Georgia Tech Ramblin Rocketeers Critical Design Review

Transcription

1 eorgia Tech Critical Design Review

2 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW PAE INTENTIONALLY LEFT BLANK eorgia Institute of Technology 2 of 196

3 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table of Contents Table of Contents... 3 Table of Figures... 7 Table of Tables Introduction School Information and NAR Section Contacts Work Breakdown Structure Launch Vehicle Summary Overview Changes since PDR Payload Summary Overview Changes Since PDR Project L.S.I.M. Overview Mission Statement Requirements Flow Down Mission Objectives and Mission Success Criteria System Requirements Verification Matrix (RVM) Mission Profile Launch Vehicle Overview Mission Criteria System Design Overview Recovery System Altimeters Arming Switches Parachute Dimensions Drift Profile Analysis Kinetic Energy of Launch Vehicle Ejection Charges Testing Structure Payload Section Stingers Connectors Rib eorgia Institute of Technology 3 of 196

4 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Booster Section Structure Mass Section Integration Manufacturing Future Testing Vespula Mk II Mass Breakdown Interfaces and Integration Interface with the round Interface with the round Launch System Launch Vehicle Operations Launch Checklist Launch Vehicle Performance Analysis Fin Design Fabrication and Materials CP Location Nose Cone Motor Selection CP and C Flight Experiment Introduction to the Experiment and Payload Concept Features & Definition Accomplishments Since PDR Science Background Overview of the Experiment Hypothesis and Premise Testing plan Design review Payload Relevance and Science Merit Experiment Requirements and Objectives REFP Flight Experiment Integration Flight Avionics Avionics Overview Avionics Success Criteria SIDES Design Approach SIDESboard SIDES Electrical Harness Master IMU eorgia Institute of Technology 4 of 196

5 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Master Clock Science Experiment Computer Telemetry De-scope Options Power Budget EM Interference Transmission Frequencies and Protocols Software Maturity De-scope Option: Flight Computer Definition round Station Purpose Function Design Considerations Choice of Antenna Choice of Camera Motor Sizing Power Systems eneral Safety Overview Launch Vehicle Safety Payload Safety Environmental Concerns Project Budget Funding Overview Current Sponsors Actual Project Cost CDR Budget Summary System-Level Budget Summary Flight Hardware Expnditures Flight Hardware Expenditure Overview Flight Hardware Cost Breakdown Project Schedule Schedule Overview Critical Path Chart: CDR to PLAR Schedule Risk High Risk Items Low-to-Moderate Risk Tasks eorgia Institute of Technology 5 of 196

6 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 10. Educational Engagement Plan and Status Overview Atlanta Makers Faire Civil Air Patrol FIRST Lego League and Tech Challenge Atlanta Middle School Outreach References Appendix I: antt Chart Appendix II: Launch Checklist Appendix III: Science Overview Appendix IV: round Test Plan Appendix V: Science MFOs and Drawings eorgia Institute of Technology 6 of 196

7 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table of Figures Figure project work breakdown structure Figure 2. Flow down of requirements Figure 3. Project L.S.I.M. mission profile Figure 4: Vespula Mk. II Overall Dimensions Figure 5: Internal Layout of the Launch Vehicle Figure 6: Drogue Parachute Assembly Figure 7: Main Parachute Assembly Figure 8: Electronic Altimeter Schematic Figure 9: Arming Switch Diagram Figure 10: Main Parachute Sizing Figure 11: Vespula Mk II Structural View Figure 12: Staggering of the Stringers Figure 13: Payload Section Configuration Figure 14: Stringer with Connection Hole Figure 15: Universal Connector Figure 16: Payload Section Rib Figure 17. Booster Section Configuration Figure 18: Testing rig Figure 19: Vespula Mk II Mass Breakdown Figure 20: Fin Sleeve Figure 21: Fin Sleeve Attached to Booster Section Figure 22: Open Rocket Aerodynamic Rocket Figure 23: Fin Approximation of Modeling in Simulation Software Figure 24: CP as a function of the Number of Fins Figure 25: 45% Scale Test Rocket and Flight Figure 26: Sub-scale Flight One Flight Data Figure 27: Sub-scale Flight Two Flight Data Figure 28: Nose Cone Profile Figure 29: Vespula Mk II Altitude Plot Figure 30: Vespula Mk II Velocity Plot Figure 31: Thrust vs Time for L2200 Motor Figure 32: L2200 Propellant Mass Burn Figure 33: Vespula Mk II Stability Margin Figure 34: LSIM testing logic, illustrating a simple relationship of information between the test sequences and emphasizing that they flow down from the pursuit of the LSIM hypothesis Figure 35: Preliminary static testing of MR fluid mixtures in magnetic fields Figure 36: Shear stress of a fluid using the two-plate test (Source: Wikipedia) Figure 37: Payload Assembly Figure 38: Payload Base with 150N of loading eorgia Institute of Technology 7 of 196

8 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 39: Factor of Safety vs. Total Load from SolidWorks SimulationXpress and generated trend line equation Figure 40: SIDES system layout Figure 41: SIDESboard bottom side view Figure 42: SIDESboard top side view Figure 43: eneric star topology diagram Figure 44: Example of an electrical harness using zip ties and connectors Figure 45: Maple board used in the Master IMU Figure 46: The connections of the clock to different nodes Figure 47: A possible camera used to analyze the payload experiment Figure 48: Xbee transceiver unit Figure 49: Antenna performance as a function of range Figure 50: eneralization of flight computer software Figure 20: Custom flight computer layout Figure 52: Diagram of a helical antenna Figure 53: Typical radiation pattern for a helical antenna Figure 54: Canon Powershot SX Figure 55: High-Level Software Process Figure 56: Updating Rocket State Figure 57: Updating Servo Position Figure 58: Updating Camera Zoom Figure 59:Transmit Rocket Location Figure 60. (a) Power budget for the A.P.E.S. computer and the Flight Computer; (b) subtotals of the A.P.E.S. computer and the Flight Computer Figure 61: Discharge characteristics of the A123 battery Figure 62: A single A123 LiFePO battery Figure 63. System expenditure summary at CDR Figure 64. Sub-system Testing/Development Breakdown Figure 65. Sub-System Flight Hardware Breakdown Figure 66. Flight Systems flight hardware breakout Figure 6. Critical Path Chart from CDR to PLAR Figure 31. Participation at the Atlanta Makers' Faire Figure 32: Previous FIRST Lego League outreach event Figure 70: FLL Regional Event at Wheel High School Figure 71: FLL Regional Straw Rocket Activity Figure 72: Plot of B field magnitude in MR fluid versus magnitude of vector μ0h, for iron volume concentrations of 10, 20, and 30 percent Figure 73: Shear stress of ideal Bingham plastic (and MR fluid model) versus shear rate dvdn, compared to ideal Newtonian liquid Figure 74: Microgravity time as a function of launch angle from horizon Figure 75: Slosh regimes and similarity parameters Figure 76. Schematic and free-body diagram of slosh dynamic model eorgia Institute of Technology 8 of 196

9 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 77. Base Plate Figure 78. Second and Top Plate Figure 79. Side view of main structure Figure 80. Trimetric view of main structure Figure 81. Top view of structure with 90 Degree L Brackets Figure 82. Side view of 90 Degree Brackets Figure 83. Test Structure with Base Plates Table of Tables Table 1. Mission Objectives and Mission Success Criteria for the L.S.I.M. mission Table 2. Launch Vehicle RVM Table 3. Flight Systems RVM Table 4. Flight Avionics RVM Table 5: Mission Success Criteria Table 6: Launch Vehicle System Requirements Table 7: Launch Vehicles Properties Table 8: Recovery System Properties Table 9: Drift Estimates Table 10: Recovery Characteristics Table 11: Kinetic Energy at Drogue Parachute Deployment Table 12 : Kinetic Energy at Main Parachute Deployment Table 13: Black Powder Properties Table 14: Black Powder Masses Table 15: Success Criteria Table 16: Failure Modes Table 17: Variable Definition Table 18. Payload Section Stringer Material Table 19: Payload Section Rib Material Preliminary Estimates Table 20: Launch Vehicle Structural Mass Budget Table 21: Payload Section Weight Budget Table 22: Booster Section Weight Budget Table 23: Overall Weight Budget Table 24: CP Variable Definitions Table 25: Fin Position vs. CP Location Table 26: Nose Cone Symbol Definitions Table 27: Methods currently available for damping slosh Table 28: Accomplishments since PDR Table 29: Elements of the theoretical modeling for the LSIM payload Table 30: Scientific method fulfillment for LSIM Table 31: Test sequences and descriptions, included options de-scoped since PDR eorgia Institute of Technology 9 of 196

10 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 32: List of MR fluid ingredients Table 33: LSIM success criteria from the Requirements Verification Matrix Table 34: LSIM Requirements Table 35: Payload Assembly Dimensions Table 36: Data from SolidWorks SimulationXpress, highlighting the data from assumptions. 110 Table 37: Avionics requirements Table 38: Avionics Success Criteria Table 39. SIDES Power Budget Table 40: Major Flight Computer Components Table 41: round station requirements Table 42: Risk Identification and Mitigation Steps Table 43. Launch vehicle failure modes and mitigation Table 44: Environmental Hazards, Risks, and Mitigation Table 45. Summary of sponsors for the Ramblin. Rocketeers Table 46. List of current sponsors of the Ramblin' Rocketeers Table 47. CDR Project Budget Summary Table 48. Design milestones set by the USLI Program Office Table 49. Identification and Mitigations for High-Risk Tasks Table 50. Low to Moderate Risk items and mitigiations Table 51: Microgravity times for fall heights Table 52: Similarity parameters for simplified flight profile of the launch vehicle eorgia Institute of Technology 10 of 196

11 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 1. Introduction 1.1. School Information and NAR Section Contacts Team Summary School Info & Project Title School Name Team Name Project Title Launch Vehicle Name Payload Option eorgia Institute of Technology Liquid Stabilization in Microgravity (LSIM) Vespula Mk II 1,2 0F1 Project Lead / Team Richard Team Information Official Safety Officer Team Advisors Tony, Joseph Dr. Eric Feron Dr. Marilyn Wolf NAR Section Primary: Southern Area Rocketry (SoAR) #571 NAR Information NAR Contacts Secondary: A Tech Ramblin Launch vehicle Club #701 Primary: Matthew Vildzius Secondary: Jorge Blanco 1.2. Work Breakdown Structure 1 The LSIM payload is applicable to both the Option 1 and Option 2 payload options listed in the USLI Handbook. On its own, the LSIM payload is intended to be an engineering payload demonstrating a novel technology; additionally, the LSIM payload can be scaled up and will be shown to meet the requirements to compete for the Option 2 payload option. eorgia Institute of Technology 11 of 196

12 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW In order to effectively coordinate design efforts, the project is broken down along technical discipline lines that emulate typical programs in the Aerospace industry. Each sub-team has a general manager supported by several technical leads and subordinate members. Team memberships were selected based on the individuals areas of expertise as well as personal interest. Figure 1 shows the work breakdown structure. Figure project work breakdown structure Launch Vehicle Summary Overview The Vespula Mk II launch vehicle has a gross-lift off weight of approximately 35 pounds and features a 75 mm L2200 solid motor. The structure of the launch vehicle features a rib-and- eorgia Institute of Technology 12 of 196

13 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW stringer design covered by a thin cellulose polymer composite skin to minimize weight and increase structural efficiency. The recovery system utilizes a 30 drogue parachute slowing the launch vehicle down to ft/s and a 120 main parachute to slow the launch vehicle down to ft/s Changes since PDR The following changes have been made since the Preliminary Design Review: The stringer-connector sub-assembly was changed to simplify manufacturing processes and payload integration. Selected motor has changed to the AeroTech L2200 motor. The 10 fiberglass tubes for the main and drogue parachute assemblies has been changed to 12 fiberglass due to cost and limited availability of 10 fiberglass. The thicknesses of the payload and booster section ribs were increased by 25 percent. Ballasted mass of 3 lbs. was added to the payload section Payload Summary Overview The will design, build, test, and fly a system for damping liquid slosh through the use of magnetorheological fluid. This fluid will be actuated with solenoids and driven to a pre-defined state in the Liquid Stabilization in Microgravity (LSIM) experiment. Further, Flight Systems will implement a network of SIDESboards for distributed sensor networks, empowering LSIM, and collecting valuable engineering data. A substantial ground station for observation and telemetry is planned to support the flight of the launch vehicle. Additionally, the will pursue the NASA payload options 1 and 2 in the design, construction, testing, and flight of a primary science experiment and Reduced ravity Education Flight Program. This payload will test the feasibility and practicality of systems to eorgia Institute of Technology 13 of 196

14 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW manipulate magnetorheological (MR) fluids in microgravity for the purpose of demonstrating possible methods for reducing propellant slosh in low-gravity environments Changes Since PDR Infrared will no longer be used as a primary means of measuring slosh in the experiment for the launch vehicle. A camera independent of the avionics apparatus may be used however the primary sensor is seen to be a vibration sensor placed into the base bolt and integrated into a SIDES node. The precise details of ground testing have been reviewed in depth and many changes as to the specifics have been made as testing platforms have been developed. These should result in high quality ground testing data. This data will be used to complete the final link in an expanded theory describing MR fluid. While operating under several assumptions and simplifications, this expanded theory should aid greatly in the development of control software for the flight experiment. eorgia Institute of Technology 14 of 196

15 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 2. Project L.S.I.M. Overview 2.1. Mission Statement The mission of the Mile High Yellow Jackets is: To maintain a sustainable team dedicated to the gaining of knowledge through the designing, building, and launching of reusable launch vehicles with innovative payloads in accordance with the NASA University Student Launch Initiative uidelines Requirements Flow Down The requirements flow down is illustrated in Figure 2. As illustrated by the requirements flow down, the Mission Success Criteria flow down from the Mission Objectives of Project A.P.E.S. All system and sub-system level requirements flow down from the either of the Mission Objectives, Mission Success Criteria, or the USLI Handbook. eorgia Institute of Technology 15 of 196

16 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 2.3. Mission Objectives and Mission Success Criteria Table 1. Mission Objectives and Mission Success Criteria for the L.S.I.M. mission MO Mission Objective MO-1 An altitude of 5,280 ft above the ground is achieved. MO-2 Create an environment in which to test microgravity payloads. MO-3 Reduction in the sloshing motion of a propellant simulatn in microgravity with a magnetic fluid. MO-4 Successful recovery of the launch vehicle resulting in no damage to the launch vehicle. MSC Mission Success Criteria Source Verification Method MSC-1 Minimum Mission Succes: Achieve an altitude of Testing MO-1 5,280 ft., with a tolerance of +320 ft./-640 ft. MSC-2 Minimum Mission Succes: Achieve a microgravitiy Testing MO-2 environment of ± 0.1 MSC-3 Minimum Mission Success:Sucessfully record video Testing of flight experiment during microgravity and start/stop the experiment without mechanical and electrical MO-3 failures. MSC-4 Full Mission Succes: Successful matching of the Testing damping ratio for ringed baffles in the wave amplitudes MO-3 experienced during flight to within ±30%. MSC-5 Minimum Mission Success: The Launch Vehicle is MO-4,USLI Testing recovered with no damage to the structure of the launch vehicle. Handbook 1.4 MSC-5.1 Full Mission Succes:The Launch Vehicle is recovered with no damage to the skin of the launch vehicle. MSC-7, MO-4 Testing 2.1. System Requirements Verification Matrix (RVM) Table 2, Table 3, and Table 4 list the requirements verification matrix for each subsystem. eorgia Institute of Technology 16 of 196

17 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 2. Launch Vehicle RVM Requirement No. Requirement Source Verification Method Design Feature Status Verification Source LV-1 The Launch Vehicle shall carry a scientific or engineering payload. USLI Handbook 1.1, MO-2 Inspection imps standardized payload interface In Progress Payload section The maximum Table 26: LV-1.1 payload weight including any supporting avionics shall not exceed 15 LV-1 Inspection Maximum Parachute Sizing In Progress Overall Weight Budget, Page 69 lbs. The Launch Vehicle Three (3) Section 2.1: shall have a sections: Overview, LV-1.2 maximum of four USLI Handbook 1.5 Inspection nosecone, In Progress Page 31 (4) independent or payload, and tethered sections booster LV-2 The Launch Vehicle shall carry the payload to an altitude of 5,280 ft. above the ground. USLI Handbook 1.1, MO-1 Testing Modified tube fins for straight flight, motor sizing In Progress Section 2.5.5: Altitude Predictions, Page Error! Bookmark not defined. eorgia Institute of Technology 17 of 196

18 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method The Launch Vehicle shall use a commercially available solid LV-2.1 motor using ammonium perchlorate composite propellant (APCP). The total impulse provided by the LV-2.2 Launch Vehicle shall not exceed 5,120 N-s. The Launch Vehicle shall remain LV-2.3 subsonic throughout the entire flight. The Launch Vehicle shall carry one commercially LV-2.4 available barometric altimeter for recording of the official altitude Design Status Feature Use of a USLI Handbook commercially Inspection 1.11 available In Progress solid motor A motor with a maximum USLI Handbook Inspection motor class of 1.12 "L" shall be In Progress used USLI Handbook 1.3 Analysis Motor Sizing In Progress Commercially USLI Handbook 1.2 Inspection available In Progress altimeter Verification Source Section 2.5.3: Motor Selection, Page 83 Section 2.5.3: Motor Selection, Page 83 Section 2.5.3: Motor Selection, Page 83 Section 2.3.1: Altimeters, Page 45 eorgia Institute of Technology 18 of 196

