University of Illinois Urbana-Champaign Illinois Space Society Student Launch Preliminary Design Review November 4, 2016

Transcription

1 University of Illinois Urbana-Champaign Illinois Space Society Student Launch Preliminary Design Review November 4, 2016 Illinois Space Society 104 S. Wright Street Room 18C Urbana, Illinois 61801

2 Table of Contents 1. PROJECT SUMMARY Team Summary Launch Vehicle Summary Payload Summary CHANGES SINCE PROPOSAL Launch Vehicle Changes Payload Changes Project Plan Changes VEHICLE CRITERIA Launch Vehicle Overview Mission Statement Success Criteria Mass Statement System Modeling Booster Subsystem Motor Selection and Justification Motor Casing Motor Mount Assembly Motor Retention Outer Airframe Fins Launch Rail Integration Booster Subsystem Mass Statement Avionics Coupler Subsystem Avionics Coupler Tubing Avionics Sled Configuration Bulkhead Configuration Avionics Coupler Subsystem Mass Statement Payload Bay and Upper Airframe Subsystems Payload Sled Configuration... 25

3 Payload Bulkhead Configuration Nosecone Payload Bay and Upper Airframe Subsystems Mass Statement Recovery Subsystem Deployment Scheme Deployment Mechanisms Avionics Hardware Avionics Electrical Setup Attachment Hardware Redundancy Mission Performance Predictions Simulation Methods Flight Profile Analysis Stability Kinetic Energy Drift SAFETY Safety Plan Overview Safety Briefings Member Requirements Equipment Training NAR Mentor Energetics Handling Motor Purchase and Storage Risk Assessment Overview Personnel Hazard Analysis Failure Modes and Effects Analysis Launch Vehicle Payload Environmental Concerns Project Risk Analysis... 82

4 4.8. Preliminary Checklists Final Assembly Launch Compliance NAR High Power Safety Code Federal Aviation Requirements Range Safety Officer Authority PAYLOAD CRITERIA Payload Overview Payload Success Criteria Mechanical Landing Subsystem Landing Subsystem Selection Process Chosen Landing Subsystem Overview Tip-over Analysis of Landing System Launch Vehicle Integration Prototyping and Testing Image Processing Subsystem Subsystem Overview Processing Unit Selection Camera Selection Camera Position Power Delivery Software Altitude Adjustment Triggering Mechanism Prototyping and Testing PROJECT PLAN Competition Requirement Verification Plan Launch Vehicle Payload General Project

5 6.2. Team Requirement Verification Plan Launch Vehicle Payload General Project Educational Outreach Update Illinois Space Day Upcoming Events Budget Funding Plan Project Timeline Appendix A: Definitions Appendix B: Acronyms Appendix C: ISS Tech Team Safety Policy Appendix D: Education Feedback Form Appendix E: ISD Schedule and Brochure Appendix F: Nar High Power Rocket Safety Code Appendix G: Federal Aviation Regulations 14 CFR, Subchapter f, Part 101, Subpart c Amateur Rockets

6 List of Figures Figure 1: Flight profile overview... 3 Figure 2: Full rocket assembly modeled in NX Figure 3: Dimensioned drawing of launch vehicle Figure 4: Booster subsection modeled in NX Figure 5: Inward view of booster subsytem Figure 6: L1390G-P thrust curve Figure 7: RMS 75/3840 motor casing Figure 8. Motor Mount Tube Figure 9: Aeropack 75mm flanged motor retainer Figure 10: Dimensioned trapezoidal fin Figure 11: Close up of rail button Figure 12: Rail button mounting hardware Figure 13: Model of avionics coupler subsytem Figure 14: Both sides of the lower avionics sled Figure 15: Payload bay and upper airframe subsytem Figure 16: Payload sled one front and back Figure 17: Payload sled 2 front and back Figure 18: Model of nosecone Figure 19: Nosecone dimensioned in inches Figure 20: Section 1 -- avionics coupler Figure 21: Section 2 -- booster tube Figure 22: Section 3 -- Payload Coupler Figure 23: Section 4 -- Upper Airframe and Nosecone Figure 24. The AltOS main menu Figure 25. Example of a graph that AltOS can generate from flight data Figure 26. AltOS s ground tracking of a vehicle Figure 27. PerfectFlite DataCap software for StratoLogger Figure 28. Jolly Logic Chute Release Figure 29: Electrical schematics for the booster section Figure 30: Electrical schematic for the payload section Figure 31: Current desing modeled in OpenRocket Figure 32: Sample design modeled in RockSim V Figure 33. Custom MATLAB flight simulator user interface Figure 34. Example plot of altitude from MATLAB simulation Figure 35. Example plot of vertical velocity from MATLAB simulation Figure 36: Simulated flight, velocity, and acceleration profiles from OpenRocket Figure 37: Rendering of the full payload system in its launch configuration Figure 38: The landing system in its fully deployed configuration Figure 39: Exploded view of the landing leg connection mechanism

7 Figure 40: Minimum tipping angle over a range of horizontal speeds Figure 41: Landing leg deployment sequence Figure 42: Upper leg technical drawing Figure 43: Lower leg technical drawing Figure 44: Prototype 3D printed legs with springs attached Figure 45: Part of the assembly for the prototype landing system Figure 46: Full electrical schematic for the image processing subsystem Figure 47: Block diagram of the fully redundant image processing subsystem Figure 48: Diagram of the payload cameras, showing the estimated available line of sight of the cameras in blue Figure 49: Camera angle optimization simulation based on tarp distance from launch Figure 50: Image processing software overview Figure 51: Simulation showing the expected tarp sizes throughout sample flight profiles Figure 52: Orbital simulator demo Figure 53: Egg drop challenge competition Figure 54: Example ISD participant survey

8 List of Tables Table 1: Vehicle and Recovery Competition Requirements... 4 Table 2: Vehicle and Recovery Team-Derived Requirements... 8 Table 3: Projected Mass Statement... 9 Table 4: L1390G-P Performance Characteristics Table 5: Airframe Material Trade Study Table 6: Fin Material Selection Table 7: Booster Subsystem Mass Statement Table 8: Avionics Coupler Subsystem Mass Statement Table 9: Payload Bay and Upper Airframe Mass Satement Table 10: OpenRocket Apogee Simulations Table 11: Flight Profile Performance Characteristics Table 12: Stability Parameters of the Launch Vehicle at Takeoff Table 13: Mass under Each Parachute Broken Down by Section Table 14: Drogue Parachute Options Table 15: Main Parachute Options Table 16: Phantom Parachute Parameters Table 17: Payload Parachute Options Table 18: Performance Characteristics of Chosen Parachutes Table 19: Terminal Velocity of Each Section of the Vehicle for Chosen Parachutes Table 20: Kinetic Energy of Each Section of the Vehicle at Impact for Chosen Parachutes Table 21: Drift Calculations for Each Section of the Launch Vehicle Table 22: Risk Assessment Codes (RACs) Table 23: Level of Risk and Member Requirements Table 24: Severity Definitions Table 25: Probability Definitions Table 26: Personnel Hazard Analysis Table 27: Referenced Material Safety Data Sheets (MSDS) Table 28: Structures and Recovery Risk Analysis Table 29: Payload Risk Analysis Table 30: Potential Hazards to the Environment Table 31: Potential Environmental Hazards to the Project Table 32: Project Risks Table 33: Payload Requirements Table 34. Success Criteria for the Image Processing Payload Table 35. Landing Mechanism Trade Study Table 36: Tip Over Angle Table 37: Landing Legs Subsystem Testing Table 38: Processing Unit Trade Study Table 39: Image Processing Camera Trade Study

9 Table 40: Image Processing System Energy Budget Table 41: Ground Imaging Subsystem Tests Table 42: Competition Safety Requirements Table 43: General Competition Requirements Table 44: General Project Team-Derived Requirements Table 45: Project Expenses Table 46: Funding Sources Table 47: Project Gantt Chart Table 48: Launch Vehicle Structure Gantt Chart Table 49: Payload Gantt Chart

10 1. PROJECT SUMMARY 1.1. Team Summary Team Name Illinois Space Society Student Launch Team Team Website Mailing Address 104 S. Wright Street Room 321D Urbana, Illinois Team Leader Stephen Vrkljan, Project Manager Phone: (630) Safety Officer Brian Hardy NAR Mentor Mark Joseph, NAR Number 76446, Certified Level Launch Vehicle Summary The vehicle will be a single stage rocket that separates at apogee. The vehicle for this project will measure 119 inches from the aft end of the motor retainer to the tip of the nose cone. The vehicle body will be constructed of inch outer diameter fiberglass. The current design mass of the vehicle is 35.4 pounds. The motor currently chosen for this vehicle is the Aerotech L1390G-P reloadable rocket motor. The lower separation section will employ a traditional dual deploy recovery system. An 18 Fruity Chutes elliptical will act as a drogue that deploys 2 seconds after separation and a 96 Iris Ultra will act as a main that deploys at 650ft AGL. The upper separation section will employ a traditional drogue-less dual deploy recovery system with a single 44 in Skyangle elliptical acting as the payload parachute. The preliminary design review flysheet is included at the end of this document Payload Summary The team is tackling the target detection and vertical landing challenge. Fold-out doublejointed aluminum legs will allow for vertical landing without the need for any electronics or controllers. A combination of Raspberry Pi hardware, open-source image processing packages, and custom coded analysis and interface software will allow the rocket to detect the target tarps during the ascent of the launch vehicle. 1

11 2. CHANGES SINCE PROPOSAL 2.1. Launch Vehicle Changes The design for the launch vehicle for the ISS Student Launch team has been changed since the proposal. One change is the type of motor. The team decided that an AeroTech L1390G motor would be a better fit for this year s rocket instead of the AeroTech L1170FJ-P. Another change is the material for the centering rings. The team decided that making the centering rings out of fiberglass would be better than aircraft plywood because of the added strength that fiberglass provides. The thickness of the fins was also changed from 1/4 to 1/8 in order to reduce weight while still maintaining structural integrity. Another change to the launch vehicle is the way the top section safely lands. Instead of the payload parachute deploying at separation, an electronic tether keeps the payload parachute bundled while also keeping the upper section from going ballistic by acting as a streamer. This change minimizes the drift distance of the upper airframe. In addition, the payload parachute will deploy at a separation between the upper airframe and the coupler instead of between the nose cone and upper airframe. This change was made to minimize the weight that the landing legs have to support Payload Changes The payload for the ISS Student Launch rocket has seen only minor modifications since the submission of the proposal. The image processing subsystem software plan has not changed since the proposal, the only hardware adjustment for this portion of the payload is a slight change in the downward facing angle of the camera to maximize the probability of sighting all of the tarps. The mechanical landing system has also changed only slightly since proposal. The bulkhead mounts that will secure the payload landing legs to the bulkhead have been modified to have bolts passing all of the way through the piece. This will allow the piece to have greater mechanical strength and removes the possibility of stripping the previously planned tapped hole while assembling and disassembling the bulkhead. The legs themselves have added a small ridge to the design that will prevent the outermost joint from bending backwards on landing Project Plan Changes Minimal changes have been made to the project plan since proposal. A full project timeline has been constructed to better structure team meetings and effort. Some parts have been sourced, which changes to the selected motor, in an effort to minimize the cost of the project. Illinois Space Day occurred on October 15 th, a major milestone towards completing the competition s educational outreach requirements. 2

12 3. VEHICLE CRITERIA 3.1. Launch Vehicle Overview The rocket designed to house the payload incorporates many design choices typical of a dual deploy high powered rocket. However, the team s design differs from a traditional dual deploy high powered rocket in one major way. In order to limit the load on the landing leg system, the rocket will physically separate at apogee into two sections. One section from the bottom of the rocket includes the booster tube and an avionics coupler. The other section from the top end of the rocket includes the payload coupler, some airframe tubing, and the nosecone. Following apogee, the bottom section follows the standard dual deploy model. After separation, an electronic tether keeps the main parachute bundle closed until it s deployment is desired. A drogue parachute deploys from the middle of the section two seconds after apogee to slow down this heavy section in preparation of main deployment. Once the booster tube reaches an altitude of 650 ft. AGL, the main parachute is untethered and fully deploys. From this position, the rocket can safely land, ready for recovery and reuse. Following apogee, the top section follows a drogue-less dual deploy with a single payload parachute which deploys two seconds following apogee. Another electronic tether keeps this parachute bundled while the bundle and this separation prevents the upper section from going ballistic. At 900 ft. the tether breaks, allowing for the parachute to fully deploy. This system slows the upper section enough so that the landing legs located on the bottom end of this section can allow the system to land upright once it reaches the ground. Figure 1 below illustrates the procession of events from launch to landing for the rocket design. Figure 1: Flight profile overview 3

13 Mission Statement The mission of the Illinois Space Society Student Launch Team is to safely launch and recover a reusable high power rocket simulating that satisfies the requirements provided by NASA as well as several of our own. In doing so the team will apply knowledge learned in the classroom. Additionally, participating in the competition facilitates further exploration of classroom topics for a practical purpose and provides an avenue for experienced members to mentor newer members while still allowing new members to make substantial contributions towards the success of the project. This experience will not be contained within the team, and frequent educational outreach events will allow the team to spread knowledge and enthusiasm for the aerospace industry to the greater community Success Criteria Table 1 below presents the requirements for the performance and design of the vehicle as set forth by either NASA, the method of verifying the satisfaction of these requirements, and the section of the report that addresses that requirement and how the current design fulfills it. Requirement The payload shall be delivered to an altitude of 5280 ft AGL. A commercially available altimeter shall record the official altitude. All recovery electronics shall be powered by commercially available batteries. The launch vehicle shall be recoverable and reusable. The launch vehicle shall have a maximum of 4 independent sections. The launch vehicle shall be limited to a single stage. Table 1: Vehicle and Recovery Competition Requirements Requirement Source Vehicle Requirements 1.1 Vehicle Requirements 1.2 Vehicle Requirements 1.3 Vehicle Requirements 1.4 Vehicle Requirements 1.5 Verification Method Analysis Rocket design and motor selection has been made around the mile high apogee requirement Demonstration Stratologger will report the official altitude. Demonstration Recovery system design requires commercially available batteries. Demonstration Recovery system has been designed to safely land the rocket. Inspection Rocket design has two independent sections. 4 Section Addressed Vehicle Requirements 1.6 Demonstration Rocket design requires a single stage. 3.2 The launch vehicle shall be Vehicle Test The team will ensure that

14 Requirement prepared for flight at launch site within 4 hours The launch vehicle shall be capable of remaining ready to launch on pad for 1 hour The launch vehicle shall be capable of being launched by a standard 12 volt DC firing system. The vehicle shall require no external circuitry to launch. The motor shall be commercially available. Pressure vessels on the vehicle shall adhere to certain criteria. Motor shall not have a total impulse greater than 5120 Ns. The vehicle shall have a minimum stability margin of 2.0 at rail exit. The vehicle shall exit the rail at a minimum velocity of 52 fps. All teams shall successfully launch a recover a subscale model of their rocket prior to CDR. All teams shall successfully launch a recover their fullscale rocket prior to FRR. Requirement Verification Method Source Requirements 1.7 the launch vehicle can be prepared for flight in less than four hours at the test launch. Vehicle Demonstration If the launch Requirements 1.8 is delayed, all systems (including batteries) can remain ready for at least 1 hour. Vehicle Requirements 1.9 Vehicle Requirements 1.10 Vehicle Requirements 1.11 Vehicle Requirements 1.12 Vehicle Requirements 1.13 Vehicle Requirements 1.14 Vehicle Requirements 1.15 Vehicle Requirements 1.16 Vehicle Requirements 1.17 Demonstration A standard 12 volt DC firing system will be used at the test launch. Inspection No parts of the rocket require any external circuitry for launch. Inspection The motor will be purchased from a commercial retailer. Inspection Pressure vessels will meet specific requirements. Inspection The purchased motor will have an impulse than the maximum. Analysis OpenRocket software shows that stability margin is at least 2.0. Analysis OpenRocket software shows that exit rail velocity is greater than the minimum. Demonstration The subscale rocket will be designed, built, and flown before CDR. Demonstration The full-scale rocket will be built and flown prior to FRR. 5 Section Addressed

15 Requirement Any structural protuberances on the rocket shall be located aft of the burnout center of gravity. The vehicle shall not utilize forward canards. The vehicle shall not utilize forward firing motors. The motor shall not expel titanium sponges. The vehicle shall not utilize hybrid motors. The vehicle shall not utilize a cluster of motors. The vehicle shall not utilize friction fitting for motors. The vehicle shall not exceed Mach 1 at any point during flight. Vehicle ballast shall not exceed 10% of total weight. Drogue event shall occur at apogee. Each team must perform ground ejection charge testing for all parachutes used in the system. Max kinetic energy of any independent section at Requirement Source Vehicle Requirements 1.18 Vehicle Requirements Vehicle Requirements Vehicle Requirements Vehicle Requirements Vehicle Requirements Vehicle Requirements Vehicle Requirements Vehicle Requirements Recovery System Requirements 2.1 Recovery System Requirements 2.2 Recovery System Requirements 2.3 Verification Method Inspection Rocket design has no structural protuberances above the burnout center of gravity. Inspection Rocket design has no forward canards. Inspection Rocket design has no forward firing motors. Inspection Motor does not expel titanium sponges. Inspection The motor is solid fuel engine. Inspection The motor uses a single motor. Inspection Friction fitting will not be used for motors. Analysis OpenRocket software shows that max velocity is under Mach 1. Analysis Any ballast used will be less than 10% of total weight. Demonstration Recovery system has drogue opening at separation event at apogee. Test The ejection charges will be tested on the ground before the flight of the rocket. Analysis Calculations show that each independent section Section Addressed

16 Requirement landing shall not exceed 75 ft-lbf. The recovery system shall be electrically independent of any payload circuits. Recovery system shall contain redundant altimeters. Parachute deployment shall not use motor ejection. Each altimeter shall be armed by a dedicated switch on exterior of rocket airframe. Each altimeters shall have a dedicated power supply. Each arming switch shall be capable of being locked in the ON position for launch. Removable shear pins shall be used for main and drogue parachute compartments. Launch vehicle shall be trackable during and after flight. Recovery system shall not suffer from any interference from other components in vehicle. Requirement Source Recovery System Requirements 2.4 Recovery System Requirements 2.5 Recovery System Requirements 2.6 Recovery System Requirements 2.7 Recovery System Requirements 2.8 Recovery System Requirements 2.9 Recovery System Requirements 2.10 Recovery System Requirements 2.11 Recovery System Requirements 2.12 Verification Method has less than the maximum kinetic energy. Inspection Recovery system is designed to not involve payload circuits. Inspection Recovery system has redundant altimeters to ensure deployment. Test Parachute deployment will occur due to black powder charges. Inspection Switch band will have a switch dedicated to each altimeter. Inspection There will be one power supply per altimeter. Demonstration The arming switch will be locked in the ON position for the test flight. Inspection Shear pins for the main and drogue parachute compartments will be removable. Demonstration The rocket will be tracked using the TeleMetrum. Demonstration Rocket design ensures that recovery system is kept safe from any interference from other components. Section Addressed Table 2 below presents the requirements for the performance and design of the vehicle as set forth by the ISS Student Launch team, the method of verifying the satisfaction of these requirements, and the section of the report that addresses that requirement and how the current 7

17 design fulfills it. The team has made it a focus to design a launch vehicle with reusability as a focus. Table 2: Vehicle and Recovery Team-Derived Requirements Requirement Team should promote sustainable involvement in the project by reusing as many RSO owned parts as possible. Couplers should have items placed in a way to minimize clutter from unnecessarily long wires. Subscale vehicle will be designed in a way that allows the team to approximate a Cd for the full scale vehicle. Drift should be minimized to allow the payload to process images on the way down as well as to time spent on recovery efforts. Rocket should look aesthetically pleasing to better the team s presence at community and member recruitment events. Verification Method Analysis Try to optimize design around expensive owned parts such as parachutes and motor casings. Demonstration - Altimeters and Raspberry Pi Zeros are placed in a way to minimize distance to rotary switches and batteries. Analysis Subscale will have the same outer airframe material as the full scale vehicle and the same paint that will be used for the full scale. Analysis Combination of Jolly Logic Chute Releases allow the team to minimize drift without failing the competition kinetic energy requirement. Inspection Team will collaborate on design for rocket and implement it for the full scale vehicle. Section Addressed N/A N?A Mass Statement Table 3 below presents the estimated mass breakdown of the total system as of this milestone. In order to more clearly present the data, the mass statement has been broken down by subsystems. A complete mass total, including all of these subsystems, is presented at the bottom of the table. Unknown masses, such as the completed payload system or epoxy totals are given built-in margins in anticipation of future mass growth. The team has accurately estimated the mass of its system for the last few years of the competition. Adding significant excess margin mass can affect the performance of the rocket once a design is constructed. Years of experience and access to prior year s mass estimates give the team great confidence in the following mass report. Should any systems change by the next design review, this statement will be revised and motor selection and structural design will be fine-tuned to optimize performance. 8

18 Table 3: Projected Mass Statement Item Total Mass [lb] Use Structure: Booster Tube 5.00 Outer Airframe for Booster Section Avionics Switch Band 0.34 Switch Band for Avionics Bay Leg Insert/ Main Parachute Bay Tubing 2.20 Leg Insert/ Main Parachute Bay Upper Airframe 2.52 Houses Payload Parachute Trapezoidal Fins (3) 2.35 Fins Motor Mount Tube 0.71 Motor Mount Tube Centering Rings (3) 1.05 Centering Rings Epoxy and Resin 0.25 Structural Joints Aeropack Motor Retainer 0.25 Motor Retainer Avionics Coupler 1.85 Avionics Coupler Tubing 6 Coupler Bulkhead (4) 0.64 Bulkhead for Coupler Tubing 6 Airframe Bulkhead (4) 0.60 Bulkhead for Airframe Tubing Fiberglass Nosecone 1.78 Nosecone Nuts, Bolts, Washers, and Screws 0.10 Connections 1515 Rail Button (2) 0.05 Connection to Launch Rail Margin 0.20 Future Growth Structures Total Mass: Recovery Equipment Stratologger CF (2) 0.05 Altimeter Telemetrum 2.0 (2) 0.05 Altimeter/Tracker 9V Battery (2) 0.20 Battery for Stratologger 9V Battery Clip (2) 0.03 Attach 9V Batteries to Sleds Telemetrum Li-Po Battery (2) 0.05 Battery for Telemetrum Payload Parachute Parachute for Payload Section Main Parachute Main Parachute Drogue Parachute Drogue Parachute 20 ft Tubular Kevlar (3) Shock Cord for all Parachutes Jolly Logic (2) 0.30 Tether for Main Parachute Quick Links (4) 1.00 Attachment Hardware Charge Cups (6) 0.10 Black Powder Container Nylon Shear Pins 0.01 Shear Pins ¼ Threaded Rods (6) 0.20 Mounts for Sleds 1/8 Plywood Sheet 0.10 Sleds for Mounting Equipment Terminal Blocks (4) 0.05 Connect Altimeters to Black Powder Charges Rotary Switches (4) 0.10 For Activating Altimeters on the Pad 9

19 Item Total Mass [lb] Use Margin 0.25 Further Growth Recovery Equipment Total Mass: 4.00 Motor Equipment L1390G-P Reload Kit 4.35 Motor Fuel Grain RMS 75/3840 Motor Casing 2.40 Motor Casing 75mm Forward Closure 1.00 Closure for Motor Casing 75mm Aft Closure 0.80 Closure for Motor Casing Motor Equipment Total Mass: 8.55 Payload (Camera System) Raspberry Pi Zero (2) Image Analysis Raspberry Pi Camera V2 (2) Image Capture Raspberry Pi Zero Camera Cable 0.02 Image Analysis (2) LiPo 12V 500mAh Battery (2) 0.22 Power to Camera System DC-DC Converter (2) 0.01 Use LiPo with Raspberry Pi Altimeter Module MS5607 (2) Image Analysis Tool Acrylic Cutout (2) 0.80 Allows Cameras to See 3D Printed Camera Mount (2) 0.10 Angle Cameras for Better FOV Wooden Sled (2) 0.15 House Equipment Angle Bracket (2) 0.10 Connect sleds to threaded rods ¼ Threaded Rods (4) 0.20 Mounts for Sleds Nuts, Bolts, Washers, and Screws 0.05 Connections Margin 0.17 Further Growth Camera System Total Mass: 1.90 Payload (Landing System) Inner Leg Segment (4) 0.42 Inner Landing Legs Outer Leg Segment (4) 0.28 Outer Landing Legs Bulkhead Leg Hinge (4) 0.05 Connects Legs to Bulkhead Nuts, Bolts, Washers, and Screws 0.15 Connections Torsion Spring ( in*lb.) (4) 0.02 Torsion Spring Torsion Spring (-4.5 in*lb.) (4) 0.02 Torsion Spring Margin 0.16 Further Growth Landing System Total Mass: 1.10 Total Mass of System:

