 Mixed $62.07 $62.07 Drogue Chute Body Tube A 1 Nylon $0.00 Drogue Assembly 0301-ASM 1 $ in Parachute A 1 Nylon $21.50 $ /4 in Quick Link A 1 Steel $0.93 $0.93 Tubular Kevlar Shock Cord A 1 Kevlar $19.00 $19.00 Kevlar Blast Sheet A 1 Kevlar $20.64 $20.64 Electronics Bay 0400-ASM 1 Mixed $ $ Coupler Tube A 1 Fiberglass $0.00 Outside Coupler Tube Ring A 1 Fiberglass $0.00 Board Assembly 0401-ASM 1 $0.00 Unit Cost Total Cost 25 25

27 Zip Ties A 1 Plastic $9.99 $9.99 Wooden Sled A 1 Wood $2.25 $2.25 Threaded Rods A 1 Steel $2.27 $2.27 9V Battery Connector A 1 Steel $1.49 $1.49 9V Battery A 1 Lithium $17.72 $17.72 Terminal Blocks A 4 Plastic $3.25 $13.00 Stratologger CF Altimeter A 1 Silicon $54.95 $ Guage Wire A 1 Copper $17.77 $17.77 Bulkhead Assembly 0402-ASM 2 $0.00 1/4 in Stand Offs A 2 Plastic $1.79 $3.58 Black Powder (8g) A 1 Black Powder $16.99 $16.99 Washers A 10 Steel $0.11 $1.10 Nuts A 10 steel $0.14 $1.40 E-Matches A 1 Various $33.60 $33.60 Single Pull Key Switch A 2 Plastic $9.46 $18.92 Blast Caps A 8 Various $3.15 $25.20 Main Parachute Bay 0500-ASM 1 Mixed $ $ Main Parachute Body Tube A 1 fiberglass $0.00 Main Assembly 0501-ASM 1 $ in Parachute A 1 Nylon $ $ /4 in Quick Link A 1 Steel $0.93 $0.93 Tubular Kevlar Shock Cord A 1 Kevlar $19.00 $19.00 Kevlar Blast Sheet A 1 Kevlar $20.64 $20.64 Lower Body Assembly 0600-ASM 1 Mixed $ $ Lower Body Tube A 1 fiberglass $0.00 Motor Mount 0601-ASM 1 $0.00 Centering Ring A 3 fiberglass $8.00 $ mm Motor Mount Tube A 1 fiberglass $ $ Motor Retainer A 1 Steel $50.00 $50.00 Welded Eye Bolt A 1 Steel $10.91 $ Motor Alternatives The motor for the team rocket was determined by analyzing data generated by OpenRocket. There were several motors that were also considered for the team's rocket. Motors considered but not chosen include the Aerotech K560W-P and the Cesaroni 2430-K661-BS-0. Both motors had similar performance parameters similar to the to the Aerotech K1000T-P. However, after running simulations of both motors, the alternative motors were found to not perform as well as the Aerotech K1000T-P. The Aerotech K560W-P was not chosen because when compared to the Aerotech K1000T-P, the apogee reached by the rocket was further away from the target apogee of one mile. The apogee of the Aerotech K560W-P was 5110 ft. This is 170ft. away from one mile. The apogee of 26 26

28 the Aerotech K1000T-P was 5355 ft. this is only 75 ft. away from one mile. The motor also required a 12 ft. launch rod in order to reach a stable velocity above 52 ft/s per NASA requirements. Similar to the Aerotech K560W-P, the Cesaroni 2430-K661-BS-0 was not chosen because it's apogee was further away from the one mile mark than the Aerotech K1000T-P. The apogee of the Cesaroni 2430-K661-BS-0 was This is 184 ft. away from one mile. The motor also required a 12 ft. launch rod in order to reach a stable velocity above 52 ft/s per NASA requirements. The ultimate deciding factor in not choosing this motor was the fact that the Cesaroni 2430-K661-BS-0 is not commercially available at this time. 3.3 Recovery Subsystem Recovery Design Alternatives CO2 Canisters This idea utilizes CO 2, or air stored in high pressure containment, that would be used to eject both the drogue and main parachutes at apogee and 700 feet respectively by separating the rocket body into multiple sections at predetermined coupling locations. The design principle of these canisters is the same as that of black powder, namely the increase in pressure within the enclosed volume of the parachute carrying sections. However, the use of an air or CO 2 canister will not result in residue accumulation within these sections of the rocket. This alternative design can also be more reliable since the canisters are of a known size and there is no risk of mistakes during the loading of black powder resulting in insufficient black powder to generate the required separation. The downfalls are that CO 2 has not been used by the club, CO 2 is more expensive, CO 2 will greatly increase the total weight of the recovery ebay, and there is no way to adjust the expelled volume size such as can be done by adding or subtracting black powder for unexpected flight conditions. Since the club does not have experience with these systems and the recovery section is paramount to a successful launch, the CO 2 ejection method will not be utilized for this year s competition Black Powder Black powder is the standard parachute deployment mechanism in high-powered rocketry; however, it represents a unique hazard during the design and generation of the rocket. First, the black powder must be bought, stored, and transported to launch sites. Any flame or spark hazards can result in extremely dangerous reactions that could harm individuals or damage equipment or facilities. To avoid these hazards sufficient safety precautions must be taken. It would be preferable to avoid these considerations if possible during fabrication, testing, and launches. In addition, black powder produces significant deposits within the rocket sections which must be cleaned after every use. It is also necessary to include Nomex fire resistant fabric on the parachutes to prevent 27 27

29 damage during deployment. These add to the cost and weight of the system, there is also an increased risk of incorrect packing resulting in damage to the parachutes during deployment or potential binding or tangling of the parachutes. However, the team has used black powder based deployment systems for all previous launches and thus has significant experience in their design and safe operation. They also have the added benefit of simple volume adjustments to increase or decrease the produced pressures in case of anomalous flight conditions. For these reasons, black powder deployment was chosen for this year s preliminary design StratoLogger Collects Flight Altimeters For this year s competition, it was decided to utilize a StratoLogger CF Altimeters based systems. The team had previously used these altimeters for multiple launches and have found them to be reliable devices. Utilizing the same altimeter will reduce the cost of the electronics bay section for the team and the familiarity with the system will ensure the team properly operates the system. While similar altimeter systems are available, due to the cost of attaining another system and the limited benefit received from doing so other systems were not seriously considered as replacements for this device Drogue Parachute Sizes Several possible drogue parachute sizes were considered for the preliminary design. The primary constraint for the drogue parachute selection was the drift distance of the rocket which must not exceed 2500 feet. A secondary constraint that must be considered is the kinetic energy present during the deployment of the main parachute. If the kinetic energy of a falling section is too high, damage may occur to the nylon shock cords connecting the parachute to the bulkhead. OpenRocket simulations were performed with a compete model of the rocket to evaluate the total drift distance and kinetic energy present at main deployment. Drogue parachutes with diameters from 24 to 32 inches were considered in the simulations. It was determined that all drogue selections would not exceed the 2500 feet maximum drift at a worse case of a 20 ft/s crosswind. However, it was discovered that the 24 inch drogue would produce a potentially dangerous level of kinetic energy at main deployment. While a 28 inch drogue would be sufficient to reduce the kinetic energy to safe levels, no commercially available drogue parachutes of that size and of the required design were found. As a result, a 32 inch drogue was selected for the preliminary design Preliminary Recovery Design In the figure below the team is utilizing a very common design for the recovery electronics bay. The electronics bay will use 4 bulkheads, two on each side. The bulkhead on the outside of the electronics bays diameter will be equal to the outer diameter of the electronics bay. The inner bulkhead diameter will be equal to the inner diameter of the electronics bay body tube. The bulkheads will then be clamped and epoxied together with a threaded rod. The altimeters will be mounted on ¼ inch stand-offs, and the 9V battery's will be zip tied into place

30 Figure 8: Recovery Sled The electronic hardware will be using two StratoLogger CF Altimeters, two four to four terminal blocks, two single pole key switches, two 9 V batteries, and four e-matches. The StratoLogger CF Altimeters, 9 V batteries, single pole key switches, and 4 to 4 terminal blocks, and e-matches will be connected via the schematic below

31 Figure 9: Recovery System Circuit Diagram Figure 10: StratologgerCF 30 30

32 3.3.3 Preliminary Analysis The club has decided to use a 32 inch drogue parachute and an 80 inch main parachute. The parachutes will be protected by a Nomex sheet. Both the drogue and main parachute will be from the Sky Angle CERT-3 Series. The main has a coefficient of drag of 1.26 and the drogue has a coefficient of drag of The parachutes will be tethered by a forged eyebolt fixed to the bulkheads with nylon lock nuts on either side of the recovery bay. The club uses tubular nylon shock cord to tether the sections of the rocket after separation. A shock cord length of 25 feet (~3.5x the length of the rocket) will be used for each parachute deployment system. For considerations of the kinetic energy for each system, the masses of each separated section are important to note. Table 3: Mass of vehicle sections Total Mass (Kg / #) Top Section (Kg / #) Mid Sec (Kg / #) End Sec. (Kg / #) Main Parachute (Kg / #) The kinetic energies (KE) for each separated section when using a 32inch drogue parachute are calculated from an Excel sheet with focus on vehicle events as simulated by OpenRocket results are listed below for the KE of the top and bottom sections under drogue at 700 feet Above Ground Level and when under main and drogue upon contact with the ground. Table 4: Kinetic Energy under 32 inch drogue Section Under Drogue Kinetic Energy (ft-lbf) Velocity (ft/sec) Top + Mid Bottom Section under Drogue and Main Top Mid Bottom Two more sets of data, for a 28 inch drogue parachute and a 24 inch drogue parachute respectively, are presented in table 6 and 7 for comparison to the selected 32 inch drogue

33 Table 5: Kinetic Energy under 28 inch drogue Section Under Drogue Kinetic Energy (ft-lbf) Velocity (ft/sec) Top + Mid Bottom Section under Drogue and Main Top Mid Bottom Table 6: Kinetic Energy under 24 inch drogue Section Under Drogue Kinetic Energy (ft-lbf) Velocity (ft/sec) Top + Mid Bottom Section under Drogue and Main Top Mid Bottom The above values describe a system that fulfills the USLI requirement of maintaining a kinetic energy of each independent section of less than 75 ft-lbf, qualifying the simulated parachute sizes of an 80in main parachute and a 32 inch drogue parachute. When simulated a smaller drogue parachute to limit the drift distances experienced, the calculated values for the KE of the Top and Mid sections, which are not yet separated from each other, reaches values approaching 1000 ft-lbf. The safety concerns of a shock cord failure lead to the decision to use a larger drogue parachute, maintaining control over a less energetic system up until the main parachute deploys. The limits of ½ inch tubular nylon strength have been modeled by some supplies as 900 ft-lbf. As the strength of tubular nylon is linearly dependent on the mass of nylon, increasing the width of the tubular nylon shock cord to an inch is a possibility to mitigate the possibility of shock cord failure. This year the team will use a dual parachute deployment using an 80inch main parachute and a 32 inch drogue parachute. The drogue parachute will deploy at apogee with a secondary charge to ignite after a 1 second delay. The main parachute will deploy at 700 feet with a secondary black powder charge to ignite at 600 feet

34 3.3.4 Redundancy The recovery system utilizes two StratologgerCf Altimeters. Each altimeter is powered by its own 9V battery. Each flight will require a new pair of batteries to ensure they have a charge, rather than risking a launch with faulty, used, low amperage, batteries. As well as replacing the batteries, the wires from the altimeters to the charges are also replaced after each launch to ensure continuity. The altimeters are activated via a pair of exterior screws which are then tightened on the pad in order to close the circuit. The primary altimeter is programmed to activate when the rocket reaches apogee. At this point the altimeter sends a charge to ignite a charge cap filled with two grams of black powder utilizing e-matches taped (with masking tape) to the top of the charge cap. The force will then separate the lower end of the rocket, releasing the drogue parachute. The backup altimeter will set off a similar charge shortly after, in case the primary altimeter or blast cap malfunctioned. The main parachute is released at the programmed altitude of 700 feet AGL during the descent by the primary altimeter, again, sending a charge to ignite the blast cap at the top of the electronics bay. The backup, again, follows this detonation at 700 feet AGL with a one second delay in case issues arise from the primary ignition system. The nose cone includes a TeleGPS tracking unit that transmits live flight data to the ground control via HAM band radio. The data is saved onto the TeleGPS which is then downloaded to locate the rocket s impact location as well to validate data collected by alternate sensors onboard. The rocket will be located using an AltusMetrum TeleGPS locator. This device tracks the rocket using GPS satellites and stores this data onboard. Live data is also transmitted by the device using HAM band radio at MHz. This signal is received by a ground station antenna and the TeleGPS software displays the position of the rocket superimposed on a satellite map of the area, along with the current velocity and altitude of rocket. This system will be crucial in tracking the rocket through the duration of the flight