19 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source The amount of LV-2.5 ballast, in the vehicle's final configuration that will be flown in Huntsville, shall be no more than 10% of the unballasted USLI Handbook 1.14 Inspection Proper motor selection for gross lift-off weight of the launch vehicle. In Progress vehicle mass. The Launch Vehicle Section 2.5.4: LV-2.5 shall have aerodynamic stability margin of 1.5 to 3 cailbers prior to leaving the LV-2 Analysis Modified tube-fins for aerodynamic stabilization. In Progress CP and C, Page 86 launch rail. Parachute Section 2.3: The Launch Vehicle Sizing and Recovery LV-3 shall be safely recovered and be MSC-7.1 Testing real time round In Progress System, Page 42 reusable. Station tracking The Launch Vehicle round Section 2.3.1: LV-3.1 shall contain redundant USLI Handbook 2.5 Inspection testing of altimeter In Progress Altimeters, Page 45 altimeters. ejection. eorgia Institute of Technology 19 of 196

20 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source Section 2.3: Recovery The recovery System, LV-3.2 system shall be designed to be LV-3 Inspection Arming Switches In Progress Page 42 armed on the pad. Launch Operations in Appendix The recovery Section 2.3.1: LV-3.3 The recovery system electronics shall be completely independent of the payload electronics. USLI Handbook 2.4 Inspection system electronics shall be entirely independent of from all In Progress Altimeters, Page 45 other systems. Each altimeter shall Section 2.3: be armed by a Recovery Recovery dedicated arming system design System, switch which is shall Page 42 accessible from the incorporate LV-3.4 exterior of the USLI Handbook 2.6 Inspection one (1) In Progress Launch vehicle airframe independent Operations in when the vehicle is arming switch Appendix in the launch for each configuration on the altimeter launch pad. eorgia Institute of Technology 20 of 196

21 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source Recovery Section 2.3.1: system design Altimeters, shall Page 45 Each altimeter shall incorporate LV-3.5 have a dedicated USLI Handbook 2.7 Inspection independent In Progress power supply. power supplies for each altimeter. The arming Section 2.3.2: Each arming switch switches will Arming shall be capable of be designed to Switches, LV-3.6 being locked in the USLI Handbook 2.8 Testing use a key to In Progress Page 46 "ON" position for change the launch. state of the switch. Arming Section 2.3.2: LV-3.7 Each arming switch shall be a maximum of six (6) feet above the base of the Launch Vehicle. USLI Handbook 2.9 Inspection switches shall be located near the booster section of the launch In Progress Arming Switches, Page 46 vehicle LV-3.8 The Launch Vehicle shall utilize a dual deployment recovery system. USLI Handbook 2.1 Inspection Utilization of a drogue and main parachute In Progress Section 2.3: Recovery System, Page 42 eorgia Institute of Technology 21 of 196

22 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source Removable shear Plastic shear Section 2.3: pins shall be used pins will be Recovery LV-3.9 for both the main and drogue USLI Handbook 2.10 Inspection installed in the recovery In Progress System, Page 42 parachute compartments compartments. LV-3.10 All sections shall be designed to recover within 2,500 ft. of the launch pad assuming 15 MPH winds. USLI Handbook 2.3 Analysis Parachute sizing will incorporate descending velocities and drift restricitions. In Progress Section 2.3.4: Drift Profile Analysis, Page 49 Properly sized Section 2.3.5: LV-3.11 Each section of the Launch Vehicle shall have a maximum landing kinetic energy of 75 ft-lb f. USLI Handbook 2.2 Analysis main parachute to ensure landing kinetic energies below 75 ft.- In Progress Kinetic Energy of Launch Vehicle, Page 50 lb f eorgia Institute of Technology 22 of 196

23 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source Proper shielding The recovery shall be system electronics incorporated LV-3.12 shall be shielded from all onboard LV-3 Testing into the design to In Progress transmitting protect the devices. electronics from payload interference. LV-4 The Launch Vehicle shall be launched utilizing standardized launch equipment LV-3 Inspection Use of standard 1515 rail buttons and 8 foot launch pad rail. In Progress Section 2.7.2: Interface with Launch System, Page 71 The Launch Vehicle Section 2.7.2: shall be capable of Interface with being launched by a Launch standard 12 volt System, direct current (DC) Use of Page 71 LV-4.1 firing system and USLI Handbook 1.9 Testing standard In Progress shall require no igniters. external circuitry or special ground support equipment to initial launch. eorgia Institute of Technology 23 of 196

24 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source The Launch Vehicle Section 2.7.2: LV-4.2 shall not require any external circuitry or special ground support equipment to initiate the launch other than what is provided by the USLI Handbook 1.10 Testing Use of standard igniters, 1515 rail buttons, and 8 foot launch rail. In Progress Interface with Launch System, Page 71 range. The Launch Vehicle Follow Section 2.3.1: LV-4.4 shall have a pad stay time on one (1) USLI Handbook 1.7 Testing manufacturers recommendati In Progress Altimeters, Page 45 hour. ons for power LV-4.5 The Launch Vehicle shall be capable of being prepared for flight at the launch site within two (2) hours from the time the waiver opens. USLI Handbook 1.6 Testing Easy assembly of the rocket structure and easy integration of the payload and avionics. In Progress Section 2.8.1: Launch Vehicle Checklist, Page 71 The Launch Vehicle Section 2.7.2: shall be compatible Interface with with either an 8 foot Utilization of Launch LV-4.6 long, 1 in. rail (1010), or an 8 feet USLI Handbook 1.8 Testing 1515 rail and rail interfaces In Progress System, Page 71 foot long, 1.5 in. rail for launch (1515), provided by the range. eorgia Institute of Technology 24 of 196

25 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 3. Flight Systems RVM Requirement Requirement Definition Source Verification Design Status Verification Number Method Feature Source Document FS-1 The flight systems team shall design and build the LSIM Payload MO-3 Inspection LSIM payload In Progress MO-3 FS-2 The LSIM payload shall be designed to fly on a SLP rocket USLI Handbook Inspection LSIM payload In Progress USLI Handbook FS-4 The Flight Systems Team shall produce a working system for manipulating MR fluid MSC-3 Testing Solenoids and Control Algorithms In Progress MSC-3 in LSIM. FS-5 The Flight Systems Team shall ensure that all avionics are properly shielded from the LSIM payload. MSC-3 Testing Faraday cages and webbing tied to ground on the harness Not Started MSC-3 FS-6 The Flight Systems Team shall design all LSIM components and avionics such that they may be easily integrated with the Modular MSC-3 Inspection Mounting system Complete MSC-3 Payload System of the payload bay in the rocket. eorgia Institute of Technology 25 of 196

26 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement Requirement Definition Source Verification Design Status Verification Number Method Feature Source Document FS-7 The Flight Systems Team shall conform to all weight, power, and dimensional MSC-3 Analysis TBD In Progress MSC-3 requirements as per the rocket design. FS-7.1 The Experiment and Avionics, with mechanical supports, shall weight no more LV-1.1 Inspection TBD In Progress LV-1.1 than 15 lbf. FS-8 The flight computer shall execute all tasks necessary to the operation of the LSIM MSC-3 Inspection Maple SIDES node In Progress MSC-3 payload and avionics. FS-9 The LSIM payload shall have a dedicated power supply. MSC-3 Inspection SIDES node In Progress MSC-3 FS-10 The Flight Systems Team shall ensure redundancy and reliability of all internal MSC-3 Inspection SIDES network In Progress MSC-3 electrical hardware. eorgia Institute of Technology 26 of 196

27 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement Requirement Definition Source Verification Design Status Verification Number Method Feature Source Document FS-11 The Flight Systems Team shall provide for payload operation with up to 1 hour of wait on the launch pad and 2 hours of wait during USLI Handbook 1.6 Inspection TBD In Progress USLI Handbook 1.6 preparation of the Rocket. FS-12 The Flight Systems Team shall provide for electrical operations to begin at the beginning MSC-3 Inspection TBD In Progress MSC-3 of the flight trajectory. FS-13 The Flight Systems Team shall ensure that the LSIM payload is shut down safely during MSC-3 Inspection TBD In Progress MSC-3 the deployment phase of the flight trajectory. FS-14 Data from the LSIM payload shall be collected, analyzed, and reported by the team using the scientific USLI Handbook 3.2 Inspection Data logging in SIDES network In Progress USLI Handbook 3.2 method. eorgia Institute of Technology 27 of 196

28 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement Requirement Definition Source Verification Design Status Verification Number Method Feature Source Document FS-15 The LSIM payload will be designed to be recoverable and be able to launch again on the same day without any USLI Handbook 3.5 Inspection Appropriate mounting to the payload interface. In Progress USLI Handbook 3.5 repairs or modifications. Table 4. Flight Avionics RVM Requirement No. FA-1 FA-2 FA-3 Requirement All Flight Avionics shall have sufficient power sources to survive 1-hour pad stay in additon to normal operation requirements The Flight Computer shall collect video of the flight experiment during microgravity The Flight Computer shall collect Launch Vehicle position data and environment conditions (e.g. acceleration). Source Verification Design Method Feature Status USLI Handbook 1.7 Testing Power Supply In Progress MSC-3 Testing Camera In Progress MO-4 Testing IMU, PS In Progress eorgia Institute of Technology 28 of 196

29 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status FA-4 The Flight Avionics shall downlink telemetry necessary to a round Station for the recovery of USLI Handbook 2.11 Teting PS, round Station, Xbee In Progress the Launch Vehicle FA-5 The PS coordinates of all independent Launch Vehicle sections shall be transmitted USLI Handbook Teting PS, round Station, Xbee In Progress to the round Station FA-6 The Flight Avionics shall operate on an independent power supply from the USLI Handbook 2.12 Inspection Power Supply In Progress recovery system Mission Profile Figure 3 illustrates the mission profile for Project L.S.I.M. In order to achieve the desired microgravity environment, the launch vehicle will continue through for one (1) second until deployment of the drogue parachute. This post-apogee delay will yield approximate 4.5 seconds of microgravity to perform the L.S.I.M. eorgia Institute of Technology 29 of 196

30 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW eorgia Institute of Technology Figure 3. Project 30 L.S.I.M. of 196 mission profile.

31 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 3. Launch Vehicle 3.1. Overview The purpose of the launch vehicle is to carry a scientific payload to one mile in altitude and safely return the vehicle to the surface of the Earth. Embracing innovative and out-of-the-box thinking, the launch vehicle will have the ability to carry a wide range of payloads, from scientific experiments to engineering flight demonstrations. The unique rib and stringer design of the vehicle incorporates a standardized payload interface within the primary structure of the rocket. This integration will allow for higher structural efficiency, a lower structural mass fraction, and an increased payload carrying capacity. In addition, the design will feature a 5.25 inch airframe and be 10 feet, 6 inches in length. The full structure of the launch vehicle is illustrated below in Figure 4. The launch vehicle is composed of three sections; the nose cone, the payload section, and the booster section. As with any unique aerospace design, extensive ground testing will be performed to verify structural integrity and successful integration of the payload into the fully assembled launch vehicle. Figure 4: Vespula Mk. II Overall Dimensions The chosen launch vehicle design is unique in many aspects including its rib and stringer internal structure and one-of-a-kind fin design. The rib and stringer design provides an abundance of space for a variety of payload designs while allowing for easy access and integration. Though a kit launch vehicle would be easier to construct, the rib and stringer internal structure will have a lower mass and cost. Similarly, the modified tube fin design is more complicated but provides a highly stable flight even in high winds. The modified tube-fin design features five equally spaced eorgia Institute of Technology 31 of 196

32 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW fins that aim to reduce drag while increasing stability by combining aspects from both tube fins and straight fins. In addition, the use of a fin sleeve to attach the five fins allows their distance from the nosecone to be adjusted thereby allowing the location of the center of pressure to be adjusted prior to each flight and accommodating changes in payloads while maintaining the desired 1.5 calibers of stability. The Vespula Mk II will utilize a dual-deployment recovery system that will minimize the drift of the launch vehicle by mitigating the effects of unpredictable wind conditions with a drogue chute descent. However, the overall purpose of the recovery system, to minimize damage to the launch vehicle from impact with the ground, will be maintained by a main chute deployed closer to the ground. The drogue parachute will be housed in the section connecting the booster and payload sections, while the main parachute will be located between the payload section and nose cone. Both parachutes are made of rip-stop nylon. To ensure successful chute deployment, redundant systems will be used. Each chute will feature two independent black powder ejection charges with corresponding redundant igniters and StratoLogger altimeters. The powder charges will be ignited using low-current electronic matches with independent power supplies at the command of the altimeters Mission Criteria The criteria for mission success are shown in Table 3. eorgia Institute of Technology 32 of 196

33 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 5: Mission Success Criteria Requirement Design feature to satisfy that requirement Requirement Verification Success Criteria The payload requires a steady, but randomly vibrating platform to test the L.S.I.M. system. Provide a suitable environment for the payload. Unsteadiness in the motor's thrust and launch vehicle aerodynamics cause vibrations. In addition, deployment of the drogue By measuring the acceleration with the payload's accelerometers. The L.S.I.M. system reduces a recordable amount of sloshing. parachute will be delayed one second to maximize time in microgravity. To fly as close to a mile in altitude as possible without exceeding 5,600 ft. A motor will be chosen to propel the vehicle to a mile in altitude. Through the use of barometric altimeters. The altimeters record an altitude less than 5,600 ft. Through finite The vehicle must be reusable. The structure will be robust enough to handle any loading encountered during the flight. element analyses and structural ground testing of The vehicle survives the flight with no damage. components System Design Overview lists the derived system-level requirements in order to meet the success criteria. The requirement numbers reference the requirements in the USLI Handbook. eorgia Institute of Technology 33 of 196

34 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 6: Launch Vehicle System Requirements Requirement No. Requirement Source Verification Method Design Feature Status Verification Source LV-1 The Launch Vehicle shall carry a scientific or engineering payload. USLI Handbook 1.1, MO-2 Inspection imps standardized payload interface In Progress Payload section The maximum Table 21: LV-1.1 payload weight including any supporting avionics shall not exceed 15 LV-1 Inspection Maximum Parachute Sizing In Progress Overall Weight Budget, Page 69 lbs. The Launch Vehicle Three (3) Section 3.1: shall have a sections: Overview, LV-1.2 maximum of four USLI Handbook 1.5 Inspection nosecone, In Progress Page 31 (4) independent or payload, and tethered sections booster LV-2 The Launch Vehicle shall carry the payload to an altitude of 5,280 ft. above the ground. USLI Handbook 1.1, MO-1 Testing Modified tube fins for straight flight, motor sizing In Progress Section 0: Motor Selection, Page 83 eorgia Institute of Technology 34 of 196

35 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source The Launch Vehicle Section 0: shall use a Motor LV-2.1 commercially available solid motor using ammonium perchlorate USLI Handbook 1.11 Inspection Use of a commercially available solid motor In Progress Selection, Page 83 composite propellant (APCP). The total impulse A motor with Section 0: LV-2.2 provided by the Launch Vehicle shall not exceed USLI Handbook 1.12 Inspection a maximum motor class of "L" shall be In Progress Motor Selection, Page 83 5,120 N-s. used The Launch Vehicle Section 0: LV-2.3 shall remain subsonic throughout USLI Handbook 1.3 Analysis Motor Sizing In Progress Motor Selection, the entire flight. Page 83 The Launch Vehicle Section 3.3.1: shall carry one Altimeters, commercially Commercially Page 45 LV-2.4 available barometric USLI Handbook 1.2 Inspection available In Progress altimeter for altimeter recording of the official altitude eorgia Institute of Technology 35 of 196

36 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source The amount of Section 3.5: LV-2.5 ballast, in the vehicle's final configuration that will be flown in Huntsville, shall be no more than 10% of the unballasted USLI Handbook 1.14 Inspection Proper motor selection for gross lift-off weight of the launch vehicle. In Progress Vespula Mk II Mass Breakdown, Page 68 vehicle mass. The Launch Vehicle Section 4.5: LV-2.5 shall have aerodynamic stability margin of 1.5 to 3 cailbers prior to leaving the LV-2 Analysis Modified tube-fins for aerodynamic stabilization. In Progress CP and C, Page 86 launch rail. Parachute Section 3.3: The Launch Vehicle Sizing and Recovery LV-3 shall be safely recovered and be MSC-7.1 Testing real time round In Progress System, Page 42 reusable. Station tracking The Launch Vehicle round Section 3.3.1: LV-3.1 shall contain redundant USLI Handbook 2.5 Inspection testing of altimeter In Progress Altimeters, Page 45 altimeters. ejection. eorgia Institute of Technology 36 of 196

37 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source Section 3.3: Recovery The recovery System, LV-3.2 system shall be designed to be LV-3 Inspection Arming Switches In Progress Page 42 armed on the pad. Launch Operations in Appendix The recovery Section 3.3.1: LV-3.3 The recovery system electronics shall be completely independent of the payload electronics. USLI Handbook 2.4 Inspection system electronics shall be entirely independent of from all In Progress Altimeters, Page 45 other systems. Each altimeter shall Section 3.3: be armed by a Recovery Recovery dedicated arming system design System, switch which is shall Page 42 accessible from the incorporate LV-3.4 exterior of the USLI Handbook 2.6 Inspection one (1) In Progress Launch vehicle airframe independent Operations in when the vehicle is arming switch Appendix in the launch for each configuration on the altimeter launch pad. eorgia Institute of Technology 37 of 196