20 System Modeling The proposed system as described in this report was modeled in NX 10 as a tool to present the design. Figure 2 below shows the completed system as it would sit on the launch rail and color coded based on the four independent sections that will land. Outer airframe tubing is colorcoded dark blue with separating switchband tubing color-coded light blue to differentiate the various sections. Figure 2: Full rocket assembly modeled in NX

21 A drawing showing some general dimensions of the design is presented below as Figure 3. The rocket has a total height of approximately 10 feet and a diameter of roughly 6 inches. Figure 3: Dimensioned drawing of launch vehicle Booster Subsystem The booster subsystem is responsible for providing thrust and basic guidance to the rocket. Without a structurally sound booster subsystem, the motor and by extension rocket cannot perform to their intended potential. The booster system contains the motor, motor casing, motor mount, motor retention, outer airframe, and fins, all integrated onto the rocket s launch rail. The mount and retainer are fitted to contain the motor properly within the outer airframe. The fins attached to the outer airframe are the only independent flight control system on the rocket. Integrated onto the launch rail, these systems allow for the force of the motor to provide a guided flight to the rocket. A model of the booster subsystem is included below as Figure 4. An inward view of the booster subsystem is visualized in Figure 5. 12

22 Figure 4: Booster subsection modeled in NX. Figure 5: Inward view of booster subsytem Motor Selection and Justification The motor the team has decided to utilize is the AeroTech L1390G-P. As per requirements, this motor is not a hybrid motor and the rocket will not employ a cluster of these motors. Additionally, the design is not utilizing any forward firing motors, and the motor selected does not expel titanium sponges. This motor is also able to be ignited using a 12-volt firing system. For high powered rocketry, a thrust to weight ratio greater than or equal 5 is generally considered to be safe. Using OpenRocket software as well as the characteristics of the chosen motor, the rocket will have a thrust to weight ratio of 8.8. This falls within the average range of vehicles of this size and as a result, the vehicle should leave the launch rail at a safe and stable velocity. In order to ensure a stable release from an 8-foot-long 1515 launch rail and fulfill competition requirements, the rocket will have to have an off-rail velocity greater than 52 ft/s. 13

23 Achieving this velocity will provide sufficient airflow over the rocket s fins to provide a correcting force to further stabilize the rocket. Based on the simulations provided by OpenRocket software, the off-rail velocity of the rocket is projected to be around 60.9 ft/s. This is greater than the 52 ft/s required velocity by a sizeable margin. Additional information regarding the specifications of the motor the team has chosen can be found at thrustcurve.org. As shown in the thrust curve provided below, the motor quickly achieves its average thrust of lbf in roughly 0.1 seconds. This quick spike in the thrust curve results in a high off-rail velocity that the current design achieves in spite of its weight. Further analysis shows that the motor continues to provide its average thrust of lbf for a total burn time of around 2.6 seconds. The total impulse that is generated by the L1390G-P motor is lbf*s. Figure 6: L1390G-P thrust curve. Based on the specifications in Figure 6 above, the main reason the team chose to use the AeroTech L1390G-P motor is to efficiently launch our rocket and payload into a safe, stable trajectory. Choosing a larger class rocket motor could unnecessarily send our rocket above the required altitude, while a less powerful motor might not provide enough thrust to safely stabilize the rocket off of the launch rail. The L1390G-P motor will allow the launch vehicle to fulfill all competition require the overall flight of our rocket. Table 4 below summarizes the characteristics and performance of the system with an L1390G-P 14

24 Table 4: L1390G-P Performance Characteristics Characteristic Value Motor Diameter 75 mm Max Thrust lbf Average Thrust lbf Total Impulse lbf*s Burn Time 2.6 s Exit Rail Velocity 60.9 ft/s Thrust to Weight Ratio Motor Casing In high powered rocketry, the motor case is a crucial component of the rocket. Essentially, the motor case is a tube that holds the rocket propellant. This container must be very strong in order to control the hot gases that are generated from the rocket motor. The casing the team will implement to house our L1390G-P motor is the RMS 75/3840. This aluminum casing provides a proper environment for the propellant to burn while also being strong enough to withstand the high temperatures and force loads. In accordance with the manufacturer, the RMS 75/3840 motor casing is the recommend and only choice in order to optimize the performance of the motor the team has selected. As a result, it is believed that the combination implemented will be efficient. The motor casing will also provide a secure, lower attachment point for the drogue parachute shock cord. A dimensioned drawing of the motor casing, provided by the manufacturer, is presented below in Figure 7. 15

25 Figure 7: RMS 75/3840 motor casing Motor Mount Assembly Housing the motor casing is a motor mount tube constructed of fiberglass that will hold the engine during launch without using adhesives on the motor casing. Fiberglass was chosen as a precaution to possible excess heat from the motor casing. A series of three fiberglass centering rings will attach the motor mount tube to the larger outer airframe integrating the two systems together while keeping the thrust pointing parallel with the rocket. Fiberglass was chosen because of its high strength to keep the motor mount positioned correctly. A model of the motor mount tube is included below as Figure 8. 16

26 Figure 8. Motor Mount Tube Motor Retention An Aero Pack 75 mm motor retainer was chosen for motor retention. The team chose Aero Pack Incorporated because they make reliable, quality motor retainers. The purpose of a motor retainer is to secure the motor inside the motor mount tube of the rocket, ensuring that the motor is locked in place and will not fall out once the rocket is vertical. This is a high strength aluminum component used to prevent the motor from shifting its position forward or aft during flight. The retainer consists of two pieces: a body and a screw on cap. The body of the retainer is permanently fixed to the lowest centering ring. After the motor case is slid into the rocket, the retainer cap securely threads on to the body of the retainer. This prevents the motor from inadvertently moving during flight and also provides a quick method of loading and removing the motor casing. Figure 9: Aeropack 75mm flanged motor retainer. 17

27 Outer Airframe The team debated between many different options of materials to use for the airframe. Several potential materials were discussed for use based on prior team experience and research into popular options within the high powered rocketry community. After discussion, airframe material was quickly narrowed down to three options: fiberglass, Blue Tube, and carbon fiber. Materials such as aluminum and other heavy metals were quickly ruled out due to large of manufacturing challenges and an increased cost that wouldn t translate into a noticeable performance benefit. To decide between the three remaining materials, pros and cons for each material had to be weighed against each other. Blue Tube is slightly less dense than fiberglass and would save roughly three pounds of weight, but it would not provide the same level of structural integrity as fiberglass would be able to provide. Fiberglass provides the payload team with a greater amount of manufacturing flexibility, being easier to make large, irregular cuts into the fiberglass for insertion of the payload components. Each material was later assessed in a trade study and evaluated for its respective advantages and disadvantages as seen in Table 5 below. A score of 5 represents the best possible score in a category, while 1 represents the poorest possible score in a category. While fiberglass was the material that was agreed upon by the team due to the higher level of strength it provides in relation to Blue Tube, Blue Tube offers a viable alternative should fiberglass prove to be ineffective for any reason. Carbon fiber was ruled out as an alternative due to its higher cost. Table 5: Airframe Material Trade Study Figure of Merit Blue Tube 2.0 Fiberglass Carbon fiber Cost: Strength: Ease of Manufacturing/Access: Fins In order to ensure that the rocket is stable and maintains its orientation and intended flight path, three fins will be attached to the body tube. These fins are each located 120 degrees apart in a symmetric fashion. The team chose to use trapezoidal fins because they provide aerodynamic stability while also not adding significant amounts of drag. The team decided to use fins with a thickness of 1/8 " as opposed to 1/4 " order to minimize weight. Thinner than 1/8 fins will not be used in order to maintain the structural integrity needed to ensure that the fins are strong enough to provide support during flight and survive impact with the ground upon landing. Full flutter analysis will be done for CDR to further optimize the thickness of the fins. The material to be used for the fins was decided upon by the team to be fiberglass, although other viable options were present. Fiberglass fins offer a large amount of strength and 18

28 stability, with minimal risk of shearing off of the body tube or of cracking, and were also considered heavily because of the team s choice to use fiberglass for the body tube. Aircraft plywood was discussed as well, and would be a good option due to the low cost and the ease of manufacturability. However, the weight of the rocket as a whole demands fins that can withstand the brunt force of a direct impact. This is something aircraft plywood simply can t provide, even if the plywood structure is supplemented with a composite coating for additional strength. Carbon fiber was also discussed, but as was the case with the body tube, the high cost diverted the team s attention from it. A trade study of the top fin materials can be found in Table 6. A score of 5 represents the best possible score in a category, while 1 represents the poorest possible score in a category. Table 6: Fin Material Selection Material: Fiberglass Aircraft Plywood Carbon Fiber Cost: Strength: Ease of Manufacturing: The fiberglass fin set is dimensioned below in Figure 10. Figure 10: Dimensioned trapezoidal fin 19

29 Launch Rail Integration The purpose of integrating rail buttons is to help guide the rocket on its initial ascent until sufficient speed is developed for the fins to further stabilize the vehicle. The rail buttons also prevent the rocket from swaying prior to launch in the case of breezy conditions. As previously stated in the Motor Selection and Justification section, the motor the team has chosen in order to achieve exit rail velocity greater than 52 ft/s is the L1390G-P. Through the use of this motor and the integration of the rail buttons our rocket will safely launch and release from the 8 ft 1515 rail. Each rail button will be attached to a mounting point secured to one of the vehicles centering rings. This mounting point will consist of a plywood block with a T-nut. This allows the rail buttons to easily screw in and out in case one needs to be replaced, but it also provides for a secure mounting configuration. Additionally, this method mitigates any damage to the structural integrity of the relatively thin centering rings. A close up of one of the rail buttons can be seen below in Figure 11. The mounting hardware can be seen below in Figure 12. Figure 11: Close up of rail button. Figure 12: Rail button mounting hardware. 20

30 Booster Subsystem Mass Statement Table 7 below presents the mass statement of the booster subsystem as a whole. Mass data has been taken from manufacturer documentation or from measurements taken by the team in prior years. A small margin has been added but in general all uncertain values such as the amount of epoxy are generous to allow for future growth. The team avoids applying a large blanket mass margins due to issues in prior year where underweight rockets have resulted in higher than desired altitudes. Table 7: Booster Subsystem Mass Statement Item Total Mass [lb] Use Structure: Booster Tube 5.00 Outer Airframe for Booster Section Trapezoidal Fins (3) 2.35 Fins Motor Mount Tube 0.71 Motor Mount Tube Centering Rings (3) 1.05 Centering Rings Epoxy and Resin 0.25 Structural Joints Aeropack Motor Retainer 0.25 Motor Retainer Nuts, Bolts, Washers, and Screws 0.10 Connections 1515 Rail Button (2) 0.05 Connection to Launch Rail Margin 0.20 Future Growth Structures Total Mass: 9.96 Motor Equipment L1390G-P Reload Kit 4.35 Motor Fuel Grain RMS 75/3840 Motor Casing 2.40 Motor Casing 75mm Forward Closure 1.00 Closure for Motor Casing 75mm Aft Closure 0.80 Closure for Motor Casing Motor Equipment Total Mass: 8.55 Total Mass of System:

31 3.3. Avionics Coupler Subsystem The avionics coupler contains all of the hardware required to track and safely recover the lower separation section of the launch vehicle. It s also responsible for serving as the separation point for the launch vehicle, with large shoulders ensuring that the vehicle stays as one piece until desired separation occurs at apogee. A model of the avionics coupler subsystem is included below as Figure 13. Figure 13: Model of avionics coupler subsytem Avionics Coupler Tubing The couple section s shoulder and switch band will be made of fiberglass. Fiberglass was chosen to keep the outer diameter of the vehicle constant. The fiberglass will offer more than enough strength for the coupler during normal flight conditions. The couple will be 15.5 inches long. 3.5 inches of this length will be the switch band leaving 12 inches for the shoulder lengths. This gives the coupler 6 inches on each side of the switch band. The shoulder length was chosen based on a standard for high-power rocket. One caliber for the shoulder lengths is usually considered long enough for a stable connection between the coupler and airframe tubing. The switch band length was chosen to ensure that there was enough room for two altimeters, wiring, and batteries. The switch band must also be long enough for a switch to be installed to turn on the altimeters Avionics Sled Configuration All avionics hardware will be mounted to 1/8 plywood sleds that are then mounted by 3D printed L-channel stock to aluminum rods and the bulkheads, where the aluminum rods hold the two bulkheads together, to ensure that the fragile and valuable altimeters will stay in place and safe throughout the duration of flight. The L-channel stock will be made longer than in previous years, so that it can better adhere to the bulkheads with screws. In previous years two shorter pieces of 3D printed L-channel stock and glue were used, which came off during flight. 3D 22

32 printed L-channel stock was chosen over metal as it adds unnecessary extra weight and the 3D printed L-channel has proven to be strong enough in the past. Figure 14 below provides a model of both sides of the lower avionics sled. The StratoLogger is colored green, the StratoLogger magenta, the 9V is colored brown, and the 4V LiPo powering the TeleMetrum is colored silver. Note that while the placement of most of the hardware is somewhat arbitrary, it is required for the proper function of the GPS function for the TeleMemtrum to be orientated in the way that it is (parallel with the roll axis of the rocket). Figure 14: Both sides of the lower avionics sled. The size of the avionics coupler is limited insomuch as a shoulder length of six inches is required on both ends to ensure proper stability for the connections between the coupler and airframes. Consequently, the space inside the coupler is relatively underutilized, a fact which may drive design in the future Bulkhead Configuration The bulkhead must be able to withstand high pressures caused by the black powder discharge and wind pressure. The material must also be amenable with screws, U-bolts, and other pieces of hardware. The choice of materials for the bulkhead are either fiberglass, aircraft plywood, or carbon fiber. The trade study for fin material selection is also applicable for the selection of the bulkhead material. Aircraft Plywood was chosen as the material because of its strength, lightness, and durability. Attached to the bulkheads are the black powder charges and E-matches which are taped onto black powder caps. Instead of using tape and unnecessary components, the black powder charges can be attached to the bulkhead with putty. Black tack is similar to blue tack but stronger 23

33 and stickier. Putty offers greater strength and stability compared to tape and allows for flexible placement on the bulkhead. Comparing U-bolts and Eye bolts, there is no significant advantage between the two. The U-bolt offers a greater distribution of force, decreasing the chance of a single point failure. However, the U-bolt takes more space and an Eye bolt. An Eye bolt on the other hand is smaller and therefore better for situations where space is limited. However, the Eye bolt must be welded closed and has a higher chance for a single point failure Avionics Coupler Subsystem Mass Statement Table 8 below presents the mass statement of the avionics coupler subsystem as a whole. Mass data has been taken from manufacturer documentation or from measurements taken by the team in prior years. A small margin has been added. The team avoids applying large blanket mass margins due to issues in prior year where underweight rockets have resulted in higher than desired altitudes. Table 8: Avionics Coupler Subsystem Mass Statement Item Total Mass [lb] Use Structure: Avionics Switch Band 0.34 Switch Band for Avionics Bay Avionics Coupler 1.85 Avionics Coupler Tubing 6 Coupler Bulkhead (2) 0.32 Bulkhead for Coupler Tubing 6 Airframe Bulkhead (2) 0.30 Bulkhead for Airframe Tubing Nuts, Bolts, Washers, and Screws 0.10 Connections 1515 Rail Button (2) 0.05 Connection to Launch Rail Margin 0.20 Future Growth Structures Total Mass: 3.16 Recovery Equipment Stratologger CF (1) 0.03 Altimeter Telemetrum 2.0 (1) 0.03 Altimeter/Tracker 9V Battery (1) 0.10 Battery for Stratologger 9V Battery Clip (1) 0.02 Attach 9V Batteries to Sleds Telemetrum Li-Po Battery (1) 0.03 Battery for Telemetrum Main Parachute Main Parachute Drogue Parachute Drogue Parachute 20 ft Tubular Kevlar (1) 0.21 Shock Cord for all Parachutes Jolly Logic (2) 0.30 Tether for Main Parachute Quick Links (3) 0.75 Attachment Hardware Nylon Shear Pins 0.01 Shear Pins ¼ Threaded Rods (2) 0.07 Mounts for Sleds 24

34 1/8 Plywood Sheet 0.10 Sleds for Mounting Equipment Terminal Blocks (4) 0.05 Connect Altimeters to Black Powder Charges Rotary Switches (2) 0.05 For Activating Altimeters on the Pad Margin 0.25 Further Growth Recovery Equipment Total Mass: 2.78 Total Mass of System: Payload Bay and Upper Airframe Subsystems The payload bay and upper airframe subsystems are responsible for the safely deploying and recovering the team s payload that will carry out the target detection and vertical landing experiment. A model of this subsystem is included below as Figure 15. Figure 15: Payload bay and upper airframe subsytem Payload Sled Configuration The payload sleds in the upper section of the aircraft also house a couple altimeters to control the deployment of the payload parachute and are pictured below is. The attachment for the payload electronic sleds is the same as the avionics sleds. The sleds have an L channel attached to them and attached to the bulkheads. Going through the L-channels and connect the bulkheads are four aluminum rods (2 for each sled). The placement of items on these sleds were decided on more deliberately due to the increased amount of hardware and decreased amount of space when compared to the avionics 25

35 sleds. It was decided that the TeleMetrum needed its own side in order to have sufficient space for the antenna. It also needed to placed parallel with the roll axis of the rocket for proper functioning of the GPS function. The cameras, as seen in the picture in pink and green, were positioned such that they are able to look through the windows of the rocket to see the outside. Lastly, the raspberry pi was placed as close to the camera as possible to minimize the amount of wires and keep the sleds as tidy as possible. Just like the raspberry pi s, the batteries were placed close to the raspberry pi to minimize wiring. Since proposal, two altimeters that will interface with the Raspberry Pi Zero to aid in analysis have been added and placed on the sled in a convenient location. Figure 16: Payload sled one front and back. 26

36 Figure 17: Payload sled 2 front and back Payload Bulkhead Configuration Two bulkheads are being used in order to create a tighter seal inside the rocket. By gluing together two bulkheads that fit into the main airframe and the coupler individually, the cap will prevent the wind from disrupting the altimeter and other instruments inside the payload. The material for the bulkhead will be aircraft plywood due to its strength, lightness, and flexibility. Other materials were researched for the bulkhead, focusing on fiberglass and carbon fiber. The strengths of fiberglass are its durability and uniformity with the airframe material, whereas the strength of carbon fiber is mainly its impressive toughness. However, the differences in strength were outweighed by the plywood s economic price tag and manufacturing ease. The placement of items on the bulkhead would follow previous year designs with the exception of the black powder charges. In the past, the charges have taped onto PVC end caps that serve no other purpose. Instead of using tape and unnecessary components, the black powder charges can be attached to the bulkhead with putty. Black tack is similar to blue tack but stronger and stickier. Putty offers greater strength and stability compared to tape and allows for flexible placement on the bulkhead while preventing the black powder charges from moving if a stray quick link hits it. Comparing u bolts and eyebolts, both would provide sufficient strength for the recovery system. The U-bolt offers a greater distribution of force, decreasing the chance of a single point failure. However, the U-bolt takes more space and an eyebolt. An eyebolt on the other hand is smaller and therefore better for situations where space is limited. However, the eyebolt must be welded closed and has a higher chance for a single point failure. Because space on bulkheads of 27

37 this size is not an issue, the recovery team has opted to utilize U-bolts due their superior durability and strength over even forged steel eyebolts Nosecone Due to lack of options at this rocket diameter, the nose cone is pre-bought and made of fiberglass. Fiberglass is favorable because it can be molded precisely, combining high strength, smoothness and ease of finishing. This combination of material characteristics allows the nosecone to operate under conditions significantly more intense than demanded by this challenge. A model of the nosecone is provide below as Figure 18. Figure 18: Model of nosecone. A dimensional drawing of the nosecone is included below as Figure

38 Figure 19: Nosecone dimensioned in inches Payload Bay and Upper Airframe Subsystems Mass Statement Table 9 below presents the mass statement of the booster subsystem as a whole. Mass data has been taken from manufacturer documentation or from measurements taken by the team in prior years. A small margin has been added but in general all uncertain values such as the amount of epoxy are generous to allow for future growth. The team avoids applying a large blanket mass margins due to issues in prior year where underweight rockets have resulted in higher than desired altitudes. Table 9: Payload Bay and Upper Airframe Mass Satement Item Total Mass [lb] Use Structure: Upper Airframe 2.52 Houses Payload Parachute 6 Coupler Bulkhead (2) 0.64 Bulkhead for Coupler Tubing 6 Airframe Bulkhead (2) 0.60 Bulkhead for Airframe Tubing Fiberglass Nosecone 1.78 Nosecone Nuts, Bolts, Washers, and Screws 0.10 Connections Margin 0.20 Future Growth Structures Total Mass: 5.84 Recovery Equipment Stratologger CF (1) 0.03 Altimeter 29

39 Item Total Mass [lb] Use Telemetrum 2.0 (1) 0.03 Altimeter/Tracker 9V Battery (1) 0.10 Battery for Stratologger 9V Battery Clip (1) 0.02 Attach 9V Batteries to Sleds Telemetrum Li-Po Battery (1) 0.03 Battery for Telemetrum Main Parachute Main Parachute Drogue Parachute Drogue Parachute 20 ft Tubular Kevlar (1) 0.21 Shock Cord for all Parachutes Jolly Logic (2) 0.30 Tether for Main Parachute Quick Links (3) 0.75 Attachment Hardware Nylon Shear Pins 0.01 Shear Pins ¼ Threaded Rods (2) 0.07 Mounts for Sleds 1/8 Plywood Sheet 0.10 Sleds for Mounting Equipment Terminal Blocks (4) 0.05 Connect Altimeters to Black Powder Charges Rotary Switches (2) 0.05 For Activating Altimeters on the Pad Margin 0.25 Further Growth Recovery Equipment Total Mass: 2.7 Payload (Camera System) Raspberry Pi Zero (2) Image Analysis Raspberry Pi Camera V2 (2) Image Capture Raspberry Pi Zero Camera Cable 0.02 Image Analysis (2) LiPo 12V 500mAh Battery (2) 0.22 Power to Camera System DC-DC Converter (2) 0.01 Use LiPo with Raspberry Pi Altimeter Module MS5607 (2) Image Analysis Tool Acrylic Cutout (2) 0.80 Allows Cameras to See 3D Printed Camera Mount (2) 0.10 Angle Cameras for Better FOV Wooden Sled (2) 0.15 House Equipment Angle Bracket (2) 0.10 Connect sleds to threaded rods ¼ Threaded Rods (4) 0.20 Mounts for Sleds Nuts, Bolts, Washers, and Screws 0.05 Connections Margin 0.17 Further Growth Camera System Total Mass: 1.90 Payload (Landing System) Inner Leg Segment (4) 0.42 Inner Landing Legs Outer Leg Segment (4) 0.28 Outer Landing Legs Bulkhead Leg Hinge (4) 0.05 Connects Legs to Bulkhead Nuts, Bolts, Washers, and Screws 0.15 Connections Torsion Spring ( in*lb.) (4) 0.02 Torsion Spring 30

40 Item Total Mass [lb] Use Torsion Spring (-4.5 in*lb.) (4) 0.02 Torsion Spring Margin 0.16 Further Growth Landing System Total Mass: 1.10 Total Mass of System: Recovery Subsystem The Recovery Subsystem is an essential part of the rocket. The Recovery subsystem ensures that the rocket lands safely at a reasonable speed at a safe distance from the launch site and the crowd of people. The team will use three total parachutes for this rocket, one referred to as a main parachute, one a payload parachute, and one a drogue parachute. The main will be an Iris Ultra 96, owned and already successfully used by the team in previous years. The drogue will be a standard 15 elliptical parachute purchased from Fruity Chutes. Finally, the payload parachute will be a SkyAngle Model C2 from b2 Rocketry that the team owns and has successfully deployed in prior years Deployment Scheme At apogee, the rocket will split into two main parts via black powder ejection charges. One section from the bottom of the rocket includes the booster tube and an avionics coupler and is modeled below as. The other section from the top end of the rocket includes the payload coupler, some airframe tubing, and the nosecone. Two Jolly Logic Chute Releases prevent the now free main parachute from fully deploying. Two seconds after apogee, the drogue parachute deploys from the lower separation section and the payload parachute deploys from the upper separation section. Another two Jolly Logics prevent the payload parachute from fully deploying until 900 ft AGL. Finally, the Jolly Logics on the main parachute untether around 700 ft and allow the main parachute to fully deploy around 650 ft. The four sections of the rocket are presented in the figures below. Figure 20: Section 1 -- avionics coupler. 31