35 3.4 Mission Performance Predictions Preliminary Simulation After running several simulations in OpenRocket to verify the simulation results, the rocket is projected to reach an apogee of 5355 ft. using an Aerotech K1000T-P motor. The motor weighs 90.4 oz. y = vy0t - 1/2gt vy0 = 0 m/s t = 18.1 s g = m/s^2 y = m = ft %Error = (Theoretical-Actual)/Theoretical = (5355ft ft)/ ft = 1.63% Figure 11: Flight Profile 34 34

36 For the thrust curve, the motor has an average thrust of 1012 N, a maximum thrust of 1140 N, and a total impulse of 2497 N*s Stability Margin Figure 12: K1000T-P Thrust Curve The rocket has a stability margin of 2.53 caliber with the selected motor and a stability margin of 3.76 caliber without the motor. The graph of stability versus time is shown below. The stability margin is above the required 2 calibers as set forth in the NASA SL Handbook and Guidelines. The rocket design also meets the stability goal of a stability caliber between 2.5 and 3 with a motor that was set forth by the club. The stability margin is low enough that unexpected weather cocking should not be an interference factor. If the need to change the stability margin arises in the future, the addition or subtraction of weight in the ballast may be used to adjust the stability as necessary

37 Figure 13: Stability Margin vs Time Figure 14: Center of Gravity and Pressure positions The above image is the OpenRocket simulation design which lists the locations of the center of gravity and center of pressure. The current center of pressure is located at inches and is shown in red on the above image. The center of gravity is located at inches and is shown in blue on the above drawing

38 3.4.3 Kinetic Energy The kinetic energy requirements for the competition is to maintain a ground contact kinetic energy value less than 75 ft-lbf for each independent section of the vehicle. One further value the team has investigated is the kinetic energy of each section during deployment of parachutes. A safe kinetic energy in each independent section should be observed so that the energy of the deployment will not be sufficient to damage the parachute s shock cords. When deploying parachutes, the shroud lines and shock cord must maintain integrity. The team has decided to utilize ½ inch shock cord-this should enough strength to keep up with the demands of the rocket. Table 7: Kinetic Energy under Main Parachute at Ground Contact Section Under Main Kinetic Energy (ft-lbf) Velocity (ft/sec) Top Mid Bottom Drift Calculation Drift calculations were all completed using simulation data obtained from OpenRocket simulations executed with the proposed rocket design and motor as the model. The simulation was performed with a vertical rocket launch (vertical launch rail) and with a 90-degree crosswind of varying velocity

39 Figure 15: Drift distance with 0 mph wind With a simulated crosswind velocity of zero mph and zero standard deviation, the total lateral drift for the current model and expected motor is less than eight feet from the launch rod position when modeled as launching vertical at a ninety-degree angle to the ground

40 Figure 16: Drift distance with 5 mph wind With a simulated crosswind velocity of 5 mph and zero standard deviation, the total lateral drift for the current model and expected motor is less than 460 feet from the launch rod position when modeled as launching vertical at a ninety-degree angle to the ground

41 Figure 17: Drift distance with 10 mph wind With a simulated crosswind velocity of 10 mph and zero standard deviation, the total lateral drift for the current model and expected motor is less than 1000 feet from the launch rod position when modeled as launching vertical at a ninety-degree angle to the ground. This drift distance is approaching 1000 feet of drift distance but still maintains a value of less than half the maximum 2500 feet drift distance

42 Figure 18: Drift distance with 15 mph wind With a simulated crosswind velocity of 15 mph and zero standard deviation, the total lateral drift for the current model and expected motor is approximately 1500 feet from the launch rod position when modeled as launching vertical at a ninety-degree angle to the ground. This simulated crosswind is 75% of the maximum wind allowable for safe launch conditions and maintains a drift distance of a 60% of the maximum drift distance allowable for calculations. Such a value suggests that further evaluation of the drogue parachute selection may be required, but the team has decided, as mentioned above, to go for the safe design of a drogue parachute with a smaller kinetic energy than with one that disallows higher, yet still manageable, drift distances

43 Figure 19: Drift distance with 20 mph wind With a simulated crosswind velocity of 20 mph and zero standard deviation, the total lateral drift for the current model and expected motor is less than 2232 feet from the launch rod position when modeled as launching vertical at a ninety-degree angle to the ground. The maximum allowable drift in the most severe allowable conditions is a modeled 2500 feet at 20 mph when modeled as the above does. The most severe conditions simulate a drift of 89.3% the maximum allowable drift, which still gives the team a margin of error for launch day conditions and launch angle while maintaining an acceptable drift distance. Table 8: Drift Distances Trial wind velocity Nominal Drift Distance 0 mph crosswind 8 feet 5 mph crosswind 460 feet 10 mph crosswind 1000 feet 15 mph crosswind 1500 feet 20 mph crosswind 2232feet 42 42

44 4 Safety 4.1 FMEA Procedure Overview The Safety Officers Duties The safety officer's duties ensure the following. Under their care all risks will be minimized. All personnel have read and acknowledged that they have a clear understanding of all rules and regulations set forth by the latest version of the safety manual. All personnel involved have undergone proper training on the equipment being used or processes being performed. Personal Protective Equipment (PPE) is used as indicated by the safety lab manual and MSDS. All procedures were correctly followed during construction of the rocket, testing, prelaunch preparations, and the launch. All components were thoroughly inspected for damage or fatigue prior to any test or launch. To ensure safe and professional behavior during all rocketry events. Risk Assessment The safety officer's duty is to post, discuss and prevent all safety concerns during all parts of construction and launch of the rocket. To ensure this the safety officer will create lists of safety concerns for every event and ensure that all members understand the concerns. A safety manual will be created by the safety officer that will include; all the pertinent Safety Data Sheets (SDS), the risks involved with each tool, material, location, weather, personnel, unexpected failures, and time constraints and their mitigation plans, a copy will be stored in the work area with the safety equipment and PPE, a copy will be kept on the team's website, a copy will be available on the team's google drive, and a copy will be brought to every meeting outside of the normal location. The safety officer will then take measures to ensure all members present avoid and mitigate all safety concerns. All safety concerns will be ranked and, regardless of classification, a mitigation plan will be developed. Then a decision to move forward or not with the mitigation plan will be made. The risk assessment plan will include, but will not be limited to; materials, weather, environments, personnel loss, failures, anomalies and time constraints. These events will be ranked by their expected Frequency and Severity, with proper training, procedure, and PPE in mind

45 Frequency The frequency ranks from Extremely Improbable (1), meaning that it is highly unlikely that this specific event would occur, to Frequent (5), meaning that the specific event would likely occur. The numbers correlate to how high of a risk these events pose. The higher the number, the higher the risk. The level of severity will be multiplied by the level of likelihood to get the total risk. The complete scale, along with the severity rankings, are in a table below. Severity The Severity Scale ranks from Catastrophic/ Total Loss of Vehicle/ Fatality (5), meaning that there would be a total loss of vehicle or massive damage to the environment or personnel, to No Injury/ No Damage (1) meaning that no harm has come to any aspect of rocket including people and the environment. The numbers correlate to how high of damage these events pose. The higher the number, the higher the damage. The level of severity will be multiplied by the level of likelihood to get the total risk. The complete scale, along with the frequency, is as follows

46 4.2 Personnel Hazard Analysis Personnel safety hazards will be accounted for in the chart below, ranked based on their Severity and Frequency, and a mitigation plan will follow for each event in case of occurrence. Severity - (S) Fatality (F) (5) Medical Attention (MA) (4) Serious Incident (SI) RE: EI: Frequency - (F) Description: F: 5 RP:4 O: This event will be likely to occur in normal operation. F : Frequent (F) (5) 80%-100% chance of occurring Reasonably Probable (RP) (4) (3) Occasional (O) (3) First Aid (FA) (2) Remote (RE) (2) No Injury (NI) (1) Extremely Improbable (EI) (1) This event will not be unusual to occur in normal operation. 60%- 80% chance of occurring This event is unlikely to occur in normal operation. 40%- 60% chance of occurring This event will only result from an unforeseen event. 20%- 40% chance of occurring This event is one so unlikely that it is not anticipated to occur during the entire operation. 0%- 20% chance of occurring MA: SI: High FA: Medium NI: Low Risk Magnitude: Risk Magnitude (RM) = Severity * Likelihood Activity/ Event Hazard Hazard Effect Uncontrolled Risk Existing Controls Controlled Risk Sanding Fumed Silica. Sanding body tubes. Drilling. Soldering. Cutting Fiberglass. Using Epoxy. Testing Electronics Bay. Testing payload. Using Blackpowder. Using a Hacksaw. Battery leakage. Exposed to fumes. Fire starts. Small, airborne, hazardous particles, carcenogenic material. Fiberglass splinter. Material being drilled may enter eyes. High temperatures may cause burns. Cutting fiberglass may cause fiberglass splinters, and release small particles that may be hazardous if inhaled. The epoxy being used dries at high temperatures. Dried epoxy does not come out of clothing and does not come off of skin easily. Electric shock may occur while testing. Fire may start due to improper testing. Electronic components may become hot or start a fire. Black powder is extremely flammable, and will ignite easily. Hacksaws have sharp, jagged edges. Battery acid can cause severe burns when in contact with skin or eyes. Fumes from motor, epoxy, fiberglass, or other objects. could make a team member sick, or affected from gases. The exhaust of the motor is at a very high temperature and may cause dry vegetation to ignite. Electronics cause a fire while launch. Fire starts during construction of rocket. S L Risk Mag Damage to lungs High Fiberglass entering skin causing extreme pain. Damage to eye(s), possible blindness. Burns cause pain, and possible time off. Painful splinters, lung damage from small particles. Possible burns, epoxy may dry on skin. Urgency S L Wear face mask to prevent fume silica from entering lungs. Risk Mag Urgency Low Medium Wear thick gloves Low Medium Wear safety glasses Low Medium Wear gloves and use extra caution Low Medium Wear face mask and gloves when cutting fiberglass Low Medium Team members will wear gloves Low Burns or fires Medium Wear gloves and use extra caution Low Burns or fires Medium Wear gloves and use extra caution Low Burns, shrapnel entering skin. Could puncture or lacerate skin Medium Medium Burns, severe pain Low Gases could make a team member sick, or lung damage. Flames could cause burns. Fumes from burnt fiberglass should not be inhaled. A fire during construction could grow and burn other parts of the construction area Medium Medium Use extreme caution, and keep black powder away from ignition sources. Wear gloves and use clamp to hold piece being cut. Wear hand and eye protection when handling batteries. Team members who are exposed to fumes will be required to wear masks, as well as required take frequent breaks if exposed to the fumes for an extended period of time. Room needs to be properly ventilated. The safety team will clear the launch sight of any dry vegetation and ensure the launch rail has an exhaust shield. If fiberglass is burned the recovery team will wear facemasks. There will be a fire Low Low Low Low Low 45 45