38 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source Recovery Section 3.3.1: system design Altimeters, shall Page 45 Each altimeter shall incorporate LV-3.5 have a dedicated USLI Handbook 2.7 Inspection independent In Progress power supply. power supplies for each altimeter. The arming Section 3.3.2: Each arming switch switches will Arming shall be capable of be designed to Switches, LV-3.6 being locked in the USLI Handbook 2.8 Testing use a key to In Progress Page 46 "ON" position for change the launch. state of the switch. Arming Section 3.3.2: LV-3.7 Each arming switch shall be a maximum of six (6) feet above the base of the Launch Vehicle. USLI Handbook 2.9 Inspection switches shall be located near the booster section of the launch In Progress Arming Switches, Page 46 vehicle eorgia Institute of Technology 38 of 196

39 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source LV-3.8 The Launch Vehicle shall utilize a dual deployment recovery system. USLI Handbook 2.1 Inspection Utilization of a drogue and main parachute In Progress Section 3.3: Recovery System, Page 42 Removable shear Plastic shear Section 3.3: pins shall be used pins will be Recovery LV-3.9 for both the main and drogue USLI Handbook 2.10 Inspection installed in the recovery In Progress System, Page 42 parachute compartments compartments. LV-3.10 All sections shall be designed to recover within 2,500 ft. of the launch pad assuming 15 MPH winds. USLI Handbook 2.3 Analysis Parachute sizing will incorporate descending velocities and drift restrictions. In Progress Section 3.3.4: Drift Profile Analysis, Page 49 Properly sized Section 3.3.5: LV-3.11 Each section of the Launch Vehicle shall have a maximum landing kinetic energy of 75 ft-lb f. USLI Handbook 2.2 Analysis main parachute to ensure landing kinetic energies below 75 ft.- In Progress Kinetic Energy of Launch Vehicle, Page 50 lb f eorgia Institute of Technology 39 of 196

40 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source Proper Section 3.3.1: shielding Altimeters, The recovery shall be Page 45 system electronics incorporated LV-3.12 shall be shielded from all onboard LV-3 Testing into the design to In Progress transmitting protect the devices. electronics from payload interference. LV-4 The Launch Vehicle shall be launched utilizing standardized launch equipment LV-3 Inspection Use of standard 1515 rail buttons and 8 foot launch pad rail. In Progress Section 3.6.2: Interface with Launch System, Page 71 The Launch Vehicle Section 3.6.2: shall be capable of Interface with being launched by a Launch standard 12 volt System, direct current (DC) Use of Page 71 LV-4.1 firing system and USLI Handbook 1.9 Testing standard In Progress shall require no igniters. external circuitry or special ground support equipment to initial launch. eorgia Institute of Technology 40 of 196

41 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source The Launch Vehicle Section 3.6.2: LV-4.2 shall not require any external circuitry or special ground support equipment to initiate the launch other than what is provided by the USLI Handbook 1.10 Testing Use of standard igniters, 1515 rail buttons, and 8 foot launch rail. In Progress Interface with Launch System, Page 71 range. The Launch Vehicle Follow Section 3.3.1: LV-4.4 shall have a pad stay time on one (1) USLI Handbook 1.7 Testing manufacturers recommendati In Progress Altimeters, Page 45 hour. ons for power LV-4.5 The Launch Vehicle shall be capable of being prepared for flight at the launch site within two (2) hours from the time the waiver opens. USLI Handbook 1.6 Testing Easy assembly of the rocket structure and easy integration of the payload and avionics. In Progress Section 3.7.1: Launch Vehicle Checklist, Page 71 eorgia Institute of Technology 41 of 196

42 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement No. Requirement Source Verification Method Design Feature Status Verification Source The Launch Vehicle Section 3.6.2: shall be compatible Interface with with either an 8 foot Utilization of Launch LV-4.6 long, 1 in. rail (1010), or an 8 feet USLI Handbook 1.8 Testing 1515 rail and rail interfaces In Progress System, Page 71 foot long, 1.5 in. rail for launch (1515), provided by the range Recovery System The purpose of the recovery system is to minimize damage to the launch vehicle from impact with the ground. The launch vehicle will use a dual-deployment recovery system to mitigate the effects of unpredictable wind conditions on drift with a drogue chute descent. The drogue parachute will be housed in the compartment connecting the booster and payload sections, and the main parachute will be located between the payload section and nose cone, as illustrated in Figure 5. The launch vehicle will be armed on the launch pad using two arming switches, one for each independent altimeter and ejection charge. For the purpose of simulation, the launch vehicle has been modeled using the Open Rocket Software, with both parachutes made of rip-stop nylon. As a linearized assumption in the limitations of the software, the five modified tube fins are modeled as ten flat fins for drag profiles during descent. eorgia Institute of Technology 42 of 196

43 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 5: Internal Layout of the Launch Vehicle During descent, Kevlar webbing will connect the parachutes to the launch vehicle. The drogue parachute will be housed in a cylindrical compartment in the rear section between the payload and booster sections as illustrated below in Figure 6. This compartment has an outer diameter of 5.25 inches and a length of 10 inches. A bulkhead in the rear payload section will house the ejection wells and also serve to take the impulse of the gun powder blast. The drogue parachute s retention mechanics includes a U-Bolt placed between the two ejection wells on the underside of the payload section, as well as a U-Bolt in the booster section thrust plate. In addition, a shock cord connecting the booster section and main rocket body together. At deployment, the ejection charges will separate the booster section from the main rocket, releasing the drogue parachute as well. eorgia Institute of Technology 43 of 196

44 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 6: Drogue Parachute Assembly The main parachute will be placed in a section above the payload bay. The section has an outer diameter of 5.25 inches and a length of 12 inches. The main parachute s ejection wells will be placed such that the impulse is imparted on the payload section and the nose cone is separated from the main rocket pulling the main parachute out. Shock cords will connect the main parachute to the nose cone and the payload section of the launch vehicle, ensuring that the all sections remain together during descent. The main parachute assembly is illustrated below in Figure 7. eorgia Institute of Technology 44 of 196

45 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 7: Main Parachute Assembly The parachute casings will be made of 12 fiberglass, and the bulkhead under the main chute will be made of aluminum. 12 was chosen as opposed to 10 due to availability and sizing. For manufacturing purposes, the available sizes of 12 airframes were a much better match for our design. Although this will increase the mass of the recovery section due to the difference in density between 10 and 12, the structural integrity of the casings will not be affected, nor will the adhesion potential of the casings be adversely affected. Two-inch stainless steel U-Bolts will be drilled into the bulkheads, and will be used to attach the shock cords. Five Nylon 2-56 screws will be used as shear pins to keep both the main and drogue chute compartments together during flight until the parachutes are deployed. PVC end-caps will be used to direct the ejection charges in order to protect the casing from thermal shock, and a NOMEX shield will protect the parachutes. The charges will be ignited using an e-match Altimeters eorgia Institute of Technology 45 of 196

46 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW To ensure successful chute deployment, redundant systems will be used. Each chute will feature two independent black powder ejection charges with corresponding redundant igniters and StratoLogger altimeters. The altimeters will ignite the ejection charges through the use of lowcurrent electronic matches using independent power supplies. The components which compose each altimeter system are independent of all payload electronics. The altimeters and all recovery electronics have a pad stay time of at least an hour. The system setup for each altimeter is shown below in Figure 8. Figure 8: Electronic Altimeter Schematic In addition, the recovery electronics wiring will be protected from transmitting devices in the rocket through faraday cages and shielding integrated into the wiring harnesses, these devices are discussed further in the Avionics. round testing will determine whether transmission interference will affect the altimeter devices directly Arming Switches The altimeters and the recovery systems will be activated on the launch pad with two arming switches. Each arming switch activates one of the two independent altimeter systems. The arming switches will be located at the base of the payload section which is approximately four feet above the bottom of the launch vehicle. The arming switches will be key activated and are capable of being locked in ON or OFF position, as illustrated below in Figure 9. eorgia Institute of Technology 46 of 196

47 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 9: Arming Switch Diagram Parachute Dimensions The sizing of the main parachute is determined by the weight of the launch vehicle and the kinetic energy constraint of the launch vehicle when it touches down. Based on the LV-3.10 and LV-3.11 requirements, the launch vehicle should not experience more than 75.0 ft-lbf of kinetic energy upon landing, this places an upper limit on the landing velocity to be approximately ft/s. The main parachute can have a diameter of 10 to 13ft depending on the weight of the payload section. Figure 10 below shows the main parachute sizing. eorgia Institute of Technology 47 of 196

48 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 10: Main Parachute Sizing Table 7 and Table 8 outline the dimensions and properties of the constraining launch vehicle properties and the properties of the parachutes. Table 7: Launch Vehicles Properties Launch Vehicle Properties Weight of launch vehicle 55 lb C D of Launch vehicle 0.75 Max Kinetic Energy 75.0 ft-lbf eorgia Institute of Technology 48 of 196

49 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 8: Recovery System Properties Properties Main Parachute Drogue Parachute Diameter(ft) 16 3 Surface Area(ft^2) Estimated C D Target Descent Drift Profile Analysis Drift profile analysis is the method used to estimate and constrain the landing site for the launch vehicle. Based on how long the launch vehicle will be in flight and the wind speed at launch, the range can be estimated. Using Equation (1) for drift, Equation (2) for time in flight, and Equation (3) for decent velocity the drift of the launch vehicle under the main and drogue parachutes can be determined. The results are shown below in Table 9. Drift = Time in flight V wind (1) Time in flight = Alt max descent speed (2) descent velocity = 2mg ρac d (3) Table 9: Drift Estimates Launch Vehicle Drift Estimates Wind Speed (mph) Drift (ft) Drogue Parachute Main Parachute Total Drift eorgia Institute of Technology 49 of 196

50 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Launch Vehicle Drift Estimates Wind Speed (mph) Drift (ft) Drogue Parachute Main Parachute Total Drift The descent velocity of the launch vehicle will be estimated using the terminal velocity. The terminal velocity is the constant speed of a free-falling object when the drag due to air resistance prevents further acceleration. The values are listed below in Table 10. Table 10: Recovery Characteristics Recovery Systems Properties Drogue Parachute Main Parachute Diameter (ft) 3 Dimensions (ft) 16 Flight Time (s) Flight Time (s) Terminal Velocity (ft/s) Terminal Velocity (ft/s) Horizontal Drift (ft) Horizontal Drift (ft) Total Drift (ft) Kinetic Energy of Launch Vehicle Kinetic energy calculations are based on (4) below. eorgia Institute of Technology 50 of 196

51 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW KE = 1 2 mv2 (4) Using the masses of the separate sections, the kinetic energies can be calculated using the velocity of the system at different points in the mission. The Kinetic Energies of separate sections after the drogue chute is deployed are given below in Table 11. After the drogue chute is deployed, the launch vehicle has separated only between the booster section and the payload section, so the payload and nosecone sections are treated as one part. The velocities listed are the terminal velocities under the drogue parachute once the Table 11: Kinetic Energy at Drogue Parachute Deployment Launch Vehicle Section Weight (lb.) Velocity (ft/s) Kinetic Energy (ft-lbf) Nose Cone Payload Booster The Kinetic Energies of the separate sections after the deployment of the main chute and landing are given below in Table 12. After the main chute is deployed all three sections have separated and their separate masses were used in the calculations. All sections will have the same velocity due to the shock cord tethers. Table 12 : Kinetic Energy at Main Parachute Deployment Launch Vehicle Section Weight (lb.) Velocity (ft/s) Kinetic Energy (ft-lbf) Nose Cone Payload Booster eorgia Institute of Technology 51 of 196

52 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Ejection Charges To eject the parachutes, redundant black powder charges will be used. The containers housing the chutes will also be pressurized in order to ensure chute deployment. Due to the different requirements for the drogue and main chutes, two sets of calculations will be needed. The amount of black powder used in the ejections charges can be calculated through Equation (5) below. Once the amount of black powder is determined the values can then be tested before flight. The equation relates weight of black powder to the ejection pressure, volume of the container, black powder combustion gas constant, and the black powder combustion temperature. The constants used are listed below in Table 13. Pressure Volume lb of Black Powder = (5) RT Using the pressurization of 10 psig and 9 psig as a structural maximum for the main and drogue chute compartments, the resulting black powder masses are calculated to be 5 grams and 2 grams for the main and drogue chutes, respectively, as illustrated below in Table 14. The masses used will depend on the final container dimensions, which were estimated at 5.25 inches in radius and 12 and 10 inches in length for the main and drogue, respectively. The force required for separation with the given number of Nylon shear pins would be 446 lb f for the main chute and 393 lb f for the drogue chute. Table 13: Black Powder Properties Constant Combustion as Constant Combustion Temperature Value ft lb f / lb m R 3307 R eorgia Institute of Technology 52 of 196

53 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 14: Black Powder Masses Main Drogue Total Pressurization 10 psig 9 psig Ejection force 446lbf 393lbf Black Powder 5 grams 2 grams Testing In order to ensure the safety and viability of the calculations made in determining the black powder masses, ground testing will be done before flying the launch vehicle recovery system. Due to the explosive nature of black powder charge testing, the tests for this launch vehicle will be coordinated with the campus security and the eorgia Tech Fire Marshal. For the black powder test, the rocket will be placed horizontally on the ground on a relatively smooth surface to minimize unwanted static friction irrelevant to a flight environment. Padding will surround the test area to protect participants and the rocket from debris. Table 15 and Table 16 illustrate the conditions for test success and failure. Table 15: Success Criteria Success Criteria Ejection charge ignites Shear pins break Launch vehicle moves half the distance of shock cord Failure Criteria Table 16: Failure Modes eorgia Institute of Technology 53 of 196

54 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW The fiberglass of the tube coupler shatters due to the charge. The shear pins don t shear, and the launch vehicle stays intact. The NOMEX/cloth shield fails and the parachute is burned. The E-matches fail to ignite the black powder Structure The structural subsystem of the launch vehicle must provide a strong, reliable frame for carrying the payload throughout the duration of the mission, allow for the easy integration of aerodynamic components, and have minimal weight. Since the first two criteria must always be met, the structure should be optimized to have a minimal weight while still satisfying the first two constraints. To achieve this minimal weight, a rib and stringer architecture was chosen over the traditional thick wall architecture. Since PDR, the structural architecture has remained the same, though some component geometries have been altered to optimize the manufacturing and quality control processes and simplify payload integration. In addition, when one stringer is removed the payload section can be accessed easily. The complete structural architecture is illustrated below in Figure 11. Figure 11: Vespula Mk II Structural View For the sake of clarity in the following discussion, the notation and variables used in the course of this section are summarized below in Table 17. eorgia Institute of Technology 54 of 196

55 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 17: Variable Definition Variable t d s n Definition Minimum thickness (in inches) for a stringer wall or a rib to prevent yield, buckling, or general failure. Outer diameter (in inches) of each stringer. A coefficient (dimensionless) related to the number of stringers in each section. Factor of safety (dimensionless). Critical loading (in pounds-force). For stringers, the loading is in compression. P cr For ribs, the loading is in shear. The critical loading is determined by the maximum expected force in flight. L E r o -r i τ ultimate The length (in inches) of each stringer. Young's modulus (in psi). The difference in the outer radius and inner radius (in inches) of each rib. The ultimate shear stress (in psi) of the rib material. In the rib and stringer architecture, long, thin stringers run parallel to the longitudinal axis of the launch vehicle and bear flight loads in compression. To minimize the weight of each stringer, cylindrical tube geometry was chosen. The specific dimensions required for structurally competent strings are given by the following equations: t = d d4 snp cr L 2 2π 3 E (6) 16 s = (Number of stringers per segment) (7) eorgia Institute of Technology 55 of 196

56 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Equation (6) gives the minimum wall thickness, t, of a given stringer assuming a hollow cylindrical stringer design. In Equation 1, the factor of safety, n, is taken to be 2.5, and the outer diameter of the stringers, d, the critical loading of the stringers, P cr, the stringer length, L, the stringer coefficient, s, given by Equation (7), and the modulus of elasticity, E, are design decisions based on the requirements for each structural section. Another component of the rib and stringer architecture is the thin ribs running along the central axis of the launch vehicle. To avoid large column lengths in the stringers and thicker stringers, the stringers are staggered such that a stringer on one surface of the rib is not coaxial with a stringer on the opposite side of the rib, as seen below in Figure 12. Figure 12: Staggering of the Stringers Because of this staggering, the ribs experience very few forces in compression; the primary force on the rib is in shearing. The minimum thickness of each rib is governed by its strength in eorgia Institute of Technology 56 of 196