41 Figure 21: Section 2 -- booster tube. Figure 22: Section 3 -- Payload Coupler 32

42 Figure 23: Section 4 -- Upper Airframe and Nosecone Deployment Mechanisms The motor in use does not contain an ejection charge, so black powder will be used. After apogee, each parachute will be deployed at the desired time. The black powder charges will be stored in canisters, ready for deployment upon receiving correlating electrical information. Black powder is used because the charge can be varied based upon ejection charge testing specific to the rocket. The testing for this quantity will be performed in the same configuration as actual in flight deployment. The charge size will be based upon the governing equations: = = The black powder is ignited by e-matches. These e-matches will be connected to altimeters. The altimeter will send current from its battery to the e-matches, and in doing so light the black powder charges. A Jolly Logic will provide additional timing control of the parachute deployment process Avionics Hardware Four altimeters will be used with two Chute Releases for the deployment of the parachutes. Two altimeters will be used on the upper airframe and two will be used on the booster section of the vehicle. The booster section and the upper airframe will each have one of PerfectFlite s StratoLoggerCF and one of Altus Metrum s TeleMetrum 2.0. Both of these 33

43 altimeters are commercially available and use a barometer to measure altitude. The StratoLogger will be used as the primary altimeter with the TeleMetrum acting as the secondary altimeter. While each of these altimeters are able to have two deployment events, two altimeters are needed for both sections to have a fully redundant system. Having a fully redundant system is desired for something that is crucial for safety. These altimeters will send current from their respective batteries to e-matches which will be connected to black powder charges. These charges will be used for the recovery events TeleMetrum The second altimeter that will be used in this system is Altus Metrum s TeleMetrum. This altimeter is powered by a lithium ion battery that can be obtained commercially. The TeleMetrum is similar to the StratoLogger in that it can be connected to a switch, emits beeps for different flight parameters, and is capable of dual deployment events. The switch will allow the TeleMetrum to be turned on at the launch pad and off before the competition altimeters altitude is reported to the NASA official. Each altimeter will have its own rotary switch so they can be turned on and off independently. Unlike the StratoLogger, the TeleMetrum has GPS capabilities. The TeleMetrum has an on-board integrated GPS receiver which will allow it to transmit its coordinates to the ground station s dongle in real time. Members of the team have experience using the TeleMetrum, giving the team confidence for its use in this competition. Two TeleMetrums will be used to track the booster section and the upper airframe as well as providing redundancy to the deployment of the parachutes. Altus Metrum offers a software called AltOS. This software allows the team to both get a live feed of the rocket s GPS coordinates and export flight profile data for later analysis. Figure 25 shows a graph generated by AltOS and from data collected using a TeleMetrum. For the TeleMetrum to acquire a GPS lock, the antenna must be oriented upright inside the vehicle. The payload sled was designed so the antenna is oriented upright. AltOS allows the TeleMetrum to communicate to a ground station via a dongle and antenna setup. The dongle interfaces with a computer via USB port. Before launches, a laptop will be connected to the dongle and ground station antenna. The ground station will then acquire a connection to the altimeter. The ground station will also ensure that the TeleMetrum has GPS lock before the vehicle is launched. Figure 26 shows last year s team s ground track. This was updated in real time during the flight of the vehicle and allowed the team to track the vehicle s drift. Figure 24. The AltOS main menu. 34

44 Figure 25. Example of a graph that AltOS can generate from flight data. Figure 26. AltOS s ground tracking of a vehicle. 35

45 Since the vehicle separates into two sections, the vehicle will use two TeleMetrums for tracking. Since the team has never used to TeleMetrums on the same vehicle, it was necessary for the team to conduct research into how two of these altimeters will interact with each other. Research conducted examined the possibility of interfering radio waves from the two altimeters being in such close proximity to each other. It was determined that despite their close proximity, the altimeters signals would not interfere with each other provided that they are operating on different frequencies. This finding was based on the altimeters usage of radio waves. Radio waves, in actuality do interfere but this interference is negligible as seen with the transmission of radio waves throughout the country. The ability of receivers to tune to a specific frequency allows for the reduction of any significant interference being transmitted during the process. This means the altimeters will not interfere with each other but they do require the use two different antennas to be able to receive the telemetry at each distinct frequency. Two laptops will be used on the ground station. One dedicated to each one of the TeleMetrums Stratologger The StratoLogger altimeter is powered by a commercial 9-volt battery. The altimeter is capable of being connected to a switch, allowing it to be turned on from the outside of the vehicle before launch. The StratoLogger has a small speaker that communicates different parameters about the flight via a series of beeps before launch. These beeps will tell the team when the parachutes will be deployed and that there is continuity to the e-matches. Continuity is important to check since it will ensure the e-matches are lit and therefor allow the vehicle to be recovered safely. After the vehicle has landed, the altimeter will emit a new series of beeps which reports the altitude of the last flight. This will allow the altitude to be recorded by the NASA official on launch day. The StratoLogger is also resistant to loss of power. The altimeter will stay on for a full two seconds without connection to the battery. This resistance adds security to the altimeter s data collection without any additional complexity. The StratoLogger was also chosen because members of the student launch team have used StratoLoggers successfully on other high power rockets. The PerfectFlite allows the team to pull flight profile, altitude, and velocity from the altimeter following launch. PerfectFlite offers a software called DataCap to export the flight data off of the StratoLogger. DataCap can be used to graph the data too. An example a flight profile can be seen graphed in DataCap below in Figure 27. The altitude is graphed in blue and the velocity is graphed in red. DataCap can also be used to perform some statistics on the data. The data can be passed through filters to make it appear smoother. The max altitude and velocity can also be displayed easily using DataCap. 36

46 Figure 27. PerfectFlite DataCap software for StratoLogger Jolly Logic Chute Release Four Jolly Logic Chute Releases will be used in the avionics system and payload system of the vehicle. These Chute Releases will allow the main parachute to be ejected at apogee but not deploy until the desired altitude and will allow the payload parachute to act as a stream after apogee. The Chute Release is an altimeter with a pin and lock mechanism as seen in Figure 28. Using the buttons on the face of the altimeter, the altitude for the deployment of the parachute can be specified from 100 ft to 1,000 ft at 100 ft increments. To use the Chute Release, the pin of the system is connected to the altimeter using a rubber band. The rubber band is then wrapped around the bundled parachute. The pin is then placed into the lock which is controlled by a servo motor. When the vehicle reaches the specified altitude, the servo will release the pin, allowing the parachute to be deployed. The Chute Release has a test setting. This will allow the team to test the test the parachute bundling and ensure that the parachute will be released without interference at the specified altitude. Two Chute Releases will be used on both the main parachute and the payload parachute. The two Chute Releases will be linked together in series. When one Chute Release deploys the parachute will be allowed to become unbundled. This will allow the Chute Releases to be fully redundant. 37

47 Figure 28. Jolly Logic Chute Release. Other systems were considered to do the unique deployment scheme of this vehicle. Tinder Rocketry offers an option called tender descenders. The tender descenders work in conjunction with an altimeter to keep the main parachute in its deployment bag until the main charge is fired using the altimeter. The tender descender was not chosen to be the method of delaying the main parachutes because of its complexity and its use of black powder. The tender descender uses black powder to release the parachute while the chute release does not. Removing black powder from this part of the recovery system of the vehicle will it to be assembled in a safer environment. Make the tender descender redundant requires a series of quick links and more black powder. The way the chute release works allows it to be redundant easily by connecting two of them in series. For these reasons, the Jolly Logic chute release was chosen. 38

48 Avionics Electrical Setup Figure 29: Electrical schematics for the booster section. Figure 30: Electrical schematic for the payload section The avionics electrical schematics for the vehicle are shown above in Figure 29 and Figure 30. These figures show the wiring from the respective batteries to the charges used for the recovery events. For the booster section of the rocket, there will be both a primary and secondary altimeter system. The primary system will be powered by a 9V battery which will turned on with 39

49 a rotary switch such that it can be accessed outside the rocket. The rotary switch is able to be locked into the on position when the rocket is waiting on the launch pad to ensure that the recovery equipment will remain on throughout the duration of flight. The primary altimeter is the StratoLoggerCF, which will trigger the ejection charges when the rocket reaches its apogee. The secondary altimeter system is used for redundancy in the case that the primary system fails to deploy the parachutes at apogee. The secondary system will consist of a Telemetrum 2.0 powered by a 4V LiPo battery. In terms of the payload section, the setup will be generally the same. However, this section will only be utilizing one ejection charge for its parachute. The altimeters will be wired to the e-matches through terminal blocks on the outside of the coupler. This will allow the e-matches to be switched connected and disconnected with ease. There are other types of switches such as a key switch. A key switch, however, requires a specific key to use. Therefore, it is easier to use a rotary switch. The rotary switch is able to be locked into the on position while the rocket is waiting on the launch pad to ensure that the recovery equipment will remain on throughout the duration of the flight Attachment Hardware Tubular Kevlar will be used as the shock cord for this vehicle. The shock cord will be used to keep the nosecone and upper airframe connected to the payload parachute and it will keep the coupler section and booster section connected to the main and drogue parachutes. The tubular Kevlar will be ½ inch in width. ½ inch tubular Kevlar can withstand forces greater than 7,200 lbs. This is strong enough to withstand the forces created by the recovery system, notably during the parachute deployments. To attach the shock cord to the parachute in a safe and secure way, the shroud lines will be passed through a loop in in the shock cord and then the parachute will be passed through the looped shroud lines. The shock cord that is attached to the motor mount will be done so by using a steel quick link and a steel eye bolt. The eye bolt will be attached to the top of the motor mount in a slot designed for this purpose. The steel quick link attaches the Kevlar shock cord to the steel eye bolt. This quick link allows for easy assembling on launch day as well as increases safety. The drogue parachute will be attached to the avionics bay. A U-bolt will be screwed into a plywood bulkhead that is attached to the avionics bay. The U-bolt will be attached using a nut and washer on each side of the bulkhead. In order to ensure that everything is structurally sound, epoxy will also be added to the nuts and washers. The main parachute will be attached to the other side of the avionics bay using the same technique of the drogue. The payload parachute will be attached to a bulkhead in the upper airframe using a quick link and U-bolt in similar fashion as the other parachutes. The bulkhead will be epoxied into place in the upper airframe. Machine screws will be used to connect the nosecone to the upper airframe, shear pins will connect the upper airframe to the payload coupler, and coupler to the booster section. The shear pins will be able to keep the rocket together during launch and the coasting sections of the vehicle s launch prior to desired parachute deployment. The shear pins will break when the black powder charges detonate, thus allowing the sections of the vehicle to separate. The main 40

50 parachute will be stored in between the coupler section and the upper airframe. The drogue parachute is housed in the booster section, under the coupler. The payload parachute is housed in the upper airframe under the nosecone. The amount of shear pins needed in each of these joints will be calculated at a later time when ejection charge testing begins Redundancy Redundancy is another critical aspect that the team strived to employ in every part of the recovery system. Redundant altitude measurement and ejection systems are vital not only to mission success, but also to the safety of observers on the ground. All ejection events are triggered by a primary and secondary charge, with primary charges triggered by one altimeter and secondary charges by a separate altimeter. Each altimeter also has its own power supply. Additionally, the secondary backup charge will be larger than the primary charge to ensure separation in the event the back charge is required. The redundancy employed in the ejection charges ensures that even with complete failure of an entire altimeter or battery, all parachutes still have a complete backup system. The Jolly Logic Chute Releases on the main and payload parachutes are also fully redundant. There are two releases joined in series. In the event of failure of one release, the other independent system should perform all the duties needed for a successful deployment of the parachute Mission Performance Predictions Through the use of a combination of software and hand calculations, the team was able to predict the performance of the current launch vehicle design, and ensure that all mission requirements were met Simulation Methods In order to get a sense as to the performance and stability of the design, the team utilized OpenRocket simulation software. While the team has a history using OpenRocket and the software is highly regarded within the high powered rocketry community, utilizes additional simulation software as the project moves forward will give the team greater confidence in the performance of the flight vehicle OpenRocket OpenRocket is one of the simulation programs that will be used to predict the flight of the vehicle. OpenRocket allows the user to create a high powered rocket design. The user can choose different body tube lengths, diameter, and material for the sections of the vehicle. The software also allows for masses of the avionics and payload to be added to the simulated vehicle. OpenRocket s fin construction offers many options including trapezoidal and freeform. The thickness, lengths, number of fins, and fin tabs are parameters that OpenRocket allows the user to set. Different stages can also be simulated using OpenRocket. This is an important option as 41

51 the vehicle in this year s competition will descend in two separate sections. Once the vehicle has been designed, it calculates the center of pressure and center of gravity and displays the stability in calibers. The software also simulates the vehicle s flight and can be used to export altitude, velocity, drift distances, and acceleration. OpenRocket has also been used by team members on previous rocket designs. It was used by team members for last year s student launch competition. OpenRocket s simulated apogee has been a good prediction in other rocket designs that members of the team have built before. OpenRocket was chosen as the main simulation software for the student launch because of the flexibility in the vehicle's design, the accuracy of the simulation based on previous years, and the team s familiarity with the software. Figure 31 below shows the team s current design modeled in OpenRocket. Figure 31: Current desing modeled in OpenRocket. Table 10 includes the apogee of the rocket at various wind speeds. All projected apogees are above the mile-high requirement set by the Student Launch handbook. Table 10: OpenRocket Apogee Simulations Wind Speed [mph] Apogee [ft]

52 RockSim As the team moves forward, the team will be supplementing data from OpenRocket with RockSim, developed by Apogee Components. In terms of functionality and appearance, RockSim and OpenRocket are very similar pieces of software. However, RockSim has more robust options in terms of controlling drag coefficients and staging and the data provided by the two software packages can be compared against each other and future test flights in order to best design the final launch vehicle. RockSim is available to the team at a reduced price for members who want to work from home and the computers in the aerospace computer laboratory come equipped with the software so no purchase is actually necessary. Figure 32 below shows a sample image of the RockSim V9 software. Figure 32: Sample design modeled in RockSim V9. 43

53 Custom MATLAB Program A custom MATLAB simulation is currently in the process of being written. This simulation will be used to predict the flight of the vehicle. The simulation takes in different parameters of the vehicle, e.g. fin shape, cd of vehicle, parachutes sizes, and motor thrust curve, and predicts the vehicle s performance. The simulation predicts altitude, velocity, acceleration, and drift distances for different wind speeds. The data from the simulation will be used to predict exit rail velocity and descent speeds. These will allow the team to confirm OpenRocket and RockSim predictions and allow the team be confident that the vehicle will have a safe flight. The custom simulation will take 2D versions of the equations of motion of the vehicle and solve them using numerical integration. MATLAB s ode45 function will be used in the simulation. Ode45 takes in the time span which to solve the equations, initial conditions of the state, and the function where the equations of motion are and then solves them over the time space. Ode45 uses a Runge-Kutta (4,5) method with a variable time step for efficient computation. This will allow the simulation to run quickly while still remain accurate to the 5th order with an accumulated error to the 4th order. Figure 33. Custom MATLAB flight simulator user interface. The first version of the simulation has been written. This version predicts an apogee that is similar to that predicted using OpenRocket. The simulation provides the user with velocity off of the launch rail and max velocity during the flight. Figure 33 shows the interface that the user of the simulation work with and Figure 34 shows and example of a graph of the booster s altitude produced by the simulation. Things that will be added and refined between version 1 and version 2 is drift prediction, refinement of the Cd of the body, and calculating and plotting vehicle 44

54 stability over the flight time. The team will also continue to make changes to the simulation when problems arise or when new information about the flight of the vehicle is wanted. Figure 34. Example plot of altitude from MATLAB simulation. Figure 35. Example plot of vertical velocity from MATLAB simulation. 45

55 Hand Calculations Rounding out the team s simulation tools will be calculations by hand using readily available equations found online. Hand calculations are used in this report to calculate terminal velocities and kinetic energies on impact of the four section of the launch vehicle. In future reports, the team will hand calculate values such as the center of pressure and apogee. These calculations will serve to strengthen the team s confidence in the design s expected performance as the project heads towards manufacturing and testing Flight Profile Analysis In order to determine maximum altitudes, velocities, and accelerations, the team utilized the OpenRocket software and the team s custom MATLAB script to generate predicted flight profiles from which the aforementioned values could be determine. A plot of the OpenRocket flight profile at typical 10 mph cross winds is displayed below as Figure 36. Figure 36: Simulated flight, velocity, and acceleration profiles from OpenRocket. The team used the OpenRocket software as the official software to determine the maximum altitudes, velocities, and accelerations. The OpenRocket data shows that apogee will occur at an altitude of about 5,350 ft (~1 mile) at a time of 17.9 seconds after the launch of the rocket. The rocket will reach a max velocity about 697 ft/s, and the maximum acceleration will be about 311 ft/s^2 or 9.67 g. To determine the maximum Mach value, the team used the equation: 46

56 where M is the maximum Mach number, a is the local speed of sound (ft/s), and V is the maximum velocity of the rocket (ft/s). The team used an estimated first order value of 1,125 ft/s for speed of sound. Using the above equation and the data provided by the OpenRocket simulation, the maximum Mach number of the rocket will be All flight profile characteristics at 10 mph cross winds are presented below in Table 11. Table 11: Flight Profile Performance Characteristics Performance Characteristic Value Max Altitude 5,350 ft Max Velocity 697 ft/s Max Acceleration 311 ft/s^2 Max Mach Number Stability Stability is an important factor to consider when designing any high powered rocket. The Cg must be located sufficiently far enough up the rocket away from the Cp in order for the fins and other aerodynamic surfaces to provide a sufficient restoring force to the rocket as it encounter disturbances such as wind during flight. At the same time, two high of a separation between Cp and Cg could result in too high of a restoring moment, resulting in a tendency for the rocket to follow the direction of disturbances. With that in mind the team aimed for a launch pad stability margin of around 2 calibers, as measured using the following equation: where D is the diameter of the rocket (in), and Cp and Cg are measured from the tip of the nosecone in inches. OpenRocket was utilized for the purposes of calculating stability margin (in calibers). According to the software, the stability of the completed design as it would sit on the launch rail is 2.06 calibers. Along with this the rocket will have a Cg from the nosecone and a Cp from the nosecone. This information is included below in Table 12. Generally speaking, a margin significantly below 2 calibers is considered understable, while a margin significantly above 2 calibers is considered overstable. This stability will only increase before it leaves the rail, insuring that the stability margin will be above 2 calibers at time of rail exit as required by the competition. The team also used OpenRocket to plot the stability margin of the rocket as a function of time during flight and found that the rocket will continue to be stable throughout the motor burn and coasting stages of flight. 47

57 Table 12: Stability Parameters of the Launch Vehicle at Takeoff Parameter Value Center of Gravity in Center of Pressure in Stability Margin 2.04 calibers Kinetic Energy It is important to consider terminal velocities and kinetic energy on impact when sizing parachutes for any high powered rocket. If a rocket falls down too fast and hard, individual components such as the fins or body tubing can become damaged if not outright unrepairable. Since the team is interested in creating a launch vehicle that is truly reusable, parachutes must be adequately sized to limit the kinetic energy of the vehicle on impact. This is also a requirement set forth by NASA, with an upper limit on kinetic energy on impact for any individual section of the rocket at 75 ft*lbf. In order to determine the kinetic energy of each section of the rocket, the terminal velocities under each parachute must first be analyzed. Knowing the weight under the parachute, the drag coefficient of the chute, and the area of the chute, the terminal velocity of the two falling sections can be determined using the following equation: With density of air approximated to a constant lbm/ft 3 and g equal to ft/s 2. Next, Table 13 gives a mass breakdown for each of the four sections that will be landing and how much mass is under each parachute. Table 13: Mass under Each Parachute Broken Down by Section Section Mass [lbm] Drogue Parachute Section 1: Avionics Coupler 6.9 Section 2: Booster Tube 14.1 Mass Under Drogue: 21 Main Parachute Section 1: Avionics Coupler 6.2 Section 2: Booster Tube 14.1 Mass Under Main:

58 Section Mass [lbm] Payload Parachute Section 3: Payload (Camera + Landing) 5.73 Section 4: Nosecone 4.30 Mass Under Payload Chute: With terminal velocities known, kinetic energy ( ) can be calculated using the following equation: where m is the mass in slugs, and is the terminal velocity in ft/s. With these equations, the recovery team set out to size parachutes that satisfy the kinetic energy requirement while maintaining a sufficiently slow decent speed. Drift calculations were also taken into account but the full analysis on that will be included in Section 0. Starting with the drogue parachute, three different size of elliptical drogue parachutes from Fruity Chutes were evaluated for their impact on kinetic energy and terminal descent speeds for the particular vehicle design. These findings are tabulated below in Table 14. Table 14: Drogue Parachute Options Parachute Cd Diameter (in) Vt (ft/s) Ek Section 1 Ek Section 2 Fruity Chutes Elliptical Fruity Chutes Drogue Fruity Chutes Drogue The values for all three sizes fall largely within what the high powered rocketry community considers a safe descent speed. These recommendations include a drogue descent between ft/s. A similar analysis was done for three potential main parachute options and those results are tabulated below in Table 15. The team already owns 60 in and 96 in Iris Ultra parachutes and those options are highlighted in the table, and an 84 in was included as another potential option. With the main parachute, the most important parameter was the kinetic energy or the individual sections, with drift a secondary parameter for selection. Table 15: Main Parachute Options Parachute Cd Diameter (in) Vt (ft/s) Ek Section 1 Ek Section 2 Fruity Chutes Iris Ultra Fruity Chutes Iris Ultra Fruity Chutes Iris Ultra

59 With regards to the recovery of the upper payload section, deviations from the standard dual deploy model had to be made. Since there is currently only one compartment for a payload parachute, the recovery team decided that a single parachute should be used. While it isn t unheard of to deploy two parachutes from a single compartment, the increased risk of entanglement is one that the team would like to avoid at all costs. At the same time, it is impossible to deploy one parachute at apogee and satisfy the kinetic energy and drift requirements for the upper payload section. In order to satisfy these requirements with a single parachute, the team has opted to follow a drogue-less dual deploy model for the upper payload section. Two seconds after apogee, the payload parachute will deploy. However, the parachute will deploy as a bundle tethered by two Jolly Logic chute releases in series and in this way act as a sort of streamer as the upper section descends. At 900 ft the chute releases will untether the payload parachute and allow the payload parachute to fully deploy. In order to model the terminal descent speed of the tumbling upper section under a bundled parachute acting as a streamer, the team consulted documentation from Apogee Rockets. From this documentation, the team found that the optimal way to model this scenario is to add a phantom parachute of zero mass that deploys when the bundled parachute deploys. The diameter of this parachute is calculated using the following formula: Where L is the length of airframe and d is the outer diameter of the airframe. The effects of drag from the nosecone were ignored for the purposes of this analysis. The parameters of this phantom parachute and resulting terminal decent speed of this system are included below in Table 16. Table 16: Phantom Parachute Parameters Parameter Value L 31 in d in D 15.6 in Cd 0.8 Terminal Velocity ft/s This terminal velocity falls within the recommended range of ft/s. With the viability of this setup confirmed, the team analyzed three different options for the payload parachute that would fully deploy at 900 ft AGL and analyzed the resulting terminal velocity, kinetic energy, and drift and tabulated these calculations (minus drift) below in Table 17. The already owned 44 Skyangle is highlighted in yellow in this table. Table 17: Payload Parachute Options Parachute Cd Diameter (in) Vt (ft/s) Ek Section 3 Ek Section 4 SkyAngle Model C