47 Premature rocket launch. Rocket does not launch when activated. Ignition of motor. #1 Ignition of motor. #2 While putting igniter in motor, rocket motor ignites. Burns Medium extinguisher on every build sight. Keep all electrical devices away from igniter as it is going into the motor. Rocket sputters and launches when recovery is in progress. Burns Medium Follow the NAR code. Wait sixty seconds before approaching rocket. Catastrophe At Take Off (CATO). Launch lug failure. Motor burn. #1 Motor Retainer failed. Motor burn. #2 Impact with Flying object. Motor burn. #3 Electronics Bay misfires. Motor Burnout. #1 Motor Burnout. #2 Coasting. #1 Coasting. #3 Coasting. #4 Apogee. #1 Apogee. #2 Apogee. #3 Apogee. #4 Apogee. #5 Descent on drogue parachute. #1 Descent on drogue parachute. #2 Rocket descends to 700 ft. #1 Rocket descends to Rocket separates due to drag from fins. Motor casing falls out Motor casing get too hot and ignites fiberglass. Fins rip off. Payload or Electronics Bay electronics overheat and start a fire onboard. Electronics Bay does not detect Apogee. E Matches fail to ignite black powder. Too much black powder is used. Shock cords are not properly installed or break. Drogue chute does not open. Fire in Eletronics Bay or payload electronics. Rocket hits flying object. Electronics Bay does not deploy main parachute. Main parachute doesn't open when deployed. Shrapnel flying into spectators and team members. Fires from motor failure. Remaining parts of rocket falling out of control. Launch lugs fail to keep rocket straight. If too extreme, rocket may become pointed at spectators. Motor may explode, or make rocket take a unsafe trajectory. Flying object changes rocket's trajectory. Rocket will separate while motor is burning, which may cause rocket to break apart and fall uncontrollably. Rocket will separate while moving at high speeds and may cause the rocket to break apart and parts to fall uncontrollably. Motor will fall at high speeds, and if not spotted could severely hurt a spectator or team member. Burnt fiberglass is hazardous to breathe in. Fires may cause structural damage. Burnt parts of rocket may fall off on descent. Parts of the fins that have fell off fall to at high speeds. Rocket may lose control after fins are sheared off. Burnt fiberglass is hazardous to breathe in. Fires may cause structural and electrical damage that may be critical to the project. The rocket does not break apart and will impact the ground and very unsafe speeds. Rocket does not separate and impacts ground at high speeds. Fiberglass body tube breaks apart and pieces fall uncontrollably. Large and heavy parts of rocket fall uncontrollably. Rocket falls at high speeds which may interrupt deployment of main parachute. Burnt fiberglass is unsafe to breathe. May cause structural damage and separation resulting in parts falling uncontrollably. Damage to main parachute or deployment system. Drogue parachute may become tangled causing the rocket to fall at a high speed. Flying object my become unable to fly and high ground at high speeds. Rocket impacts the ground at a high speed. Rocket impacts the ground at a high speed Medium Low Low Inspect motor for problems when purchased. Buy from trusted motor providers. Ensure launch lugs are straight and will not fall out easily. Propulsion Team reviews motor casing before use Low Low Low Low Low Low Check skies before launching Low Medium Test Electronics Bay, purchase altimeters from trustworthy manufacturers Low High Use shear pins Low Medium Low Low High High High Medium Medium Medium Test the motor mount so motor casing does not fall out. Inspect motor for problems when purchased. Buy trusted motor providers. Check motor casing for problems. Calculate minimum strength needed for the fins to stay attached. Test Electronics Bay and Payload Bays. Install a backup altimeter. Test Electronics Bay. Test Matches for continuity before installing. Test Electronics Bay. Test, measure, and use the appropriate amount of blackpowder. (2g) The recovery team will double check shock cord length and strength, and double check that nylon and shock cords are properly attached. Experienced parachute folder will practice and fold parachute on launch date Low Low Low Low Low Low Low Low Low Medium Test Electronics Low Low Check skies before launching Low Medium Test Electronics Low Medium Have an experienced parachute folder practice and fold the Low 46 46

48 700 ft. #2 parachute on launch day. Rocket descends to 700 ft. #3 Rocket descends to 700 ft. #4 Rocket descends to 700 ft. #5 Rocket descends on main parachute. Rocket lands. #1 Rocket lands. #2 Recovery. #1 Recovery. #2 Writing parties. Launch events. #1 Launch events. #2 Launch events. #3 Launch Event. #4 Launch Event. #5 Education. #1 Education. #2 Education. #3 Education. #4 Allergic reaction. Burns. Cuts. Electric shock. Insufficient amount of black powder is used and main parachute does not eject. Burnt fiberglass is unsafe to breath. May cause structural damage and separation resulting in parts falling uncontrollably. Damage to main parachute or deployment system. Too much black powder is used. Electronic fire. Rocket lands on spectator. Rocket lands on power line. Electronics Bay misfires when a recovery attempt is made. Motor casings will have high temperatures. High temperatures in computer labs. High temperatures at venue. Cold temperatures. Rain. Fog. Ice / snow. Children start a fight, not pay attention, or purposefully try to hurt someone. Education subscale goes ballistic. Education CATO`s Child has an allergic reaction. A team member may have an allergy to a building material, or a food allergy at the event. Ejection charges, motor burn, and electronics can get hot. Knives, drills and other sharp objects are used. When wires are live, the risk of electrical shock exists. Rocket impacts the ground at a high speed. Large and heavy parts of rocket fall uncountably. Breaks apart the fiberglass and shrapnel falls uncontrollably. Burnt fiberglass is unsafe to breathe. May cause structural damage and separation resulting in parts falling uncontrollably. Rocket may cause concussion, broken bones, or cuts. If an attempt to recover is not done by a professional then electric shock is possible. Misfire may cause burns or cuts Medium Medium Medium Test Electronics Bay, calculate appropriate amount of blackpowder. Recovery team will double check the connections and use a cord that has a high tensile strength. Use correct amount of blackpowder. (2g) Low Low Low Medium Test Electronics Low High Medium Medium Burns Low Temperatures may cause dehydration and possible fainting Low Control flight path so rocket is not above spectators. Have air horn at launch site to alert spectators. The safety team will ensure that professionals recover the rocket in the event that the rocket lands on a powerline. Test Electronics. Before recovery attempt is made listen to altimeters to ensure that both altimeters have fired. Control flight path so rocket is not above spectators. Have air horn at launch site to alert spectators. Members are encouraged to take breaks and stay hydrated and fed. Safety team will inform team Team members may become dehydrated and become sick Medium members of water and snacks before launch. Team members may experience frostbite. Team members may become drenched in water Launch fails due to poor viability Team members may slip and fall or develop hypothermia. Children could hurt each other or themselves. Subscale rocket will impact ground at very high speeds. Rocket is uncontrolled and very dangerous. If an allergic reaction occurs child may have to be sent to the hospital. Member may have to be seek medical attention. If a team member touches a hot object they may have severe burns. Minor cuts could require special care. Deep cuts require going to the hospital. If shock is severe, team member may have to seek medical attention Medium Medium Medium Medium Medium Medium Medium Safety team will inform team members to dress for the weather. Safety team will inform team members to dress for the weather and to bring an umbrella. Safety will wait until the backup day to launch Safety team will inform team members to dress for the weather. Constant vigilance and keep children under control at all times. Test fly subscale before launching at education events. Test fly subscale before launching at education events Low Low Low Low Low Low Low Low Low Low Low Low Low High Do not let children touch rockets Low Medium Medium Medium Medium The safety team will account for any allergies and alert the person(s) when that material is being used or an allergen is near. The safety and recovery teams will inform the team where and when they can touch the rocket. Wear gloves when cutting, be trained and use caution. The Safety Team will be taught how to handle injuries. Make sure no live wires are touched after batteries are connected to systems. Check all wires' protective coatings for cracks. The highest voltage used on the rocket will be 12 VDC. Caution will be taken when using devices wired into 120 VAC electrical outlets Low Low Low Low 47 47

49 Car accident. During travel a car accident occurs. If serious, members may have to go to hospital Medium Only properly grounded wires will be used. Drive carefully and follow all traffic laws Low 4.3 Failure Modes and Effects Analysis Failure of components on the rocket and the effects on the whole rocket's safety hazards will be accounted for in the chart below. These will be ranked based on their severity and frequency, and a mitigation plan will follow for each event in case of occurrence. Severity - (S) Likelihood - (L) Description: F: 5 RP:4 O: 3 RE: 2 EI: 1 Risk Magnitude: Total loss of vehicle (TL) (5) Severe damage to vehicle (DV) (4) Frequent (F) (5) Reasonably Probable (RP) (4) This event will be likely to occur in normal operation. 80%-100% chance of occurring This event will not be unusual to occur in normal operation. 60%- 80% chance of occurring TL : DV: Risk Magnitude (RM) = Severity * Likelihood Repair required (RR) (3) Occasional (O) (3) This event is unlikely to occur in normal operation. 40%- 60% chance of occurring RR: High Superficial Damage (SD) (2) Remote (RE) (2) This event will only result from an unforeseen event. 20%- 40% chance of occurring SD: Medium No Damage (ND) (1) Extremely Improbable (EI) (1) This event is one so unlikely that it is not anticipated to occur during the entire operation. 0%- 20% chance of occurring ND: Low Activity/Event Hazard Hazard Effect Uncontrolled Risk Existing Controls Controlled Risk Ignition of Motor. #1 Ignition of Motor. #2 Ignition of Motor. #3 Motor failure causes CATO. Vehicle body breaks during flight. Launch buttons do not fit rail, or disconnect Rocket is destroyed, unable to complete mission criteria. Rocket is destroyed, unable to complete mission criteria. Rocket comes loose from rail, unable to launch. S L Risk Mag Urgency S L Risk Mag Medium Medium Medium Use and implementation of a motor from a known production line with a good and clear history of few to no issues. Structural considerations in development of the vehicle body make sure there is no possibility of breaking under normal flight circumstances. Launch buttons fit a 1010 rail, are attached firmly to Urgency Low Low Low 48 48

50 Ignition of Motor. #4 Motor Burn. #1 Motor Burn. #2 Motor Burnout. Coasting. #1 Coasting. #2 Coasting. #3 Apogee. #1 Apogee. #2 Apogee. #3 Apogee. #4 from rocket. Centering rings are unable to withstand motor force during launch. Motor casing fails. Centering rings are unable to withstand motor force during launch. Motor casing falls out. Rocket impacts object during flight. Drag from fins separates rocket after burnout but before apogee. Fins shere off. Altimeter does not fire at apogee. E-match does not ignite. E-match comes out of the charge cap, unable to ignite black powder. Batteries do not provide enough electricity to altimeters or e- matches. Motor and motor casing will fly through rocket and destroy insides of rocket. Motor CATO mission failure. large parts of rocket is falling uncontrollably, launch is a failure. Motor casing fall, mission is unsafe and deemed a failure. Impact alters flight profile. Launch will be a disaster, severe damage to rocket. Mission fails, rocket is recoverable. Drogue parachute fails to deploy, rocket falls ballistic. Parachutes do not deploy, rocket goes ballistic. Parachutes do not deploy, rocket goes ballistic. E-matches do not ignite, parachutes do not deploy leading to damage to the rocket Medium Medium Medium Medium Low rocket body and have been tested. Body team uses stress analysis to confirm that motor mount is secure. Safety team will inspect the motor casing for damage. Stress analysis. Ues correct amount of shock cord. Test motor retainer for strength. Sky will be confirmed clear before launching rocket Low Low Low Low Low Low Use shear pins Low Low High Medium High High The body team will calculate how strong the fins have to be. Secondary altimeter backing up drogue deployment. E-matches have been tested and flown on previous flights. Redundant altimeters provide backup. E-matches are taped into position with masking tape. E-match leads are trimmed to prevent snagging. Redundant altimeters provide backup. New batteries are used for every flight and are tested beforehand Low Low Low Low Low Apogee. #5 Drogue Rocket enters a High Experience recovery Low 49 49

51 Apogee. #6 Apogee. #7 Apogee. #8 Apogee. #9 Apogee. #10 Apogee. #11 Decent on drogue. #1 Decent on drogue. #2 Rocket Descends to 700 ft. on drogue. #1 parachute does not fully deploy. Shear pins do not break during recovery deployment. Excess amount of black powder is used. Insufficient amount of black powder is used. Shock cords are not properly installed or break. Main parachute deployed. Too much masking tape over black powder charges. Rocket hits flying object on decent. Electrical fire. Altimeter does not fire at main deploy. ballistic trajectory. Rocket is destroyed on impact or when the main parachute is deployed. Parachutes do not deploy, damaging rocket Rocket does not separate. Rocket goes Ballistic. Rocket does not separate. Rocket goes Ballistic. Parts of rocket fall uncontrolled without a parachute. Mission failed Medium High High High team members will pack parachutes. Altimeters and other components set up as described. Shear pins have been tested and flown on previous flights. Recovery will calculate and measure the correct amount of black powder. (2g) Recovery will calculate and measure the correct amount of black powder. (2g) Recovery team will double check connection, and inspect shock cords for damage before launch. Calculate the tensile strength and select a shock cord that is appropriate Low Low Low Low Greatly increases drift distance Medium Test electronics Low Black powder charges do not Recovery team only separate rocket High uses one layer of Low Rocket goes masking tape. ballistic. Drogue chute rips, or becomes tangled. Falls at The safety team will high speed which check the skies Medium will affect the before launch for Low deployment of the flying objects. main parachute. Mission failed. Electronics Bay and/or Payload electronics have a major problem Medium Test electronics Low and ignite. Mission Failed. Main parachute Secondary altimeter does not deploy. backing up Main Rocket falls under Medium parachute drogue only, deployment. damaging rocket Low 50 50