57 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW shearing. Equation (8) gives the minimum thickness of the rib as a function of the critical loading, P cr, the factor of safety, n, the radial thickness, r o -r i, and the shear strength, τ ultimate. t min = P cr n 2(r o r i )τ ultimate (8) Payload Section Stingers In order to have sufficient space to contain the scientific payload and to allow for easy access to said payload, the payload section will be comprised of three segments of four stringers as illustrated below in Figure 13. Figure 13: Payload Section Configuration Each segment will be 12 inches in length. Due to the difficulty of manufacturing complex, threedimensional circular parts to a quality standard, it was decided that the stringer geometry would be modified to a square rod. By adjusting the geometry accordingly, the stringers can be cut to length with computer-controlled machinery without complicated tooling. Initially, carbon fiber was being considered the most likely material to be used for the stringers due, to its high strength to weight ratio. However, more materials were also considered and compared with the carbon fiber, as listed in Table 18. eorgia Institute of Technology 57 of 196

58 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 18. Payload Section Stringer Material Material Minimum thickness (in) Actual thickness (in) Stringer mass (lb m ) Aluminum rade 4130 Alloy Steel /arolite Carbon Fiber Copper Alloys Composite materials such as garolite and carbon fiber were not chosen due to manufacturing and lifecycle risks including delamination and material tolerance issues. As the stringers will have holes on each end to provide connection points with the ribs, as seen below Figure 14, the material must not fatigue around the pinning holes such that a material failure could occur. Figure 14: Stringer with Connection Hole eorgia Institute of Technology 58 of 196

59 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Because of the aforementioned difficulties in using composite materials, aluminum 6061 was chosen as the stringer material in lieu of carbon fiber. The aluminum 6061 stock selected for manufacturing has a flatness tolerance of ±0.004 inches per foot and a thickness tolerance of ± Since aluminum is not a composite material delamination around the holes in the stringers is not a concern. In addition, aluminum s reasonable strength to weight ratio, price, and availability from the supplier made aluminum the ideal stringer material. Delamination could cause a structural liability and possible failure of structural integrity. In addition, the straightness tolerance of carbon fiber from the material supplier was not rated. iven the internal requirement that the launch vehicle should be within 1 degree of vertical, a low straightness tolerance for the stringers is essential. After researching and comparing alternative materials, Aluminum 6061 was chosen as the stringer material in lieu of carbon fiber. Aluminum 6061 has a straightness tolerance of ±0.010 inch per foot and an outer diameter and inner diameter tolerance of ±0.004 inches for a tube. Since aluminum is not a composite material delamination around the holes in the stringers is not a concern. In addition, aluminum s reasonable strength to weight ratio, price, and availability from the supplier made aluminum the ideal stringer material Connectors Fundamentally, the architecture of the connections remains identical to the connections established at the time of the Preliminary Design Review, though the drilling location has been modified to constrain the connector to the rib. The connector design is simplified such that the features only exist on two orthogonal faces, and that these features could be cut with precision computer-controlled machinery. The simplified, two-face design reduces the need for tooling and eorgia Institute of Technology 59 of 196

60 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW decreases the rate of quality issues in the components. These design parameters yield a critical loading of the stringers of 87.5 lbf. The revised connector is illustrated below in Figure 15. Figure 15: Universal Connector Rib In designing the ribs for the payload section, the shear loading was taken to be 87.5 lb f to match the loading experienced by each stringer. That is, given the predicted mass of the launch vehicle at lift-off and the maximum acceleration due to thrust, each stringer in the payload section will transmit 87.5 lb f to the surface of the rib. To reduce mass, the center section of the rib was removed, as most of the forces collect on the outer surface of the rib, rendering the core a nonload-bearing element. Furthermore, eight points along the outer diameter were notched, as seen in Figure 16, to allow the avionics group easy accessfor wiring the interior of the launch vehicle. eorgia Institute of Technology 60 of 196

61 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 16: Payload Section Rib This design feature eliminates the tangling of wires on the interior of the launch vehicle and allowes for easy removal of avionics components. Several materials were examined in determining the optimal material and thickness for the ribs. The results of this study are found below in Table 19. Table 19: Payload Section Rib Material Preliminary Estimates Material Minimum thickness (in) Actual thickness (in) Rib mass (lb m ) Aluminum Carbon steel Copper / arolite* Plywood* eorgia Institute of Technology 61 of 196

62 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Material Minimum thickness (in) Actual thickness (in) Rib mass (lb m ) Stainless steel Titanium 6Al-4V It should be noted that after finite element analysis, it was determined that severe plastic deformation would occur in the materials at the given thicknesses because the thickness was computed using the ultimate shear strength instead of the yield strength. Because all calculations used the ultimate shear strength, a reasonable material choice may be made from the projected masses. Despite the light weights of composite materials such as -10 fiberglass (garolite) and plywood, composites tend to perform poorly in shear along layer boundaries and tend to have poor bearing strengths. The failure modes of composite materials include delamination, which is difficult to detect but can have devastating consequences, and shattering, which compromises the rest of the launch vehicle. Metals, on the other hand, experience a plastic deformation before ultimately failing. While this deformation ultimately destroys the structural integrity of the rib, the deformation does not cause the immediate failure of the rest of the launch vehicle structure. Due to these failure modes, aluminum 6061, a ductile metal which was also chosen for the stringers, has been chosen for the rib material. Initial studies in the SolidWorks simulation environment indicate that the chosen thickness is not sufficient to prevent plastic deformation, so the thickness of the rib was increased to inches. In the interest of verifying that the plastic deformation was not the result of a simple mathematical error, the same simulations were run on the rib component, but using a variety of the materials previously considered. These trials confirmed that an increase in material thickness would be required. eorgia Institute of Technology 62 of 196

63 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Booster Section The structure for the booster section of the launch vehicle is based on the structure of the payload section. Similar to the payload section structure, the structure of the booster section will be a skeleton composed of long, thin stringers connected to thin ribs. Due to concerns of integrating aerodynamic subsystems with the booster section structure, the booster section structure has a specific set of requirements to adapt to aerodynamic design. Because of the launch vehicle s unique five-fin design and the fact that the positioning of the fin sleeve can be changed with respect to the longitudinal axis of the launch vehicle, the connection points in the structure must match those of the fin sleeve. Since the fin sleeve connection points are directly above and below each fin, connection points in the fin sleeve are 72 apart. This geometry requires that the connection points in the booster section structure are also 72 apart. The connection separation angle requires each subsection of the booster skeleton to contain five equally-spaced stringers as opposed to the four required by the payload section skeleton, hence the coefficient s from Equation (7) is 3.2 for the booster section. The coefficient s relates the distribution of the critical loading to the number of stringers in a structural segment. Since each booster section stringer has a length of 11.0 inches, the minimum side dimension for the booster section is smaller than for the payload section. However, this difference in the required side dimension for the payload and booster sections is not sufficiently significant to obtain new material stock specifically for the booster stringers. Additionally, stock with a smaller side dimension is not readily available and would be wasteful to manufacture. For this reason, the booster section stringers will use the same material, and have the same side dimension of inches, as the payload section stringers. The radial dimensions of the booster section are not significantly different from those of the payload section, and because booster section components will experience the same loading as payload section components, ribs and connectors will maintain nearly the same dimensions and eorgia Institute of Technology 63 of 196

64 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW mass properties as their payload section counterparts; thus, it is not necessary to reexamine the material selections for each component. A model of the booster section can be seen below in Figure 17. Figure 17. Booster Section Configuration Structure Mass A summary of the masses and the current mass breakdown of the launch vehicle structure can be found in Table 20 below. Table 20: Launch Vehicle Structural Mass Budget Component Component Mass (lb m ) Quantity Total Mass (lb m ) Payload stringer Booster stringer Payload rib Booster rib Connector screws #8 lock nut Payload skin Booster skin Total Mass 8.25 eorgia Institute of Technology 64 of 196

65 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Due to significant structural design changes, the primary airframe structural mass (which includes the masses of the payload and booster structures) has increased from 4.03 lb m at the time of the Preliminary Design Review to 8.25 lb m at the current time. The mass increase for the ribs is due to an increase in thickness and the mass increase for the stringers is due to a change in geometry. These mass increases were compensated with the larger decrease in mass due to the simplification of the stringer-to-rib connector pieces for an overall decrease in structural mass Section Integration The three sections of the rocket, namely the nose cone, payload, and booster sections, will be separated by two parachute bays made of -12 fiberglass. These bays, one for the drogue parachute and the other for the main parachute, will serve as structural elements as well as sealed compartments for recovery purposes. At the end of each section is a sealing bulkhead with a U- bolt to which adjacent sections of the launch vehicle are tethered, in addition to recovery devices Manufacturing There are three main parts that make-up the internal structure of the launch vehicle; the ribs, stringers, and connectors. Due to the change in design of both the stringers and connectors they will be manufactured differently. The stringers cross section has changed from cylindrical to rectangular. The design of the connectors has been simplified to interface with the now rectangular prism stringers and be manufactured using a water jet. The material of the connectors is now aluminum The ribs have the simplest geometry of all the components, as a thin circular-shape cut from a flat plate. Because of this simple geometry, the ribs will be manufactured as a two-dimensional object. Using the two-dimensional drawing of the rib, a part is cut out from a flat plate of eorgia Institute of Technology 65 of 196

66 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW aluminum 6061 using a water jet. The water jet, which cuts a material using a computercontrolled cutting tool, produces precision parts with minimum wasted material while maximizing both accuracy and precision. The stringers are cut from a rectangular-prism-shaped rod of aluminum using a water jet. Each stringer for the payload section is 12 inches long while each stringer in the booster section is 10 inches long. In order to securely fit each stringer to a connector, a thru hole will be drilled ¼ of an inch from the top and bottom ends of each stringer. The thru hole is in the long side of the stringer and both holes are in the same plane. The purpose of the thru holes is to insert a #8-32 McMaster screw to hold the stringer securely to the connecter. The connector is designed solely for the purpose of mechanically attaching the ribs and the stringers. The design and manufacturing for the connectors have been simplified. Each connector will be manufactured from a plate of aluminum 6061 which is 9.53 mm, inches, thick. The connectors will be cut from the water jet. Each connector has a thru hole going from the top to the bottom in the center of the left and right hooks of the U and a thru hole going from the front to the back in the center of the straight part connecting the hooks. The connector is one part and not an assembly of parts. The purpose of the thru holes is to attach the connectors to the ribs and the stringers to the connectors using #8-32 McMaster screws Future Testing round testing will be performed to ensure structural integrity while loads up to a factor of safety of 2.5 are applied. The test rig used is designed to perform dynamic and static loading tests and is illustrated below in Figure 18. The testing rig features a rail-mounted impact machine capable of holding different mass that can be lifted to a max height of five feet. The ability to hold different mass at specific heights simulates different impact energies. eorgia Institute of Technology 66 of 196

67 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 18: Testing rig For each test a known mass will be dropped at a certain height that correlates to a specific design impulse. The impulse value will estimated from the maximum acceleration of the launch vehicle in the program simulation Open Rocket. The height for the mass in Equation (11) will be found using the Conservation of Energy and Conservation of Momentum principles shown in Equation (9) and Equation (10), where m t is the drop mass. E = m t gh m t V 2 (9) Momentum = m t V 2 m t V 1 (10) h = 1 I 2g M In addition to the dynamic loading testing static loading testing will be conducted on the ribs and thrust plates to determine their maximum allowable loading and yield stresses. The static loading 2 (11) eorgia Institute of Technology 67 of 196

68 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW testing will be conducted utilizing an Instron loading machine and Vic-3D stress/strain analysis image capturing software. In addition, future testing will possibly include a vibration test to check for resonance frequencies. This test is important in that if the vibrations oscillate with the components natural frequency structural failure will occur Vespula Mk II Mass Breakdown Mass breakdown for the booster and payload sections are summarized in Table 21 and Table 22 with systems level summary shown in Table 23. The values obtained for the booster section were mostly estimated utilizing Solidworks. The values for nose cone, drogue chute, main chute, shock cords, imps structures, and the motor case are actual weights obtained from a scale. Note that the ribs make up the most mass for the imps structure. This component is overbuilt structurally in order to have a satisfactory fastener edge clearance. Additionally, the mass breakdown is also presented in terms of mass fractions, as illustrated in Figure 19. In addition, there is a ballasted mass in the payload section. Table 21: Payload Section Weight Budget Payload Section Weight (lb.) Quantity Total Weight (lb.) Rib Stringer Connector Bulkhead Skin Epoxy/Paint screws eorgia Institute of Technology 68 of 196

69 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Payload Section Weight (lb.) Quantity Total Weight (lb.) #8 lock nut Ballast Total Table 22: Booster Section Weight Budget Booster Section Weight (lb.) Quantity Total Weight (lb.) Rib Stinger Connector Epoxy/Paint skin Fasteners U-Bolts Motor Case Total 7.9 Table 23: Overall Weight Budget Component Weight (lbs) Nose Cone 1.7 Drogue Chute + Shock Cords 1.43 Main Chute + Shock Cords 2.76 Avionics System 5 Payload 10 Payload & Recovery Structure eorgia Institute of Technology 69 of 196

70 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Booster Structure 7.9 AeroTech L Total Figure 19: Vespula Mk II Mass Breakdown 3.6. Interfaces and Integration The interfaces between the launch vehicle and the ground, and ground launch system, shall be described such that the operation of interfacing the launch vehicle with these systems can be correctly carried out to ensure optimal launch vehicle performance, with maximum safety to the USLI team, and so that a sustainable architecture can be developed to show new members the necessary action items of launch vehicle/ground/ground launch system integration Interface with the round The launch vehicle will have a PS tracking system that will deliver real-time telemetry, as well as the launch vehicle s landing location, to the ground tracking station via an XBEE radio eorgia Institute of Technology 70 of 196

71 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW transmitter. When the power system is locked to the ON position on the launch pad, the XBEE will begin transmitting telemetry data Interface with the round Launch System The launch vehicle will have attached large launch lugs, so that it can fit within a launch rail with an aluminum 1515 T-slotted extrusion, of a minimum length of 8 feet. The launch vehicle will be placed on a launch stand designated by the LCO after being inspected and certified flight-worthy by the RSO. After proper assembly and insertion of the motor, inspection and certification, and attachment to the launch stand, the electronics necessary for the payload and recovery system, will be activated and locked into position. The altimeter will announce the readiness of the electronics and payload system via a series of beeps. The launch vehicle will be launch using standardized launch equipment including a standard 12 volt direct current firing system Launch Vehicle Operations It is the responsibility of Launch Operations to create comprehensive guides and checklists to ensure proper operation of the launch vehicle and the safety of the USLI team. Proper operation of the launch vehicle requires that certain protocols and procedures are observed by the Ramblin Rocketeers team during assembly and launch Launch Checklist The Launch Checklist ensures that all tasks necessary for a successful launch are completed and completed in the most efficient order. The Launch Checklist has both a performer and an inspector to ensure all tasks are completed correctly. In addition, there is a Troubleshooting Chart to address common problems when preparing and launching rockets. The Launch Checklist remains largely unchanged from the previous year in which the launch vehicle was prepared for launch in one hour. Because of this the are confident that the time needed to prepare the launch vehicle for launch will remain well below the two hour eorgia Institute of Technology 71 of 196

72 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW requirement, LV-4.5. The Launch Vehicle Packing List can be found in Appendix I and the Launch Checklist can be found in Appendix II. eorgia Institute of Technology 72 of 196

73 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 4. Launch Vehicle Performance Analysis 4.1. Fin Design Performance characteristics as related to the aerodynamics of the launch vehicle primarily involve the management of stability factors throughout the ascent phase. Knowledge of the center of gravity (C) location with respect to the location of the center of pressure (CP) is critical in designing for static stability. Because the location of the C and CP change throughout flight, evaluation of these changes was completed through software simulation and analysis. Vespula Mk II utilizes a new stabilizing fin concept, as shown in Figure 20. Figure 20: Fin Sleeve The new fin concept is a modified tube fin design intended to provide more stability with less drag than a traditional tube fin. Additionally, the fins will be mounted to an adjustable sleeve for eorgia Institute of Technology 73 of 196

74 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW the purpose of altering fin position on the launch vehicle with respect to the nosecone. The fin sleeve design allows for the CP location to be altered before flight. Figure 21: Fin Sleeve Attached to Booster Section For such a design, flight simulation tools can not accurately predict performance with regard to stability factors. Evaluation of the CP location for the fin configuration was completed in Open Rocket. The Open Rocket aerodynamic model used for these approximations is illustrated below in Figure 22. Figure 22: Open Rocket Aerodynamic Rocket eorgia Institute of Technology 74 of 196

75 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Approximations were completed to account for the unique shape of the fins through modeling each fin as having a span equal the surface length from the fin root chord to a tip chord lying at the mid-point of the semi-circle, as shown in Figure 23. Figure 23: Fin Approximation of Modeling in Simulation Software In addition, each Vespula Mk II fin was modeled as two fins in the Open Rocket software program. Since the maximum number of fins available for modeling purposes is eight in Open Rocket, the CP location was plotted as a function of number of fins, as displayed in Figure 24. eorgia Institute of Technology 75 of 196

76 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 24: CP as a function of the Number of Fins A trend-line was formed from the data plot in order to develop equation X, which provides the CP location, in inches, with respect to the nose cone tip with x representing the number of fins. CP location =. 062x x x (12) From equation (12), the CP location is calculated to be 81.4 inches from the nose cone tip. The CP and C locations are separated by a designed distance to allow for 1.75 calibers of static stability at lift-off as illustrated previously in Figure Fabrication and Materials The modified tube fins will be composed of a carbon fiber/kevlar hybrid. These fins will be fabricated using a vacuum bagging technique over a mold made of an appropriately-sized steel rectangular block and a steel half-cylinder welded together. The adjustable sleeve on which the fins will be mounted will be a hollow carbon fiber tube fabricated in a similar manner. The sleeve will have holes drilled into it on either end, which will be used to fasten the sleeve to the eorgia Institute of Technology 76 of 196