60 SkyAngle Fruity Chutes Iris Ultra After analysis of the above options, the team found no significant benefit to be gained from buying payload and main parachutes when the Skyangle and Iris Ultra were sufficient for satisfying the competition requirements. For the drogue the team found the 18 Fruity Chutes ellipital to provide a good balance between terminal velocity and drift. Table 18 below gives performance data for the chosen parachutes. Table 19 presents the calculated terminal velocity values for the payload under its own shoot and the bottom half of the rocket under both drogue and main. Table 20 below gives a breakdown of the impact kinetic energy of the four sections of the rocket. All calculated kinetic energy values fall comfortably below the 75 ft*lbf limit set by the competition. Table 18: Performance Characteristics of Chosen Parachutes Recovery Device Model Cd Diameter Main Parachute Iris Ultra Drogue Parachute Fruity Chutes Elliptical Payload Parachute Skyangle Model C Table 19: Terminal Velocity of Each Section of the Vehicle for Chosen Parachutes Section Booster + Avionics Coupler (bottom half) Under Drogue Parachute: Under Main Parachute: Payload + Nosecone (upper half) Under Bundled Payload Parachute: Under Deployed Payload Parachute: Terminal Velocity ft/s ft/s ft/s ft/s Table 20: Kinetic Energy of Each Section of the Vehicle at Impact for Chosen Parachutes Section Kinetic Energy [ft*lbf] 1. Avionics Coupler Booster Tube Payload (Camera + Landing) Nosecone + Upper Airframe

61 Drift While drift distance is no longer an explicit requirement in the handbook for the competition, drift is still on a high level of priority. This is to help retrieval efforts, respect the boundaries of the launch field provide to the Student Launch competition, and to ensure the payload will have a good view of the detection tarps during its descent. With this in mind the team set a max drift distance of 2500 ft in worst case scenario cross winds of 20 mph winds. All analysis was done assuming a launch angle of zero degrees as specified by the Student Launch handbook, although the actual launch angle will be five degrees. Analysis of realistic drift distances for this five-degree launch angle will be done for CDR. With that, the drift distance of each section using the chosen parachutes at a variety of different wind speeds are tabulated below in Table 21. Table 21: Drift Calculations for Each Section of the Launch Vehicle Section Upper Payload Section Lower Booster Section Drift in 0 mph winds [ft] Drift in 5 mph winds [ft] Drift in 10 mph winds [ft] Drift in 15 mph winds [ft] Drift in 20 mph winds [ft] ,225 1, ,300 2,080 With an idealized zero-degree launch angle, no drift distances exceeded the 2,500 ft drift requirement. This further justifies the team s choice of parachute and parachute sizing decisions. 52

62 4. SAFETY 4.1. Safety Plan Overview The safety of all team members is of the absolute highest priority for the ISS Student Launch Competition team. Should a situation arise in which a project-critical choice needs to be made, safety is considered before the success of the project. The safety officer this year will be Brian Hardy, who will oversee a small team that will conduct a thorough analysis of any hazards the team may encounter this year when building or launching the rocket. Brian and the safety team will also implement plans and procedures to minimize the risk of these hazards. This year, using a combination of safety briefings, online classes and thorough documentation, the team will actively encourage participation in the adherence to safety procedures. Safety training will be required for any member that wishes to participate in construction sessions or attend a launch. By keeping lists of safety-trained members and having experienced members actively involved at every build session, the team can ensure that everyone working in lab spaces understand safety protocol in situations when everything is going well and in the rare event when something goes wrong Safety Briefings In order to better facilitate the spread of knowledge on safe practices, the safety officer will brief the team on hazard recognition and accident avoidance any time he sees fit, especially before launching, testing, and construction. These safety briefings will be led by the safety team and include the acknowledgement of all hazards and risks associated with the relevant work. Mandatory safety protocol will also be emphasized to the team: wearing safety glasses, wearing respirators when working with fiberglass, being properly dressed for construction, and being knowledgeable about rocket launch safety protocol. The team will not only address what to do in case of emergency, but more importantly how to avoid emergencies in the first place. More details on team-wide safety requirements can be found in the next section Member Requirements All members of the team are required by the University of Illinois to complete mandatory safety training. The required safety courses are Electrical Safety for Labs and General Laboratory Safety and are provided by the University. The Electrical Safety for Labs course provides an awareness of basic electrical safety concepts involving household-level voltages that should be followed in laboratories to avoid electrical shock, damage to sensitive equipment, and the ignition of combustible materials. The General Laboratory Safety course provides information regarding standard laboratory safety guidelines, laboratory signs and labels, personal protective equipment, working with biological/chemical/radiological materials, waste disposal, and emergency preparedness. Additionally, all members of the team will be required to read, sign, and date a general safety contract written by the safety officer prior to participating in any construction activities. This contract will include all general safety hazards and the NAR High 53

63 Power safety code. If the safety officer sees that any of these requirements are not met, or if the safety contract is breached in anyway by a member of the team, he will have the power to prohibit that member from doing any project-related work. Finally, the Student Launch team will be required to read and abide by all of the rules in the ISS Tech Team Safety Policy, which can be referenced in Appendix C: ISS Tech Team Safety Policy Equipment Training In order to provide team members with the benefit of knowledge on how to operate a wide array of equipment and tooling, the safety team will provide tutorial sessions on all machinery and tooling that may be used during the course of construction. To that end, the safety team will write up or adapt manufacturer operating procedures for these tools and upload them to the team s shared drive for easy access. Example of such tooling that will require tutorial sessions include: Laser cutter 3D printer Disk sander Diamond saw Down draft table Table saw Dremel tool In order to operate this machinery, team member will have to attend these tutorial sessions or receive training separately from a member of the safety team or team management. In providing and requiring these sessions, the team not only reduces the risk of mishaps due to misuse of equipment, but also ensures redundancy in knowledge of construction techniques NAR Mentor Mark Joseph will be the NAR team mentor for the ISS Student Launch team for this year s competition. Mark has worked with the team for several years now and is familiar with competition and its structure. Mark s primary role outside of general design guidance will be to handle all handling of energetics and the explosive motor fuel grains the team will need to utilize Energetics Handling In addition to providing valuable design input and accompanying the team to Huntsville, the NAR team mentor will also be tasked with the handling of all energetics. This includes the motors used in both the full scale and subscale vehicle, as well as the e-match ejection charges that will separate the rocket at apogee and deploy the three parachute. 54

64 Motor Purchase and Storage Motor storage, transportation, and preparation will be in accordance with the National Fire Protection Agency, specifically NFPA code The motor shall be stored in a Type 3 or Type 4 indoor magazine because the chosen rocket motor is under 50 lbs. Transportation of the motor will comply with 49 CFR Subchapter C Hazardous Materials Regulation, which covers the packaging, handling, and transportation of high-power rocket motors. The operations manual for the motor will be posted on the team website as soon as the motor arrives. When purchasing motors, ISS will purchase from the particular vendor using Mark Joseph s NAR number. Once the Student Launch team receives the motors from the storefront, Mark will store the motors until needed. This ensures that Mark will be the only one to interact with the explosive fuel grain. Only the team mentor will handle, purchase, store, and transport all explosives and motors. There will also be fire extinguishers on hand in all locations where construction or storage will take place. Safety officer will brief the team on launch procedure etiquette, as well as accident avoidance and hazard recognition. All team members will be required to review and sign a team safety agreement and abide by the terms within, which include all pertinent laws and regulations. Environmental regulations will be referenced during the course of this project to ensure compliance. The group s safety officer is responsible for finding these relevant regulations for the handling and proper disposal of hazardous or environmentally harmful materials Risk Assessment Overview To better prepare for issues that inevitably arise during any project of large scale and to prioritize the team s time, the safety team conducted a thorough risk analysis based on the severity of these issues. The safety team analyzed risks to the project, the environment, and above all, the health of team members during the construction process. The team used Risk Assessment Codes (RACs) to evaluate the various hazards to both personnel and the project. Table 22 introduces the risk matrix and the risk assessment codes that will be used to classify risks throughout the rest of the safety section. Risks are color-coded based on the severity and Table 23 discusses the team s response to these various levels. Table 24 defines the levels of severity as it relates to personnel, project, and environmental health. Finally, 55

65 Table 25 defines individual instance probability and probability of occurrence of these risks throughout the entire project timeline. Table 22: Risk Assessment Codes (RACs) Severity Probability 1 Catastrophic 2 Critical 3 Marginal 4 Negligible A - Frequent 1A 2A 3A 4A B - Probable 1B 2B 3B 4B C - Occasional 1C 2C 3C 4C D - Remote 1D 2D 3D 4D E - Improbable 1E 2E 3E 4E Table 23: Level of Risk and Member Requirements Level of Risk High Risk Moderate Risk Low Risk Minimal Risk Level of Training and Supervision Required Highly undesirable. The risk factor will be compared with the importance to the success of the project. Procedure or equipment operation must be done by the safety officer or team lead, or under their direct supervision. Undesirable. Procedure or equipment operation requires documented approval from safety officer and team lead in form of training and proof of online safety training completion. Procedure or equipment operation requires supervision. Acceptable. Procedure or equipment operation requires training, but no direct oversight is necessary. Acceptable. Procedure or equipment operation require almost no training and no direct oversight. Instruction is highly recommended for new members. 56

66 Table 24: Severity Definitions Description Personnel Safety Project Success Environmental 1 Catastrophic Loss of life or permanent injury. 2 Critical Severe injury or illness requiring hospitalization. 3 Marginal Minor injury or operationrelated illness. No hospitalization required. 4 Negligible Very minor injury or operation-related illness. Loss of or irreparable damage to launch vehicle. Failure to meet critical mission goals. Severe but reparable damage to launch vehicle. Failure to meet critical mission goals. Reparable damage to launch vehicle. Failure to meet noncritical mission goals. Minor or cosmetic damage to launch vehicle. All mission goals successfully met. Irreversible and severe damage that violates law or regulation. Loss of project. Reversible damage to environment that violates law or regulation. Significant damage to project. Reversible damage to environment that does not violate law or regulation. Minor damage done to project. Minimal effect on environment or project. 57

67 Table 25: Probability Definitions Description Definition A Frequent Highly likely to occur in any individual session and likely to be experienced continuously throughout the course of the project. B Probable Likely to occur in any individual session and expected to occur regularly throughout the course of the project. C Occasional Expected to occur a few times throughout the course of the project. D Remote Unlikely to occur in any individual session but expected to occur once throughout the course of the project. E Improbable Highly unlikely to occur in any individual session and not expected to occur throughout the course of the project. 58

68 4.4. Personnel Hazard Analysis The safety of all project personnel is paramount to the ISS Student Launch team. Any potential for human injury, however minor, must be considered in order to fully understand all involved risks and develop mitigation strategies to prevent harm to participants. To this end, the safety team has put together a comprehensive personnel hazard analysis to identify all known personnel risk factors and how best to address any safety concerns. The results of this analysis are presented in Table 26. For each identified risk, the team has considered four factors: its likely cause, its impact on personnel, its assigned Risk Assessment Code, and an optimal mitigation strategy. This hazard analysis will help the team anticipate risks for human injury well in advance, with the Risk Assessment Codes allowing the safety officer to pay particular attention to regulating higher-risk activities. The ultimate goal is to preemptively implement mitigation strategies before, not after, incidents occur. In completing the personnel hazard analysis, the safety team also referenced a variety of Material Safety Data Sheets (MSDS). These sheets allowed the team to better understand the risks of working with various materials, as well as the standard safety protocols associated with each. The safety team will become familiar with all of the materials that will be used during the construction process and will create summarized hazards and harm mitigation techniques for each of them. The safety officer will also instruct members working with hazardous materials on the dangers of a given material, how to safely work with it, and what to do if something go wrong. A list of all referenced MSDS is presented in Table 27, following the personnel hazard analysis. 59

69 Table 26: Personnel Hazard Analysis Risk Cause Impact Mitigation Trips, slips, and miscellaneous accidents Exposure to high voltage Cluttered lab space and distractions in the workspace Contact with exposed electrical equipment and outlets Ranges from minor cuts and scrapes to severe injuries including concussions and broken bones 1) Possible fire hazard 2) Electric shock can cause severe burns, muscle pains, seizures, and death. Work area should be properly maintained and organized. 1) Keep work area clean and well-lit. 2) Place clearly-visible signs where tripping or slipping hazards exist. Spills should be mopped up immediately. 3) Do not operate any power tools in volatile environments. 4) Do not operate machinery that poses any threat of danger in the presence of distractions. 1) Make sure team knows proper grounding procedures. 2) Ensure safety of equipment and workspace before handling live circuit boards, power tools, or electrical cords. RAC w/ Mitigation 3D 1E 60

70 Risk Cause Impact Mitigation Injury from use of diamond cutting saw 1) Injuries can occur if hands slip or if they are placed too close to the saw. 2) Kickbacks can occur if blade height is not correct or if the blade is not maintained properly. 3) The cutting action of the blade may throw wood chips or other material splinters. 1) Minor irritation up to severe injury to eyes due to flying debris 2) Severe lacerations to hands if placed in running diamond saw 1) Use a guard always and use a push stick to cut small pieces of material. Keep hands out of the line of cut. 2) Make sure blade height is correct before cutting, maintain and sharpen blade, and stand at the side of the saw blade to avoid injury in case of kickback. 3) Remove cracked blades from service, maintain sharp blades, and always wear eye protection when using the table saw. RAC w/ Mitigation 2D Injury from use of 3D printer Looking at laser cutter station during operation 1) Hot surfaces, namely the printer head block and UV lamp 2) Printing materials such as thermoplastics can be flammable. Lasers emit high levels of energy. Minor burns to skin Possible damage to vision, with the severity depending on duration of viewing Only those with proper training will be permitted to operate the table saw. Keep body parts away from 3D printer when it is in use. Only those with proper training will be permitted to operate the 3D printer. Team members will be instructed not to look at the laser cutting station when it is in use. Only those with proper training will be permitted to operate the laser cutter. 3E 4D 61

71 Power drill slippage Risk Cause Impact Mitigation Pricking from needles and syringes (used for application of epoxy) Contact with operating Dremel tool bits Skin injury and respiratory issues from working with Blue Tube 2.0 Dull drill bits are hard to handle and prone to slippage. Mishandling of needles and syringes 1) Sharp, fast rotating objects 2) Bits can fail and create shrapnel when cutting hard material. 1) Blue Tube can have sharp edges when cut. 2) Any dust generated during cutting can be an eye irritant. Mild lacerations or puncture wounds to hands and other extremities 1) Slight cuts or puncture wounds 2) Needle stick injuries can cause exposure to blood and other infectious materials. 1) Cuts, scrapes, and other mild skin injuries 2) Eye injuries from flying debris Mild eye irritation or skin damage 1) Always keep drill bits sharp. 2) Drill small pilot holes before drilling large holes. 3) Make sure the chuck is securely tightened before use. 1) Exercise proper care when dealing with needles. 2) Dispose of all needles properly in sharps containers. 1) Keep bystanders away while operating this tool. 2) Do not operate in explosive environments, such as in the presence of flammable liquid or dust. 3) Make sure members wear eye protection and are aware of the proper bits to use for different types of materials. 1) Use gloves when handling Blue Tube that has been recently cut, before it is sanded. 2) Wear surgeon s mask when sanding Blue Tube. RAC w/ Mitigation 3D 3E 4B 3E 62

72 Risk Cause Impact Mitigation Skin injury and respiratory issues from working with fiberglass Accidental dispersion or ignition of black powder 1) Splinters from freshly cut fiberglass can puncture skin. 2) Fiberglass dust is very hazardous and hard to avoid when cutting or sanding. Improper handling of black powder or storage in nonapproved containers 1) Can cause irritation of the eyes and skin. 2) Dust is dangerous to inhale and can cause respiratory issues when cutting or sanding down. 1) Possible bodily injuries such as skin burns 2) Respiratory issues if inhaled 1) When cutting or sanding fiberglass, gloves, goggles, and a respirator must be worn at all times. 2) Fiberglass will only be cut and sanded in properly ventilated areas that are approved by the safety officer. 1) Black powder should be stored only in approved containers. 2) Only the minimum amount of powder needed should be kept in the open. The main container should be kept separate from the primary lab space. 3) Open flames and heat sources are strictly prohibited in the lab space when black powder is present. RAC w/ Mitigation 3D 2E Black powder will only be handled by the team mentor or any other member with the proper certification to work with black powder. 63

73 Risk Cause Impact Mitigation Exposure to Proline epoxy LiPo battery explosion Chemical burns from ignition of the rocket motor Smoke Inhalation Epoxy is a toxic substance and not safe to touch directly. Its fumes are harmful as well. Can explode if punctured Mishandling and/or faulty installation of the rocket motor, leading to premature ignition of the motor s fuel grains Members working on ejection charge testing could inhale smoke from the charges. Can cause irritant contact dermatitis and allergic reactions if it comes in contact with skin. Hands, wrists, and eyes are the most exposed areas. 1) Mild to severe skin and eye damage if in contact with battery contents 2) Possible burns Second to third-degree skin burns Prolonged irritation to respiratory system 1) Gloves will always be worn when working with epoxy and changed at any sign of damage. 2) Goggles will be worn when mixing epoxy to avoid splashing into the eyes. 3) Uncured epoxy is to be treated as hazardous waste and disposed of as such. Batteries will be properly stored when not in use to prevent any possible structural damage. 1) Ensure that the team mentor will be working with all components related to the motor, as per regulations. 2) Ensure that the minimum distance table is consulted before all launches and tests of the motor. 1) Ensure all members are wearing personal protective equipment including masks and goggles. 2) Make sure all charge testing is done with the assistance of the team mentor. RAC w/ Mitigation 3D 3E 2E 3E 64

74 Risk Cause Impact Mitigation Hearing damage from overuse of power tools Inhalation of spray paint propellant Collision of descending vehicle component with spectators Personnel could get sunburn Improper use of hearing protection Spray painting a component with no ventilation or little to no airflow 1) Improperly angled launch rail 2) Recovery system failure or insufficiently large parachute, causing rapid descent rate Intense UV rays Potentially permanent hearing damage if exposure is long enough Dizziness, loss of consciousness, and potential brain damage Blunt force trauma and possible concussion Personnel could get severe sunburn, sun poisoning, and or skin cancer in the future. Ensure that all members using (or in the vicinity of) power tools wear ear plugs or safety headphones. Ensure that there is proper ventilation and airflow in painting areas. All painting should ideally be done in an outdoor environment. 1) Monitor launch day wind speed and rail angle to ensure vehicle descends sufficiently far from spectators. 2) Size parachute appropriately to ensure safe descent speeds. 3) Implement redundancy in all aspects of the vehicle recovery system. 4) Keep a small noisemaker (e.g. an air horn) on hand at launch to alert spectators of any imminent hazards. Plenty of sunscreen will be made available to anyone outside. The danger that the sun can cause will be stressed to all team members RAC w/ Mitigation 1E 3E 2E 3C 65

75 Table 27: Referenced Material Safety Data Sheets (MSDS) Material PLA Plastic 3D Filament Fiberglass Tubing Black Powder Proline Epoxy 4100 System Resin & Hardener Aerotech Rocket Motors Rust-Oleum Spray Paint MSDS Source Sheet.pdf

76 4.5. Failure Modes and Effects Analysis Launch Vehicle In order to anticipate any safety or operational hazards that could arise during the launch process, the structures and recovery team conducted a thorough analysis to identify specific vehicle-associated risks and their consequences. Here, vehicle-associated risks are considered to be those that directly impact the flight path or structural integrity of the rocket. Any risks to the team or spectators on launch day are addressed separately in the previously-discussed personnel hazard analysis. When considering the failure modes of the launch vehicle, the goal of the study was to identify all anomalies that could occur during each stage of flight, including ignition, ascent, and recovery. Similar to the personnel hazard analysis, the vehicle risk analysis associates each identified risk with its cause, impact on the vehicle, Risk Assessment Code, and optimal mitigation strategy. A tabulation of the entire launch vehicle risk analysis can be found below in Table 28. Under the safety officer s supervision, the ISS Student Launch team will be implementing all identified mitigation strategies both on and leading up to launch day. 67

77 Table 28: Structures and Recovery Risk Analysis Risk Cause Impact Mitigation RAC w/ Mitigation Motor retainer fails. 1) Insufficiently secured motor retainer 2) Cap not properly tightened Unrestrained motor falls out the bottom of rocket. 1) Secure motor retainer to lower centering ring with epoxy and machine screws. 2) Ensure aft retainer cap is secure before launch. 1E Motor mount tube separates from outer airframe during launch. 1) Improperly installed motor mount tube 2) Lack of epoxy on centering rings Unrestrained motor travels through rocket body or falls out the bottom. Adequately coat centering rings with epoxy when affixing motor mount inside booster tube. 1E Motor ignition fails. 1) Defect in motor 2) Igniter not securely placed Vehicle fails to launch. 1) Ensure igniter is placed as far up the motor as possible. 2) Review launch sequence beforehand. 3B Motor ignition is delayed or the motor chuffs before full ignition. 1) Defect in motor 2) Ignitor not securely placed Non-optimal vehicle performance off the launch rail 1) Ensure adequate placement of igniter. 2) Inspect propellant grains for obvious defects before motor assembly. 3) Buy motors from trusted manufacturers. 2D 68

78 Risk Cause Impact Mitigation RAC w/ Mitigation Motor ignites prematurely. 1) Igniter prematurely triggered 2) Fuel grains exposed to open flames or other heat source 1) Vehicle launches unexpectedly. 2) Some electronic components may have yet to be activated. 1) Keep open flames and heat sources away from rocket at all times during setup. 2) Only arm igniter immediately before launch. 1E Motor backfires or experiences a severe internal anomaly during flight. Motor defect Loss of the vehicle 1) Inspect propellant grains for obvious defects before motor assembly. 2) Buy motors from trusted manufacturers. 1E Rail buttons scrape excessively along launch rail. 1) Improperly sized rail buttons for the given launch rail 2) Insufficiently sanded rail buttons Decreased vehicle performance off the launch rail 1) Ensure rail buttons are a proper fit for the launch rail being used. 2) Avoid use of excessive rail buttons. 3D One or more fins separate from the vehicle during launch. Insufficiently secured fins Loss of stability 1) Attach fins with multiple epoxy fillets on both the motor mount tube and outer airframe. 2) Insert fins in between centering rings for additional structural rigidity. 1E 69

79 Risk Cause Impact Mitigation RAC w/ Mitigation Vehicle spins excessively during launch. Improperly aligned fins Possible damage to sensitive internal components Use a fin jig during construction to maintain vertical alignment of fins. 2E Main parachute ejection charges fail. 1) Continuity not achieved 2) Defect in e-match 1) Booster and payload sections do not separate. 2) Main parachute cannot deploy. 1) Test ejection charge systems for continuity beforehand. 2) Use redundant charges and altimeters. 1E Drogue parachute ejection charges fail. 1) Continuity not achieved 2) Defect in e-match 1) Drogue parachute cannot deploy. 2) High descent speeds will prevent deployment of main parachute. 3) Loss of booster section 1) Test ejection charge systems for continuity beforehand. 2) Use redundant charges and altimeters. 1E Main parachute tether systems fail. Defect in Jolly Logic chute release 1) Main parachute cannot deploy. 2) Damage likely to booster section 1) Test Jolly Logic system beforehand. 2) Use redundant chute releases tethered together. 1E 70

80 Risk Cause Impact Mitigation RAC w/ Mitigation Rocket sections do not separate after ejection charges detonate. 1) Insufficiently sized black powder charges 2) E-match failure 3) No continuity to charge 1) Parachute stored between given sections cannot deploy 2) Possible loss of vehicle section 1) Calculate appropriate shear pin and black powder charge sizes. 2) Conduct ground tests of complete ejection system (rocket sections, parachutes, charges, and shear pins). 1E Altimeters lose power during flight. 1) Battery failure 2) Battery not sufficiently charged No flight data is recorded. 1) Test altimeter systems beforehand. 2) Confirm sufficient battery voltage prior to launch. 2E Outer airframe fractures during flight. Chosen material not strong enough to withstand aerodynamic forces 1) Loss of stability 2) Possible damage to internal components 3) Possible loss of vehicle Perform adequate material studies prior to selecting vehicle components. 1E Vehicle sections drift excessively far or into hazardous terrain. 1) Main parachute deployment at apogee 2) Payload parachute deployment at apogee Vehicle components are unrecoverable or require excessive time to locate. 1) Model the vehicle s flight characteristics and size parachutes appropriately to minimize drift distance. 2) Cancel launch in the event of excessive wind. 2D 71

81 Risk Cause Impact Mitigation RAC w/ Mitigation One or more vehicle components fracture upon landing. Vehicle sections impact ground at excessive speeds. Time-consuming repairs required before a second launch can take place 1) Use simulations to predict and implement sufficiently low descent velocities. 2) Ensure all outer components (particularly fins) are securely fixed to vehicle body with epoxy. 2E 72