52 Rocket descends to 700 ft. on drogue. #2 Rocket decends to 700 ft. on drouge. #3 Rocket descends to 700 ft. on drogue. #4 Rocket descends to 700 ft. on drogue. #5 Rocket descends to 700 ft. on drogue. #6 Rocket descends to 700 ft. on drogue. #7 Rocket descends to 700 ft. on drogue. #8 Rocket decent to 700 ft. on drogue. #9 Rocket lands. #1 Rocket lands. #2 Main parachute does not fully deploy. E-match does not ignite. E-match comes out of the charge cap, unable to ignite black powder. Shear pins do not break during recovery deployment. Excess amount of black powder is used. Insufficient amount of black powder is used. Shock cords are not properly installed or break. Too much masking tape over black powder charges. Rocket lands in water. Rocket lands on power lines. Rocket lands at high velocity, potentially damaging rocket. Parachutes do not deploy, damaging rocket. Parachutes do not deploy, damaging rocket. Parachutes do not deploy, damaging rocket. Black powder explosion shaders the fiberglass body tube. Mission failed. Rocket does not separate. Impacts the ground while only on ground parachute. Parts of rocket fall uncontrolled without a parachute. Mission failed. Black powder charges do not separate rocket. Rocket lands with high kinetic energy. Electronics in the rocket will be ruined. Mission is a failure. Rocket will be stuck on power lines for extended Medium Medium Medium Medium High High High High High Medium Experience recovery team members will pack parachutes. Altimeters and other components set up as described. E-matches have been tested and flown on previous flights. Redundant altimeters provide backup. E-matches are taped into position with masking tape. E-match leads are trimmed to prevent snagging. Redundant altimeters provide backup. Shear pins have been tested and flown on previous flights. Recovery will calculate and measure the correct amount of black powder. (2g) Recovery will calculate and measure the correct amount of black powder. (2g) Recovery team will double check connection, and inspect shock cords for damage before launch. Calculate the tensile strength and select a shock cord that is appropriate. Recovery team only uses one layer of masking tape. Inspect launch field. If there is a large pond, river, or puddle then No-Fly. Change flight path op avoid power lines. If power lines Low Low Low Low Low Low Low Low Low Low 51 51

53 Travel injuries. Travel damage. Miswiring of Arming Switches. Rocket loses contact with GPS. Conversion errors. Drogue parachute is burnt. Main parachute is burnt. Theft of equipment. Motor or motor casing is not available by launch site retailer. Large part of rocket falls without a parachute. Accidents when loading or unloading vehicles. Rocket damage during transportation or loading. Switches do not work on launch day. Rocket may not be found. Members use incorrect conversion or wrong units. Ejection charge burns drogue parachute. Ejection charge burns main chute. Parts or tools gets stolen. Motor or casing is unavailable. Section of rocket is not attached properly. period of time. Electronics will lose data due to loss of power. Injure team members. Rocket unable to fly or require repairs. Have to rewired in field, if not possible then the launch will be scrapped. Mission failed if rocket is not found. Numbers for important values may be off. If errors are not errors could cause a mission failure. When rocket descends to 700 feet the rocket may have too much kinetic energy to correctly deploy the main parachute. If the main parachute is burnt, then upon landing the rocket will have a high kinetic energy. If part of tool is mandatory mission fails. If not available mission is failed. Mission failed Low Low can not be avoided the No-Fly. Team members will wear proper safety gear when handling rocket or materials. Rocket will be stored securely during travel Low Low Medium Test electronics Low High High Medium Medium High High Medium Charge GPS and check to make sure that GPS is communicating with receiver. Team lead of each team will double check any conversions that take place. Parachute will be inspected by the recovery and safety teams. Nomex will be used in between the parachute and ejection charge. Parachute will be inspected by the recovery and safety teams. Nomex will be used in between the parachute and ejection charge. The team will be observant of all tools and parts. The work space will be locked at all times when unattended. The team will confirm with retailers on sight about stock. Recovery team double checks connections Low Low Low Low Low Low Low 52 52

54 Rocket lands with high kinetic energy. Rocket goes ballistic. Roll Control. #1 Roll Control. #2 Roll Control. #3 Roll Control. #4 Roll Control. #5 Roll Control. #6 Rocket falls to the ground on drogue parachute. Rocket does not separate at apogee and impacts ground at extremely high speeds. Lithium battery becomes punctured or overloaded during the flight or while loaded on the pad. Battery harness becomes severed and may short out battery or board. Erratic motion of fins due to payload not properly detecting different states of the rocket. Batteries, wiring, or board catch fire before or during flight. Upon crash, the lithium battery becomes punctured and ignites a potential fire. Parload fins rip off. Mission failed Medium Mission failed. Payload will be severely damaged and will not be able to fly. Payload may function erratically or not at all. Potential battery fire. Rocket will spin more or less than expected. Payload will not function. Damage to structure of rocket and object/area that rocket crashes into. Roll control experiment fails Medium Medium Low Low Low Medium Medium Recovery team checks Electronics Bay. Recovery team checks Electronics Bay. The 3d printed enclosure for the battery has no sharp points that could puncture the battery and fuses are used to prevent overloading the battery. Fuses will be used on the battery to prevent fire. All connections will be properly terminated with solder and heat shrink tubing. The payload is designed to move all fins in unison which only allows the payload to control to roll. Payload must also detect the high G's from launch meaning if the payload resets, the fins are most likely to return to 0 degrees. Since the airframe is fiberglass, the small potential fire will be contained by the airframe preventing damage to the environment or flight of rocket. The enclosure and rocket wall will partially protect the flammable components from being punctured. Payload team ensures roll control fins are strong enough Low Low Low Low Low Low Low Low 53 53

55 Roll Control. #7 Roll Control. #8 Roll Control. #9 Roll Control. #10 Environmental Impact on Rocket. #1 Environmental Impact on Rocket. #2 Environmental Impact on Rocket. #3 Environmental Impact on Rocket. #4 Environmental Impacts on Rocket. #5 Corrupted files. Data recovery High Detects motor burn out. MPU does not initiate. Payload does not start. Mission failed. Payload does not start. Wire Payload does not connections are work, mission not properly failed. connected. Launch pad is not put on level ground, or launch pad sinks into ground. Snow or freezing rain. Rain. High winds. High Temperature. Rocket will not fly straight up resulting in loss of altitude. Ice build up on ventilation holes. Altimeters do not detect apogee and rocket goes ballistic. Electronic damage resulting in possibility of rocket going ballistic. Wind cocking and excessive drift may cause loss of altitude and possible loss of rocket. Fiberglass may melt or become weaker. Electronics my overheat and malfunction. Electronic malfunction could lead to rocket going ballistic. Weakened fiberglass could lead to complete mission failure High Low Payload team properly formatted SD card. Setting proper threshold, running test. Fins will go to 90 degrees to alert payload team to perform power cycle Low Low Low High All clip connections Low Low Low Low Medium Low Use a plywood board to prevent sinkage. No-Fly if launch pad is not level. No-Fly if ice is forming on rocket. No-Fly if rain can not be overcame. GPS in nosecone. No-Fly if wind is greater than 20 MPH. Limit exposure to sunlight Low Low Low Low Low 54 54

56 4.4 Environmental Hazard Analysis Environmental safety hazards will be accounted for in the chart below, ranked based on their Severity and Frequency, and a mitigation plan will follow for each event in case of occurrence. Severity - (S) Likelihood - (L) Description: F: 5 RP:4 O: 3 RE: EI: Risk Magnitude: 2 1 Catastrophic (CA) (5) Frequent (F) (5) This event will be likely to occur in normal operation. CA: %-100% chance of occurring Risk Magnitude (RM) = Severity * Likelihood Critical (CR) (4) Reasonably Probable (RP) (4) Marginal (M) (3) Occasional (O) (3) Negligible (NE) (2) Remote (RE) (2) No Injury (NI) (1) Extremely Improbable (EI) (1) This event will not be unusual to occur in normal operation. 60%- 80% chance of occurring This event is unlikely to occur in normal operation. 40%- 60% chance of occurring This event will only result from an unforeseen event. 20%- 40% chance of occurring This event is one so unlikely that it is not anticipated to occur during the entire operation. 0%- 20% chance of occurring CR: M: High NE: Medium NI: Low Activity/Event Hazard Hazard Effect Uncontrolled Risk Existing Controls Controlled Risk Littering. Fire on locations. Rocket motor explosion. Battery leaking. Litter on the ground from the team or rocket debris. Electronics or engine fire. Fragments of the rocket could hit humans and litter launch area. Battery acid burns. Cause harm to humans, vegetation or wildlife. Burn vegetation, release potentially toxic fumes, may spread and destroy property, damage personnel, or damage the environments or buildings. Damage to property, environment, or people. Battery acid could harm the environment, personnel, and objects in the area. S L Risk Mag Medium High Medium Low Urgency S L Provide trash/recycling recepticles. Fasten items that may blow away. Clean up during and after events. Keep fire extinguisher on hand. Clear launch area of debris. Follow correct fire prevention protocols. Follow procedural checklists before launch. Proper protective clothing such as; long thick fire resistant pants, closed toed shoes, long sleeved shirts with tight cuffs. Inspect all used and new batteries. Never use a battery that appears damaged or fails any tests performed. Proper protective clothing Risk Mag Urgency Low Low Low Low 55 55

57 Excess drift. Recovery system failure. Tree landing. Controlled and uncontrolled explosions in rocket. Main parachute does not deploy. Rocket landing. Leaking of toxic chemicals. Can cause loss of rocket. Rocket can go ballistic. Landing in a tree and remain there. Too large of an explosion may cause the exterior fiberglass to break apart. Rocket impacts ground at high speeds. Shatters fiberglass body tube and fins. Rocket lands on wildlife or spectator. Leakage of battery fluid or excess fuel. Rocket damages an uncontrolled environment, personnel, wildlife, or property in the area. Rocket damages an uncontrolled environment, personnel, wildlife, or property in the area. Cause damage to tree, potential safety risk to remove. Small parts of fiberglass will litter the environment, potentially damaging wildlife. Fragments of the rocket could hit humans and litter launch area. May disturb vegetation or strike animal. Large mass of rocket can cause pain and possible death to small wildlife. Could possibly injure or kill spectators and personnel. Heavy damage to environment and objects in the area. Poison ground water, ponds or streams. Could injure personnel Medium High Low High Medium Medium Low such as; long thick fire resistant pants, closed toed shoes, long sleeved shirts with tight cuffs. Follow recovery bay system checklist. Use onboard GPS to track rocket. Follow recovery bay system checklist. Report any issues to Recovery Team Lead. Ensure backup altimeter is operational. Warn others in area of any concerns. Point launch rail in proper direction according to wind. Use correct amount of blackpowder, ground test electronics. Experienced parachute packer will practice and fold the parachute on launch date. Ensure that rocket's landing speed is within NAR code. Use an air horn to warn spectators of rocket landing. Electrical components provided extra separation from environment within body tube; rocket recovered quickly to minimize exposure time; launch site chosen away from bodies of water Low Low Low Low Low Low Low 56 56

58 Miscellaneous Payload Electrical Fire. Fire caused by Ballistic Crash. Batteries, wiring, or board catch fire before or during flight. Upon crash, the lithium battery becomes punctured and potentially ignites a fire. Payload will not function Low Damage to environment nearby Low Since the airframe is fiberglass, the small potential fire will be contained by the airframe preventing damage to the environment or flight of rocket. The enclosure and rocket wall will partially protect the flammable components from being punctured Low Low 57 57