77 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW rocket body. The fastening points on the rocket body will be placed one half inch apart, allowing the CP to be adjusted by moving the fin sleeve over a range of four inches Modified Tube Fin Sub-Scale Test Flight Verification of the evaluated results from flight simulation was completed through the development of a 45% scale test vehicle. The test flight served to verify the driving concept behind the experimental fins and visually inspect the vehicle s response to flight conditions in order to validate CP location approximations. Figure 25 displays the 45% scale launch vehicle on the launch rail and during ascent. Figure 25: 45% Scale Test Rocket and Flight eorgia Institute of Technology 77 of 196

78 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW The sub-scale model was successfully launched and recovered twice. For both flights, altitude and velocity were recorded using a Featherweight altimeter. Figure 26 and Figure 27 below display altitude above sea level and velocity as a function of time for sub-scale flight one. Figure 26: Sub-scale Flight One Flight Data eorgia Institute of Technology 78 of 196

79 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 27: Sub-scale Flight Two Flight Data Sub-scale flight one and two both utilized a 350 Ns solid rocket motor for thrust. Data from the two flights was reduced to determine vehicle acceleration during the initial coast phase of flight. Based on knowledge of vehicle mass at burnout, weather conditions, and acceleration, a drag coefficient for the sub-scale model was derived with 95% accuracy. For flight one and two, a CD of and was calculated, respectively. Sub-scale flight testing was successful as it met the goals of proving the experimental fin design and collecting flight data to yield a CD value CP Location Center of pressure calculations by use of flight simulation software was not permitted due to the launch vehicle fin design. An algorithm was created to solve the launch vehicle CP based on fin sleeve location. The method was created by calculating the projected area of each component eorgia Institute of Technology 79 of 196

80 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW and the CP of each component with respect to a reference line, in particular, the launch vehicle nose cone tip. Table 24 below defines the variables used in the following equations. Table 24: CP Variable Definitions Variable D BT L BT SD TC RC K R 1 R 2 Definition Body tube diameter Body tube length Sweep Distance Tip Chord Root Chord boat tail length Boat tail base diameter Boat tail forward end diameter The equations below were formed to locate the center of pressure on the vehicle. A BT = D BT L BT inches 2 A Fins = (# of fins) Span 1 SD TC + 1 [RC (SD + TC)] inches2 2 2 A T = 2(R 2 )K + (R 1 R 2 )K inches 2 A NC = 82.3 inches 2 A Total = A BT +A Fins + A T + A NC CP location for each component from aft end of launch vehicle: eorgia Institute of Technology 80 of 196

81 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Body Tube CP = L BT 2 + K 7 Fin CP = D F + 6 (TC) (RC) 8 3 (SD) RC Boat tail CP = K(R R 1 R 2 + 3R 2 2 ) 4(R R 1 R 2 + R 2 2 ) Launch Vehicle CP location from aft end: Nose Cone CP =.5L NC + K + L BT CP From Aft = Body Tube CP A BT + Fin CP A Fins + Boat tail CP A T + Nose Cone CP A NC A Total Launch Vehicle CP location from nose cone tip: CP From Forward = K + L BT + L NC CP From Aft Table 25 below displays the CP location with respect to the nose cone tip as a function of fin location from the aft end of the boat tail. Table 25: Fin Position vs. CP Location Fin Base Distance CP Location from from Aft End (in) Nose Cone Tip (in) eorgia Institute of Technology 81 of 196

82 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Fin Base Distance CP Location from from Aft End (in) Nose Cone Tip (in) Nose Cone The nose cone style selected in termed a Von Karman nose cone. Von Karman nose cones are designed for a theoretical minimum drag, and described mathematically by the following equations: θ = cos 1 1 2x L (13) y = D/2 π θ sin 2θ 2 (14) The variables are defined below in Table 26. Table 26: Nose Cone Symbol Definitions Symbol θ x L y D Definition Surface Turning Angle Incremental Length from Nose Cone Tip Overall Nose Cone Length Incremental Distance from Nose Cone Centerline Maximum Nose Cone Diameter These equations yield a nose cone 25 inches in length with an outer diameter of 5 inches. The shape of the nose cone is illustrated below in Figure 28. eorgia Institute of Technology 82 of 196

83 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 28: Nose Cone Profile 4.4. Motor Selection Current mass budget predictions list a gross lift off mass of kg, or lbs. with added margins. Simulations were performed using Open Rocket to calculate expected performance measures. Based on mass estimates and C/CP locations, an L2200 motor has been selected for the competition flight. Simulated flights in Open Rocket list a max altitude of 5,305 ft. with the current mass budget and launch vehicle design. Figure 29 and Figure 30 are plots of altitude and velocity versus time. As shown, the launch vehicle velocity remains sub-sonic as required throughout the flight. eorgia Institute of Technology 83 of 196

84 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 29: Vespula Mk II Altitude Plot Figure 30: Vespula Mk II Velocity Plot eorgia Institute of Technology 84 of 196

85 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 31 and Figure 32 display thrust and propellant mass versus time for the flight using an L2200 motor. Figure 31: Thrust vs Time for L2200 Motor Figure 32: L2200 Propellant Mass Burn eorgia Institute of Technology 85 of 196

86 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW As shown in the above figure, the L2200 motor has an average thrust of 504 lbs and maximum thrust of 695 lbs. Additionally, the motor has a 2.3s burn time and is rated at 5104 Ns CP and C Windy flight conditions can cause the launch vehicle s angle of attack, and hence CP location, to change during flight. An increase in angle of attack causes the CP location on the rocket to move forward. Simultaneously, in response to the change in angle of attack, the launch vehicle changes pitch in attempt to return to a zero angle of attack, returning the CP to its original location. The pitch and pitch rate of the launch vehicle in response to angle of attack perturbations are determined by the distance between the C and CP locations Stability versus time was plotted with C/CP margin of 1.75 at lift off. Figure 33 below shows the simulated stability margin for the flight. As shown in the figure, the launch vehicle maintains approximately 3 calibers of stability after motor burnout. Figure 33: Vespula Mk II Stability Margin eorgia Institute of Technology 86 of 196

87 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 5. Flight Experiment 5.1. Introduction to the Experiment and Payload Concept Features & Definition With the rise of entrepreneurial space flight, many new exotic spacecraft are being designed for the purpose of finding a profit in space. Many of these spacecraft will be equipped with liquid fuel propulsion and attitude control systems, or will seek to store large quantities of liquid propellant. These liquids present difficulties in the design and operation of a spacecraft because in low gravity, the fluids will be dominated by a combination of capillary/inertial/gravity gradient forces and will respond to perturbations. The response of stored liquids to such perturbations is termed slosh, and slosh is known to 1) alter the inertia matrix of a spacecraft and 2) to hamper the use of vents and propellant feed lines. Some methods of controlling slosh are listed in Table 27. Table 27: Methods currently available for damping slosh. Damping Method Description Tank geometry The choice of tank geometry (cylindrical, spherical, toroidal, etc) is known to have an impact on slosh damping through viscous effects. Ring baffles Annular disks along the circumference of a tank that impede slosh and may be given various camber geometries. Lids and mats Lids and mats float on a free surface of the liquid and impede slosh. Floating cans Cans impede slosh by absorbing and dispersing the kinetic energy of the liquid. Expulsion bag or diaphragm Bags and diaphragms reduce slosh by containing the propellant and forcing it into propulsion feed lines. Non-ring baffles Non-ring baffles are baffles that do not necessarily follow a tank circumference, e.g. cruciform baffles. Flexible baffles Flexible baffles are baffles made of flexible materials that deform under the inertia of sloshing liquids. While present methods of reducing slosh may be very effective in some flight regimes, there are design issues inherent to some of these systems. For baffles perhaps the most effective eorgia Institute of Technology 87 of 196

88 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW dampers for the additional inert mass instabilities can occur during launch if propellant levels are below the lowest baffle as in the case of the Saturn I. Similarly, such problems could occur in low gravity situations where the baffles are rendered ineffectual from lack of contact with the liquid. However, with the expense of mechanical complexity and inert mass, expulsion bags and diaphragms can be used to avoid such instabilities. The intend to provide another alternative solution by demonstrating the use of magnetorheological (MR) fluid as a moveable, deformable baffle and potentially a diaphragm equivalent Accomplishments Since PDR The science section of Flight Systems has made progress towards several milestones since PDR, and these accomplishments are listed in Table 28. Table 28: Accomplishments since PDR Accomplishments since PDR 1 round testing design development and design review 2 Testing redefinition and re-scoping 3 Theoretical modeling and simulation advances 4 Acceptance to REFP round testing designs were the emphasis of much of the science team work since PDR. iven the importance of experimental design on data collected, the team has thus far accepted delays in development for the purpose of enhancing designs. Emphasis on effectively testing what is desired with cost/time optimization for manufacturing has been the emphasis and has produced promising results for the spring semester. A lengthy process of design iteration is finally ending and manufacturing as well as testing should begin soon. Also, advances in theoretical modeling towards a more unified theory of slosh and MR fluid were also pursued. These developments give results that should enhance the development of control systems for MR fluid during ground testing. Finally, the science team was accepted into the Reduced ravity eorgia Institute of Technology 88 of 196

89 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Education Flight Program, and will be further developing a proposed up-scale LSIM and feasibility test in the coming months Science Background A complete science background is included in Appendix XXX and references are included in Appendix XXX Important Highlights In the science background, Appendix XXX, several important relationships are developed. These relationships are given in Table 29. Table 29: Elements of the theoretical modeling for the LSIM payload LSIM element Response of MR fluid to a magnetic field (motion as it rigidifies) Longitudinal Slosh Model Lateral Slosh Model Damping of MR Fluid if it is a rigid baffle Damping of Container Relationship L = g b 2 m ΔL k m L gk θ = (kl + mg) θ b 1 θ m By experiment and in correspondence with reference material tabulated data and plots From reference material, δ = 4.98ν 1/2 R 3/4 g 1/ Overview of the Experiment eorgia Institute of Technology 89 of 196

90 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Hypothesis and Premise The hypothesis posed in the LSIM experiment is that If a baffle can be manipulated during the flight of a spacecraft, then unstable slosh can be actively damped. The experiment will apply radial magnetic fields to the propellant tank to manipulate and rigidify the MR fluid during the microgravity phase of the launch vehicle trajectory to perform Liquid Stabilization in Microgravity LSIM. The launch vehicle ascent will provide a high vibrational intensity environment in which to test the anti-slosh system. Furthermore, the use of diaphragms and propellant bags are eliminated with the assumption that: Trading mechanical complexity for electrical complexity is preferable from a reliability standpoint. Therefore, the will implement a design to apply these concepts to both the launch vehicle and REFP Experimental Method and Relevance of Data The experimental method for LSIM requires a multi-step approach for ground testing, flight testing, and REFP. The purpose of ground testing will be to characterize the shear stress behavior of MR fluid of different composition and magnetic field configuration, the manipulation of MR fluid, and preliminary data on slosh damping ability. Flight testing will provide actual data on the capability of the MR fluid system to dampen slosh, especially in the microgravity environment. REFP would seek to explore a big-picture system that actively attempts to remove any stray MR fluid as propellant simulant is pumped out of the tank. In any of the test cases, an optimal mixture of MR fluid will enable an application of active control to maneuver MR fluid into position in flight. The testing cases are organized by the team testing matrix for LSIM, which is designed to enable comparative analysis of the results and to verify eorgia Institute of Technology 90 of 196

91 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW completion of the data set. Following the scientific method, the test matrix outlines control experiments and baseline comparisons to develop a qualified understanding of MR fluid in the context applicable to LSIM. A summary of scientific method fulfillment is given in. Table 30. Table 30: Scientific method fulfillment for LSIM Method step Fulfillment Question What are options for electrically damping slosh? Research Study of MR fluid and a review of The Dynamic Behavior of Liquids in Moving Containers Hypothesis If a baffle can be manipulated during the flight of a spacecraft, then unstable slosh can be actively damped. Test round testing plan and test matrix, flight test, REFP Analysis Data examination, post-processing, and analysis Communicate SLP documentation and VTC Furthermore, in an improvement over previous experimental design, the team intends to fly a control experiment as part of the flight test, permitting greater validating capability for the effectiveness of the damping system Testing plan Overview To accomplish the objectives of LSIM, several distinct testing sequences are necessary. Key to the success of LSIM is ground testing, where MR fluid mixtures will be characterized and manipulated with solenoids. Following on these tests are the USLI flight test and separately the REFP project. However, at this juncture of the project some testing has been de-scoped, namely shake-table testing of the bench test platform to demonstrate slosh reduction in 1-gee. This test was de-scoped due to the complexity and time constraints of procuring an appropriate shake-table. eorgia Institute of Technology 91 of 196

92 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW eorgia Institute of Technology 92 of 196

93 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 31: Test sequences and descriptions, included options de-scoped since PDR Test Sequence Explanation Purpose MR fluid characterization A two-plate test To determine experimentally the viscosity performance of MR fluid mixtures. Bench testing Solenoid operations To develop a method of control for rigidifying and raising MR fluid within a canister. Descoped sequences Shake table testing Formerly, to test slosh reduction in 1-gee. Launch Vehicle test USLI flight test Control and experiment test inside the launch vehicle to determine comparative reduction in slosh. REFP Up-scaled testing and feasibility Microgravity University study of MR fluid slosh cooperative project reduction. A brief description of each testing sequence is given in Table 31 above and the relationship between these testing sequences is illustrated in Figure 34 below. eorgia Institute of Technology 93 of 196

94 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 34: LSIM testing logic, illustrating a simple relationship of information between the test sequences and emphasizing that they flow down from the pursuit of the LSIM hypothesis. round testing will serve four general purposes: (1) the creation of MR fluid, (2) the verification and validation of theory and control systems, (3) the characterization of MR fluid, and (4) the development of a working model for flight testing. For the successful completion of ground testing, the team will create an optimal mix of MR fluid. An optimal mix will depend on the fluid's balance between rigidity and fluidity for manipulation under a magnetic field, such that the MR fluid is easily moved to an appropriate location in the tank. Verifying the Ramblin' Rocketeers' solution and theory of using MR fluid as a baffle to dampen unstable slosh will go through two phases. During phase one, only MR fluid will be subjected to a magnetic field. Phase two will include water along with MR fluid being subjected to a magnetic field. The results from these phases will indicate whether the solution is feasible by observing the controllability of MR fluid by a magnetic field as well as observing differences between MR fluid and the propellant simulant. By characterizing the MR fluid, the team will understand the eorgia Institute of Technology 94 of 196

95 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW various properties of the MR fluid such as its exerted shear force and how it changes under a magnetic field. The characterization process will include testing the force and viscosity of the MR fluid and observing preliminary slosh damping. Further testing for infrared reflectance of MR fluid relative to water may be conducted so that position information may be derived from a camera installed within the flight model. Finally, a working ground model will be developed using the results from (1), (2), and (3) with constraints for flight experimentation MR Fluid Creation and Validation of Theory MR fluid can be created from three ingredients: carrier oil, magnetic particles, and surfactant. Table 32 provides example MR fluid ingredients in the design space. Table 32: List of MR fluid ingredients Carrier Oil Magnetic Particles Surfactant Mineral Oil IRON100 Powder Citric Acid Nanometer particulate ferrofluid IRON325 Powder Oleic Acid FE Powder Soy Lecithin Fe304 M1 Powder For a preliminary ground test in search of better understanding the behavior of MR fluid thereby making more informed decisions on the design space the team opted to use mineral oil, IRON325 powder, and oleic acid. By trial and error testing, the team created a stable MR fluid mixture using the aforementioned ingredients. The team created two mixtures of differing viscosities. While some sources had presented the iron concentration as 60% by mass, the preliminary tests found it necessary to increase this percentage. The first mixture resulted to be too fluid with 17 grams of mineral oil, 1 gram of oleic acid, and 56 grams of IRON325 powder (76% by mass). The second mixture resulted to be too viscous with 16 grams of mineral oil, 1 gram of oleic acid, and 56 grams of IRON325 powder (77% by mass). The ingredients were eorgia Institute of Technology 95 of 196

96 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW measured using a scale accurate to a gram. Future measurements will use a more accurate scale. From trial and error testing, the team created an MR fluid testing matrix that will test every possible combination between ingredients as well as small deviations from the trial and error test. For example, the team will gradually decrease the iron percentage by mass while gradually increasing the mineral oil percentage until the optimal mixture a mixture that appears to be rigid enough to act as a baffle and manipulative enough to move readily has been attained. Each mixture will be static tested by neodymium magnets and good mixtures may be tested with solenoids as ground testing improves. Validation of theory and control of MR fluid will occur if there is a change in the MR fluid's viscosity under a magnetic field. Figure 35: Preliminary static testing of MR fluid mixtures in magnetic fields From the results of preliminary testing, the composition of MR fluid is likely to be changed to using carrier oil made of ferrofluid. Ferrofluid is a mixture nanometer-scale ferromagnetic particles in oil with a surfactant. However, unlike MR fluid, ferrofluid does not have as high a percentage of pure iron and does not rigidify in the same manner as MR fluid. As carrier oil, the team hypothesizes that ferrofluid will increase the mobility and useability of the MR fluid mixture; even with 60% and greater mass ratios of iron powder. Furthermore, smaller iron particulates may also increase the mobility of the MR fluid. reater mobility than the initial mixtures is preferred such that the MR fluid may be moved to the final baffle location using solenoids, and eventually for the mobility desired for REFP. eorgia Institute of Technology 96 of 196