82 Payload The team understands that the proposed payload is ambitious and will require significant analysis and testing. Before finalizing designs and beginning the manufacturing process, the team is currently prototyping essential competition-critical technologies and plans to continue to do so. This prototyping and testing phase will allow the team to realize unforeseen safety hazards and reevaluate currently known risks. To facilitate this iterative design process, the payload team has conducted a payload risk analysis similar to those put together for personnel and launch vehicle safety. Top risks to the successful completion of the imaging and lander payload have been identified and can be found in Table 29. As in previously-detailed studies, each risk is also listed with its respective cause, expected impact on the payload, Risk Assessment Code, and mitigation strategy. The payload team will be implementing these mitigation strategies both now, during the prototyping phase, as well as later during final construction and operation. 73

83 Table 29: Payload Risk Analysis Risk Cause Impact Mitigation All four landing legs do not fully deploy. Landing legs fail at deployment. Landing legs fail when the payload reaches the ground. Payload parachute does not deploy. Camera movement or obscuring of view during flight 1) The legs may get stuck inside of the body tube. 2) The springs and hinges may fail. Legs deploy too fast and fracture when snapping into place. 1) Excessive descent speeds cause the payload to violently impact the ground. 2) An improperly deployed leg adds additional loads to remaining legs. 1) The parachute may get tangled up. 2) Vehicle could have a failed separation event. 1) The camera may move during flight due to vehicle vibration and movement. Payload will not land vertically. Payload will not land vertically. Payload will not land vertically. 1) Payload will not decelerate. 2) Uncontrolled descent in the air 3) Upright landing unlikely Blurred or obstructed images will make target identification 1) Properly size leg deployment slots. 2) Run multiple deployment tests, including during charge testing and the test flight. 1) Size deployment springs appropriately. 2) Repeatedly test leg deployment. 1) Work with structures and recovery subteam to size payload parachute appropriately. 2) Conduct impact tests by dropping payload from short heights under parachute. 1) Work with structures and recovery subteam to ensure the recovery system is well tested and reliable. 2) Use redundant recovery electronics. 1) Redundant cameras have been included. 2) Testing will occur in RAC w/ Mitigation 2D 2E 2E 1E 2D 74

84 Raspberry Pi memory full Depleted payload batteries 2) The camera may come loose if not securely attached. There is too much data from the ground images. 1) Overuse of the electronics 2) Insufficiently charged battery less likely. Inability to prove identification of the target Failure to identify targets LiPo battery explosion Can explode if punctured Loss of power to camera and data collection system Inaccurate altimeter reading The altimeter does not get a proper pressure reading because the pressure inside the body tube may not be stable or equal to the external pressure. 1) Altitude of the payload will not be accurate. 2) Image processing system may fail to identify correct targets. 75 flight-like conditions for this segment of the system. 3) The camera will be securely attached. 1) Images will only be stored when the targets are identified positively. 2) A failsafe will be included in the code to prevent errors due to lack of memory. 1) The battery has been sized to greater than 90 minutes of runtime. 2) The system will also undergo long-duration testing. Batteries will be properly stored when not in use to prevent any possible structural damage. 1) Significant ground testing of altimeter system. 2) Test flight will show any high velocity specific issues. 3) Code modifications to avoid large errors due to outlier data points may be required. 2E 1E 1E 2E

85 Loss of data connection between payload components (e.g. camera & Raspberry Pi) Software failure in flight 1) Improperly soldered connection points 2) Wire slips during flight. Improperly written or implemented code Data will cease being recorded. Data will cease being recorded. 1) Ensure any soldered components are properly attached with sufficient amounts of solder. 2) Tighten all connection points where wires can come loose. 3) Thoroughly test complete electronics system to isolate and repair any faulty intercomponent connections 1) Run and debug all code pre-flight. 2) Test code on Raspberry Pi and verify proper data output 2E 2E 76

86 4.6. Environmental Concerns Due to the nature of high-powered rocketry and the challenging payload, this project has several environmental concerns that must be addressed. Several tests such as the ground imaging test on a quadcopter, payload vertical landing test, black powder charge separation ground test, and full scale flight test will be done outdoors, some in public spaces. In addition, the subscale flight and competition flight will also be completed outdoors. It is important that the team obeys all federal, state, and local laws including University policies regulating activities outdoors and in public spaces. The safety officer will ensure that the team follows all regulations by reviewing any pertinent legal documentation regarding outdoor activities, creating a document with all relevant and important rules, and briefing relevant team members about the important rules. For all outdoor activities including tests and rocket launches, the team will formulate a detailed document with step-by-step procedures, safety precautions, and methods to minimize environmental impact. The safety officer and safety team will review the procedures prior to the activity to ensure that the team is following all regulations and minimizing environmental impact. It is important to minimize the chances of harming the environment, to be respectful of both others and nature. 77

87 Table 30: Potential Hazards to the Environment Risk Cause Impact Mitigation RAC w/ Mitigation Items free falling from high altitudes 1) Failed deployment of the recovery system on rocket 2) Failed control of a quadcopter 3) Items becoming loose during flight Depending on the size and mass of the item, it may damage the ground and any other area where the item lands. Ensure that all components are secured as designed and that recovery systems are assembled correctly. 3D Vehicle landing in trees, roofs, or other high locations Public interference Littering Causing fires Hazardous materials being spread into the The rocket drifts unexpectedly far or in a different direction than intended. Loud noises and startling other people If small parts come loose or if a part goes missing Rocket motors, black powder charge, failed batteries. Non-natural materials such as fiberglass, Possible damage to shingles, tree cover, or residential lawns Due to lack of knowledge on the subject, people may be afraid. The parts may not be biodegradable and could affect the natural wildlife. Fires will damage the area and can be very dangerous. Depending on the material the impact can 78 1) Make drift predictions before launches 2) Ensure that the launch site and conditions are optimal to avoid these situations Speakers will be used at all launches to alert the public to an upcoming rocket launch. All components will be checked over before launch to ensure they aren t loose. No flammable material will be located near a launch site or test site. All non-natural materials will be regulated to 4D 4D 3D 2D 1E

88 Risk Cause Impact Mitigation RAC w/ Mitigation environment epoxy, and batteries getting in unwanted places vary anywhere from negligible effects to poisonings. ensure they do not get into the environment. Significant greenhouse gas emissions Transportation with motor vehicles releases carbon dioxide emissions which are harmful to the environment Greenhouse gas emissions and burning of fossil fuels pollutes the air and is said to cause climate change. The team will minimize the number of vehicles being used by carpooling 2D 79

89 Table 31: Potential Environmental Hazards to the Project Risk Cause Impact Mitigation Rocket drifts to far under parachute Equipment could get struck by lightning The rocket lands in a difficult place to retrieve Parts of the rocket could overheat or warp Moving components could freeze due to ice buildup High wind speed Lightning nearby Adverse terrain High temperatures Low temperatures The rocket drifts outside the defined recovery zone. Death or serious injuries can occur from direct or nearby lightning strikes. Equipment could also be damaged. The rocket may not be able to be retrieved without causing damage to the environment. Parts of the body tube may not fit together as designed. In extreme heat the electronics could overheat. The landing legs may not deploy, and the sections may not separate cleanly. Analysis on parachute along with the weight of the rocket was completed to minimize the drift distance for a maximum, safe wind speed. In the case that a thunderstorm approaches as much equipment as possible should be packed up before the storm arrives. Once lightning is seen, or thunder is heard, all personnel should be in a safe location indoors or in a vehicle. The rocket will be designed so that it will land within the designated recovery zone. A strong effort will be made to keep all the components cool if it is a very hot day. If the electronics are on for a long period of time, there will be adequate airflow to keep them cool. Components of the rocket will be inspected for ice before the launch if it is cold enough for ice buildup. RAC w/ Mitigation 4C 1E 3D 2E 2E 80

90 Risk Cause Impact Mitigation Sections of the rocket may swell and cause failed separation events If there is low visibility in the sky, a plane could be hit by a rocket High humidity Low cloud cover The sections of the rocket do not break apart and the rocket comes down without a parachute. An unsafe cloud top could put aviators in grave danger. Tests will be performed on the launch day before the launch to ensure all the components separate successfully. Talcum powder and other methods can be used to reduce the friction between components. No rockets will be launched if they could fly high enough to go through the clouds. RAC w/ Mitigation 2D 2E 81

91 4.7. Project Risk Analysis As with any project of this size and scope, there are a variety of risks and concerns when it comes to the viability and affordability of the project itself. The project requires a certain level of capital both in manpower and dollars that if not met, could results in setbacks or even in the failure of the project. In order to best mitigate these kinds of risks, the team manager has performed a detailed analysis into factors that could affect the project and ways to alleviate these risks. This analysis is presented below as Table 32. In all aspects of the project, the team lead and subteam leads will strive to mitigate project-related risks by implementing the suggestions discussed in this analysis. Generally speaking, experience in previous years has shown that the key to project success lies in strong leadership, effective organization, and committed team members. 82

92 Table 32: Project Risks Risk Cause Impact Mitigation Team loses knowledge in some particular area The team runs out of money for the project The team does not have enough time to finish the project The full scale test flight is not successful Membership decreases substantially Team member leaves project The anticipated funding does not come in Underestimation of what is left to complete Instability or failure of the recovery system Interest in the project fades, school becomes more difficult Team must divert efforts from current design problems onto relearning material. The team will not be able to buy all the components necessary to compete in the competition. The vehicle or payload may not be able to be completed in enough time for a full scale launch before FRR. The vehicle may not be able to be fixed or rebuilt in time to fly another test flight before the competition. There may not be enough people working on the project to successfully complete all the papers and the vehicle. Ensure more than one member is knowledgeable about each section of the project and in doing so ensure a redundancy in team knowledge. The team applies for more funding then what is needed to complete the project in case some of it doesn t come through. The team will start building as early as possible to make sure it is finished in time. Care will be taken to make sure the rocket is stable at all times and that the recovery system is properly constructed. The team lead works to keep everyone interested in and involved with the project. RAC w/ Mitigation 3C 2E 3D 2D 3D 83

93 Risk Cause Impact Mitigation The team loses its lab space to construct the rocket The CIA loses its launch location The university takes away the workspace available to us The owners of the properties decide to stop allowing the launches New projects or construction cause the team to lose their lab space to build the rocket and payload. There will be no local locations for the full scale test launch. There are no known future plans for construction or projects that would cause the team to lose access to the lab. The team will have to travel further away to locations such as Princeton Illinois or Northern Indiana RAC w/ Mitigation 2E 4D 84

94 4.8. Preliminary Checklists In order to ensure that all of the proper steps are taken in the days leading up to launch to ensure a safe and successful flight, the team has devised the following preliminary checklists. While some of these steps are menial in nature and unlikely to be forgotten, it s possible that a trivial, but important, step is missed in the hustle and bustle of launch preparation that could have large ramifications on the quality of the flight or cause significant delays Final Assembly The following checklists detail the steps that must be taken when preparing the launch vehicle for launch in order to ensure that the launch vehicle will perform as expected Recovery Preparation 1. Prepare Recovery Electronics a. Assemble avionics bay payload sleds, check that all connections are secure i. Altimeters should be wired to switches, batteries, and two terminal blocks ii. Insert and connect fresh batteries b. Lock altimeter switches in off position c. Attach e-matches to altimeters via terminal blocks d. Turn on altimeters and check continuity, then turn off altimeters (3 beeps for continuity) e. Slide avionics sleds into couplers and attach bulkheads i. Insert sleds so that the altimeters face the key switches ii. Thread a nut and washer on each bulkhead on each rail (4 total nuts and 4 washers) f. Check that altimeters are off, and attach ejection charges to terminal blocks 2. Pack drogue parachute a. Packing Procedure to be determined through assembly, testing, and practice 3. Insert Drogue parachute into booster airframe a. Attach quick link to eyebolt on motor case, confirm quick link is closed b. Insert wrapped shock cord into booster, pack down firmly c. Push packed drogue parachute into booster, with the protector facing upwards d. If the parachute is too tight, adjust packing to make the package wider or longer as necessary e. Attach the upper quick link to the bottom eyebolt of the avionics bay, ensure the link is closed 4. Pack main parachute and flame retardants a. Packing procedure to be determined through assembly testing and practice 5. Insert main parachute into airframe a. Attach quick link to avionics bay, confirm quick link is closed b. Insert wrapped shock cord into airframe, pack down firmly 85

95 c. Push packed main parachute into airframe, with the protector facing downwards d. If the parachute is too tight, re-adjust packing to make the package wider or longer as necessary Motor Preparation Adapted from Manufacturer's instructions 1. Assemble Forward Closure a. Apply lubricant to all threads and O-rings b. Insert the smoke charge insulator into the smoke charge cavity c. Lubricate one end of the smoke charge element and insert it into the smoke charge cavity 2. Assemble Case a. Deburr the inside edges of the liner tube b. Insert the larger diameter portion of the nozzle into one end of the liner, with the nozzle liner flange seated against the liner. c. Install the propellant grains into the liner, seated against the nozzle grain flange. d. Place the greased forward seal disk O-ring into the groove in the forward seal disk. e. Insert the smaller end of the seal disk into the open end of the liner tube until the seal disk flange is seated against the end of the liner. f. Push the liner assembly into the motor case until the nozzle protrudes approximately 1-3/4 from the end of the case. g. Place the greased forward (1/8" thick X 2-3/4" O.D.) O-ring into the forward (bulkhead) end of the case until it is seated against the forward seal disk h. Thread the forward closure assembly into the forward end of the motor case by hand until it is seated against the case i. Place the greased aft O-ring into the groove in the nozzle j. Thread the aft closure into the aft end of the motor case by hand until it is seated against the case Launch The following checklists represents the steps that need to be taken the day before and the day of launch in order to ensure a safe and productive flight Day Before Launch 1. Check that mentor has: a. Correct Aerotech L1390G-P b. Correct charge size for each separation event. Charge sizes to be determined 2. Check that all flight hardware is stored for transportation to launch site 86

96 a. Booster Airframe b. Motor casing c. Motor Adapter (three pieces) d. Motor Forward Seal Disk e. Main and Drogue Parachutes f. Main and Drogue shock cords g. 8 quick links h. Coupler (assembled with sled, end-cap bulkheads, and altimeters) i. Flat Screwdriver for Rotary Switches j. Upper Airframe k. Screws for upper Airframe attachment l. Payload fairing (assembled with electronics) m. Shear Pins n. Motor retainer ring 3. Check that all backup equipment and tools are prepared to complete any necessary final fixes or alterations a. Phillips screwdriver for screws b. Small screwdriver for altimeter contacts c. Adjustable wrenches d. Allen wrenches 4. Check that all ground support equipment is packed a. Ground Station Antenna b. Laptop with ground station software c. Micro USB Cable d. GPS tracker e. Binoculars 5. Check that all team members have read or heard safety briefing and are informed of their responsibilities Day of Launch Pre-Launch 1. Pack equipment for travel, as listed above 2. Travel to launch location 3. Unpack equipment at launch site 4. Assemble avionics bay and hatch payload sleds, check that all connections are secure a. Altimeters should be wired to switches, batteries, and two terminal blocks each b. Insert and connect fresh batteries 5. Lock altimeter switches in off position 6. Attach e-matches to altimeters via terminal blocks 7. Turn on altimeters and check continuity, then turn off altimeters (3 beeps for continuity) 8. Slide avionics sled into coupler and attach bulkheads 87

97 a. Insert sled so that the altimeters face the key switches b. Thread a nut and washer on each bulkhead on each rail (4 total nuts and 4 washers) 9. Check altimeters are off, and attach ejection charges to terminal blocks 10. Pack drogue parachute Packing procedure to be determined through assembly testing and practice 11. Insert Drogue parachute into booster airframe a. Attach quick link to eyebolt on motor case, confirm quick link is closed b. Insert wrapped shock cord into booster, pack down firmly c. Push packed drogue parachute into booster, with the protector facing upwards d. If the parachute is too tight, adjust packing to make the package wider or longer as necessary e. Attach the upper quick link to the bottom eyebolt of the avionics bay, ensure the link is closed 12. Attach coupler and insert shear pins for drogue parachute a. Refer to labeling on coupler (This end down, shear pin alignment marks) b. Push coupler into booster airframe c. Rotate as necessary to line up shear pin holes d. Insert three shear pins 13. Pack main parachute and flame retardants. Packing procedure to be determined through assembly testing and practice 14. Attach upper airframe to coupler a. Refer to alignment marks b. Insert three screws 15. Insert main parachute into airframe a. Attach quick link to avionics bay, confirm quick link is closed b. Insert wrapped shock cord into airframe, pack down firmly c. Push packed main parachute into airframe, with the protector facing downwards d. If the parachute is too tight, readjust packing to make the package wider or longer as necessary 16. Attach nosecone to upper airframe a. Refer to alignment marks b. Insert four shear pins 17. Insert motor into booster airframe a. Attach the adapter rings (three pieces) b. Insert into motor mount c. Screw on retainer ring, confirming the motor is secure 18. Bring rocket to RSO for safety inspection 19. Make changes as specified by RSO 88

98 Day of Launch Launch Time 1. After RSO approval, wait for range clear 2. When range is clear, move rocket to pad 3. Lower launch rod and mount rocket on the rod a. Ensure team members are supporting the weight of the rocket b. Rail button should slide easily along rail. If not, don't apply pressure, rather rotate the rocket 4. Raise rod and rocket to upright position, be sure to support the rocket while lifting 5. One at a time, turn the key switches; listen for continuity, settings check a. Payload altimeter: i. Verify the altimeter turns on b. Payload computer: i. Verify ground station receiving from transmitter c. Primary altimeter: i. 3,1, 10, 10,10 beeps for main deployment altitude ii. Series of beeps for last flight data iii. Series of beeps for battery voltage (Volts, tenths of Volts) iv. Three quick beeps for continuity d. Secondary Altimeter i. 4,9,0,0 beeps for main deployment altitude ii. 5 second siren for apogee delay iii. Series of beeps for last flight data iv. Series of beeps for battery voltage (Volts, tenths of volts) v. Three quick beeps for continuity 6. Check pad power is off and attach igniter to pad controller 7. Insert igniter into motor and plug 8. Leave range and wait for launch 9. Acquire signals from GPS transmitters and camera system before launch 10. Launch rocket 11. At apogee, wait for separation 12. Wait for rocket to land 13. Upon range clear: retrieve rocket, check for undetonated charges and remove 14. Return to safe area Day of Launch Post Launch 1. Remove altimeters from coupler and collect data 2. Turn off all avionics and store for transport When travel is finished, clean all dirty components, remove power sources from avionics, and store all materials for future flights. 89

99 4.9. Compliance The team agrees to comply with the following safety related codes and requirements described in this section. The safety team will brief the team as a whole on these codes and requirements, and all members will have to individually sign a document stating that they agree to comply with these safety related measures. These documents will be included as an appendix in the following design reviews, and members will not be allowed to help with construction until their name is on these documents NAR High Power Safety Code The team will comply with the High Power Rocket Safety Code provided on the NAR website that has been effective since August The 13 step code and Minimum Distance Table on the website will be reviewed by the safety officer. All members on the team will be required to read the safety code online as it is a relatively short list of codes. The rules set forth by the NAR High Power Rocketry Code will always be respected and followed as they are set to ensure the safety of people and the environment. The safety officer, team manager, and subteam managers will always make sure to comply with the safety code and ensure the rest of the team is properly complying. A copy of the NAR High Power Rocketry Code is included in this report as Appendix F: Nar High Power Rocket Safety Code Federal Aviation Requirements The team will comply with all laws and regulations set forth by the FAA in terms of using airspace for test launches and quadcopter flights. The team s safety officer will be responsible for educating all involved members of the regulations regarding the use of airspace: Federal Aviation Regulations 14 CFR, Subchapter F, Part 101, Subpart C; Amateur Rockets, Code of Federal Regulation 27 Part 55: Commerce in Explosives; and fire prevention, NFPA1127, Code for High Power Rocket Motors. ; as well as all applicable laws. ISS will be contacting the FAA before any test flights are done, but only after having approval from the local RSO. All of the flights will be suborbital, remain in the United States, and be evaluated and deemed safe for all members of the team and community. A copy of this code is located on the shared team google drive and will be posted on the team website once that is completed Range Safety Officer Authority The team will comply with the range safety officer at the competition launch in Huntsville, AL and at the test launches for the vehicle. All team members present at the test launches and competition launch will be instructed to listen to all instructions given by the range safety officer and will understand that the officer has the final say on whether or not the final rocket flies. Members will also be given a briefing before attending launches by the safety officer and other experienced rocketry personnel about launch field expectations and safety protocol. Also, team members participating in the preparation of the rocket on launch day will create a procedures list prior to the date. 90

100 5. PAYLOAD CRITERIA 5.1. Payload Overview The team will develop a payload capable of meeting the requirements of the Landing Detection and Controlled Landing topic for the NASA Student Launch competition. This technical challenge has relevance to recent developments in reusable launch vehicles as well as the field of planetary entry, descent, and landing. The payload designed by the team will test image processing techniques onboard an aerospace vehicle. After the images have been collected, a technique will be tested to land the payload section upright (in the same orientation it was at launch). Requirements for the successful completion of this payload have been identified from the NASA Student Launch handbook. Additional requirements have been placed on the payload internally, by the team, to create a clear set of goals moving forward. All of these requirements are summarized in Table 33 below. Requirement Each team shall choose one design experiment option. An additional experiment may be flown. If the team choose to fly additional experiments, they shall provide appropriate documentation. The payload must be capable of identifying three tarps on the launch field. The payload must be able to differentiate these three tarps based on color. All data processing must occur in real time during flight. The output of the data processing may be collected after recovery of the rocket. Table 33: Payload Requirements Requirement Source Payload Requirements Payload Requirements Payload Requirements Payload Requirements Payload Requirements Payload Requirements Verification Method Inspection- Team has chosen to develop a payload for experiment option 1. Inspection Team has not opted to fly two experimental payloads. Inspection Team has not opted to fly two experimental payloads. Post-flight human analysis of the images. Post-flight human analysis of the images. Software will write text file with in-flight altitudes of each image to verify processing occurs during flight. Section Addressed 5.1 N/A N/A The software performing the Payload The team lead and payload 5.4.7

101 Requirement data processing may make use of open source image processing libraries, but otherwise should be custom made. The launch vehicle section containing the camera shall land upright. Payload must be able to function for at least 90 minutes. The payload system shall fit in a 6 inch or smaller diameter rocket. The payload system shall have a weight of less than 3 lb. Requirement Source Requirements Payload Requirements Vehicle Requirements 1.8 Internal Internal Verification Method manager will oversee the creation of the software to ensure custom software is produced following open source guidelines. A combination of onboard video and day of flight observations will be used to ensure an upright landing takes place. Testing prior to launch date of the full system operating for 90 minutes. The system will be designed to this constraint, will be verified by team lead at each design milestone to ensure compliance. Payload will be weighed after construction completes, weight will be reported to team lead and structures manager. Section Addressed The system designed to meet these requirements will be located above the main coupler in the rocket body, and will deploy at apogee. Two cameras built into each side of the rocket will image downwards during flight, allowing them to capture the ground targets both on ascent and under parachute. The cameras will feed their images to the set of onboard Raspberry Pi Zeros, where they will be analyzed using custom software, constructed by the team, to locate the tarps within the images. Altimeters for the Pi Zeros will be used to determine relative tarp sizing by correlating the altitude with the images taken. This will work as to not confuse tents or other objects for the tarps. This software will make use of image processing algorithms to identify regions of the images captured that share similar color characteristics. Testing by the team using the samples provided will help define a range of colors that the system will positively identify. This section of the rocket will complete an upright final landing making use of four spring loaded landing legs. These landing legs will be contained within a coupler section of tubing, which will be loaded into the body tube to ensure they do not deploy during ascent. They will 92

102 extend immediately after the section is jettisoned at apogee by the ejection charges. A rendering of the full system can be seen in Figure 37 below. Figure 37: Rendering of the full payload system in its launch configuration. The payload section is designed to have a total weight of less than three pounds, according to the internal team requirement. The payload section is comprised of the mechanical landing legs and the camera system. The total electronics hardware and camera system will weigh about 1.73 lbs. The landing leg system will consist of 4 mechanical legs, each comprised of an inner and outer segment with torsion springs, held together with various nuts and screws. The landing legs have a total weight of 0.94 lb., leading to an estimated system weight of 2.67 lb. This leaves a 12% margin available for mass growth below the 3 lb. subsystem requirement. The full component level mass budget for the current design can be seen in Section of the Structures and Recovery section of the proposal. The structural mass of the rocket encapsulating this system (i.e. body tubes, bulkheads) has been accounted for separately in the structures and recovery mass. The specifics of the current design for both the mechanical landing subsystem and the image processing subsystem are included in the following sections. 93