59 5 Payload 5.1 Design Rationale Payload Mission Statement For this year s payload, the team will design and construct a rover that will be deployed from within the rocket body after landing. A remote trigger will activate the deployment procedure after which point the deployment and rover control routine will proceed autonomously. The rover will be able to move a minimum of 5 feet from the nearest part of the rocket, stop, and deploy a folded set of solar panels. In the spirit of exploration, the rover and deployment system will be designed as though the tests were being performed on a celestial body. This payload will also serve to illustrate the team s creativity and competencies in numerous engineering disciplines Payload Success Criteria The requirements to deem the mission a success include following all NASA ULSI requirements and team derived criteria: Table 9: Payload Success Criteria Criteria The team will design a custom rover that will deploy from the internal structure of the launch vehicle. At landing, the team will remotely activate a trigger to deploy the rover from the rocket. After deployment, the rover will autonomously move at least 5 ft. (in any direction) from the launch vehicle. Once the rover has reached its final destination, it will deploy a set of solar cell panels. The payload will be able to detect the transmitted signal from at least 2,500 feet away. The incoming radio signals will be filtered so that the response of the deployment system will only occur for the team s generated signal and at safe Action Plan The designed rover will utilize geared wheels and a rack system located within the diameter of the rocket to drive itself out regardless of orientation. A 433 MHz transceiver will be located on the rover with a suitable antenna to receive potential low powered signals. Filtering will be performed to determine appropriate activation time. Four servos will independently drive each geared wheel. In the rocket the geared wheels will be meshed with a rack to provide sufficient driving traction. GPS and IMU signals will be combined to evaluate the rover position to ensure proper driving direction and that the final location is at least 5 feet from the rocket. Upon achievement of the previous criteria, an ultra nano-servo will drive a set of two doublesided solar panels from either side of the rover. A high sensitivity transceiver will be selected, and the transmitting antenna will be supplied sufficient power while staying with regulated power levels. A filter will be designed and constructed on the payload to process incoming signals to locate the desired signal, likely a desired pattern, before the 58 58

60 operating conditions. The rover s locking mechanism (servos, wheels, rack) will be able to hold the rover position for a rocket acceleration of up to 15g. The rover s battery life will last at least 2 hours over the average expected operating conditions. The solar panels will completely slide out from their position on the rover. The nose cone will be ejected sufficiently far from the rocket body so as not to interfere with the deploying rover. Utilization of the GPS and the onboard Inertial Measurement System will be able to determine rover position to within 4 feet. The Rover will be able to transmit its GPS coordinates back to receivers near the ground station during and after flight. The gas discharge used to free the nose cone will not dislodge the rover from its position within the payload bay. The electrical coupling system from the rover to the CO 2 cylinder s e-match will successfully separate during rover deployment. deployment mechanism is activated. Further, data from the IMU and GPS will be analyzed to determine that the rocket has come to a stop after launching to prevent potential activations before reaching a safe deployment position. Servos with sufficiently high stall torques will be fixed to geared wheels and a rack that can resist the forces experienced at 15g. Careful consideration of the servos fastening system will be considered. The retaining rail design will be evaluated for strength within the potential operating conditions whilst retaining sufficient traction on the geared surface. A high capacity, low form factor battery will be selected for use on the payload. Preliminary current measurements will be made and evaluated to determine the average expected current draw and compared to the selected battery size. The rails the solar panel rides on will be rigidly constructed and any material residue will be removed to ensure the fit is snug during deployment. Tests will be performed with the servo to ensure proper deployment time is observed for the expected operating conditions. A properly sized CO 2 cylinder will be utilized for the given volume of the payload bay. Shear pins will be utilized to fasten the nose cone so that the energetics can properly eject the nose cone. A GPS receiver will be located on the rover alongside an IMU containing gyroscopes and accelerometers. A sensing protocol (e.g. Kalman Filter) will be used to improve the positional accuracy of the sensor suite. A high sensitivity receiver will be used at the ground station to pick up signals from the payload to account for its inherently low power design. The rover will be rigidly fixed by a combination of the servos, rack-pinion system, and the lower guide rail. The concentrated discharge will be offset from the rover s position so that most of the discharge passes harmlessly around the rover. The coupling system will be rigid, so as to resist the flight conditions, but will be susceptible to the force output by the rover in the desired orientation

61 5.1.3 Payload Design Alternatives Solenoid Deployed Solar Panels A solenoid controlled solar panel system was proposed in the Project Kronos Proposal. This design used a solenoid to hold back a spring loaded solar panel system. There were key disadvantages to this system that lead to the determination that another design would be needed. One of the reasons the solenoid idea was scrapped was because the power required to actuate the solenoid would have required additional electronics to control the power applied to it. Additionally, there were concerns that a spring-loaded system may not be robust enough to deploy the solar panels if the rover was positioned in a non-optimal way on the terrain. The solenoid position also limited the ground clearance of the rover due to the size of the solenoid itself. For these reasons the solenoid deployed solar panels were abandoned Flexible Solar Panels Figure 20: Solenoid Deployed Solar Panel Design Flexible solar panels were considered for use as the foldable panels on the rover. There were some concerns with this design, including availability of appropriately sized panels. Without developing custom made solar panels, the team was unable to locate a flexible 60 60

62 solar panel that fit the requirements for the rover. There were also concerns with creating a viable mechanism to deploy the panel. For these reasons this alternative design was abandoned early in the design process. Figure 21: Flexible Solar Panel Clam Shell Solar Panels A variation of the solenoid deployed solar panels was considered that utilized a solar panel rack that enabled the solar panels to spring out into either an X or diamond arrangement. Similar to the solenoid deployed system, the rack would have been spring loaded, with the individual solar panels spring loaded with torsion springs. This design was abandoned for the same reasons as the solenoid deployed system. Additionally, there were concerns that the solar panels could strike the ground when deploying. However, this design was more of a true foldable design, as compared to the solenoid deployed panels Sliding Solar Panels The following solar panel design is the design eventually selected for the Preliminary Design Review, which is comprised of solar panels on sliding trays that extend from the sides of the rocket. This design is explained in detail in section An ultra-nano servo drives mirrored rack and pinions to extend the solar panels. The solar panel trays are fixed to the rack of the rack and pinions, and will be extended in unison by the single gear. This system has a lower profile than other solar panel designs, which allows for higher ground clearance for the rover. This design also allows for the largest change in exposed solar panel surface area as compared to the alternatives designs, which exposed either one or two solar panels, with this design exposing three panels. The solar panels being deployed by the rack and pinion means there is a positive mechanical system deploying the panels, as compared to a passive spring system as in previously described alternatives. For these reasons the sliding solar panel design was selected

63 Gear Wheels The preliminary design continues to utilize the gear wheel design that was described in the Project Kronos Proposal. The wheels are nylon gears attached directly to continuous servos. The servos will drive a rack and pinion system in which the wheels act as the pinions and are meshed to two sets of gear racks attached to the base of the payload bay. This ensures there is sufficient traction while the rover is within the payload bay, enabling the rover to drive itself out of the payload bay on its own power. Though gears are not traditional tires, due to the size of the gears and the teeth on them, it was determined they will provide sufficient traction on the terrain at the launch field. These will be detailed in section Figure 22: Gear Wheel Non-Gear Wheels An alternative to pure gear wheels was to use standard wheels with gears attached next to the wheels to drive a similar rack and pinion design. Some concerns with this design were the amount of space taken up by a wheel and a gear, as this could limit the width of the rover body. Additionally, as wheel designs were discussed, most still used many of the features present in the proposal's gear wheel design, so it was decided to remain with a pure gear wheel design through the preliminary design review Payload Bay Orientation System The rover design of both the Project Kronos Proposal and Preliminary Design Review utilize a rover that functions independent of orientation. A payload bay that would reorient after landing was considered. Such a design would have used bearings or rollers to enable an internal structure inside the payload bay to rotate until level with the ground before allowing the rover to drive out. A locking mechanism would have held the internal structure in place during ascent to prevent any mass from shifting during flight. After landing, the locking mechanism would have unlatched, and the internal structure would have rotated either under power or passively into position. This design was abandoned due to complexity. Bearings or rollers for the internal structure would need to be low friction to allow the structure to rotate purely on 62 62

64 passive gravity. Additionally, this design required a latching mechanism to allow the structure to rotate which is not required by an orientation-independent rover, in turn decreasing the total weight of the payload system. For these reasons, the payload bay orientation system was not selected for the preliminary design review Orientation-Independent Rover The orientation-independent rover design of the proposal continues onto the preliminary design review. The rover is designed such that all operations it performs can be performed regardless of orientation, including deployment, solar panel deployment, and driving. The wheels of the rover extend out past the body of the rocket on both sides, allowing the wheels to have clearance and traction on either side of the rover. The rover is attached in the payload bay in a way that allows it to have traction against the rack and pinion regardless of orientation. As this is the preliminary design, it is further detailed in section Solenoid Payload Latch Multiple designs for opening a payload Latch to let the rover drive out where considered. The design in the Project Kronos Proposal used four solenoids to create a latch to allow the nose of the rocket to disconnect from the payload bay. The rover would then drive itself out and push the nose cone out of the rocket. Powering four solenoids would have required additional batteries and circuitry due to them operating at 24 volts and their large current requirements. Additionally, there were concerns the rover would not have enough torque to move the nose cone. The solenoid power was a significant enough issue to prompt the team to consider alternative designs. Figure 23: Solenoid Payload Latch 63 63

65 Rotary Payload Latch An alternative to the solenoid payload Latch design was conceptualized that used a mechanical system to move the latching pins. A disk with spiral tracks would have rotated on a servo, as the disk spun, pins would have followed along the tracks and have been pulled inwards, freeing the nose cone of the rocket. This was beneficial as a servo could be powered using the same voltage source as the electronics without the need for additional power supplies and circuitry. Ultimately, this design was abandoned for similar concerns as the solenoid system: there was a risk the mass of the nose cone would be too much for the rover to move, or the nose cone could bind as it was pushed out. Figure 24: Rotary Payload Latch 64 64

66 Black Powder Payload Hatch In order to overcome any risk of the nose cone being too massive for the rover to push, or having the nose cone bind up during extraction, using energetics were proposed. Initially, black powder was considered, but this was quickly dismissed as black powder would leave residue throughout the payload bay and likely damage the electronics of the payload. Additionally, residue would be deposited onto the solar panels, limiting their effectiveness. For these reasons black powder was abandoned quickly as an alternative CO2 Payload Hatch To overcome the issues brought on by using black powder as an energetic, CO 2 was proposed. Commercially available CO 2 systems intended for recovery deployment could be repurposed to deploy the nose cone of the payload bay, allowing the rocket to drive out of the bay without having to push the nose cone out. As compared to black powder, the CO 2 canister will not leave residue on the rover or the solar panels. This allows for multiple deployment tests without the need to continually clean the rover and the surfaces of the payload bay. This design has become the preliminary design, and is detailed completely in section Figure 25: Peregrine CO 2 Kit 65 65

67 Sonic Distance Measurement A team derived goal is to measure how far the payload has moved by calculated the distance using two audio emitters and receivers. One emitter would remain at the rocket body and start sending signals to the rover, after which the rover would replay with a signal back. Both emitters would be using different frequencies to avoid cross talk. This method would require a module on the rover to measure temperature and pressure of the environment to calculate the speed of sound at that given time to increase distance accuracy. This method would require large amount of time to fine tune and research to ensure the goals could be met, and given the requirement that the rover be five feet from any part of the rocket it was determined that this mechanism would not be effective at determining the required distance. This approach was therefore abandoned GPS Distance Measurement GPS was considered as a measuring system to achieve the teams derived goal of measuring how far the payload has moved. GPS requires clear line of sight to the sky as the rover should have after deployment making GPS an easy option to incorporate onto the rover. GPS is a well matured technology to locate objects across the Earth's globe. The rover will pull in coordinates from the GPS and run them through the Haversine formula to determine its location on the globe. The microcontroller will handle both running the coordinates through the formula and determining how far the rover has moved. The disadvantage to using a GPS is that GPS is not always accurate, often within only 20 feet. GPS accuracy can also be affected by environmental conditions, including atmospheric conditions and nearby obstructions or reflection points. However, additional sensors can be combined, such as an accelerometer and gyroscope, with the GPS to form a more accurate overall measurement. This system is commonly found in autonomous robotics and so a large body of expertise may be utilized to assist in the design of this systems. The GPS based distance measurement has been selected for the preliminary design and will be described further in section Preliminary Payload Design Rover Design The rover will utilize four Adafruit continuous rotation micro servos as seen in figure 36 in order to give power to each of the four wheels independently. These servos will be able to rotate at a maximum speed of 130 RPM. When no load is applied the servos will draw 120 ma of current and the servo will stall out when it is using 650 ma of current. At 6 volts the servos will produce oz/in of torque as a maximum. Each servo is 1.3" x 1.2" x 0.5" and weighs ounces. They will all be attached to the rover base via the mounting holes on the servos and machine screws, a small amount of epoxy will be added to ensure an especially stable attachment to the rover. Each of the servos will be attached to a gear made of nylon that will act as the wheels when driving outside of 66 66