97 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW eorgia Institute of Technology 97 of 196

98 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW MR Fluid Shear Stress Characterization: Two Plate Test The team will determine the shear stress MR fluid exerts inside and outside magnetic fields to better understand how to manipulate the MR fluid as desired. To determine the shear stress, the team will perform a two-plate test with and without magnetic field acting upon the MR fluid. This test was chosen because of its simplicity; other tests such as a barometer test were considered for measuring the MR fluid's viscosity and force they turned out too complicated to realize. The two-plate test consists of two plates: a bottom plate, which is fixed to the ground and a top plate, which is free to move. A load sensor will be placed on the top plate to measure the reaction force that is generated. The current plate choice is acrylic. Figure 36: Shear stress of a fluid using the two-plate test (Source: Wikipedia) A control test will be performed by only having two plates together with a load sensor on the top, moving plate to calculate the frictional force by the plates themselves. For accurate and consistent results, a mechanical pulling device will be used to pull the top plate. Once a control has been measured, a quantity of MR fluid will be placed between the two plates and the same procedure will repeat with and without the MR fluid under a magnetic field. These tests will characterize the force that MR fluid will generate when it is under a magnetic field and when it is eorgia Institute of Technology 98 of 196

99 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW free of a magnetic field. A complete ground testing plan description is included in the round Test Plan appendix Detection of motion The IR detection scheme discussed in PDR has been de-scoped. A digital video camera will recoding video of slosh from above while a piezoelectric element placed in the mounting bracket will provide information on sloshing motion via vibrations in the structure. Control tests will permit the use of this data in understanding slosh Working round Model The team will develop three methods of MR fluid control: an array of solenoids, a movable solenoid, and a fixed solenoid. For the launch vehicle and REFP, solenoid arrays appear to be the best current option Sensors A camera and piezoelectric sensor are planned for the canisters flown inside the launch vehicle Design review Viscosity Test Rig A two-plate test rig was designed to characterize the viscosity of MR fluid. The design of this test rig underwent many revisions in order to meet measurement and budget requirements. The objective of this rig is to measure the force of fluid acting on the plate and ultimately, the viscosity value of the MR fluid. The reaction force and the viscosity constant are related by the following equation: F fluıd A = ηv 0 D (1) eorgia Institute of Technology 99 of 196

100 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW In Eq.(1), A is the surface area of the plates, V 0 is the pulling velocity of the top plate, D is the distance between the plates, and η is the viscosity coefficient. The surface area of the plate is in 2, and the distance between the plates is estimated to be about.25 in. A motor is used to control the velocity, so V 0 is also known. F fluid is measured with a nichrome wire device. This device measures force through changes in resistivity induced by small changes in cross-sectional area resulting from tension in the wire. The measurement limits for this device are set by the accuracy of electrical interface hardware and calibration testing. One of the design requirements is to simulate a non-frictional surface. The aim is to ensure that the only force acting on the plate is the fluid s shear force. This is accomplished by adding railings between the plates and the base of the structure. Although the railings do not completely remove friction, they minimize it enough so that friction is negligible. Part drawings for the test rig are available in the Appendix. The drawings shows locations of the railings attachment points and the tap needed as well. The railings require a #8 type screw, and a ¼ in tap is needed for hard woods and acrylic sheets with this type of screws. The dimensions and locations of the features on these parts are mainly driven by two parameters: solenoid strength and test time. Because the magnetic strength of the solenoid decreases drastically as the distance increases, the plates need to be close to the solenoid base. This requires short railings. The group also wants to maximize the contact time between the plates and the MR fluid, so the plates are designed to be long and skinny. Solenoids are aligned along the length of the plate to produce a uniform magnetic field during testing. Construction of the test rig depends on a number of assumptions and is subjected to revision for alternative methods if necessary. First, super wood glue will be used to connect wood pieces. If this is not sturdy enough, elbow brackets will be used to connect the corners of the wood pieces. The motivation for using glue instead of brackets is saving money. Second, wood pieces and acrylic sheets will also be glued together with epoxy. This method is the norm for connecting wood to acrylic sheet, and it saves space and money. A #8 type screw will be used to tighten the eorgia Institute of Technology 100 of 196

101 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW railings to the woods and acrylic sheets. The railings attachment holes are designed to provide a bit of a leeway. Attachment holes on the structure pieces and test plates can be off by about.05 in. A motor will be used to pull the top plate. The motor will stand on a piece of wood that has been designed with a height that will perfectly align the motor with the top plate. The tolerance for the height offset is ±.04 in. String and hook will be used to round out the pulling mechanism. This testing rig was delayed due to the design process and the winter break. Ordering and construction are planned to begin coincident with CDR and testing could begin around the time of CDR VTC in late January round Test MR fluid production and manipulation In order to meet the system functional requirements, the MR fluid must have certain properties and also adhere to certain standards. The fluid must be sufficiently rigid when magnetized to not shear significantly or break due to fluid slosh; however, the fluid must also not be excessively resistant to motion when moving and shearing against a wall, so that it may be moved into position by the magnets. It is known that the size of iron particles makes the largest difference in the rheometry of the fluid. For instance, one batch of low quality fluid that was created earlier during preliminary testing, with larger iron particles than is typical of MR fluids, was found to be extremely resistant to motion. Therefore, to find a high-quality fluid with intermediate properties, it is wished to test iron powders with particles of mean diameter between 0.1 µm and 10 µm. In addition, to ensure purity, we shall attempt to purchase all powders from well-known sources. For example, one option being explored is a purchase of carbonyl iron powder from BASF. For the carrier fluid, mineral oil or hydraulic oil are both known to be fairly typical choices; the properties of the fluid should not be significantly affected by which is chosen. The final choice of components and their proportions will be made based on the results of the two-plate testing, as well as qualitative experience from attempts to move the fluid by manually moving magnets. eorgia Institute of Technology 101 of 196

102 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW The MR fluid will then be manipulated by solenoids in a ground testing platform that permits the placement and use of solenoids to control the MR fluid. The foundation of the bench test is built from Maker Beam parts. This allows for great configurability and flexibility. Four 30 cm Maker Bars are the corner stands of the test rig; they are placed vertically about 20 cm apart from each other. Four 20 cm Maker Bars are placed horizontally in between the vertical 30 cm Maker Bars and are attached to give support. There are three thin acrylic plates: one acts as the main base and the other two have holes cut for the beaker to fit inside the plates. The base plate sits on 4 flat L Maker Beam brackets about 2 cm above from the base. The 4 L Maker Beam brackets are attached to the four 20 cm horizontal Maker Bars. The beaker is placed in the middle of the base acrylic plate and the solenoids are placed radially around the beaker on the base plate. Four 90 Degree brackets are used to hold the second plate 2 cm above the first plate. The second plate has a hole cut in the middle to allow for the beaker to pass through the middle. More solenoids can be placed around the second acrylic plate. The same is done with the third acrylic plate as was done with the second acrylic plate, but 4 cm above the first, base acrylic plate. This design using Maker Beam parts allows for future design modification and addition of parts. Once more data has been collected, the team can attach a vibration motor underneath the first base plate to simulate slosh. Solenoids can be added, moved, or removed from each of the three acrylic plates during testing as test results shed more light on what is needed for more accurate testing. The top two acrylic plates with holes in them can be moved up or down the 30 cm maker beam bars to adjust the height at which the solenoids interact with the beaker Payload Relevance and Science Merit The top priority of the Flight Systems team during project development was to create a payload concept leveraging team expertise while pursuing achievable and NASA-relevant experiments. Previously, the project investigated moving oxygen gas with an electromagnet eorgia Institute of Technology 102 of 196

103 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW essentially a steady-state siphon for paramagnetic materials. The USLI team investigated active platform electromagnetic stabilization, developing control algorithms for magnetic levitation during flight. After review by Flight Systems and the eorgia Tech, the team decided that the most relevant primary payload would be to demonstrate the use of MR fluids in anti-slosh applications using technology development from the and eorgia Tech USLI experiments. Combining technologies from the previous projects, the new LSIM payload will demonstrate a possible method to combat propellant sloshing. The benefits of such an anti-slosh system would be most applicable in deepspace long-duration missions. In such missions, large quantities of fuel must be stored and/or transported with cargo/personnel. A major issue in low-gravity environments for propellants is sloshing, where fluid begins to float freely in space relative to the propellant tanks. Sloshing may cause loss of pressurization in propellant feed systems, potentially creating dangerous propulsion failures. The current solution is to create a moveable and deformable baffle from MR fluid. Using electromagnets, the controlled fluid may then be used to dampen the propellant oscillations. Systems might be needed to insure that the fluid is removed from the propellant, and a magnetic siphon could be used if the mixing between fluid and propellant is minimal. This is the basis for the REFP experiment discussed later in this document. enerally however, the LSIM experiment is a science and engineering payload that involves phenomena from several fields, primarily magnetism, rheology and viscous flow, as well as nearinviscid fluid dynamics. Among the goals of LSIM is to develop a scientific model encompassing all of the above fields in order to understand the interactions between the various components of the system. This will be achieved by combining theory with experimentation and testing. Data will be collected for variables such as MR fluid position, MR fluid shear stress, and simulant position and acceleration as a function of time, rocket acceleration, and electromagnet currents and positions. Collecting this experimental data will enable changes in the applied control scheme to be made according to the observed data, as well as allowing for refinement of the dynamic and scientific model of the MR fluid-propellant simulant system. A full explanation eorgia Institute of Technology 103 of 196

104 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW of the science of slosh and MR fluid relevant to LSIM is included in the Science Background appendix. For MR fluids, the primary focus of research in current years has been on the properties of the MR fluids themselves, and on their interactions with solid objects or containers, rather than on their interactions with other fluids. Therefore, the LSIM experiment should give insight into this less-studied subject. In addition to the above modeling, there are other scientific benefits of this experiment. The behavior of MR fluids in microgravity has been of significant interest, with the InSPACE experiment on the International Space Station being a large-scale investigation on this topic. However, engineering applications of the fluid specifically in microgravity do not seem to have been investigated to the same extent. Microgravity is one of the places where MR fluid is likely to be most effective, as settling of iron particles and thus degradation of integrity does not occur in the near-absence of gravity. Therefore, the LSIM experiment allows for investigation of actual applications of MR fluids in microgravity, as well as scientific modeling of the MR fluidsimulant dynamics Experiment Requirements and Objectives Success Criteria Minimum and maximum success criteria have been defined for the LSIM payload. Table 33 lists these criteria. Table 33: LSIM success criteria from the Requirements Verification Matrix Minimum Maximum LSIM Success Criteria Successfully record video of flight experiment during microgravity and start/stop the experiment without mechanical and electrical failures. Successful matching of the damping ratio for ringed baffles in the wave amplitudes experienced during flight to within ±30%. eorgia Institute of Technology 104 of 196

105 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirements The requirements set for the LSIM experiment to satisfy both the goals of and the USLI requirements are listed in Table 34. Flight systems (experiment and avionics) are now budgeted to be 15 lbf. Table 34: LSIM Requirements Requirement Requirement Source Verification Design Status Verification Number Definition Method Feature Source Document FS-1 The flight systems team shall design and build the LSIM MO-3 Inspection LSIM payload In Progress MO-3 Payload FS-2 The LSIM payload shall be designed to fly on a SLP rocket USLI Handbook Inspection LSIM payload In Progress USLI Handbook FS-4 The Flight Systems Team shall produce a working system for manipulating MR MSC-3 Testing Solenoids and Control Algorithms In Progress MSC-3 fluid in LSIM. FS-5 The Flight Systems Team shall ensure that all avionics are properly shielded from the LSIM payload. MSC-3 Testing Faraday cages and webbing tied to ground on the harness Not Started MSC-3 eorgia Institute of Technology 105 of 196

106 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement Requirement Source Verification Design Status Verification Number Definition Method Feature Source Document FS-6 The Flight Systems Team shall design all LSIM components and avionics such that they may be easily integrated with MSC-3 Inspection Mounting system Complete MSC-3 the Modular Payload System of the payload bay in the rocket. FS-7 The Flight Systems Team shall conform to all weight, power, and dimensional MSC-3 Analysis TBD In Progress MSC-3 requirements as per the rocket design. FS-7.1 The Experiment and Avionics, with mechanical supports, shall weight no more LV-1.1 Inspection TBD In Progress LV-1.1 than 15 lbf. eorgia Institute of Technology 106 of 196

107 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement Requirement Source Verification Design Status Verification Number Definition Method Feature Source Document FS-8 The flight computer shall execute all tasks necessary to the operation of the MSC-3 Inspection Maple SIDES node In Progress MSC-3 LSIM payload and avionics. FS-9 The LSIM payload shall have a dedicated power MSC-3 Inspection SIDES node In Progress MSC-3 supply. FS-10 The Flight Systems Team shall ensure redundancy and reliability of all MSC-3 Inspection SIDES network In Progress MSC-3 internal electrical hardware. FS-11 The Flight Systems Team shall provide for payload operation with up to 1 hour of wait on the launch pad and 2 hours of USLI Handbook 1.6 Inspection TBD In Progress USLI Handbook 1.6 wait during preparation of the Rocket. eorgia Institute of Technology 107 of 196

108 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement Requirement Source Verification Design Status Verification Number Definition Method Feature Source Document FS-12 The Flight Systems Team shall provide for electrical operations to begin at MSC-3 Inspection TBD In Progress MSC-3 the beginning of the flight trajectory. FS-13 The Flight Systems Team shall ensure that the LSIM payload is shut down safely during the MSC-3 Inspection TBD In Progress MSC-3 deployment phase of the flight trajectory. FS-14 Data from the LSIM payload shall be collected, analyzed, and reported by the team using the USLI Handbook 3.2 Inspection Data logging in SIDES network In Progress USLI Handbook 3.2 scientific method. FS-15 The LSIM payload will be designed to be recoverable and be able to launch again on the same day without any repairs USLI Handbook 3.5 Inspection Appropriate mounting to the payload interface. In Progress USLI Handbook 3.5 or modifications. eorgia Institute of Technology 108 of 196

109 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 5.9. REFP The were selected for REFP. Reporting on the progress of this portion of LSIM will be to the JSC Microgravity University and TEDP documentation Flight Experiment Integration The payload includes all experimental components. A possible configuration for the payload is shown in Figure 37. The assembly is made of four parts: the base bolt, the base, the payload plug, and the payload. eneral dimensions for the payload are listed in Table 35.. Payload Plug Base Bolt Payload Base Figure 37: Payload Assembly Table 35: Payload Assembly Dimensions Parameter Value Base Diameter 4.97 Total Height Payload Height 8.95 Base Thickness 0.1 The experiment is housed in a PVC plastic pipe that is connected to a base. The payload base is designed to be the only load bearing component of the payload assembly. eorgia Institute of Technology 109 of 196

110 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 38: Payload Base with 150N of loading The base rests in the rib of the structure and holds all of the weight of the payload and any sensors used. It is made of Delrin plastic and manufactured using injection molding. The payload base can support roughly 60.85lbs of load before failure. It is designed to support an assumed maximum load of lbs with a factor of safety of 2. This load comes from the assumption that the payload weighs no more than 3lbs accelerated at 10 times the acceleration due to gravity. Figure 38 shows the stress distribution through the base, using SolidWorks SimulationXpress Wizard. Table 36: Data from SolidWorks SimulationXpress, highlighting the data from assumptions Trial Total Load (lb f ) Max Stress (psi) Factor of Safety eorgia Institute of Technology 110 of 196

111 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Trial Total Load (lb f ) Max Stress (psi) Factor of Safety Table 36 shows data taken from SolidWorks SimulationXpress for the payload base. This data was interpolated to find the maximum load of the payload base. Figure 39 shows the factor of safety plotted versus the total load on the payload base. The graph and equation allow the approximate maximum load to be determined mathematically before constructing the first prototypes. eorgia Institute of Technology 111 of 196

112 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Factor of Safety vs. Total Load Payload Base (Delrin 2700) Factor of Safety Factor of Safety 5.00 y = x Total Load (lbs) Figure 39: Factor of Safety vs. Total Load from SolidWorks SimulationXpress and generated trend line equation 6. Flight Avionics Feedback is essential to any meaningful design work. In recent years, the have implemented a number of unique launch vehicle designs, each with the intention of finding solutions to particular problems. However, with only limited visual feedback available, it is difficult, if not impossible to gauge the success of a design or to detect any unanticipated failure modes. A system that could accurately describe the state of the rocket throughout its flight would then be enormously valuable. To be effective, such a system would have to be capable of not only recording data from multiple sources but also able to temporally connect the data. This would provide the user insight into the interactions between different factors in addition to the individual measurements. Due to the potential complexity of such a design, the system also needs to be tolerant to the potential failure of any singular functional unit. This would ensure that even if some information is lost, the system will still yield meaningful feedback from tests. Finally, it would be helpful for such a system to be extensible. It is impossible now to envision eorgia Institute of Technology 112 of 196