103 5.2. Payload Success Criteria The team has identified a clear set of requirements to define the success of the image processing payload. These requirements have been defined to be as specific as possible, calling for specific thresholds to be met. Table 34 below lists both the requirements the team has chosen as the success criteria for the project, as well as the verification process that will be used on the day of launch to determine if these requirements are met. Table 34. Success Criteria for the Image Processing Payload Requirement The section of the rocket between the coupler and upper airframe shall contain the camera and land upright. Upright is defined as in the same configuration as during launch. Landing at an angle is permitted, but at least three of the four landing legs must make contact with the ground. This upright landing must be maintained for at least 10 seconds after touching down. The system must obtain at least two clear images of each of the three tarps. The system will not be penalized for missing the tarps in some of the captured frames. During flight (ascent or descent) the system will identify the three separate target tarps on the launch field. Images designated as containing these tarps must be accurate (in terms of existence, location, size, and number of tarps) at least 60% of the time. The system must be capable of differentiating between the three colored tarps on the launch field. Of the images correctly designated according to the above criteria, at least 90% of them must correctly label all of the imaged tarps by color. Verification Strategy The onboard imagers will capture images from the final phase of descent. These images will be checked to ensure the payload did not roll upon landing. Upon recovery of the system, images will be taken by the team of the landing configuration prior to retrieval. Should the payload be pulled over by the parachute after landing, the onboard video will be studied to verify successful landing. Video from the onboard cameras will be started below an altitude of 200 feet upon descent, allowing the images during landing to serve as proof. The footage will be analyzed after flight to ensure the system stays upright for at least 10 seconds. The onboard computer will save all images in which it detects tarps. The software will save both the original image, and an image with the borders of the tarps (as identified by the image processing software) marked. This border will also be color coded to indicate which tarp is which. The true existence and correct identification of the tarps in the images will be validated by the University of Illinois Student Launch team lead and the payload subteam manager. If these team members are in disagreement, the image will be considered to not have correctly detected a tarp. 94

104 5.3. Mechanical Landing Subsystem Landing Subsystem Selection Process The landing subsystem has the responsibilities of maintaining the camera payload section in a vertical orientation during landing, fitting into a six-inch body tube during flight, having the durability to survive multiple flights, and having a low mass. Landing system concepts have been evaluated on these requirements in addition to the criteria of (1) compact storage during boosted flight, (2) deployment complexity, (3) landing stability on slopes up to 40 degrees, and (4) drag produced by leg system. Three separate concepts were developed by the team and their relative merit was studied. These concepts are summarized below: Option 1: The Legs will be stored externally with torsion springs for deployment. This would decrease the required internal space, but would cause a small amount of drag. They would be held closed by the coupler, when the stages separate the legs will be free to rotate. Option 2: Legs fold and slide into rear of payload bay tube. Increased payload bay size required for storage with benefit of no additional drag during boost. Compression springs would push the leg assembly downward past the bottom of the upper stage. Then a pivot with a torsion spring would be required to double leg length. This would require the most amount of internal space. Option 3: Legs would be stored internally with two pivots. They would fold through body tube held in place by the coupler. When the stages separate the legs are free to pivot out of the tube. The second pivot is internal of the legs and allow them to unfold and double in length. Much less complex than the second option while providing the same benefits. Table 35. Landing Mechanism Trade Study Option Internal Storage Complexity Leg Length Drag Produced 1 Minimal Minimal Limited to Payload Bay Height Limited 2 Large Feasible Limited to Payload Bay Height None 3 Small Minimal Limited to Coupler Length None Chosen Landing Subsystem Overview Option 3 has been chosen as the baseline design for the payload landing system. This design provides excellent stability and reliability while still maintaining a relatively simple construction. The current design can be seen in its deployed configuration in Figure

105 Figure 38: The landing system in its fully deployed configuration. The landing leg system is composed of 4 legs, each with a pivot mounted to a bulkhead acting as the first joint. The bulkhead will act as a stable surface to mount the legs to, and serve as a barrier to protect the leg system from the parachute ejection charges. Both pivot points will consist of a bolt held in place with a nut (refer to Figure 39); this will make it easy to replace the legs in case one is bent or damaged in any way. Each leg will also have an extrusion to prevent the legs from over extending. The inner pivot mounts will be mounted to the bulk head by two pins going through the whole piece. The option of having these pivot mounts made from aluminum with two threaded holes at the bottom is being explored. this would make it easier to design around the spring. When folded into the upper body tube there will be one contact point between the outer leg and the interior wall of the coupler body tube. This surface will be polished to ensure minimal friction during leg deployment. The legs will fold out through the payload stage coupler; requiring 4 slots 0.5 wide in the coupler for the legs to fold through. These slots will be covered by the booster stage body tube during launch. After separation at apogee, the coupler slots will be clear of booster tube, allowing the legs to deploy. When folded, the legs will have a total width of 0.5. There will be a torsion spring located inside each pivot to assist the unfolding process and maintain an open position during descent/impact. The leg section attached to the bulkhead will be 5 ¾ in length and the distance between each pivot will be 5. The second leg section will be 5 ¾ in length. The inner leg will have a consistent height of ¾ and the outer leg will have a height of ¾ with a downward taper to approximately ¼ at leg-tip. The radius of the deployed legs will be 12 from the center of the rocket. As the outer leg unfolds there will be a contact surface on both the inner and outer legs to prevent the leg segment pivot from overextending beyond 180 degrees. The legs will be custom machined of aluminum, and testing will determine if other aluminum is adequate or if other materials are required. The total weight of the leg system will be approximately 0.91 lbs. 96

106 The payload/landing system will not face interference from the tethered nose cone during descent if the tether (shock cord) is of sufficient length. Figure 39: Exploded view of the landing leg connection mechanism Tip-over Analysis of Landing System With the selected landing system, the possibility exists that the landing legs will not be sufficient to keep the landing stage upright during the landing process. An increase in the landing section s horizontal velocity could mean failure of the landing system. This possibility was analyzed in-depth, and a simulation was designed in Python to model the various possible horizontal drift speeds, and compare them to the angle need for the landing system to fail. Figure 40: Minimum tipping angle over a range of horizontal speeds. The angle of tipping represents the minimum possible angle that the landing stage s centerline forms with the ground s normal in which the landing system would fail. Any angle 97

107 below the graphed line means that the landing stage will not tip over. At a horizontal velocity of approximately 17 ft/s, or 5.2 m/s, the landing system will fail when the landing stage is perpendicular to the ground. The lowest estimate for the downward velocity component was used, to simulate a worst case scenario. Table 36: Tip Over Angle Horizontal Velocity [m/s] Horizontal Velocity [ft/s] Tip Over Angle Launch Vehicle Integration The upper stage has four slots in the coupler, for each of the legs to unfold through. The legs are folded into coupler as it is loaded into the booster stage. They will be held in place by hand as the booster body tube is placed over the upper stage coupler. The reaction force from the booster body tube will keep the legs folded and compressed. There will be one contact point between each leg and the booster stage body tube. That contact point will be a polished smooth surface to minimize the friction between the legs and the body tube. The team will test to ensure friction forces will not interfere with stage separation. An alternate retaining solution will be developed if there is danger of stage separation jamming. The landing leg subsystem will be deployed as the booster stage separates from the upper stage. The separation will occur at apogee as the black powder ejection charge is ignited and parachutes are released. The legs are attached to a bulkhead to protect them from the black powder charges. As the booster body tube separates from the upper stage, the torsion springs will cause the legs to unfold. These legs are then held extended by spring tension. This process is shown in Figure 41 below. Future testing will determine if a locking mechanism, besides spring tension, will be necessary to prevent collapse. 98

108 Figure 41: Landing leg deployment sequence Prototyping and Testing Testing the hardware of the payload will be done to verify the landing capabilities of the upper airframe of the vehicle. There will be two parts to testing the payload landing system: preliminary testing and flight testing. Preliminary testing of the system will be done separate from the vehicle while the flight testing will be done during the test flight of the vehicle. Preliminary testing will validate the different system components separately and then the system as a whole. Preliminary testing will be considered a success if all components operate as desired and the system can complete the tasks in a test environment. Flight testing will be considered a success if the system deploys fully at apogee and lands upright. Table 37 below shows the tests that will and have been performed and the results from the completed tests. As a preliminary step, the designed landing system has been created using additive manufacturing. The technical drawings used to machine the aluminum legs can be seen below in Figure 42 and Figure 43. While the legs of the final version will be made of aluminum, the team tested torsion springs of various strengths on a PLA 3D printed prototype of the subsystem. These 3D printed prototypes can be seen in Figure 44 and Figure 45 below. The PLA landing legs deploy properly under the power of the torsion springs. The first iteration of the printed legs was unable to lock in the extended position as desired. The second iteration of the landing legs had extra material added to the first joint new the hinge between the two leg segments. This material made it impossible for the second segment of the leg to bend past the first segment. The prototype subsystem will be stress-tested by dropping it from various heights with estimated payload weights before the Critical Design Review. Drop heights will be chosen for calculated impact velocity. These tests will allow the team to validate some parts of the design, especially the ability of the system to withstand impact, and to land in the correct orientation on large slopes. Mass will be added as necessary to ballast the system to the expected flight configuration. 99

109 Figure 42: Upper leg technical drawing. Figure 43: Lower leg technical drawing 100

110 Figure 44: Prototype 3D printed legs with springs attached. Figure 45: Part of the assembly for the prototype landing system. Since the maximum stress of aluminum is 45,000 psi and the weight of the upper airframe is less than 10 lbs., stress testing the aluminum of the landing legs will not be necessary. The hinges used to extend the landing legs will have forces applied to them to test the joint rigidity and strength. The springs used for the deployment of the landing legs will be attached to a force gauge and extended to the designed length to ensure they produce enough force for deployment. The joint between the legs and the upper airframe of the vehicle will also be tested for impact resistance. The hinge attached to the upper airframe will also have forces applied to ensure it has the impact resistance needed during the landing of the payload section. 101

111 Once the individual components have been tested and the system has been constructed, systems of the payload will be tested. The deployment of the legs will be tested first. The upper airframe will be lowered into another six-inch diameter body tube with the legs folded inside. The upper airframe will then be pulled free of the lower section, allowing the landing legs to deploy. The legs must deploy to their full length and not interfere with the separation from the body tube for this test to be considered a success. The upper airframe will then be subjected to a vertical stability test. The system will be balanced at different angles to determine the angle at which the system can land before it tips over. The system will also be dropped from a sufficient height to replicate its velocity under parachute onto sloped surfaces to ensure the system will not tip over due to hazards in the field. The next test will test system landing ability under the descent of a parachute. The system will be dropped from a height that will allow full parachute inflation. This will be the last test before the flight test. For the flight test, the landing system will be loaded into the vehicle during launch setup. The system will be launched and the payload will be deployed at apogee. It will then descend under a parachute and land. The flight will be used to validate the successful deployment of the legs after drogue parachute ejection, as well as the ability of the system to perform a fully upright landing under flight conditions Testing Deliverables The results of the tests will be recorded in Table 37 as tests are completed. If the test is unsuccessful, the outcome will still be recorded so improvements to the system can be made. After improvements have been made, the tests will be rerun until the whole system works as desired. Since the preliminary testing will start at the component level, testing may be done in a different order than what is listed below. Table 37: Landing Legs Subsystem Testing Test Requirements Results Plastic Leg Prototyping Extended Leg Strength 3D printed legs must deploy from a folded position stay extended under the power of a spring The extended landing legs should withstand the expected landing forces. Hinge Strength The hinge connecting the landing legs to the body tube The leg prototypes showed that the spring deployment scheme and the locking method on the second hinge works as desired. Date Completed 10/15 /

112 Test Requirements Results Spring Strength Leg Deployment Vertical Stability Test Parachute Descent Complete Flight Test must withstand the force of landing. The deployment spring should supply a sufficient force to extend landing legs immediately after drogue ejection. The landing legs should separate from a 6 inch body tube and deploy to their full length. The payload must return to a stable upright orientation after tilting as much as 40 degrees from vertical. The payload must descend under a parachute and land upright. The payload must deploy landing legs to full length at apogee, and land upright. Date Completed 5.4. Image Processing Subsystem Subsystem Overview The image processing subsystem consists of two cameras each connected to an independent Raspberry Pi Zero. This subsystem will be activated by an external switch which will control power to all of the components. The cameras will be positioned on opposite sides of the payload bay angled downward to increase the ground area captured in frame. These cameras will be capturing images at an estimated rate of one frame per second, beginning at activation. The Raspberry Pi Zeros will then run those captured images through the team s image processing software to determine the presence or absence of the targets. Altimeters specific to this subsystem will be used, one for each camera, to determine the expected size of the tarps in the images. If the software detects the targets in an image, it will store the image and relevant proof of detection onto the SD card storage device. The proof of identification and differentiation can later be presented on any machine capable of interfacing with the SD card. Each of the redundant systems will have its own battery, which will power the Raspberry Pi Zero, camera module, and 103

113 altimeter. An electrical schematic and subsystem hardware diagram can be seen in the two figures below. Figure 46: Full electrical schematic for the image processing subsystem. 104

114 Figure 47: Block diagram of the fully redundant image processing subsystem Processing Unit Selection The processing unit in the image processing system has the responsibilities of receiving images from the camera, running the image processing software in real-time, and storing a form of proof of identification and differentiation of the targets. To this end, the processing unit possibilities have been evaluated on the criteria of (1) image processing capability, (2) difficulty of implementing image processing, (3) difficulty of interfacing with a camera, and (4) difficulty of interfacing with memory to store the proof. Cost, size, and weight were also considered, but fell well under the limits for each option, and are therefore considered negligible in this comparison. The team has evaluated a few hardware options that could potentially fulfill all requirements: the Trenz Electronic TE0711 FPGA, the Arduino UNO board, and the Raspberry Pi Zero board. Table 38 below shows the comparison of each of these options with more detail immediately following. 105

115 Table 38: Processing Unit Trade Study Processor Unit Processing Implementation Camera Memory TE0711 Sufficient Difficult Difficult Difficult Arduino Uno Not Possible Not Feasible Simple Simple Raspberry Pi Zero Sufficient Simple Trivial Trivial The TE0711 FPGA would have a fixed and parallel hardware pipeline for image processing which would be beneficial, as each captured image requires the same processing. Because of the lack of native support, implementation would become dramatically more difficult for the programming, camera interface, and memory interface. The Arduino Uno would allow a simple interface with a camera due to the available shields, but it fails to be acceptable in any of the other categories. The board does not have the processing power required for the complex image processing the system will be performing, and memory storage is smaller than would be required to store relevant proof. The Raspberry Pi Zero has sufficient processing power for the team s purposes. On top of that, the team will be able to utilize available image processing libraries to simplify the implementation. The hardware requirements of interfacing with a camera and memory are also built into the board, rendering each of these tasks trivial. Using the criteria the team has created, the best option is the Raspberry Pi Zero. It was chosen because it is able to interface with cameras, has the ability to process images on-board, can run image processing libraries, and has optional large storage capacity. In addition, the small size of the board (2.6" x 1.2" x 0.2") coupled with the light weight (0.3 oz) make this an excellent choice. While both other options have promising aspects, they were unable to satisfy all requirements. The team discussed using a single Raspberry Pi with a multiple camera module adapter instead of two Raspberry Pi Zeros. However, the multiple camera module adaptor does not allow operation of cameras simultaneously. The Raspberry Pi could perhaps be programmed to alternate activating the two cameras, but this would increase complexity of the software, and also decrease rate at which images can be analyzed from each camera. In addition, the weight of two Raspberry Pi Zeros is still less than the weight of the other, bigger Raspberry Pi models. Another factor in favor of the two separate Raspberry Pi Zeros is that two of them without the multi camera module adaptor is still less expensive than a single Raspberry Pi with the multi camera adaptor Camera Selection The two cameras selected will be located opposite of each other in the payload bay of the rocket. Both of these cameras will be positioned looking out of the rocket, at a downward angle of 45 degrees, increasing the amount of ground area searched. Even if the rocket does not fly straight, the field of view will be large enough to account for variations in the flight path of the 106

116 rocket. This angle will be optimized upon further analysis to provide the largest probability of sighting the ground targets. The cameras chosen for this system are identical Raspberry Pi Camera Module V2s. These cameras provide images with a resolution of up to 8 MP to the Raspberry Pi Zero. This system interfaces well with the Pi, allowing the team to focus efforts on developing the required code. Different models of the Raspberry Pi camera module were considered, but the Camera Module V2 provides a relatively large field of view, while not causing distortion due to a fish-eye lens. The full trade between these options can be seen in Table 39 below. Table 39: Image Processing Camera Trade Study Camera Name Resolution Field of View Raspberry Pi 5MP Camera Board Module Photo: 5 MP Video: 1080p30 Horizontal: Vertical: Linksprite Raspberry Pi 5MP Wide Angle Photo: 5 MP Video: 1080p Horizontal: 160 Vertical: 160 Raspberry Pi Camera Module V2 Photo: 8 MP Video: 1080p60 Horizontal: 62.2 Vertical: 48.8 Both cameras will be located safely inside the rocket with no parts protruding from the payload compartment, per competition rules. To see out of the rocket, acrylic windows will be installed into this compartment. This will be done by cutting 2 x 2 windows out of the body tube for the cameras to look out of. Pieces of the acrylic tube will cover the windows from the inside of the body tube. To not compromise that section of the rocket, the windows will have rounded corners to reduce the chance of fractures occurring. The acrylic will be bolted to the body tube in order to rigidly attach it to the rocket. The acrylic will extend further than the area of the window, both to strengthen where material was removed and to not cause too much additional weakness from where it is bolted. A diagram of this configuration can be seen on the next page in Figure

117 Figure 48: Diagram of the payload cameras, showing the estimated available line of sight of the cameras in blue Camera Position Each Raspberry Pi Zero system will be equipped with an altimeter to enable an expected tarp size to be calculated by the image processing system. This will ensure that the system does not mistake similar objects nearby - namely team tents or vehicles - for tarps and incorrectly categorize them. The specific altimeter to be used will be an Altimeter Module MS5607 due to its low cost, size, and weight, as well as accuracy and power consumption. Each altimeter will be connected directly to the Raspberry Pi Zero and not to the battery. This ensures consistent voltage to the altimeter, so as limit the chance of an incorrect reading from the altimeter. The team is also considering using the altimeter as a means of triggering the activation of the image processing software, should the team deem the current plan to be infeasible. See Section for details on the currently planned triggering mechanism. The camera module has a viewing angle of 48.8 x 62.2 ; to maximize the time that the rocket can see the tarps, the camera is placed such that the 62.2 angle is placed to scan the ground moving away from the rocket (that is, the Camera Field of View shown in Figure 28 is 62.2 ). Simulations performed using Python show that the optimal angle for the center of view is between 31 and 34 and is shown in the following figure. 108

118 Figure 49: Camera angle optimization simulation based on tarp distance from launch. This data follows a lot of assumptions, such that the rocket is launched and remains in a perfectly vertical flight path from launch, and that the rocket does not sway while descending. These assumptions are reasonable enough for a baseline idea of how sensitive the camera angle is for tarp sighting. While there is some concern about how sensitive the data appears to be, the team believes that even with the swaying, being able to see the tarps for even 50% of the time should be sufficient given how powerful the software being used is. Even with the sway, if the rocket sways one side it must sway back to other, giving the camera a chance to still catch and image of the tarp regardless. Drift is going to minimized due to a low deployment of the main parachute, and identification of the tarps should be successful by this point. As for the choices in what angular values to simulate over, 31 is the smallest theoretical angle the camera could be set at, which would be that the field of view would be flush with the side of the rocket. The upper limit of 45 is more or less arbitrary, but based on the idea that with a reasonable assumption that the rocket could sway about 15 - the camera angle would be viewing far past the 300 ft tarp distance, in fact the upper bound of the viewing would be parallel with the horizon, which is essentially useless to the mission. 109

119 Geometrically, it will be impossible to position the camera such that the field of view is flush with the side of the rocket, so in lieu of that, along with the information gathered from the simulation, the camera will be placed as closely to 31 as possible to maximize viewing time for the tarps even with some sway Power Delivery Using the Raspberry Pi Zero system with images captured via the camera attachment, along with the power needed for the altimeter, an 11.1 V 500 mah battery will be sufficient for each half of the redundant system. The energy required over 90 minutes from the time the system is powered on at the pad until after landing is significantly less than the amount provided by this battery. If this proves to be an issue after further design revisions, the team is considering changing the triggering mechanism to be activated by the altimeters of the image processing subsystem instead of being turned on immediately at the pad. The following table contains a summary of the energy required by this system during its 90 minutes of operational time. Table 40: Image Processing System Energy Budget Module Energy Required [mwh] Raspberry Pi Zero and Camera Module 1,725 Altimeter Module 10 Discharge and Distribution Inefficiency [85%] 306 Voltage Conversion Inefficiency [90%] 227 Total Energy Required 2,268 Battery Capacity 5,550 Energy Margin 3,282 The Raspberry Pi Zero will directly provide power to both the altimeter and camera via its connectors. The Pi itself will be powered by the battery, however, the input voltage on the Pi is 5 V, while the battery supplies power at a nominal voltage of 11.1 V. A DC-DC voltage converter will be used to lower the voltage. The team has past experience using these converters and has full confidence that they will fulfil the role required and function successfully during flight. The above energy budget includes an assumed 90% efficiency for this voltage conversion process Software The image processing subsystem will run onboard a Raspberry Pi Zero using the Raspbian Jessie Lite operating system, which will be imaged onto a microsd card. Raspbian is the Raspberry Pi Foundation's official supported operating system based on the stable distribution of Debian Linux. Raspbian comes pre-installed with vital dependencies for the image processing such as Python-based software and libraries, the language in which the software will be written. Raspbian Jessie in particular is an updated version of the operating system image that boasts 110

120 improved performance and flexibility in system processes. Commonly installed tools and applications come readily incorporated such as general screenshot and camera capture functionality. It also has added support for Pi peripherals, such as the camera module. Jessie features easy accessibility as a standard user to the 40 GPIO lines on the Pi Zero hardware through Python. This includes the newer and streamlined GPIO Zero libraries. GPIO stands for General Purpose Input and Output and this feature offers the possibility of connecting to motors, sensors, and other external components to extend the system's functionality or integrating it with the avionics subsystem. Raspbian Jessie Lite is the distribution's minimal headless image version, meaning that that the operating system boots directly and only into the command line rather than a heavy graphical user interface (GUI) familiar to modern day computing. The installation of this image has multiple advantages. Jessie Lite is a much more lightweight version of the full Jessie distribution. It is roughly 1 GB smaller in size and requires less accessible RAM at any given time. This allows for more ideal allotment of the Pi's resources to the necessary processes for the image processing. The backbone of the image processing system will be built using the Linux open source computer vision framework SimpleCV, a wrapper that uses Python and the open source computer vision and machine learning software library, OpenCV. With more than 2,500 optimized algorithms, OpenCV is focused in its infrastructure toward real-time vision applications. In both syntax and structure, the SimpleCV wrapper makes more abstract functionalities of the OpenCV library approachable and easy to implement in practice. The framework internally calls harder to use functions from the OpenCV software, simplifying the user interface. It provides the necessary techniques that will be used in the team's custom codebase such as image transformations, morphological filters, image arithmetic, color spaces, feature extraction for object detection, and recognition. SimpleCV also has the ability to calibrate camera modules, which provides the function to measure the dimensions of objects in computer vision using both intrinsic and extrinsic parameters such as focal length and determining the location of two images relative to each other, respectively. In order to ensure a robust and flexible installation of SimpleCV, the framework requires access to a few additional libraries. These libraries will allow the computer to have an interactive shell, do engineering computations in Python, package and process data, and more. After the dependencies have been properly set up, SimpleCV is then installed using pip from its Github repository. SimpleCV s features package includes the Blob module, a series of functions that use what is referred to as a blob as a parameter. SimpleCV s documentation defines a blob as a a cluster of pixels that form a feature or unique shape that allows it to be distinguished from the rest of the image. The identification of the three separate tarps will hinge on analyzing and processing various properties of pinpointed blobs in the camera s field of view. 111