68 the rocket, and act as way to extract the rover from the rocket after landing. The gears themselves are 2.25" in diameter and 5/16" wide. They have a pitch of 16 and have 36 teeth. They mesh with a 5/16" wide rack of the same pitch that will be mounted to a 3D printed base attached to the inner diameter of the payload tube. All gears on the rover and rocket utilize a 14 1/2 pressure angle that maximizes contact between the mating teeth for smooth operation during gear on gear power transmission. The rover base will be 3D printed on a stereolithography printer to allow for complex and smooth parts that a regular fused deposition modeling printer cannot offer the team. The base has cutouts and mounting points to allow for easy attachment of the various components of the rover. There is also a T channel on the bottom of the rover base to attach the rover to the payload base and to keep the rover secure during flight. Finally, the rover will have 3D printed parts attached to holes in the gear wheels. These 3D printed parts, called wheel balls, will ensure that even if the rover exits the rover on its side, it will not be able to balance on the large flat surfaces of the wheel gears, and instead be forced to tip to either side allowing the rover to drive away after settling on the ground. Figure 26: Isometric view of payload bay 67 67

69 Figure 27: Micro Servo Figure 28: Top view of Payload Bay 68 68

70 Figure 29: Isometric view of the rover Figure 30: Front view of the rover 69 69

71 Figure 31: 1/8 Inch Rack Figure 32: 5/16 Inch Rack 70 70

72 Figure 33: Panel Carrier drawing Figure 34: Payload Tube Base drawing 71 71

73 Figure 35: Rover Base Drawing Solar Panel Design The original design for the rover's solar panels was multiple 0.866"x0.276"x0.076" solar panels that would be arranged to be roughly 2.3"x3"x0.076". This was changed to solar panels measuring 1.378"x1.654"x0.079" when the club transitioned from the solenoid deployed design as shown in section It was changed to one larger panel in each section for ease of installation and added simplicity of the construction of the rover solar panel system. These panels output 223mW of power max at 6.3 V and 50 ma. It is a monocrystalline panel produced by IXYS Solar as a part of their IXOLAR series of small solar panels. It will be bought though DigiKey as they offer good savings when buying in bulk, along with quick shipping. Two of the solar panels will be directly mounted to the rover body and will not actively move during the solar panel deployment process. The other four solar panels will be mounted to solar panel carriers which will allow the rover to deploy the solar panels out of the sides of the rover. The panels carriers will be moved into their extended position with an ultra-nano servo shown in the appendix. This servo offers oz/in of torque at a speed of 0.1 sec/60. It is also a very small servo, thus the name, at 0.737" x 0.631" x 0.275". That allows the servo to fit well into the small space available on the rover. The servo will turn a 16 tooth gear that will mesh with two racks stacked opposingly on the gear so when the servo turns the gear, the two racks will be moved in opposite directions. The racks are 1/8" wide, 1/8" tall, and 3.09" long. These racks will be attached to the solar panel carriers and move the carriers as they are moved by the gear and servo. Once the carriers have been fully deployed, they will bind with rover base and be locked in place preventing further movement

74 Once fully deployed the rover will have a total surface area of square inches of total solar panel area exposed to the sun. Figure 36: Solar panel mechanism detail Figure 37: Rover with deployed solar panels 73 73

75 CO2 Deployment System The preliminary design utilizes a CO 2 canister to eject the nose cone from the payload bay, allowing the rover to deploy itself. The canister is a Peregrine 8g canister manufacture by Tinder Rocketry. The Peregrine CO 2 canister is intended for use in high powered rocketry recovery bays, so it is already suited for the rigors of flight. Using a commercially available system reduces design complexity and the risk of a custom-built system. Canisters come pressurized from the vendor, and are single use, so their pressurization history is only the initial fill. This follows section 2.14 from the USLI handbook. The canister will be located in a coupler at the rear of the payload bay. A turn switch will be accessible to the exterior of the rocket and will serve as the arming switch for the ejection system, per NASA's requirements. The rover will receive the signal to deploy the rocket and then activate the e-match inside of the CO 2 canister assembly. The e- match will ignite a small black powder charge, which will propel the canister into a pin that will puncture the throat of the canister. The electronics that control the canister are detailed in section The pressure caused by the CO 2 release will eject the nose cone from the payload. The nose cone will be attached to the body tube of the payload bay by nylon shear pins, a similar construction to the recovery system. As described in section , the use of CO 2 over black powder prevents soot build up on the rover, payload bay and electronics. Additionally, this ensures the nose cone is safely out of the path of the rover during deployment and when driving on the terrain. Figure 38: Peregrine CO2 Canister 74 74

76 Figure 39: Rover Deployment Coupler 75 75

77 Figure 40: CO 2 Canister drawing Payload Electronics The rover will consist of a Microcontroller, inertial measurement unit (IMU), GPS module, and radio transceiver. Once a signal is received from the ground station, the rover will begin its deployment process. The majority of the payload s electronics will be housed on the rover body itself. This reduces the overall weight of the payload system as only a single power source and controller would be required, while also keeping the overall functionality simple. To facilitate the operation of the deployment system, a coupling system has been devised that allows free motion of the rover while retaining the ability to form strong electrical contacts between the rover s electronics and the e-match circuitry on the rocket

78 Figure 41: ATMega32U2 Development Board Front and Back 77 77

79 Figure 42: ATMega32U2 Development Board Circuit The microcontroller the team has elected to control the rover is the Atmel ATMega 32U2 microcontroller. The selected microcontroller has key features that makes it perfect for the objective. First, the microcontroller incorporates Atmel's DFU USB bootloader that allows the team to program the microcontroller without an eternal programmer such as the AVR ISP MK2. The microcontroller also incorporates USB in hardware which allows the microcontroller to communicate with a computer for debugging purposes. The microcontroller is responsible for controlling 5 servo motors using PWM and includes the necessary 5 PWM ports 2 of which are 8 bit and the remaining 3 are 16 bit. The controller is also capable of communicating over SPI to allow data logging to an onboard SD card, and reading of data from the onboard IMU

80 Figure 43: SAM-MBQ-O Development Board Front and Back A SAM-MBQ-0 model GPS chip will be housed onboard the rover to provide positional accuracy during movement, while also ensuring this section of the payload will not be lost during or after flight. This GPS should not be confused with the main GPS unit located in the nose of the cone used to detect the rocket s position. The GPS will be disabled during ascent as it incorporates a G force lock out at 4G that will prevent further operation if exceeded. Upon reception of the signal, a control signal from the microcontroller will activate the GPS. Coordinates are read from the GPS and run through the Haversine equation to calculate the location of the rover on the Earth's surface. As the rover moves, the microcontroller will continue to pull coordinates from the GPS and will establish the distance travelled. To account for the inherent inaccuracy of GPS, the IMU s sensor readings will be combined with the GPS to produce greater accuracy. This is a standard control approach used in robotics as the GPS has a significant local distance error and a low temporal resolution, but it does have good absolute accuracy. Combining GPS with the faster responding accelerometer allows a greater temporal resolution and will provide relatively accurate positional changes over short periods of time. Combining the two allows for greater temporal and local accuracy while maintaining the accuracy of the system over prolonged periods of time. An MPU axis IMU containing a 3-axis accelerometer, gyroscope, and magnetometer will be installed onboard the rover. The team obtained experience with this IMU during last year s competition and its low cost, small form factor, and relative simplicity make a clear leading choice for this year s design. The IMU will facilitate the completion of multiple payload objectivities. First, the IMU will act as the primary sensor input that determines when the rocket has landed after launch. This mechanism ensures that deployment will only occur when the rocket is in a safe location. In the event of the transceiver receiving an anomalous signal interpreted as a deployment 79 79

81 command, the IMU will provide an override to the activation of the deployment mechanism if the deployment prerequisites are not met. If deployment pre-requisites are met, the IMU will then perform a vital role in the deployment of the rover. Since the rover has been designed to operate correctly regardless of operation and no orientation methods are utilized within the rocket, in order to properly control the rover, the orientation must be known. This is because if the rover is deployed upside-down then the wheels will have to rotate in the direction opposite of the direction they would travel if right-side-up. The orientation of the rover can easily be detected from the IMU s accelerometers, potentially assisted by the magnetometer readings. In addition, the accelerometer and gyroscope can be used to detect when the rover releases itself from the deployment mechanism. It will also be possible to use the IMU to provide additional positional accuracy to the rover by incorporating the accelerometer and gyroscope readings into a Filtering algorithm within the microcontroller. By combining the IMU readings with the GPS a greater positional accuracy can be obtained as the GPS has good absolute positioning, but slow temporal resolution and spatial precision, while the IMU has poor absolute positioning, especially over large intervals, but high temporal resolution and is sufficient to monitor local movements. Figure 44: AX5043 RF Transceiver Circuit 80 80

82 Figure 45: AX5043 RF Transceiver Development Board Front and Back To trigger the ejection of the rover from the rocket, a 433MHz RF transceiver will be used. The chip that will be integrated into the rover will be the ON Semiconductor AX5043. Th AX5043 is a General ISM band RF transceiver with selectable RF frequency. The team chose 433 MHz as the frequency because it allows use of a low cost and simple quarter wavelength monopole antenna that is calculated to be 6.81 inches long. Quarter wavelength monopole antennas have a characteristic impedance of 36.8 ohms when mounted above a good ground plane. This allows the team to trim a single piece of the above-mentioned length and build a pi match filter that matches the 50 ohm output of the AX5043 and 50 ohm transmission line to the above mentioned 36.8 ohms of the quarter wavelength monopole antenna. The AX5043 chip offers greater calculated range than the XBee Pro Series 1 that the team has used in the past. The calculated range of the AX5043 with the abovementioned quarter wavelength monopole antenna is 292 miles compared to the calculated range of the XBee Pro S1 which is 2 miles. In testing, the XBee Pro S1 attained a maximum range of 408 feet which is 1/39 of the calculated range. As a result of this, when determining the range of the AX5043, it is assumed the transceiver has a real-world range of 1/39 of the calculated range. Therefore, the team has deducted that the real-world range of the AX5043 is approximately 7.5 miles. While this number may seem excessively high, the AX5043 transmission power is 16 dbm whereas the XBee Pro Series 1 transmission power is 18dBm. With a lower transmit power, the range of the AX5043 is better than the XBee Pro Series 1 because the AX5043 operates at a lower frequency (433MHz vs 2.4GHz) than the XBee. Also, the AX5043 has a higher sensitivity than the XBee Pro Series 1 (-137dBm vs 100dBm). Radio transceivers will be located on the rover and at a command station for telemetry, data logging and rover deployment. A Python script will be used to monitor serial data received from the command station transceiver. This data will be saved locally and displayed in real time

83 The CO 2 ejection mechanism requires an e-match to trigger the ejection. To ignite the e-match an N-channel MOSFET will complete the circuit by connecting the e-match directly to the 7.4V 2S Lithium Polymer battery on the rover. Since the e-match and CO 2 canister will be fixed to the rocket body, a coupling system must be utilized so that the rover can separate completely from the rocket body while retaining a mechanism to activate the canister. To this end the team intends to incorporate spring-loaded pogo pins mounted to the main electronics board. These pins will contact another PCB containing large contact pads for each pogo pin that will connect to the CO 2 deployment system through screw terminals. Two pads will be used to trigger the e-match and two will be used to ensure continuity has been achieved before triggering. This system requires no mechanical coupling, only the ability to drive the rover into the contacts which is facilitated by the rack-pinion system. Four servo motors are fastened directly to the nylon gear wheels and will be used to propel the rover and maintain the rover position during flight. The motors that are being used are the FITEC FS90R Continuous Rotation Micro Servo. The lithium polymer battery will be able to supply the necessary voltage and current for the servo motors. Based on a worse case condition of stall across all 4 servos when operating at 6V, a total of 2.6A could be pulled which is well within the range of the selected LiPo battery. However, the operating voltage of the servos is 6V maximum, which is well below the nominal voltage of the 2S LiPo (7.4V). Thus, a step-down voltage converter (buck) will be utilized to ensure correct operating voltages. Each servo will be allocated its own PWM pin on the microcontroller so that all-wheel drive can be utilized. A fifth servo will be located on top of the rover to deploy the solar panels. The selected servo for this purpose is the Hitec HS-35D Ultra Nano Size Servo. This servo is limited to 4.8V and so a separate voltage regulating circuit is required for this servo. The fifth and final PWM pin available on the ATMega 32u2 will be connected to this servo. Due to the design of the sliding panels, it will be critical to know how long to apply the control signal to this servo to ensure complete deployment of the sliding panels. The necessary time will be evaluated by a series of ground tests performed at the expected operating conditions