113 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW all of the potential use cases for such a system. Designing it to be easily adapted to the needs of future projects would help ensure its success and longevity Avionics Overview The avionics are designed to accommodate the primary science payload LSIM, in addition to supporting structural and aerodynamic analysis of both the advanced fin design and the rib and stringer fuselage design. To accomplish this goal, SIDES (Simultaneous Independent Data Logging & Experiment System) is being developed to maximize the data extracted from each flight while reducing the risk of failure of a larger avionics system. SIDES architecture allows for a flexible, complex, and fault tolerant distributed data collection system for the Ramblin Rocketeers launch vehicle. eorgia Institute of Technology 113 of 196

114 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement Number Table 37: Avionics requirements Requirement Definition Source Verification 1. The flight avionics shall collect data required for a successful payload experiment. 2. Key elements of the flight systems shall operate on independent power supplies. 3. Power supplies should allow for successful payload operation during launch vehicle flight with up to 1 hour of pad stay and 2 hours of standby time during launch vehicle preparation. 4. The flight avionics shall be capable of being attached to the launch vehicle structure. USLI Handbook 1.7 Method Testing Design Feature Data logger MSC-3 Testing SIDES nodes MO-4 USLI Handbook 2.11 Testing Testing Battery systems and power management Mechanical Interfaces Status In Progress In Progress In Progress In Progress Verification Source Document eorgia Institute of Technology 114 of 196

115 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Requirement Requirement Definition Source Verification Design Status Verification Number Method Feature Source Document 5. PS coordinates of the launch vehicle shall be transmitted to a ground station. USLI Handbook Testing PS, round Station, Xbee In Progress 6. Each avionics node shall be capable of data logging with or without a clock pulse. USLI Handbook 2.12 Inspection Flight Software and Data Logger In Progress 7. Each avionics node shall operate at some equal or reduced functionality during RS485 communication failure USLI Handbook Inspection Flight Software and Redundant Node Hardware In Progress 6.2. Avionics Success Criteria The success of the avionics team will be defined in two ways: minimum success criteria that will be accomplished if the requirements are accomplished, and maximum success criteria that will be met if everything goes according to plan. Maximum success will include collecting diagnostic data for the launch vehicle, such that design feedback is available for iterating the most effective launch vehicle design, while minimum success is limited to successfully collecting and storing the LSIM payload data for recovery and analysis of the data. eorgia Institute of Technology 115 of 196

116 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 38: Avionics Success Criteria Requirement Requirement Definition Source Verification Design Status Verification Numbers Method Feature Source Document 1. The avionics system is Analysis, In functional throughout Testing Progress the flight and if failures do occur the entire system does not go down. 2. The ground station Analysis, In should be capable of Testing Progress receiving supplementary data transmitted from the launch vehicle. 3. The ground station Analysis, In should detect the Testing Progress location of the launch vehicle throughout the flight, and track the location of the landing for recovery purposes SIDES Design Approach SIDES utilizes a distributed network of microcontrollers to accomplish diverse tasks. Each node in the distributed network is capable of operating independently of other nodes. To support this, each node has a self-contained power supply and data logging capability. This approach reduces risk by preventing the failure of any node from propagating through the SIDES network. eorgia Institute of Technology 116 of 196

117 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Distributed data logging presents a synchronization challenge when compiling distributed data. The integration of the data when clock skew is present becomes much more difficult and often involves resampling and interpolating the data to obtain useful results. By providing a synchronization clock signal, the local data logging rates can be easily adjusted to prevent clock skew. In ideal operating conditions, the individual nodes of the SIDES network will be able to communicate over a bus. For noise immunity, the bus will be a differential pair. To optimize the trade between failure tolerance and weight, electrical harness weight will be reduced by using a one-to-many, multi-drop bus rather than a point-to-point solution. Software control of the multi-drop bus nodes and the use of a synchronized clock signal will reduce the risk associated with centralized communication while maintaining the weight advantages of a multi-drop bus. eorgia Institute of Technology 117 of 196

118 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW SIDESboard The SIDESboard standardizes the nodes, and helps ease implementation of the electronics. The SIDESboard contains all the features necessary at each avionics node to successfully complete the mission. The SIDESboard has a standard harness connector, data logging SD (secure digital) card, battery monitoring circuit, isolated clock input and a standard mechanical footprint. The SIDESboard firmware incorporates a standard set of libraries. These libraries allow programmers to focus on the function of the specific node rather than having to code the same functionality each time. The communication bus for the SIDESboard is handled by an RS485 transceiver. The RS485 format is differential for noise rejection, bidirectional to save weight in harness wiring, and multi-drop to reduce wiring complexity while also saving weight. Risk of communication failure is considered to be acceptable for the purposes of saving weight, because the consequences are low-impact by design. Figure 41 and Figure 42 depict the SIDESboard PCB (Printed Circuit Board) design, supporting the features listed above. Figure 40: SIDES system layout eorgia Institute of Technology 118 of 196

119 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 41: SIDESboard bottom side view Figure 42: SIDESboard top side view SIDES Electrical Harness The SIDES electrical harness connects all the SIDES nodes together, but improper harness design risks significant weight penalties. Architecture decisions mentioned previously have already helped to reduce the number of wires in any particular harness run, leaving distribution of the communication lines, clock synchronization lines, and any other signals as the main concern. The two main busses could either be distributed as a daisy-chain configuration to save weight, or using as star topology as illustrated in Figure 43. Considering increased risk has already been allowed by the choice of RS485 bus configuration, the trade-off for less weight is increased risk of failure. With daisy-chain topology, a broken wire in the harness affects all the nodes past the breakage, an unacceptable additional risk to the system. This leaves the star topology as the ideal configuration. eorgia Institute of Technology 119 of 196

120 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Each wire run will have up to 9 wires and use zip ties to keep the wires together. The zip ties placed approximately every 6 inches along the wire will keep the runs bundled and organized as well as anchor the runs to the structure. Figure 44 depicts a similar idea to the electrical harness for SIDES. The weight estimation for the harness is 115% of the planned connectors and wires making up the harness for the estimated harness lengths. Figure 43: eneric star topology diagram Figure 44: Example of an electrical harness using zip ties and connectors Master IMU The master IMU utilizes a Maple board for increased computing power over the SIDESboard triple axis accelerometer, gyro and magnetometer IMU, and RS485 hardware. The Maple board, depicted in Figure 45, is the communication bus master and allows for the different nodes to eorgia Institute of Technology 120 of 196

121 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW communicate with each other if desired. In particular, the Master IMU will facilitate sending data of interest to the Telemetry node to be forwarded to the ground station Master Clock Figure 45: Maple board used in the Master IMU The Master Clock will send a synchronization pulse to all the other nodes. Each microcontroller has a clock input and return so that the clock can run to each node. The clock has a strong drive circuit to ensure that the nodes are consistently receiving clock signals. Figure 46 shows the connection layout of the Master Clock to all the nodes. eorgia Institute of Technology 121 of 196

122 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 46: The connections of the clock to different nodes Science Experiment Computer The LSIM payload requires several LEDs (light emitting diodes), multiple vibration sensors, and a camera in order to record data, such as the camera shown in Figure 47. The LEDs will be necessary to light up the magnetorheological fluid (MR fluid) so that the camera can show the fluid slosh. The microcontroller at this payload node will also actuate solenoids that will be used to control the MR fluid during the flight. 2 Figure 47: A possible camera used to analyze the payload experiment1f Telemetry 2 eorgia Institute of Technology 122 of 196

123 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW The Telemetry node fulfills the requirement 5 of transmitting the PS data from the launch vehicle to the ground station. The Telemetry node will make use of an Xbee PS transceiver and a SIDEDboard to log the PS data while the Xbee is transmitting the data. An example of the Xbee is depicted in Figure Strain auge Figure 48: Xbee transceiver unit The strain gauge node makes use of a Wheatstone bridge, to determine the deflections the launch vehicle is experiencing at various points in the launch vehicle structure De-scope Options As part of the Flight Systems package for the previous competition cycle, a computer handling telemetry and PS was built and flown. This computer has the capability to run a solenoid driver and read the vibration sensors. Should the SIDES network need to be de-scoped, this substitute hardware already exists and can be inserted into the system design with minor modification. More info on this computer may be found in De-scope Option: Flight Computer Definition Power Budget Table 39 detail the power budget for SIDES. eorgia Institute of Technology 123 of 196

124 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 39. SIDES Power Budget. eorgia Institute of Technology 124 of 196

125 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 6.6. EM Interference Faraday cages and shielding webbing may be used to mitigate the risk of EM Interference from both the telemetry devices and experiment solenoids. More analysis will be needed to determine the amount of shielding require for example, placement of electrical harness off-axis of the experiment solenoids can negate much of the EMI risk inherent in a magnetics experiment. The precise placement of harness has not been determined as of this point in the launch vehicle development. Redundancy and robustness is key to the SIDES network and any node failures should be survivable further the ground station will provide an added redundancy through more accurate communications should signal strength from the launch vehicle experience dramatic fluctuations Transmission Frequencies and Protocols The telemetry system is designed to utilize two Xbee PRO 900-XSC modules for one-way communication from the launch vehicle to the ground station. Using a simple, loss-tolerant protocol with reliable delivery ensures the data is received if at all possible and that the information is correct. The SIDES node controlling the Xbee module on-board the launch vehicle will utilize a 900MHz monopole-monopole vertically polarized rubber duck antenna with 2 dbi gain and 10W of power. This antenna s performance is depicted graphically in Figure 49. Receipt of PS data via radio to the ground station will satisfy the recovery requirement and bolster kinematics data of the launch vehicle trajectory. eorgia Institute of Technology 125 of 196

126 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 49: Antenna performance as a function of range 6.8. Software Maturity Flight software for the SIDES nodes is pending finalization of SIDES hardware and standards. round station software is in development and the progress achieved thus far is discussed below De-scope Option: Flight Computer Definition The following text is pulled from the 2012 FRR documentation regarding the flight computer, planned as a de-scope option for the SIDES network. Flight Computer The flight computer will run the ATMEA 2560AU processor with the Arduino bootloader and other necessary components for ease of programming. The chip has sufficient I2C, serial, and analog inputs to read data from all sensors and log to an SD card based on Sparkfun s OpenLog break-out board. Additionally, the chip will run the Fastrax UP501 PS module and send the data to an Xbee PRO for transmission to the ground station. An OpenLog board will provide logging capabilities. The chip will be programmed in the Arduino language, a subset of C++ eorgia Institute of Technology 126 of 196

127 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW with some additional libraries. Figure 50 provides a generalization of proposed flight computer software. Figure 50: eneralization of flight computer software The flight computer must accomplish several tasks and handle multiple responsibilities. The main goal of this system is to collect and monitor all the relevant data from the environment around it such as the strain on the launch vehicle, environmental factors such as temperature, stray magnetic flux from the A.P.E.S system, launch vehicle acceleration, and PS position. During flight, the flight computer must also monitor the payload's control system and data through a serial bus and provide an emergency secondary disengage for the A.P.E.S. system in the case of a necessary emergency shutdown. During flight the avionics will log all data to a SD card. Solid state memory should allow recovery of flight data if a recoverable failure occurs. Post-recovery, the flight computer must switch to location and communication systems to transmit a PS signal through the telemetry system to the ground station. Figure 20 and Table 40 provide the flight computer layout and major components listing respectively. eorgia Institute of Technology 127 of 196

128 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 51: Custom flight computer layout. eorgia Institute of Technology 128 of 196

129 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 40: Major Flight Computer Components Part Number Component Picture Description 1 The flight computer microprocessor, the ATmega The PS receiver, the Fastrax UP501 PS module 3 The Xbee PRO 900-XSC module for communication between launch vehicle and ground station 4 The OpenLog board will provide logging capability round Station Amateur rocketry is a test bed for novel aerospace designs; however, normal launches provide little feedback beyond basic feasibility. This open loop makes it difficult to refine ideas and identify meaningful or effective designs. While in many cases, acquiring such feedback could be eorgia Institute of Technology 129 of 196

130 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW prohibitively expensive, many performance criteria for vehicles can be acquired through relatively cheap means with some effort. Detailed visual observation of a launch vehicle can provide meaningful insight into launch vehicle stability and other important design considerations. Today even cheap digital cameras can provide levels of detail necessary to give meaningful vehicle feedback. Past missions flown by the have encountered interesting performance anomalies and have fallen victim to speculation due to limited data collection and some provocative still camera images. By visually tracking the launch vehicle, unusual flight and structural characteristics can be positively documented and close the design loop by providing feedback for the next design iteration Purpose The ground station is designed to ensure communication with and visual observation of the launch vehicle. Communication quality will be ensured through the use of a high-gain directional antenna. A digital video camera will be used to observe the launch vehicle throughout its flight. The ground station will also feature a detachable PS unit used to make recovery of the launch vehicle easier Function Both the antenna and camera will be mounted on an alt-azimuthal mount. The mount will have motors enabling automated rotation of the platform in both of its degrees of freedom. The motion of the mount will be controlled by a microcontroller that will also be part of the ground station. In addition to controlling the motors, the controller will also perform the wireless communication that will receive signals from the launch vehicle via the antenna. To effectively accomplish its objectives, the ground station must actively track the launch vehicle throughout its flight. This will be accomplished in one of two ways. The first would use telemetric data received from the launch vehicle to create a model of the vehicle s motion. The eorgia Institute of Technology 130 of 196

131 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW second would use a stereo camera system to create disparity maps of the launch vehicle s motion and translate these into a series of distance measurements. This could then be used to create a similar model of motion. The camera zoom will also be adjusted throughout the flight to account for the changing distance between the base station and the launch vehicle and attempt to maintain a near constant level of detail. Table 41: round station requirements Requirement Design Feature Satisfying Requirement Accurately receive High-gain telemetric from launch direction antenna vehicle Maintain constant visual High optical tracking of launch camera, vehicle motorized mount and control algorithm Provide relative position Detachable PS information of launch module vehicle for recovery Design Considerations Choice of Antenna Requirement Verification Analysis of received signals Review of captured video Successfully locate launch vehicle Success Criteria Sufficient information for modeling motion and retrieving launch vehicle is received Launch vehicle remains in FOV through apogee Successfully locate launch vehicle eorgia Institute of Technology 131 of 196

132 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 52: Diagram of a helical antenna Deciding on the proper type of antenna requires two opposing design characteristics: the directionality and gain of the antenna. Choosing a higher gain antenna will allow for a greater range of operation but would give a smaller beam width. This would increase dependence on the tracking algorithm for ensuring signal quality. A helical antenna offers a good compromise between these two considerations, with typical examples offering a half power beam width of and boresight gains of 8-22 db. This beam width would give some cushion for latency in the tracking algorithm. The gain would also be sufficient to ensure good signal quality even under non-line-of-sight propagation at considerable distance, such as might be the case after landing. Figure 53: Typical radiation pattern for a helical antenna Choice of Camera The choice of video camera posed a similar design decision. Much like an antenna, a camera provides a certain angular window of coverage. For a fixed number of pixels, increasing this window will decrease the detail of the captured images. Unlike an antenna, however, these parameters can be a dynamically changed through the use of zoom. A high optical zoom would eorgia Institute of Technology 132 of 196

133 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW then allow for fairly high detail throughout the flight. Camera choice is further complicated by the need to algorithmically adjust the zoom of the camera during flight. While this functionality is built in to most digital cameras, it is seldom available to users programmatically. Models supporting this functionality often do so at prohibitively high costs. Figure 54: Canon Powershot SX260 The Canon Powershot SX260 seems to satisfy all of these requirements. The camera is capable of recording video at 24FPS with an image size of 1920x1080 pixel. The camera also offers and 20x optical zoom. Assuming a 30 vertical field of view or a 60 horizontal field of view, these parameters mean that at its furthest point, each pixel would correspond to 1.7inches of the launch vehicle. This camera also offers access to a user-supported firmware known as the Canon Hack Development Kit which provides direct access to camera operations not offered by factory firmware. This will considerably simplify gaining direct electronic control of zoom Motor Sizing The ability of the platform to track the launch vehicle is inherently limited by the speed and accuracy at which it can rotate. The rotational speed necessary will be dependent on the angular velocity of the launch vehicle from the station s reference frame. Assuming the launch vehicle s path is completely vertical from its Launchpad, the angular velocity of the launch vehicle is given by: eorgia Institute of Technology 133 of 196

134 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW dθ dt = xy y 2 + x 2 (15) Where x is the distance from the base station to the launch pad and y is the altitude of the launch vehicle. The maximum angular velocity of the launch vehicle will occur during the burn of the motor, which will occur over the first two seconds of flight. At the end of this acceleration the launch vehicle will be travelling at 177m/s. This design will be used at events where participants will likely use at most class M motors. For this size motor NAR requires a minimum personnel distance of 500 feet2f3, or approximately 150 meters. Assuming this distance for x and constant acceleration over the motor burn yields the following equation: dθ dt = t 1 (16) t s This function takes a value of approximately 0.62radians/s at t=1.4seconds. The motor must then be capable of rotating the mount at a minimum of this speed. Once the moment of inertia for the mounted camera and antenna has been decided, this value can be used to find the required torque for the motor Software Maturity The software operation of the ground station can be broken into a number of logical components. The process begins by configuring the unit, which consists of initializing contact with the rocket and initializing the state variables for the rocket and ground station. Once this is done, the station will then enter normal operation. This state consists of a loop which processes incoming telemetry information, updating state information for the rocket, deciding whether to update servo position, and deciding whether to update the zoom of the camera. During this 3 eorgia Institute of Technology 134 of 196