121 After the system has fully booted, the competition code will direct the camera to collect an image and query the altimeter for the current altitude. The custom algorithm designed by the team will then take in both pieces of data and attempt to find regions of color similar to a preprogrammed set of RGB values for each target. This set of values will be established during the testing and verification of the system. The altimeter will be used to bound the expected size of the ground targets. Once the image processing software has completed its run, if one or more of the ground targets was identified in the image, it will be saved to the onboard MicroSD card, along with a second image highlighting and differentiating the region(s) that the software identified. If no targets were identified the image will be discarded. This process will repeat until the system is powered down upon retrieval. A block diagram of the planned code can be found in Figure 50. Figure 50: Image processing software overview Altitude Adjustment The changing altitude is a problem with determining the tarp size in the images picked up by the camera, as it is not a static matter. The number of pixels that each tarp should occupy is dependent upon the altitude and camera angle. The whole point of determining this rough estimate of pixel size is to differentiate the tarps from any other object that may be in the surrounding area that could be mistaken for a tarp the most likely object being a team s tent. This information will be kept in mind when programming the Raspberry in conjunction with the altimeter to ensure that any images processed from the camera module are in fact images of the tarps and not images of another object. The altimeter will feed altitude data to the 112

122 Raspberry, which will interpret this information to search for objects within a range of pixel sizes, determined by uncertainty in drift and camera angle affected by sway. A simulation was created in MATLAB to determine the expected tarp size for the given camera field of view and resolution over a large range of potential positions. Trajectories of the rocket flight have been superimposed over this contour plot to provide examples of the expected tarp sizes throughout a typical flight in Figure 51. Figure 51: Simulation showing the expected tarp sizes throughout sample flight profiles. Assessing the estimated size of the tarps over a range wind speeds and tarp locations will allow the team to bound the expected tarp size at each altitude. This analysis will be performed prior to flight and will be proved during the test flight Triggering Mechanism The image processing software will be activated at the launch pad via rotary switch on the exterior of the rocket. In one position, the switch will not allow power to flow from the power source to the rest of the image processing system. In the other position, two power sources (one for each Raspberry Pi Zero) will both be allowed to supply power to the two cameras and the rest 113

123 of the system. This means that once the switch is triggered, the image processing software will be on and working from setup on launch rail, until the team recovers the rocket post-landing. This method poses no problem for the system, as the tracking software does not store any images into the memory unless the tarps are located. In the case of incorrect identification, failsafes will be put into place to prevent the system from exceeding its memory capacity. The only concern of running the software from setup is power supply, and assuming that the total time from launch setup to recovery is at maximum 90 minutes, the team is confident that the batteries used in the system could run continuously for this duration of time. The power requirements for the system are discussed in Section below Prototyping and Testing The testing of the ground imaging package will initially be done separately from the testing of the landing system because the two systems operate independent of each other. Because a lot of the ground imaging system is software based, the subsystems can easily be tested. Ground tests and a flight test will be conducted to identify which specific components need to be improved and which components function as desired. A lot of specific tests can be simulated easily without a full scale test flight with the vehicle, so they can be done earlier without waiting for the completion of the rocket or the landing leg system. For example, the camera operation and visual quality test can be done immediately when the camera is purchased and in the possession of the team. This allows early and meticulous testing of the ground imaging and target identification functions. Tests will continually be conducted to identify errors and areas in need of improvement in both the script and electronics. As the ground imaging subsystems testing continues and all the different components have been tested independently, the integration between the different components will also be tested. The main function that will be tested is the compatibility between the Raspberry Pi Zero, camera, altimeter, and power source. Tests will be conducted to make sure that each of the electronic components work successfully with each other to fulfill the competition requirements. Both the hardware and software of the ground imaging system will be tested in depth to ensure that it successfully works repeatedly and in various situations. The various environments that the system will be tested in include different types of motion, speed, lighting, background color, and other parameters that gives the system more flexibility to operate at the launch. In addition, as the other subsystems of the vehicle are completed, the integration between the subsystems will gradually be tested. First, the integration of the ground imaging subsystem with the vehicle body tube will be tested. The main criteria for this test is to confirm that the electronic hardware will fit comfortably inside the rocket on the avionics sled, that the camera can take clear footage out the acrylic windows, and that the electronics inside the body tube can be easily and securely positioned. Second, the integration with the landing legs will be tested. The ground imaging and landing leg subsystems are completely independent of each other, but a few things can be tested between them. This includes that both the landing leg and ground imaging components can fit comfortably into the rocket body tube and that the legs in their 114

124 deployed state do not interfere with the target identification. Final tests of the ground imaging subsystem will be done aboard a quadcopter and the rocket at the test flight. These two tests will allow the team to have an idea on how the system will operate in a similar environment as the competition. The aerial tests of the target identification function will give the team a better idea on how well the hardware and software operates, and where improvements can be made. Specific testing procedures will be formulated as the project continues and the team receives parts and begins construction. The procedures will include step by step instructions on how to operate the tests and the specific requirements and goals for each of them. The critical technologies for competition success will be tested on multiple occasions to ensure repetitive success. Stress tests will be conducted to ensure that the ground imaging subsystem will be able to safely withstand forces during the launch. Through the tests and the results, the team will identify areas that need changes or that can have improvements for the hardware and software aspects of the ground imaging subsystem. Testing will be continued until the team is confident that the system can successfully identify the targets on the ground with accuracy and flexibility. Table 41 identifies the specific tests that will be conducted and the goals of the test for the ground imaging payload. Most of the tests are designed to be independent from each other so each task will be conducted separately. The tests are ordered roughly in chronological order considering the timeline of the project. Each test and its result will be documented carefully through pictures, videos, and post-test written reviews. The results and date completed columns of the table below will be filled in as the testing phase proceeds. The team will analyze the test results and identify where improvements to the ground imaging subsystem can be made. Table 41: Ground Imaging Subsystem Tests Test Main Requirements Results Camera footage quality Camera operation Raspberry Pi The camera shall capture quality video and images to be able to identify solid colors from up to one mile away. The footage should have an adequate enough resolution to allow multiple pixels for the tarps on the ground. The camera shall be able to analyze video and take screenshot captures of the footage. The Raspberry Pi Zero shall be able to The camera module has been used and images have been captured. A test image was saved onto the microsd and was viewed on a computer Date Completed 10/24/

125 Test Main Requirements Results hardware Raspberry Pi software and target identification Altitude vs tarp size comparison Camera view through acrylic window Ground imaging hardware integration with avionics sled Landing leg subsystem integration test Stress tests Quadcopter flight test analyze images retrieved from the camera, data received from the altimeter, save images to the SD card, and be powered from the battery. Utilizing the algorithm, the Raspberry Pi Zero shall be able to identify when the targets are in its field of vision and command the camera to take a screenshot. The camera and altimeter shall communicate the size of the target it is searching for and the altitude/distance from target to the Raspberry Pi Zero. The camera shall be able to take quality footage through the acrylic window on the body tube. The effect of the curvature of the window will be identified and addressed. Also, the camera shall be able to capture a wide image of the ground below it. The camera and other hardware shall be able to fit comfortably and securely on the avionics sled. The position on the camera will be carefully adjusted to optimize view of the ground. The deployed landing legs shall not interfere with the target identification function of the system. The leg will be in the field of view of the camera but it should not be identified as a target. The entire ground imaging subsystem will be inserted into the rocket body tube on an avionics leg and shall be able to withstand motion, rotation, and vibration in all directions. The ground imaging equipment will be attached to a quadcopter and flown. Date Completed 116

126 Test Main Requirements Results Complete flight test Sample targets will be placed on the ground and the system shall capture images of the targets when they are identified. Various conditions can be tested because quadcopters are very versatile. The ground imaging equipment will be integrated with the rest of the payload and vehicle. Sample targets will be placed on the ground and a full function test flight will be conducted to simulate competition conditions. Date Completed 117

127 6. PROJECT PLAN 6.1. Competition Requirement Verification Plan In addition to safety, a top priority of the team is to satisfy the requirements of the competition to the highest level. The following sections presents the method in which the team will fulfill these requirements to the highest possible level Launch Vehicle A full analysis of competition requirements for the vehicle and recovery systems and the verification method for those requirements has been done and is included as Table 1 in Section Payload A full analysis of competition requirements for the vehicle and recovery systems and the verification method for those requirements has been completed and is included as Table 33 in Section General Project Table 42 below presents the safety requirements as included in the Student Launch handbook as well as the method in which the team will satisfy these requirements. Table 43 presents the general team requirements as included in the Student Launch handbook as well as the method in which the team will satisfy those requirements. Table 42: Competition Safety Requirements Requirement Each team shall use a launch and safety checklists. The final checklists shall be included in FRR and used during LRR and any launch day operations. Each team must identify a student safety officer who shall be responsible for all items in section 4.3. The safety officer shall monitor all team activities with an emphasis on safety. Requirement Source Safety Requirements 4.1 Safety Requirements 4.2 Safety Requirements Verification Method Demonstration Preliminary checklists are included in this document and will be updated in subsequent documents. Inspection- Brian Hardy will be this year s safety officer. Demonstration Safety officer is tasked with supervising all team activities where safety is a concern. Section Addressed

128 Requirement The safety officer shall implement procedures developed by the team for construction, assembly, launch, and recovery activities. The safety officer shall manage and maintain current revisions of the team s hazard analysis, failure mode analyses, procedures, and MSDS/chemical inventory data. The safety officer shall assist in the writing and development of the team s hazard analysis, failure mode analyses, procedures, and MSDS/chemical inventory data. Each team shall identity a mentor who is currently certified by the NAR or TRA. Teams shall abide by the rules and guidance of the local RSO during flights. Teams shall abide by all rules set forth by the FAA. Requirement Source Safety Requirements Safety Requirements Safety Requirements Safety Requirements 4.4 Safety Requirements 4.5 Safety Requirements 4.6 Verification Method Demonstration Preliminary checklists for launch and final assembly have been made; Assembly procedures will be formalized in the coming month leading up to subscale construction. Demonstration- Hazard analysis has been done and is included in this document. Demonstration- Hazard analysis has been done and is included in this document. Inspection Mark Joseph is this year s mentor and is level two certified by the NAR. Demonstration- Team has agreed to follow the discretion of the local RSO during flights. Demonstration Team has agreed to abide by the rules set by the FAA. Section Addressed Table 43: General Competition Requirements Requirement Students on the team shall do 100% of the project. Requirement Source General Requirements 5.1 Verification Method Inspection- Mentor will provide only advice and assist in any energetics handling. Section Addressed

129 Requirement The team shall provide and maintain an included project plan. Foreign National team members shall be identified by PDR. The team shall identify all team members attending launch week activities by CDR. Team shall engage 200 participants in educational, hands-on STEM activities by FRR. The team shall develop and host a website for project documentation. Team shall post, and make available for download, the required deliverables to the team web site by their respective due dates. Requirement Source General Requirements 5.2 General Requirements 5.3 General Requirements 5.4 General Requirements 5.5 General Requirements 5.6 General Requirements 5.7 Verification Method Demonstration: Budget and funding plan have been devised and included in this report. Educational outreach updates will be given in each report. Risks and mitigation have been provided and will be updated throughout the course of the competition. Checklists have been included and will be updated throughout the course of the competition. Demonstration List will be sent in with PDR deliverables. Demonstration List will be sent in with CDR deliverables. Demonstration Team will include ed-out updates in each report and fill out an activity report for every event. Inspection Team has developed a website that will continue to develop throughout the course of the competition. ISS SLI Website Demonstration Team will upload deliverable and host them in the deliverables section of the website. Deliverables Webpage Section Addressed Budget: 6.4 Funding Plan: 6.5 Ed-Out PDR update: 6.3 Risks and Mitigation: 4.7 Checklists: 4.8 N/A N/A 6.3 N/A N/A 120

130 Requirement All deliverables must be in PDF format. In every report, teams shall include a table of contents. In every report, the team shall include the page number at the bottom of the page. The team shall provide and computer equipment necessary to perform a video telecom. All teams will be required to use the launch pads provided by Student Launch s launch service provider. Teams must implement the Architectural and EIT Accessibility Standards (36 CFR 1194) Requirement Source General Requirements 5.8 General Requirements 5.9 General Requirements 5.10 General Requirements 5.11 General Requirements 5.12 General Requirements 5.13 Verification Method Inspection All documentation will be in PDF format. Inspection Table of contents is included in report. Inspection Page numbering starts with section 1 and is included at the bottom of the page. Demonstration Team has access to all equipment necessary for milestone presentations. Analysis All analysis has been done assuming the 8ft 1515 rail provided by Student Launch s launch service provider. Demonstration The team only utilizes equipment that satisfies the requirements of 36 CFR Section Addressed N/A Table of Contents All N/A 3 N/A 6.2. Team Requirement Verification Plan In addition to satisfying the requirements included in the Student Launch handbook, the team has derived a complete set of additional requirements to tackle throughout the course of the competition. In doing so the team has better defined how the project design should be shaped going forward Launch Vehicle A full analysis of team derived requirements for the vehicle and recovery systems and the verification method for those requirements has been done and is included as Table 2 in Section

131 Payload A full analysis of the team derived requirements for the payload and the verification method for those requirements has been done and can be found in Table 33 and Table 34 in Section General Project In addition to satisfying the requirements from the handbook as expected of any team competing in this year s Student Launch competition, the team has derived a set of self-imposed requirements in order to maximize the benefit of participation for both members of the team and the surrounding community. Table 44: General Project Team-Derived Requirements Requirement Active team members will have the opportunity to learn how to operate all machinery and tools Custom printed feedback will be used to improve future ed-out events Safety officer will teach team management on safe practices to allow for a redundancy in knowledge. Website will include a page with member profiles. Website will include a page with team progress updates. Website will include a page with ed-out events and their location/times. Verification Method Demonstration The safety team will prepare tutorial sessions on all machinery that will be used throughout the competition before construction begins. Demonstration Feedback was taken from ISD and will be used to improve future ISD s. Demonstration Safety officer will teach team management so that the safety officer is not require at every build session. Demonstration Web admin will add a page with member profiles sent in by active members Demonstration Web admin will add a page with team updates sent in by team management. Demonstration Web admin will add a page with ed-out updates. Section Addressed N/A N/A N/A 6.3. Educational Outreach Update Throughout the duration of Student Launch competition, the ISS Student Launch team intends to actively engage educators and students throughout Champaign/Urbana and the greater central Illinois area. The team s educational outreach strategy includes significant presences at 122

132 large education outreach events hosted by the Illinois Space Society and the University of Illinois College of Engineering in addition to providing demonstrations to small groups of K-12 students at local schools or extra-curricular groups. This combination of small and large events will allow the team to satisfy the educational outreach requirements of the competition. The purpose of these activities will be to not only teach the community about the basic physical principles governing rocketry and flight, but also to inspire support and participation in the future of spaceflight technologies. Since a lot of the theory behind rocketry is conceptually too abstract and advanced for younger students, engagement activities will revolve around hands-on demonstrations of basic physical principles such as Newton s laws. Activities such a constructed orbital simulator and exploration of the properties of a space shuttle title will give opportunities to teach these principles. A subteam led by Lui will oversee developing activities and displays for educational outreach events. These activities will be performed throughout the duration of the project. To maximize the team s impact on the community, the outcome of these activities will be evaluated to improve future events. Students, educators, and team members will be asked to respond to surveys requesting feedback on events. The focus of this feedback will be determining interest levels of those involved, and measure understanding of material demonstrated by the team. This will allow the team to adjust presentations to better educate the community in future activities. An initial draft of this feedback form is given in Appendix D: Education Feedback Form. The team website will also implement a contact system wherein participants of outreach events may request further information or demonstrations from the team Illinois Space Day Illinois Space Day (ISD) was planned by the team s educational outreach manager Lui and hosted by the Illinois Space Society on October 15 th, Illinois Space Day is a day long, annual educational outreach event hosted by the Illinois Space Society on the University of Illinois campus. In the morning, students were taught scientific principles such as the nature of gravity and Newton s laws through a variety of exhibits and demonstrations hosted by the Illinois Space Society. These exhibits included an orbital simulator where kids threw marbles to see how two bodies in space interact and rocketry booths where Newton s law were shown via short demos. One such Newton demo included having two volunteers push off one another in rolling chairs to show Newton s third law at work. 123

133 Figure 52: Orbital simulator demo. In the afternoon, the students used this knowledge to help construct containers for eggs to compete in an egg drop challenge designed by Lui. Other activities throughout the day included a presentation from an alumnus that works on propulsion systems at ULA. The full brochure, which includes a daily schedule, is included as Appendix E: ISD Schedule and Brochure. Figure 53: Egg drop challenge competition. All participants were asked to fill out an ISD participant survey written by Lui to collect feedback on how to make future ISD s better. This survey was a variant of the education feedback form the team used, with a focus on ISD specific feedback. 124

134 Figure 54: Example ISD participant survey. Participants were asked to rate from 1-5 (5 being the best score) ISD overall, the amount they learned at ISD, and whether they would return for ISD next year. Overall, ISD was graded at a 4.4 average with a median of 4. Amount learned had an average rating of 4.17 with a median of 4. And returning had an average response of 4.7 and a median of 5. Participants were also asked to write down their favorite part of ISD and what they would like to see done differently. By an overwhelming majority, the participants favorite part of ISD was the egg drop vehicle construction and subsequent competition. This result was expected since this activity was the most hands-on component for the attending students. There was also a lot of praise for the liquid nitrogen demo were various items such as marshmallows and flowers were frozen and subsequently shattered. With regards to room for improvement, many students asked for an opportunity to build their own rockets. With some additional funding, the team can purchase small Estes kits or something similar and Lui can design an activity involving building and launching those at future ISD s. 125

135 Upcoming Events The Illinois Space Society and the College of Engineering offer numerous opportunities for small-scale educational engagement activities. Particularly, the Illinois Space Society features an Educational Outreach team which has established relationships with many local schools. This offers a convenient starting point for engagement. The team intends on contacting schools in Mahomet, Urbana, and Champaign, Illinois to offer educational services to students. The team intends on offering hands on demonstrations to students at local high and grade schools. These activities typically take the form of optional after school classes for students, or interaction with school science clubs. Additionally, the ISS Tech Team has contacts with local Boy Scout groups through previous engagements, and the team plans on capitalizing on these opportunities for additional engagement opportunities. Another major opportunity for engagement is the College of Engineering s Open House on March 10th and 11th. As this is several days after the educational engagement deadline (FRR), the team will have completed the required engagement activities before this time. Nevertheless, the team still intends to participate in the Engineering Open House. This is a large event held every year attended by thousands of K-12 students and community members. Although not all of these attendees will be directly engaged by the ISS Tech Team, the Open House still provides an important opportunity to interact with the community and inspire the next generation of engineers and scientists Budget Table 45 below gives a detailed outline of the expected costs associated with the project. Items marked with a * are already owned by the team and will incur no expected costs. Items with parenthesis next to the name mark the quantity of that item needed for the project. Costs of items are rounded to their nearest dollar value due to fluctuations in price that will occur between this proposal and the purchase of these materials. In response, a margin will be added to the grand cost total to account for these fluctuations. Totals at the bottom will highlight the value of materials used and the expected cost of the project that will be covered by the funding plan covered in 6.5. Table 45: Project Expenses Item Total Cost [$] Use Structure: 6 x 48 G12 Tubing Fiberglass 415 Outer Airframe (2) ¼ Thick 24 Square Fiberglass 137 Fins Sheet (3) 3 x 48 G12 Fiberglass Tubing 92 Motor Mount Tube 75mm to 6 Centering Rings (3) 27 Centering Rings Epoxy and Resin* 69 Structural Joints 126

136 Item Total Cost [$] Use Aeropack Motor Retainer 54 Motor Retainer 16 Length Fiberglass Coupler for 6 Airframe Tubing (2) 170 Avionics and Payload Coupler Tubing 6 Coupler Bulkhead (4) 36 Bulkhead for Coupler Tubing 6 Airframe Bulkhead (4) 36 Bulkhead for Airframe Tubing Fiberglass Nosecone 105 Nosecone Nuts, Bolts, Washers, and 15 Connections Screws* 1515 Rail Button (2) 10 Launch Rail Connection Structures Total Value: 1,166 Structures Cost to Team: 1,082 Recovery Equipment StratoLogger CF (2) * 110 Altimeter TeleMetrum 2.0 (2) * 642 Altimeter/Tracker 9V Battery (2) 8 Battery for StratoLogger 9V Battery Clip (2)* 2 Attach 9V Batteries to Sleds TeleMetrum Li-Po Battery (2) * 22 Battery for TeleMetrum SkyAngle Model C2 44 * 66 Parachute for Payload Section Iris Ultra 96 * 404 Main Parachute Fruity Chutes 18 Elliptical 53 Drogue Parachute 20 ft Tubular Kevlar (3) 61 Shock Cord for all Parachutes Jolly Logic* 130 Tether for Main Parachute Jolly Logic (3) 390 Tether for Main Parachute Quick Links (4) * 10 Attachment Hardware Charge Cups (6)* 1 Black Powder Container Nylon Shear Pins* 5 Shear Pins ¼ Threaded Rods (6) 20 Mounts for Sleds 1/8 Plywood Sheet 12 Sleds for Mounting Equipment Right Angle Brackets* 5 Attach Sleds to Rods Terminal Blocks (4) 5 Connect Altimeters to Black Powder Charges Rotary Switches (4) 40 For Activating Altimeters on the Pad Recovery Equipment Total 1,986 Value: Recovery Equipment Cost to

137 Team: Item Total Cost [$] Use Motor Equipment L1390G-P Reload Kit (2) 320 Motor Fuel Grain RMS 75/3840 Motor Casing* 385 Motor Casing 75mm Forward Closure* 102 Closure for Motor Casing 75mm Aft Closure* 80 Closure for Motor Casing Motor Equipment Total Value: 887 Motor Equipment Total Cost to Team: 320 Payload (Camera System) Raspberry Pi Zero (2) 10 Image Analysis Raspberry Pi Camera V2 (2) 50 Image Capture Raspberry Pi Zero Camera Cable 14 Image Analysis (2) LiPo 12V 500mAh Battery (2) 13 Power to Camera System Altimeter Module MS5607 (2) 60 Image Analysis Tool Acrylic Cutouts 30 Allows Cameras to See 3D Printed Camera Mounts* 5 Angle Cameras for Better FOV DC-DC Converter* 2 Use LiPo with Raspberry Pi Camera System Total Value: 184 Camera System Total Cost: 177 Payload (Landing System) Inner Leg Segment (4) 308 Inner Landing Legs Outer Leg Segment (4) 308 Outer Landing Legs Bulkhead Leg Hinge (4) 160 Connects Legs to Bulkhead Nuts, Bolts, Washers, and Screws 15 Connections Torsion Spring (-.552 in*lb.) (4) 5 Torsion Spring Torsion Spring (-4.5 in*lb.) (4) 4 Torsion Spring Landing System Total Value: 800 Landing System Total Cost to Team: 800 Miscellaneous Costs Subscale Vehicle

138 Item Total Cost [$] Use Educational Outreach 100 Travel Expenses 2,500 Total Misc. Costs to Team: 2,840 Margin: 500 Account for Price Fluctuations and Additional Parts Total Value of Project: 8,363 Total Cost to Team: 6, Funding Plan All funding will come through ISS; the organization this project is under. Funds will primarily come from five sources: EC, SORF, AE Department, corporate sponsors, and ISS. Engineering council is an umbrella organization for all engineering registered organization at the University of Illinois at Urbana-Champaign. EC provides funding for projects and conferences separately. The maximum funding allocated is $500 per category per quarter. ISS will request funding for Student Launch in the winter quarter for part orders through project funding and in the spring quarter for the trip through conference funding. EC accepts funding requests 4 times a year through an online form where the specific use of the funds is explained. ISS is confident that the full amount requested will be awarded from past experiences. The Student Organization Resource Fee is a funding opportunity provided to all registered student organizations on campus. All students at the University of Illinois pays a small $5 fee included in their tuition. This fee goes directly to support student organization activities. SORF funding works by reimbursing up to half of part purchases and trip expenses with a maximum of $6,000 allocation to each organization. ISS will initially cover all the part purchases and trip expenses. The organization then will request funding for those expenses through SORF. The $2,000 estimate from SORF is based off past Student Launch project estimates and a balanced use of the maximum allocation amount distribution among the other ISS expenses. Other funding sources for the Student Launch project include the aerospace engineering department at the University of Illinois, corporate sponsors, and ISS. The aerospace engineering department funds aerospace student organizations and ISS has requested funding for various projects. ISS has requested $1,500 from the department specifically designated for Student Launch part orders and travel expenses. ISS also receives funding from corporate sponsors for the competition and other societal activities. Corporate sponsorships will be pursued from aerospace industry companies and vendors from where the part orders come from. ISS will also contribute to provide funding for the Student Launch project. The organization has funds in its university student organization accounts from past sponsorships, donations, and funding sources to be able to provide funding. Because the funding amounts from the aerospace engineering 129