84 Figure 46: Electrical Block Diagram Subsystem Interfaces The payload bay is comprised of body tube forward of the main parachute compartment. The entirety of the payload bay and related mechanisms are held within the body tubes of the vehicle. The nose cone of the rocket is attached to the payload tube by four nylon shear pins, that will release the nose cone when the CO 2 ejection system is activated. The CO 2 ejection system is housed at the base of the payload bay inside of a couple section, similar to the construction of the recovery bay. This will be rigidly screwed to the payload body tube. The aft bulkhead of this coupler is where the main parachute eyebolt will be located. This coupler will mate with the main parachute compartment of the rocket, which completes the payload bay interface

85 Figure 47: Payload Integration 84 84

86 6 Project Plan 6.1 Requirements Verification Table 10: Requirements Verification Requirement Requirement Description Verification Method 1.1 Students on the team will do 100% of the project, including design, construction, written reports, presentations, and flight preparation with the exception of assembling the motors and handling black powder or any variant of ejection charges, or preparing and installing electric matches (to be done by the team s mentor). The University of Toledo Rocketry club is comprised only of active students at the University of Toledo. The team has members with experience with high powered rocketry, along with L1 and L2 certifications. The team will be instructed on how to assemble and prepare the rocket. 1.2 The team will provide and maintain a project plan to include, but not limited to the following items: project milestones, budget and community support, checklists, personnel assigned, educational engagement events, and risks and mitigations. The project plan for Project Kronos is detailed in section Foreign National (FN) team members must be identified by the Preliminary Design Review (PDR) and may or may not have access to certain activities during launch week due to security restrictions. In addition, FN s may be separated from their team during these activities. There are no Foreign Nationals to report to NASA The team must identify all team members attending launch week activities by the Critical Design Review (CDR). Team members will include: Students actively engaged in the project throughout the entire year. One mentor (see requirement 1.14). No more than two adult educators. The team will report members attending launch week by CDR. The team has one mentor and one educator from the University of Toledo

87 1.5 The team will engage a minimum of 200 participants in educational, hands-on science, technology, engineering, and mathematics (STEM) activities, as defined in the Educational Engagement Activity Report, by FRR. An educational engagement activity report will be completed and submitted within two weeks after completion of an event. A sample of the educational engagement activity report can be found on page 31 of the handbook. To satisfy this requirement, all events must occur between project acceptance and the FRR due date. 1.6 The team will develop and host a website for project documentation. 1.7 Teams will post, and make available for download, the required deliverables to the team Web site by the due dates specified in the project timeline. The education team has already educated 352 youths with hands on projects. These are detailed in section 6.2. The website is active at and has been updated regularly over the past two years. Documents are hosted on utrocketry.com and will be ed as pdf attachments to the leadership of USLI. 1.8 All deliverables must be in PDF format. All deliverables will be PDF format and available on the team website and ed to NASA. 1.9 In every report, teams will provide a table of contents including major sections and their respective sub-sections In every report, the team will include the page number at the bottom of the page. The table of contents is located at the beginning of every report, along with a table of tables and list of figures. Page numbers are located in the footer of every page of the report The team will provide any computer equipment necessary to perform a video teleconference with the review panel. This includes, but is not limited to, a computer system, video camera, speaker telephone, and a broadband Internet connection. Cellular phones can be used for speakerphone capability only as a last resort All teams will be required to use the launch pads provided by Student Launch s launch service provider. No custom pads will be permitted on the launch field. Launch services will have 8 ft rails, and 8 and 12 ft rails available for use. The university of Toledo has numerous conferences rooms with conference phone systems that are available for the rocketry team's use. Kronos is designed to use rail buttons for either 1010 or 1515 rails, dependent on speed off rail calculations

88 1.13 Teams must implement the Architectural and Transportation Barriers Compliance Board Electronic and Information Technology (EIT) Accessibility Standards (36 CFR Part 1194) 1.14 Each team must identify a mentor. A mentor is defined as an adult who is included as a team member, who will be supporting the team (or multiple teams) throughout the project year, and may or may not be affiliated with the school, institution, or organization. The mentor must maintain a current certification, and be in good standing, through the National Association of Rocketry (NAR) or Tripoli Rocketry Association (TRA) for the motor impulse of the launch vehicle and must have flown and successfully recovered (using electronic, staged recovery) a minimum of 2 flights in this or a higher impulse class, prior to PDR. The mentor is designated as the individual owner of the rocket for liability purposes and must travel with the team to launch week. 2.1 The vehicle will deliver the payload to an apogee altitude of 5,280 feet above ground level (AGL). 2.2 The vehicle will carry one commercially available, barometric altimeter for recording the official altitude used in determining the altitude award winner. Teams will receive the maximum number of altitude points (5,280) if the official scoring altimeter reads a value of exactly 5280 feet AGL. The team will lose one point for every foot above or below the required altitude. 2.3 Each altimeter will be armed by a dedicated arming switch that is accessible from the exterior of the rocket airframe when the rocket is in the launch configuration on the launch pad. 2.4 Each altimeter will have a dedicated power supply. The rocketry club will follow the standards of EIT. The team's mentor is Art Upton, who is L3 certified with NAR and has more than 2 high powered flights in his history. The rocket is his for liability purposes. The predicted altitude is detailed in section 3.3. The rocket will fly with two StrattologgerCF Altimeters, one will be marked for judging after the flight. Each altimeter will be activated with a turn switch that will only be activated when the rocket is prepared on the pad. Each altimeter will be powered by their own, securely fastened, 9 volt battery

89 2.5 Each arming switch will be capable of being locked in the ON position for launch (i.e. cannot be disarmed due to flight forces). 2.6 The launch vehicle will be designed to be recoverable and reusable. Reusable is defined as being able to launch again on the same day without repairs or modifications. 2.7 The launch vehicle will have a maximum of four (4) independent sections. An independent section is defined as a section that is either tethered to the main vehicle or is recovered separately from the main vehicle using its own parachute. 2.8 The launch vehicle will be limited to a single stage. 2.9 The launch vehicle will be capable of being prepared for flight at the launch site within 3 hours of the time the Federal Aviation Administration flight waiver opens The launch vehicle will be capable of remaining in launch-ready configuration at the pad for a minimum of 1 hour without losing the functionality of any critical onboard components The launch vehicle will be capable of being launched by a standard 12-volt direct current firing system. The firing system will be provided by the NASA-designated Range Services Provider The launch vehicle will require no external circuitry or special ground support equipment to initiate launch (other than what is provided by Range Services) The launch vehicle will use a commercially available solid motor propulsion system using ammonium perchlorate composite propellant (APCP) which is approved and certified by the National Association of Rocketry (NAR), Tripoli Rocketry Association (TRA), and/or the Canadian Association of Rocketry (CAR). The altimeters' arming switches are turn switches that cannot be rotated by forces during flight. The rocket is built with a fiberglass body, and all parts of the rocket are fully reusable. The rocket has only two independent body parts, that will descend under parachute tethered with tubular nylon shock cord. The rocket is propelled by a single stage using a commercially available solid fuel motor. The rocket is designed such that it can be prepared within three hours. Additionally, the team will be trained and practice set up of the vehicle, so that assembly and preparation of the rocket during launch day can be completed in a timely manner. All electronics systems on the rocket have adequate battery life to have a pad stay time of at least 1 hour. The rocket utilizes a commercially available solid fuel motor that is designed to operate on typical 12 volt ignition systems. The rocket requires only the motor igniter to launch the vehicle. The vehicle utilizes an Aerotech motor, which is approved and certified by NAR, TRA and CAR

90 2.14 Pressure vessels on the vehicle will be approved by the RSO and will meet the following criteria: The minimum factor of safety (Burst or Ultimate pressure versus Max Expected Operating Pressure) will be 4:1 with supporting design documentation included in all milestone reviews. Each pressure vessel will include a pressure relief valve that sees the full pressure of the valve that is capable of withstanding the maximum pressure and flow rate of the tank. Full pedigree of the tank will be described, including the application for which the tank was designed, and the history of the tank, including the number of pressure cycles put on the tank, by whom, and when The total impulse provided by a College and/or University launch vehicle will not exceed 5,120 Newton-seconds (L-class) The launch vehicle will have a minimum static stability margin of 2.0 at the point of rail exit. Rail exit is defined at the point where the forward rail button loses contact with the rail The launch vehicle will accelerate to a minimum velocity of 52 fps at rail exit All teams will successfully launch and recover a subscale model of their rocket prior to CDR. Subscales are not required to be high power rockets All teams will successfully launch and recover their full-scale rocket prior to FRR in its final flight configuration. The rocket flown at FRR must be the same rocket to be flown on launch day Any structural protuberance on the rocket will be located aft of the burnout center of gravity. The details of the pressure vessel are listed in section The pressure vessel is a commercially available CO 2 canister designed for recovery ejection systems for high powered rockets, and is suitable for the payload ejection purpose. The preliminary motor chosen for the vehicle is a K-class motor, which is below an L-class motor. The static stability margin is detailed in section and is above 2.0 calibers. The velocity of the rocket is detailed in section 3.3 The team has a motor selection form that ensures the rocket will have a rail exit velocity of 52 ft/s. A subscale model of the vehicle is currently being designed and a launch window identified that is before CDR is due. After CDR approval of the design, the full-scale rocket will be constructed and flown. Launch windows have been identified before FRR is due. There are no structural protuberances on the rocket, with the exception of the fins which are located aft of the burnout center of gravity

91 Vehicle Prohibitions: The launch vehicle will not utilize forward canards. The launch vehicle will not utilize forward firing motors. The launch vehicle will not utilize motors that expel titanium sponges (Sparky, Skidmark, MetalStorm, etc.) The launch vehicle will not utilize hybrid motors. The launch vehicle will not utilize a cluster of motors. The launch vehicle will not utilize friction fitting for motors. The launch vehicle will not exceed Mach 1 at any point during flight. Vehicle ballast will not exceed 10% of the total weight of the rocket. 3.1 The launch vehicle will stage the deployment of its recovery devices, where a drogue parachute is deployed at apogee and a main parachute is deployed at a lower altitude. Tumble or streamer recovery from apogee to main parachute deployment is also permissible, provided that kinetic energy during drogue-stage descent is reasonable, as deemed by the RSO. 3.2 Each team must perform a successful ground ejection test for both the drogue and main parachutes. This must be done prior to the initial subscale and full-scale launches. 3.3 At landing, each independent sections of the launch vehicle will have a maximum kinetic energy of 75 ft-lbf. 3.4 The recovery system electrical circuits will be completely independent of any payload electrical circuits. As detailed in section 3 and the following subsections, the vehicle design does not utilizes any of the prohibitions in requirement As detailed in section 3.2, a drogue will be deployed at apogee, followed by a main parachute at 700 feet. The kinetic energy under both main and drogue are detailed in the same section and in the Milestone Review Flysheet A ground test will be performed before every flight to ensure proper performance of the recovery system. Flights will be grounded unless a successful ground test is performed. The kinetic energy of each segment is detailed in section 3.2 as well as the milestone review flysheet and will be checked before each flight against the simulation to ensure they are below the 75 ft-lbf requirement. The recovery electronics are located in a separate compartment from the payload, separated by the main parachute compartment. The electronics systems are entirely separate