139 department and corporate sponsors are less reliable sources, funds from ISS will back up and cover any additional expenses for the project. The budget for the Student Launch project will continuously be considered and updated as the project proceeds. ISS and the team will also continue to search for funding sources to assist in the parts and travel expenses of the project. Table 46: Funding Sources Funding Source Requested Amount Engineering Council Project $500 Engineering Council Conference $500 Student Organization Resource Fee $2,000 Aerospace Engineering Department $1,500 Corporate Sponsors $1,000 Illinois Space Society $1,000 Total $6, Project Timeline The path to completion of the project, as well as the design and construction of the rocket structure and the payload are visualized on the Gantt charts included on the following pages. Table 47 presents the major milestones and deadlines for the project as well as miscellaneous team objectives such as those involved with the website and the team s educational outreach efforts. Table 48 details the project plan for the structures and recovery subteam, which handles the design and construction of the subscale and full scale flight vehicle. Table 49 details the project plan for the payload subteam, including the development and construction of vertical landing and image processing subsystems. As the project continues, these tables will be updated with new information. Generous time estimates have been given to minimize the risk of delays to the health of the project. 130

140 Table 47: Project Gantt Chart 131

141 Table 48: Launch Vehicle Structure Gantt Chart 132

142 Table 49: Payload Gantt Chart 133

143 APPENDIX A: DEFINITIONS Computer Aided Design: Computer software that allows the design, assembly, and annotation of rocket and payload components. Critical Design Review: A design review that shows that the design is ready for full-scale production and fabrication. Chlorofluorocarbons: Commonly used in aerosol cans until the 1980 s and were determined to be damaging to the ozone layer. Central Illinois Aerospace: A local rocketry club that assists the team with test launching the rockets. They also provide their expertise during the design and building phase of the competition. Final Design Review: A design review that proves the full-scale design is successful and ready for scoring. Illinois Space Society: The parent group of the team competing in the Student Launch competition. Liquid Propane Gas: The most common propellant used in spray paint cans, and is less harmful to the ozone than CFC s. National Association of Rocketry: Governs the use of high powered rocketry to ensure the safety of the participants, spectators, and the environment. Preliminary Design Review: A design review that shows a feasible concept that will be the subject of future work. Proposal: A response to NASA s Student Launch challenge that presents a coherent design and provides proof of its viability. Registered Student Organization: A student group recognized by the university of Illinois and eligible for funding from university sources. The Illinois Space Society is a registered student organization. 134

144 APPENDIX B: ACRONYMS AGL: Above Ground Level APCP: Ammonium Perchlorate Composite Propellant CAD: Computer Aided Design CDR: Critical Design Review CFC: Chlorofluorocarbons CG: Center of Gravity CIA: Central Illinois Aerospace CP: Center of Pressure EIT: Electronics and Information Technology FAA: Federal Aviation Administration FN: Foreign National FRR: Flight Readiness Review HEO: Human Exploration and Operations ISD: Illinois Space Day ISS: Illinois Space Society LCO: Launch Control Officer LRR: Launch Readiness Review MSDS: Material Safety Data Sheet MSFC: Marshall Space Flight Center NAR: National Association of Rocketry PDR: Preliminary Design Review PLAR: Post Launch Assessment Review PPE: Personal Protective Equipment RFP: Request for Proposal RSO: Range Safety Officer RSO: Register Student Organization SLI: Student Launch Initiative SME: Subject Matter Expert SOW: Statement of Work STEM: Science, Technology, Engineering, and Mathematics TRA: Tripoli Rocketry Association 135

145 APPENDIX C: ISS TECH TEAM SAFETY POLICY Illinois Space Society Tech Team Safety Policy All students are to sign and date the present document indicating that they read, understand, and will abide by the contained policy before they enter the Illinois Space Society (ISS). These requirements apply to day to day meetings, construction in and outside of the Engineering Student Projects Lab (ESPL), testing, and any additional meetings that may occur as part of ISS Tech Team activities. The signed forms are to be collected by the team safety officer, recorded, and submitted to the Technical Projects Manager. I. ESPL Rules: Required training to gain access to ESPL General Lab and Electrical Safety training through the U of I Division or Research Safety is mandatory for all individuals before they enter ESPL and participate in Design Council supported projects. Both interactive training modules are online and available at the following link: Upon completion of the training modules the students must print, sign, date each form and give to the designated safety officer who will keep record of their training and then give promptly to ESPL Laboratory Supervisor. It is also required that all students read the present document and sign and date it. Card access to ESPL will be granted after the ESPL Laboratory Supervisor has the General Lab and Electrical Safety training forms and the present document signed and dated on file. Required training to use any tools/equipment in ESPL Students must receive training from The ESPL Laboratory Supervisor and fill out the ESPL General Use Compliance Form and the ESPL Machine Shop use Compliance Form before they use any tool/equipment on the respective forms or any potentially dangerous tools/equipment. Tools shall not be brought into ESPL without the consent of the ESPL Laboratory Supervisor. Any potentially dangerous tools or equipment not listed on the forms should be added to the ESPL General Use Compliance Form list. Students may not work on equipment until the ESPL Laboratory Supervisor has signed and dated the pertinent compliance forms. A student must not use tools/equipment she/he was not trained for. Each student group must designate a safety officer. The name, , and cell phone number of the safety officer must be distributed to each team member. The safety officer must: Make sure that all individuals in the team are working in a safe manner and in compliance with the Design Council Safety Policy. They will keep up to date record of the signed Safety Policy forms for each team member Be familiar with the daily activities of the team Maintain a complete list of MSDS sheets for all potentially hazardous materials and their respective quantities 136

146 All students must abide by the following ESPL General Use Rules: 1. A Laboratory Supervisor will oversee the Engineering Student Project Laboratory, including the Machine Shop. 2. Students may not operate any power tool unless there is somebody else in the same work area of the laboratory or shop. 3. Each student must wear safety glasses with side shield at all times while in any of the ESPL work areas. 4. Hearing protection is required by anyone near loud equipment. 5. When in the work areas one must wear appropriate clothing: closed toed shoes, pants, no loose clothing, jewelry, or hair is allowed that can potentially be caught in equipment. Do not wear ties, rings, or watches. 6. Students must not lift heavy objects without the aid of an appropriate lifting device and hold heavy objects in place using appropriate equipment such as jack stands. 7. When using power tools to cut materials, all parts must be properly clamped in a vise or clamped to a table. Never hold a piece by hand when attempting to cut or drill it. 8. Never leave any tool or equipment running unattended. This includes electronic equipment, soldering irons, etc. When you finish using anything, turn it off. 9. People welding or assisting in welding operations must wear welding masks or yellow tinted safety glasses. You may only watch the welding process if you are wearing a mask. Students who are welding or using grinders must use appropriate shields to protect others. 10. Compressed gases used for welding or other purposes pose several hazards. Users of compressed gases must read and follow the recommendations of Compressed Gas Safety available at Shop doors must not be propped open. 12. Waste chemicals must be properly discarded, See the Laboratory Supervisor. 13. Store potentially hazardous liquids, chemicals and materials in appropriate containers and cabinets 14. Students are responsible for the order and cleanness of their work space and benches according to the rule: If you make a mess, clean it up. The same rule will apply to the common areas of the laboratory including the designated dirty space, paint booth, and welding areas. 15. Work in a clean, uncluttered environment with appropriate amounts of work space and check tools and workspace for problems/hazards before working with them. 16. Know the location of all fire extinguishers, emergency showers, eye rinse stations, and first aid kits. 17. If you fill the garbage can, empty it in the dumpster outside. 18. The Laboratory Supervisor will decide how to proceed in the case of any situations not covered by the preceding rules. 137

147 ESPL Machine Shop Rules (for all students using the ESPL Machine Shop): 1. Any user of the ESPL Machine Shop must read, understand, and abide by the ESPL General Use Rules. 2. The Laboratory Supervisor controls card access to the ESPL Machine Shop. No student can use any machine tool until he/she has demonstrated competence on that machine to the Laboratory Supervisor. 3. No student may enter or remain in the Machine Tool Workshop unless accompanied by the Laboratory Supervisor or a student who is authorized to use the Shop. The authorized user is responsible for the visitor while he/she remains on the Shop. 4. Students may not operate any machine tool unless there is somebody else in the Machine Tool Workshop. 5. Each student must wear safety glasses at all times. 6. When operating machine tools, long hair, long sleeves, or baggy clothing must be pulled back. Do not wear gloves, ties, rings, or watches in the ESPL Machine Shop. 7. When using power tools to cut materials, all parts must be properly clamped in a vise or clamped to a table. Never hold a piece by hand when attempting to cut or drill it. 8. Be aware of what is going on around you. 9. Concentrate on what you're doing. If you get tired while you're working, leave the work until you're able to fully concentrate don't rush. If you catch yourself rushing, slow down. 10. Don't rush speeds and feeds. You'll end up damaging your part, the tools, and maybe the machine itself. 11. Listen to the machine, if something doesn't sound right, turn the machine off. 12. Don't let someone else talk you into doing something dangerous. 13. Don't attempt to measure a part that's moving. 14. Before you start a machine: a. Study the machine. Know which parts move, which are stationary, and which are sharp. b. Double check that your workpiece is securely held. c. Remove chuck keys and wrenches. 15. If you don't know how to do something, ask someone who does. 16. Clean up all messes made during construction a. A dirty machine is unsafe and difficult to operate properly. b. Vacuum or sweep debris from the machine. c. Do not use compressed air. 17. Do not leave machines running unattended. 18. The Laboratory Supervisor will decide how to proceed in the case of any situations not covered by the preceding rules. 138

148 APPENDIX D: EDUCATION FEEDBACK FORM Illinois Space Society Student Launch Educational Feedback Form How interesting was the demonstration? (1 Boring, 10 Extremely Interesting) How much did you learn from this demonstration? (1 Nothing, 10 A Lot) How interesting was the presentation? (1 Boring, 10 Extremely Interesting) How much did you learn from this presentation? (1 Nothing, 10 A Lot) What did you enjoy from your time with us? What was your least favorite part of your time with us? 139

149 APPENDIX E: ISD SCHEDULE AND BROCHURE 140

150 APPENDIX F: NAR HIGH POWER ROCKET SAFETY CODE NAR HIGH POWERED ROCKET SAFETY CODE EFFECTIVE AUGUST Certification. I will only fly high power rockets or possess high power rocket motors that are within the scope of my user certification and required licensing. 2. Materials. I will use only lightweight materials such as paper, wood, rubber, plastic, fiberglass, or when necessary ductile metal, for the construction of my rocket. 3. Motors. I will use only certified, commercially made rocket motors, and will not tamper with these motors or use them for any purposes except those recommended by the manufacturer. I will not allow smoking, open flames, nor heat sources within 25 feet of these motors. 4. Ignition System. I will launch my rockets with an electrical launch system, and with electrical motor igniters that are installed in the motor only after my rocket is at the launch pad or in a designated prepping area. My launch system will have a safety interlock that is in series with the launch switch that is not installed until my rocket is ready for launch, and will use a launch switch that returns to the off position when released. The function of onboard energetics and firing circuits will be inhibited except when my rocket is in the launching position. 5. Misfires. If my rocket does not launch when I press the button of my electrical launch system, I will remove the launcher s safety interlock or disconnect its battery, and will wait 60 seconds after the last launch attempt before allowing anyone to approach the rocket. 6. Launch Safety. I will use a 5-second countdown before launch. I will ensure that a means is available to warn participants and spectators in the event of a problem. I will ensure that no person is closer to the launch pad than allowed by the accompanying Minimum Distance Table. When arming onboard energetics and firing circuits I will ensure that no person is at the pad except safety personnel and those required for arming and disarming operations. I will check the stability of my rocket before flight and will not fly it if it cannot be determined to be stable. When conducting a simultaneous launch of more than one high power rocket I will observe the additional requirements of NFPA Launcher. I will launch my rocket from a stable device that provides rigid guidance until the rocket has attained a speed that ensures a stable flight, and that is pointed to within 20 degrees of vertical. If the wind speed exceeds 5 miles per hour I will use a launcher length that permits the rocket to attain a safe velocity before separation from the launcher. I will use a blast deflector to prevent the motor s exhaust from hitting the ground. I will ensure that dry grass is cleared around each launch pad in accordance with the accompanying Minimum Distance table, and will increase this distance by a factor of 1.5 and clear that area of all combustible material if the rocket motor being launched uses titanium sponge in the propellant. 8. Size. My rocket will not contain any combination of motors that total more than 40,960 N-sec (9208 pound-seconds) of total impulse. My rocket will not weigh more at liftoff than one-third of the certified average thrust of the high power rocket motor(s) intended to be ignited at launch. 141

151 9. Flight Safety. I will not launch my rocket at targets, into clouds, near airplanes, nor on trajectories that take it directly over the heads of spectators or beyond the boundaries of the launch site, and will not put any flammable or explosive payload in my rocket. I will not launch my rockets if wind speeds exceed 20 miles per hour. I will comply with Federal Aviation Administration airspace regulations when flying, and will ensure that my rocket will not exceed any applicable altitude limit in effect at that launch site. 10. Launch Site. I will launch my rocket outdoors, in an open area where trees, power lines, occupied buildings, and persons not involved in the launch do not present a hazard, and that is at least as large on its smallest dimension as one-half of the maximum altitude to which rockets are allowed to be flown at that site or 1500 feet, whichever is greater, or 1000 feet for rockets with a combined total impulse of less than 160 N-sec, a total liftoff weight of less than 1500 grams, and a maximum expected altitude of less than 610 meters (2000 feet). 11. Launcher Location. My launcher will be 1500 feet from any occupied building or from any public highway on which traffic flow exceeds 10 vehicles per hour, not including traffic flow related to the launch. It will also be no closer than the appropriate Minimum Personnel Distance from the accompanying table from any boundary of the launch site. 12. Recovery System. I will use a recovery system such as a parachute in my rocket so that all parts of my rocket return safely and undamaged and can be flown again, and I will use only flame-resistant or fireproof recovery system wadding in my rocket. 13. Recovery Safety. I will not attempt to recover my rocket from power lines, tall trees, or other dangerous places, fly it under conditions where it is likely to recover in spectator areas or outside the launch site, nor attempt to catch it as it approaches the ground. MINIMUM DISTANCE TABLE Installed Total Impulse (Newton- Seconds) Equivalent High Power Motor Type Minimum Diameter of Cleared Area (ft.) Minimum Personnel Distance (ft.) Minimum Personnel Distance (Complex Rocket) (ft.) H or smaller I , J , , K , , L , , , , M N , , O Note: A Complex rocket is one that is multi-staged or that is propelled by two or more rocket motors 142

152 Revision of August 2012 APPENDIX G: FEDERAL AVIATION REGULATIONS 14 CFR, SUBCHAPTER F, PART 101, SUBPART C AMATEUR ROCKETS Applicability. (a) This subpart applies to operating unmanned rockets. However, a person operating an unmanned rocket within a restricted area must comply with (b)(7)(ii) and with any additional limitations imposed by the using or controlling agency. (b) A person operating an unmanned rocket other than an amateur rocket as defined in 1.1 of this chapter must comply with 14 CFR Chapter III. [Doc. No. FAA , 73 FR 73781, Dec. 4, 2008] Definitions. The following definitions apply to this subpart: (a) Class 1 Model Rocket means an amateur rocket that: (1) Uses no more than 125 grams (4.4 ounces) of propellant; (2) Uses a slow-burning propellant; (3) Is made of paper, wood, or breakable plastic; (4) Contains no substantial metal parts; and (5) Weighs no more than 1,500 grams (53 ounces), including the propellant. (b) Class 2 High-Power Rocket means an amateur rocket other than a model rocket that is propelled by a motor or motors having a combined total impulse of 40,960 Newton-seconds (9,208 pound-seconds) or less. (c) Class 3 Advanced High-Power Rocket means an amateur rocket other than a model rocket or high-power rocket. [Doc. No. FAA , 73 FR 73781, Dec. 4, 2008] General operating limitations. (a) You must operate an amateur rocket in such a manner that it: (1) Is launched on a suborbital trajectory; (2) When launched, must not cross into the territory of a foreign country unless an agreement is in place between the United States and the country of concern; (3) Is unmanned; and (4) Does not create a hazard to persons, property, or other aircraft. (b) The FAA may specify additional operating limitations necessary to ensure that air traffic is not adversely affected, and public safety is not jeopardized. [Doc. No. FAA , 73 FR 73781, Dec. 4, 2008] Operating limitations for Class 2-High Power Rockets and Class 3-Advanced High Power Rockets. When operating Class 2-High Power Rockets or Class 3-Advanced High Power Rockets, you must comply with the General Operating Limitations of In addition, you must not operate Class 2-High Power Rockets or Class 3-Advanced High Power Rockets (a) At any altitude where clouds or obscuring phenomena of more than five-tenths coverage prevails; (b) At any altitude where the horizontal visibility is less than five miles; (c) Into any cloud; (d) Between sunset and sunrise without prior authorization from the FAA; 143

153 (e) Within 9.26 kilometers (5 nautical miles) of any airport boundary without prior authorization from the FAA; (f) In controlled airspace without prior authorization from the FAA; (g) Unless you observe the greater of the following separation distances from any person or property that is not associated with the operations: (1) Not less than one-quarter the maximum expected altitude; (2) 457 meters (1,500 ft.); (h) Unless a person at least eighteen years old is present, is charged with ensuring the safety of the operation, and has final approval authority for initiating high-power rocket flight; and (i) Unless reasonable precautions are provided to report and control a fire caused by rocket activities. [74 FR 38092, July 31, 2009, as amended by Amdt , 74 FR 47435, Sept. 16, 2009] ATC notification for all launches. No person may operate an unmanned rocket other than a Class 1 Model Rocket unless that person gives the following information to the FAA ATC facility nearest to the place of intended operation no less than 24 hours before and no more than three days before beginning the operation: (a) The name and address of the operator; except when there are multiple participants at a single event, the name and address of the person so designated as the event launch coordinator, whose duties include coordination of the required launch data estimates and coordinating the launch event; (b) Date and time the activity will begin; (c) Radius of the affected area on the ground in nautical miles; (d) Location of the center of the affected area in latitude and longitude coordinates; (e) Highest affected altitude; (f) Duration of the activity; (g) Any other pertinent information requested by the ATC facility. [Doc. No. FAA , 73 FR 73781, Dec. 4, 2008, as amended at Doc. No. FAA , 74 FR 31843, July 6, 2009] Information requirements. (a) Class 2 High-Power Rockets. When a Class 2 High-Power Rocket requires a certificate of waiver or authorization, the person planning the operation must provide the information below on each type of rocket to the FAA at least 45 days before the proposed operation. The FAA may request additional information if necessary to ensure the proposed operations can be safely conducted. The information shall include for each type of Class 2 rocket expected to be flown: (1) Estimated number of rockets, (2) Type of propulsion (liquid or solid), fuel(s) and oxidizer(s), (3) Description of the launcher(s) planned to be used, including any airborne platform(s), (4) Description of recovery system, (5) Highest altitude, above ground level, expected to be reached, (6) Launch site latitude, longitude, and elevation, and (7) Any additional safety procedures that will be followed. (b) Class 3 Advanced High-Power Rockets. When a Class 3 Advanced High-Power Rocket requires a certificate of waiver or authorization the person planning the operation must provide the information below for each type of rocket to the FAA at least 45 days before the proposed operation. The FAA may request additional information if necessary to 144

154 ensure the proposed operations can be safely conducted. The information shall include for each type of Class 3 rocket expected to be flown: (1) The information requirements of paragraph (a) of this section, (2) Maximum possible range, (3) The dynamic stability characteristics for the entire flight profile, (4) A description of all major rocket systems, including structural, pneumatic, propellant, propulsion, ignition, electrical, avionics, recovery, wind-weighting, flight control, and tracking, (5) A description of other support equipment necessary for a safe operation, (6) The planned flight profile and sequence of events, (7) All nominal impact areas, including those for any spent motors and other discarded hardware, within three standard deviations of the mean impact point, (8) Launch commit criteria, (9) Countdown procedures, and (10) Mishap procedures. [Doc. No. FAA , 73 FR 73781, Dec. 4, 2008, as amended at Doc. No. FAA , 74 FR 31843, July 6, 2009] 145

155 Milestone Review Flysheet Institution University of Illinois at Urbana Champaign Milestone Preliminary Design Review Vehicle Properties Total Length (in) 119 Diameter (in) Gross Lift Off Weigh (lb) 35.4 Airframe Material Fiberglass Fin Material Fiberglass Coupler Length (in) 14 Motor Properties Motor Designation L1390G Max/Average Thrust (lbf) 370.8/312.8 Total Impulse (lbf s) Mass Before/After Burn 35.4lb /31.05lb Liftoff Thrust (lb) Motor Retention AeroTech 75 mm Stability Analysis Ascent Analysis Center of Pressure (in from nose)(in) Maximum Veloxity (ft/s) Center of Gravity (in from nose) Maximum Mach Number Static Stability Margin 2.04 Maximum Acceleration (ft/s^2) Static Stability Margin (off launch rail) 2.06 Target Apogee (From Simulations) Thrust to Weight Ratio 8.8 Stable Velocity (ft/s) Rail Size and Length (in) 96 Distance to Stable Velocity (ft) Rail Exit Velocity (ft/s) Recovery System Properties Drogue Parachute Manufacturer/Model Fruity Chutes Size 15" Altitude at Deployment (ft) ~5350 (Apogee + 2s) Velocity at Deployment (ft/s) ~0 Terminal Velocity (ft/s) Recovery Harness Material Tubular Kevlar Harness Size/Thickness (in) 0.5 Recovery Harness Length (ft) 20 Recovery System Properties Main Parachute Manufacturer/Model Fruit Chutes Iris Ultra Size 96" Altitude at Deployment (ft) 650 Velocity at Deployment (ft/s) Terminal Velocity (ft/s) Recovery Harness Material Tubular Kevlar Harness Size/Thickness (in) 0.5 Recovery Harness Length (ft) 20 Harness/Airframe Interfaces Zinc Plated U Bolts and Stainless Steel Quicklinks Harness/Airframe Interfaces Zinc Plated U Bolts and Stainless Steel Quicklinks Kinetic Energy of Each Section (Ft lbs) Section 1 Section 2 Section 3 Section 4 Kinetic Enerfy Section 1 Section 2 Section 3 Section 4 of Each N/A N/A Section (Ft lbs) N/A N/A Altimeter(s)/Timer(s) (Make/Model) Recovery Electonics 2 Perfect Flite Stratologgers; 2 Atlus Metrum Telemetrums Rocket Locators (Make/Model) Recovery Electonics 2 Atlus Metrum Telemetrum Redundancy Plan Pad Stay Time (Launch Configuration) In Each Coupler: Primary: 1 Stratologger Secondary: 1 Telemetrum 2+ hours Transmitting Frequencies Black Powder Mass Drogue Chute (grams) Black Powder Mass Main Chute (grams) ***Required by CDR*** To be determed with ejection charge testing To be determed with ejection charge testing

156 Milestone Review Flysheet Institution University of Illinois at Urbana Champaign Milestone Preliminary Design Review Capture Mechanism Container Mechanism Autonomous Ground Support Equipment (MAV Teams Only) Overview N/A Overview N/A Overview Launch Rail Mechanism ***Include Description of rail locking mechanism*** Igniter Installation Mechanism Overview N/A Payload Overview Payload 1 Payload 2 Landing leg and image processing subystems to complete the target detection and vertical landing experiment. Landing legs are mechanical deployable alumuninum legs. Image processing subsystems includes a pair of Raspberry Pi Zeros, open source image processing software, and custom coded interface software. Overview N/A Test Plans, Status, and Results Ejection Charge Tests Sub scale Test Flights Tentatively sheculed for Fbruary 13, Charge sizes and shear pin amounts will be calculated and tested. Tentatively scheduled for December 11, A subscale vehicle constructed to match the aerodynamics and scaled properties of the full scale vehicle will be flown. This test will test the stability of the full scale design and help the team adjust simulation data. Full scale Test Flights Tentatively scheduled for February 18, Full scale stability will be tested and the payload will be tested for robustness. Tether systems will also be tested for robustness.

157 Milestone Review Flysheet Institution University of Illinois at Urbana Champaign Milestone Preliminary Design Review Additional Comments Section 1 : Avionics Coupler Section 2: Booster Tube Section 3: Payload (Camera+Landing) Section 4: Nosecone + Upper Airframe