92 3.5 All recovery electronics will be powered by commercially available batteries. 3.6 The recovery system will contain redundant, commercially available altimeters. The term altimeters includes both simple altimeters and more sophisticated flight computers. 3.7 Motor ejection is not a permissible form of primary or secondary deployment. 3.8 Removable shear pins will be used for both the main parachute compartment and the drogue parachute compartment. 3.9 Recovery area will be limited to a 2500 ft. radius from the launch pads An electronic tracking device will be installed in the launch vehicle and will transmit the position of the tethered vehicle or any independent section to a ground receiver The recovery system electronics will not be adversely affected by any other on-board electronic devices during flight (from launch until landing). 4.3 If the team chooses to fly additional experiments, they will provide the appropriate documentation in all design reports, so experiments may be reviewed for flight safety. The team utilizes Duracell 9 bolt batteries for powering the recovery system, new batteries are used for each flight. The vehicle utilizes two StrattologgerCF altimeters that work as a redundant system. One altimeter will be set to activate shortly after the primary altimeter, as detailed in section 3.2. The motor used will be plugged or have the ejection charge dumped, the recovery system utilizes drogue and main charges activated by the altimeters. Nylon shear pins have been selected as shear pins and will be used to secure any sections that are designed to separate during flight. Drift distances have been calculated and will be updated as the design progresses, and are detailed in section 3.3.4, the maximum drift distance in 20 mph winds is less than 2500 ft. The vehicle utilizes as TeleGPS system which will be located in the nose of the rocket. This will transmit the location of the rocket over HAM band radio to the team's ground station. The recovery system altimeters are located in an independent segment of the rocket and will not be adversely affected by the other electronics located in the payload bay and nose. The vehicle will only operate the rover experiment

93 4.5.1 Teams will design a custom rover that will deploy from the internal structure of the launch vehicle. At landing, the team will remotely activate a trigger to deploy the rover from the rocket. After deployment, the rover will autonomously move at least 5 ft. (in any direction) from the launch vehicle. Once the rover has reached its final destination, it will deploy a set of foldable solar cell panels. 5.1 Each team will use a launch and safety checklist. The final checklists will be included in the FRR report and used during the Launch Readiness Review (LRR) and any launch day operations. 5.2 Each team must identify a student safety officer who will be responsible for all items in section The role and responsibilities of each safety officer will include, but not limited to: Monitor team activities with an emphasis on Safety during: Design of vehicle and payload Construction of vehicle and payload Assembly of vehicle and payload Ground testing of vehicle and payload Sub-scale launch test(s) Full-scale launch test(s) Launch day Recovery activities Educational Engagement Activities 5.4 During test flights, teams will abide by the rules and guidance of the local rocketry club s RSO. The allowance of certain vehicle configurations and/or payloads at the NASA Student Launch Initiative does not give explicit or implicit authority for teams to fly those certain vehicle configurations and/or payloads at other club launches. Teams should communicate their intentions to the local club s President or Prefect and RSO before attending any NAR or TRA launch. As detailed in section 5, the rover will meet these requirements. The verifications for these requirements and additional team-derived requirements are explained in section Launch and safety checklists will be developed that are representative of the Project Kronos vehicle and will be a part of FRR and utilized during fullscale test flight following CDR. The safety officer for UT Rocketry is Tori Raber. The Safety Officer for the team has been briefed on her duties, and is a nursing student with experience in safety. The responsibilities and expectations for the safety officer are detailed in section 4 of this report. Multiple members in the team are L1 and L2 certified, experienced with obeying NAR regulations. The team has developed relationships with local NAR chapters, who have worked with the team to meet team needs for launches and receive proper approval

94 5.5 Teams will abide by all rules set forth by the FAA. The team has L1 and L2 certified members who are aware of FAA regulations and will ensure the team abides by them appropriately. 6.2 Team Derived Requirements Table 11: Team Derived Requirements Team Derived Requirement The payload will be able to detect the transmitted signal from at least 2,500 feet away. The incoming radio signals will be filtered so that the response of the deployment system will only occur for the team s generated signal and at safe operating conditions. The rover s locking mechanism (servos, wheels, rack) will be able to hold the rover position for a rocket acceleration of up to 15g. The rover s battery life will be at least 2 hours under the average expected operating conditions. The solar panels will completely slide out from their position on the rover. Requirement Verification A high sensitivity transceiver will be selected, and the transmitting antenna will be supplied sufficient power while staying with regulated power levels. A filter will be designed and constructed on the payload to process incoming signals to locate the desired signal, likely a desired pattern, before the deployment mechanism is activated. Further, data from the IMU and GPS will be analyzed to determine that the rocket has come to a stop after launching to prevent potential activations before reaching a safe deployment position. Servos with sufficiently high stall torques will be fixed to geared wheels and a rack that can resist the forces experienced at 15g. Careful consideration of the servos fastening system will be considered. The retaining rail design will be evaluated for strength within the potential operating conditions whilst retaining sufficient traction on the geared surface. A high capacity, low form factor battery will be selected for use on the payload. Preliminary current measurements will be made and evaluated to determine the average expected current draw and compared to the selected battery size. The rails the solar panel rides on will be rigidly constructed and any material residue will be removed to ensure the fit is snug during deployment. Tests will be performed with the servo to ensure proper deployment time is observed for the expected operating conditions

95 The nose cone will be ejected sufficiently far from the rocket body so as not to interfere with the deploying rover. Utilization of the GPS and the onboard Inertial Measurement System will be able to determine rover position to within 4 feet. The Rover will be able to transmit its GPS coordinates back to receivers near the ground station during and after flight. The gas discharge used to free the nose cone will not dislodge the rover from its position within the payload bay. The electrical coupling system from the rover to the CO 2 cylinder s e-match will successfully separate during rover deployment. All subsystems must be accurately represented in simulations. The rocket must assemble in a way that allows all subsystems to work as planned. Minimize chances of system failure wherever possible. Maintain a straight flight path. Educate 750 youth during the competition A properly sized CO 2 cylinder will be utilized for the given volume of the payload bay. Shear pins will be utilized to fasten the nose cone so that the energetics can properly eject the nose cone. A GPS receiver will be located on the rover alongside an IMU containing gyroscopes and accelerometers. A sensing protocol (e.g. Kalman Filter) will be used to improve the positional accuracy of the sensor suite. A high sensitivity receiver will be used at the ground station to pick up signals from the payload to account for its inherently low power design. The rover will be rigidly fixed by a combination of the servos, rack-pinion system, and the lower guide rail. The concentrated discharge will be offset from the rover s position so that most of the discharge passes harmlessly around the rover. The coupling system will be rigid, so as to resist the flight conditions, but will be susceptible to the force output by the rover in the desired orientation. All masses and dimensions of components will be measured once the components are obtained, then they must match the OpenRocket model. After each component is installed and secured, it will be remeasured. Have the vehicle team leader confirm each part has been properly attached. Measure the angle between each fin and use a level to make sure the fins are straight. The club has reached a total of 352 as of PDR. Additional events are planned for the future

96 6.3 Education The team would like to educate 750 youth on the topics of rocketry and space exploration. Currently, the club has educated 352 youth with hands on rocketry events. There are several events that the team will be participating in to accomplish this goal. The team participated in a Cub Scout recruiting day. They assisted children age 5-7 in building their model rockets and taught kids the science behind how a rocket works. The team also had a table to display the rockets used in the past two Student Launch competitions. The team taught the scouts how the rockets worked. When building the model rockets the kids learned the basic parts of a rocket. In addition to that they also learned how a rocket launches. The team went to an event at the University of Toledo called Hallow-engineering. Kids from the Toledo area came to learn about engineering. The team had a table for the main station. They constructed balloon rockets with the goal of moving as many paperclips as possible up to the ceiling. Kids learned how differences in mass can change how high and fast the balloons can rise. They also learned how the amount of air in the balloon can help the speed and height of the launch. The kid's competed to see who could get the most paperclips. This event made the kids excited about rockets and space. The team will be going to Clay High School to give a presentation to upper physics classes about rocketry basics, go a little in depth on some physics equations such as F = ma, trajectory equations, how to find center of mass of a rocket how to apply them in situations. The team will also talk about famous rockets in history and what can be accomplished with them, as well as what the team is currently working with the competition. The presentation will end with the club launching a rocket for the students on an F motor. The team will be going to an event in Akron, Ohio to teach Space Exploration merit badge to Boy Scout troops 390 and 310. In teaching Space Exploration merit badge the scouts will learn about, space pioneers, space travel, parts of a rocket, and how to build a rocket. They will also build their own model rockets and launch them. In addition to that they will be learning about the history of space travel, famous astronauts, and current events in space travel. With these events the team plans to educate roughly 750 youth about rocketry and space exploration. More events may be attended as the year goes on, allowing greater opportunity to expand and reach out to more people every year

97 Figure 48: Education Progress 96 96

98 6.4 Budget Projected Budget Part Quantity Payload Electronics Unit Cost Printed Circuit Boards NA NA NA Expected Cost Turnigy 1000mAh 2S 1 $8.98 $8.98 Syworks SE4150L-R 1 $3.20 $3.20 Single Pull Key Switch 1 $9.46 $9.46 STM LSM6DS3USTR 1 $3.01 $3.01 ON AX5043 Transceiver 1 $3.09 $3.09 Atmel ATmega32U2 1 $3.05 $3.05 Payload Hardware $30.79 Large 36 T Nylon Gear 2.25" dia 4 $15.83 $63.32 Small 16 T Nylon Gear 1/3" dia 1 $7.51 $7.51 Continuous Rotation Micro Servo 4 $7.50 $30.00 IXYS SLMD121H10L Panel 6 $13.13 $78.78 Hitec HS-35HD Ultra Micro Servo 1 $24.95 $ /8" Rack Gear 2 $5.82 $11.64 Mounting Hardware 1 $20.00 $20.00 Wheel Ball 4 NA NA Rover Base 1 NA NA Panel Carrier 2 NA NA Peregine Raptor CO2 kit 1 $ $ Payload Base 1 NA NA 5/16" Rack Gear 1 $8.88 $8.88 Body Hardware $ Nosecone 1 $99.00 $ " Fiberglass Tubing 6 $36.00 $ mm Fiberglass Tubing 2 $25.00 $

99 Fiberglass Centering Ring 3 $7.00 $21.00 Bulkhead 1 $7.00 $7.00 2/56 Shear Pin 8 $2.95 $23.60 Fiberglass Fin 4 $8.05 $32.20 Screws 8 $1.00 $8.00 AeroPak Motor Retainer 1 $38.00 $38.00 AeroPoxy 1 $53.30 $53.30 JB Weld 2 $22.76 $45.52 Welded Eye Bolts 4 $2.50 $10.00 Washers 4 $1.00 $4.00 Nuts 2 $0.50 $1.00 Rail Guide 1 $5.00 $5.00 Recovery $ Recovery Bay 1 $80.00 $80.00 CERT 3 DROGUE 24" CHUTE 1 $27.50 $27.50 Braided Nylon Shock Cord 28 $1.50 $42.00 SkyAngle Cert 3 Large 1 $ $ Nomex Blanket 4 $8.95 $35.80 Cable Ties 1 $9.99 $9.99 Terminal Blocks 4 $3.25 $13.00 Ejection Canister 4 $3.00 $12.00 Rotary Switch 4 $9.46 $ /4 in Quick Link 4 $0.93 $3.72 9V Battery 1 $17.72 $ V Battery Connector 2 $1.49 $2.98 Propulsion $ AeroTech K1000T 2 $ $ CTI H100 1 $35.00 $ mm Motor Mount Tube 1 $ $ Motor Retainer 1 $50.00 $50.00 Education $ F-Class Model Rocket Motors 3 $27.99 $83.97 Estes Pro Series II 3 Model Rocket 1 $44.99 $44.99 Estes A3-4T Motors 3 $10.29 $30.87 $

100 Travel 15-Person Van 1 $ $ Fuel 1 $ $ Hotel 3 $ $1, $1, Total Rocket Cost $1, Total Material Cost $3, Funding Plan Funding Plan Source Amount Marathon Petroleum Inc. $ 2, Pilkington $ UT MIME Department $ 1, Dassault Systems $ DTE Energy $ Rotary Club $

101 6.5 Timeline

102 7 Appendix Figure 49: Solar Panel drawing Figure 50: Servo

103 Figure 51: Gear Wheel drawing Figure 52: Solar Panel Gear drawing

104 Figure 53: Wheel Ball drawing Figure 54: Rover Deployment Bulkhead drawing