135 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW process, the station will also characterize the state using two Boolean variables, LAUNCHED and LANDED. Once both of these variables become true, the loop will break, and the station will transmit the resting coordinates of the rocket to the PS Pendant. The figures below show this process and the sub-processes involved in each of these steps. Figure 55: High-Level Software Process eorgia Institute of Technology 135 of 196

136 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 56: Updating Rocket State eorgia Institute of Technology 136 of 196

137 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 57: Updating Servo Position eorgia Institute of Technology 137 of 196

138 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 58: Updating Camera Zoom eorgia Institute of Technology 138 of 196

139 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 59:Transmit Rocket Location Effects of Excess RF Radiation on the Recovery Avionics A simple testing procedure was implemented to ensure the safety of using the e-matches in proximity to the transmitter. An Xbee transmitter operating at 100mW, with an omnidirectional antenna was placed next to an e-match at several points of high transmission power along the antenna and in the near field. The transmitter then sent a variety of packets varying in length from a single byte to the entire ASCII alphabet. At no point during transmission did the e-match ignite. This result was expected given the low output power of the transmitter. eorgia Institute of Technology 139 of 196

140 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Power Systems Power Budget The power budget for both the A.P.E.S. computer and the Flight Computer are illustrated in Figure 60. The duty cycle is representative of one (1) flight, or 140 seconds. For the A.P.E.S. computer and components, it is assumed that the hardware is active for only 40 seconds of the entire flight from T-20s to T+20s. eorgia Institute of Technology 140 of 196

141 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW (a) (b) Figure 60. (a) Power budget for the A.P.E.S. computer and the Flight Computer; (b) subtotals of the A.P.E.S. computer and the Flight Computer. eorgia Institute of Technology 141 of 196

142 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Power Supply The avionics system, including computers and sensors, will be powered by a 9V battery. The supply will be attached to the Avionics Computer Board which is designed to have a voltage regulator circuit providing 3.3V and 5V rails. The supply will provide 1200mAh of power. Figure 61: Discharge characteristics of the A123 battery The avionics use a negligible amount of power in sleep mode providing a minimum of 5 hours of wait capability for the launch pad. Upon launch the system is activated and will have greater then needed power capacity to perform its duties until the launch vehicle is retrieved. Separately, the A.P.E.S. system will utilize a four-pack of A123 lithium iron phosphate (LiFePO) rechargeable batteries, one of which is shown in Figure 62. These batteries have a per-unit nominal capacity and voltage of 2.3 ampere-hours and 3.3V, respectively. Furthermore, the A123 batteries provide a maximum discharge rate of 70 amperes. Figure 61 illustrates the discharge characteristics of the A123 at four discharge rates. The ability of the A123 to provide a large current is critical to the A.P.E.S. system, which will rely on pulse-width modulation to change magnetic field intensity via manipulation of a root-mean-square current. eorgia Institute of Technology 142 of 196

143 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 62: A single A123 LiFePO battery eorgia Institute of Technology 143 of 196

144 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 7. 0eneral Safety 7.1. Overview Ensuring the safety of our members during building, testing and implementation of the payload experiment is an ideal condition. Procedures have been created and implemented in all of our build environments to ensure safety requirements are met and exceeded. A key way the Ramblin' Rocketeers ensure team safety is to always work in teams of at least two when using equipment or during construction. This guarantees that should an incident occur with a device the other member could provide immediate assistance or quickly get addition help if required. The Invention Studio where the team does a majority of its work is equipped with safety glasses, fire extinguishers, first aid kits, and expert personnel in the use of each of the machines in the area. All the members of the payload and flight systems teams have been briefed on the proper procedures and proper handling of machines in the labs. Table 42: Risk Identification and Mitigation Steps Step Name Step Definition 1. Hazard Identification The first step is to correctly identify potential hazards that could cause serious injury or death. Hazard identification will be achieved through team safety sessions and brainstorming. 2. Risk and Hazard Assessment Every hazard will undergo extensive analysis to determine how serious the issue is and the best way to approach the issue. 3. Risk Control and Elimination After the hazards are identified and assessed a method is produced to avoid the issue. 4. Reviewing Assessments As new information becomes available the assessments will be reviewed and revised as necessary. eorgia Institute of Technology 144 of 196

145 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW The steps outlined above in Table 42 are being used to develop a set of standard operating procedures for launch vehicle construction, payload construction, ground testing, and on all launch day safety checklists Launch Vehicle Safety Failure modes for the launch vehicle were developed to better ensure success of the entire project. Possible modes, resultant problem, and mitigation procedures are given for each failure mode. These modes will continue to evolve and expand in scope as the project progresses. The mitigation methods will be continuously incorporated into preflight checklists. The mitigation items detailed therein will be incorporated into the preflight checklist. Launch vehicle failure modes and mitigation are listed in Table 43. Table 43. Launch vehicle failure modes and mitigation Potential Failure Effects of Failure Failure Prevention Fins Structural ribs buckle on take off Thrust retention plate Skin zippering Launch vehicle flight path becomes unstable Launch failure, launch vehicle destroyed, possible injury from shrapnel Motor casing falls out Internal components are exposed to flowing air currents, launch vehicle Test fin failure modes at connection to launch vehicle to ensure sufficient strength Wear eye wear protection, test the internal structure to ensure a factor of safety against buckling Test reliability of thrust retention plate Test skin adhesion reliability eorgia Institute of Technology 145 of 196

146 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Potential Failure Effects of Failure Failure Prevention becomes unstable Launch buttons Drogue separation Main shoot Land directly on fins Launch vehicle becomes fixed to launch rail, or buttons shear off Main shoot takes full brunt of launch vehicle inertia, launch vehicle becomes ballistic Launch vehicle becomes ballistic, severe injury, irrecoverable launch vehicle Fins break, and launch vehicle cannot be flow twice without fixing Ensure buttons slide easily in launch rail, ensure rail is of the proper size Do a ground test of drogue separation as well as a flight test Do a ground test of main shoot deployment, as well as a flight test. Test fin failure modes at connection to launch vehicle to ensure sufficient strength Ignition failure Motor failure Launch vehicle does not launch Motor explodes, possibly compromising launch Follow proper procedure when setting up launch vehicle ignition system Install motors properly according to manufacturer eorgia Institute of Technology 146 of 196

147 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Potential Failure Effects of Failure Failure Prevention vehicle and payload instructions Payload Safety blah blah 7.4. Environmental Concerns As already mentioned in Section 7.1, the same methodology to identify and assess risks for vehicle and payload safety will be used to identify hazards for constructing various flight and testing components. A Material Safety Data Sheet (MSDS) is on hand for all materials used in the construction of components, and team members have been briefed on best practices for creating a safe workplace. Table 44 lists possible environmental safety concerns. Table 44: Environmental Hazards, Risks, and Mitigation Hazard Risk Assessment Control & Mitigation Electrocution Serious Injury/death Do not touch wires that are hot and not insulated. Wear rubber gloves when the device is in operation. Handle leads to the power supply with care. Use low voltage settings whenever possible. Electromagnetic Fields Interfere with electronic devices inside the body round test equipment, keep people with electronic components in them away from the coil when the electromagnetic coil is in use. eorgia Institute of Technology 147 of 196

148 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Hazard Risk Assessment Control & Mitigation Epoxy/glue Toxic fumes, skin irritation, eye irritation Work in well ventilated areas to prevent a buildup of fumes. loves face masks, and safety glasses will be worn at all times to prevent irritation. Fire Burns, serious injury and death Keep a fire extinguisher in the lab. If an object becomes too hot or starts to burn, cut power and be prepared to use a fire extinguisher. Soldering Iron Burns, solder splashing into eyes Wear safety glasses to prevent damage to eyes. Do not handle the soldering lead directly only touch handle. Do not directly hold an object being soldered. Drills Serious injury, cuts, punctures, and scrapes Only operate tools under supervision of team mates. Only use tools in the appropriate manner. Wear safety glasses to prevent debris from entering the eyes Dremel Serious injury, cuts, and scrapes Only operate tools under supervision of team mates. Only use tools in the appropriate manner. Wear safety glasses to prevent debris from entering the eyes Hand Saws Cuts, serious injury Only use saws under supervision of team mates. Only use tools in the appropriate manner. Wear safety glasses to prevent debris from entering the eyes. Do not cut in the direction of yourself or others. eorgia Institute of Technology 148 of 196

149 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Hazard Risk Assessment Control & Mitigation Exacto Knives Cuts, serious injury, death Only use knives under supervision of team mates. Only use tools in the appropriate manner. Do not cut in the direction of yourself or others. Hammers Bruises, broken bones, and serious injury Be careful to avoid hitting your hand while using a hammer. Power Supply Electrocution, serious injury and death Only operate power supply under supervision of team mates. Turn of power supply when interacting with circuitry. Batteries Explode Eye irritation, skin irritation, burns Wear safety glasses and gloves. Make sure there are no shorts in the circuit. If a battery gets too hot stop using it an remove any connections to it. Improper Dress during construction Serious injury, broken bones Wear closed toe shoes, clothing that is not baggy, and keep long hair tied back. Exposed construction metal Punctures, scrapes, cuts, or serious injury Put all tools band materials away after use. Neodymium Magnets Pinching, bruising, and snapping through fingers. Do not allow magnets to fly together from a distance, do not play with powerful magnets, keep free magnets away from powered solenoids. RF Interference with the Recovery System Pre-mature firing of the ejection charges potential causing significant damage to the Launch Vehicle, payload, and all RF Testing has verified that, at maximum power output, the on-board XBee transmitter will not unintentionally ignite our e-matches from excess RF radiation. eorgia Institute of Technology 149 of 196

150 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Hazard Risk Assessment Control & Mitigation supporting systems Maximum output power is limited to 100 mw eorgia Institute of Technology 150 of 196

151 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 8. Project Budget 8.1. Funding Overview In order to fund the Competition year, the have sought sponsorships from academic and industry sources. The current sponsors of the Ramblin Rocketeers and their contributions can be found in Table 45. As of CDR, the Ramblin Rocketeers have received $5,700 in funding. Furthermore, the Team has also received a dedicated room in which the Team can construct and store their rocket and non-explosive components. All explosive components (i.e. black power) are properly stored in Fire Lockers in either the Ben T. Zinn Combustion Laboratory or the Center for Space Systems Flight Hardware Laboratory. Table 45. Summary of sponsors for the Ramblin. Rocketeers Sponsor Contribution Date Unused Funds from $1,000 Aug 2012 eorgia Space rant Consortium $2,500 Sept 2012 eorgia Space rant Consortium $500 Sept 2012 eorgia Space rant Consortium $1,000 Dec 2012 eneration Orbit $300 Dec 2012 eorgia Tech $1,000 (est) Feb 2013 School of Aerospace Engineering eorgia Tech $1,000 (est) Feb 2013 Student overnment Association ATK Travel Stipend $400 (est) Apr 2011 ATK Motor Stipend $200 (est) Apr 2011 Total $7,900 The team is currently pursuing the following sponsors: Virgin alactic, eorgia Tech College of Engineering, eorgia Tech SA, as well as private donations. eorgia Institute of Technology 151 of 196

152 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 8.2. Current Sponsors Table 46 lists the current sponsors of the and their contributions. Table 46. List of current sponsors of the Ramblin' Rocketeers. Sponsor eorgia Space rant Consortium Advanced Circuits eneration Orbit Contribution Financial contribution for general project expenses Financial contribution for Outreach-specific expenses Financial contribution for REFP-related activities. Manufacturing of the SIDES boards throughout the design process Financial contributions for general project expenses Actual Project Cost CDR Budget Summary Table 47 illustrates the budget breakdown as of the CDR Milestone. The summary is broken down into four (4) main categories: Launch Vehicle, Flight Systems, Operations, and Motors. The Launch Vehicle and Flight Systems categories are further broken down into two (2) subcategories: Flight Hardware and Testing. Operational expenses are broken down into four (4) sub-categories: Safety, eneric Supplies, Tooling, and Physical Capital. Lastly, while motors are specific to the Launch Vehicle subsystem, they are critical component to the architecture and as such are tracked separately from the Launch Vehicle subsystem. eorgia Institute of Technology 152 of 196

153 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 47. CDR Project Budget Summary. Category Amt Spent Amt Remaining Launch Vehicle $ $ Motors $ $ Flight Systems $ $ Operations $ $ 1, Testing/Dev $ $ System-Level Budget Summary Figure 63 illustrates the system-level expenditure summary for Project LSIM at the CDR milestone. Cost reduction techniques, such as proper resource utilization has resulted in lower Flight Systems costs. It is important to note that both the Launch Vehicle and Flight Systems include both Flight Hardware costs in addition to Test/Development costs. Additionally, Figure 64 illustrates the breakdown System Expenditure Breakdown Launch Vehicle $ Flight Systems $ Operations $ Motors $ Testing/Development $ Outreach $ Total $ Figure 63. System expenditure summary at CDR. eorgia Institute of Technology 153 of 196

154 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Sub-system Breakdown Testing/Development Aerodynamics $ Structures $ Recovery $ 0.00 Avionics $ Payload/round Testing $ round Station $ 0.00 Total $ Figure 64. Sub-system Testing/Development Breakdown Flight Hardware Expenditures Flight Hardware Expenditure Overview Figure 65 summarizes the overall expenditures for all Flight Hardware purchased up to the CDR milestone. In order to account for uncertainties in motor price, $300 has been allotted for the purchase of the flight motor. As illustrated by Figure 65, only hardware for the Aerodynamics and Mechanical Integration sub-systems has been purchased. eorgia Institute of Technology 154 of 196

155 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Sub-system Flight Hardware Breakdown Aerodynamics $ Structures $ 0.00 Recovery $ 0.00 Motors/Motor Hardware Mechanical Integration Electrical Integration Flight Avionics $ $ $ 0.00 $ 0.00 Flight Payload $ 0.00 Total $ Figure 65. Sub-System Flight Hardware Breakdown Flight Hardware Cost Breakdown Figure 66 lists the flight hardware breakout for Flight Systems. It is important to note that the materials purchased for the Launch Vehicle flight hardware has not been used to fabricate any parts, therefore no breakout is available at this time for the launch vehicle. eorgia Institute of Technology 155 of 196

156 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Flight Experiment Item Description Unit Price Qty Cost L.S.I.M. Hardware $ $ Propellant Simulant (Water) $ $ Diameter PVC Pipe $ $ Balance Solenoids $ $ Camera Assembly $ $ Piezo Vibration Sensor $ $ 5.90 Base Plate $ $ 1.00 Payload Bottom $ $48.00 Total Flight Experiment Costs $ Flight Avionics Item Description Unit Price Qty Cost SIDES Network $ $ SIDES Board $ $ Electrical Harness $ $ ClockDrive Board $ $ LSIM Board $ $ Telemetry Board $ $ MasterIMU $ $ Strain age Board $ $ SIDES Node Battery $ $ LSIM Battery $ $ Total Flight Avionics Cost $ Figure 66. Flight Systems flight hardware breakout. eorgia Institute of Technology 156 of 196

157 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 9. Project Schedule 9.1. Schedule Overview The Mile High Yellow Jacket s project is driven by the design milestone s set forth by the USLI Program Office. The design milestones are listed in Table 48. The project antt Chart for Project A.P.E.S. located in Appendix I contains only high-level activities due to the unique launch vehicle and payload designs. A more detailed Critical Path chart is located in Section 9.2. Table 48. Design milestones set by the USLI Program Office. Milestone Proposal Team Selection Web Presence Established PDR Documentation PDR VTC CDR Documentation CDR VTC FRR Documentation FRR VTC Rocket Week PLAR Documentation Date 26 SEP 17 OCT 4 NOV 28 NOV 6 DEC 23 JAN 2 FEB 26 MAR 2-11 APR APR 7 MAY eorgia Institute of Technology 157 of 196

158 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 9.2. Critical Path Chart: CDR to PLAR The critical path chart illustrated by Figure 6 demonstrates the highly integrated nature of Project A.P.E.S. The critical path chart identifies: High Risk Tasks Low-Moderate Risk Tasks Earned Value Management (EVM) oal Tasks Looping Tasks Critical and Alternate Paths Major Inputs to Tasks eorgia Institute of Technology 158 of 196

159 eorgia Tech Critical Design Review Figure 67. Critical Path Chart from CDR to PLAR

160 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 9.3. Schedule Risk High Risk Items Two (2) items have been identified as High Risk Items. These are: Launch Vehicle Structure Design Recovery System Design Table 49 lists the mitigations for these items. Table 49. Identification and Mitigations for High-Risk Tasks. High-Risk Task Potential Impact on Project L.S.I.M. Mitigation 1) Ensure personnel have direct and free access to experienced personnel on and off of the team. Launch Vehicle Design, Fabrication, & Testing 1) Schedule Impact 2) Budgetary Impact 3) Not qualifying for Competition Launch 2) Ensure personnel have knowledge on to effectively utilize simulation and analysis tools. 3) Ensure personnel have direct and free access to the simulation and analysis tools. Recovery System Design, Fabrication, & Testing 1) Excessive kinetic energy during landing resulting in damage to the rocket. 2) Failure to deploy the drogue and/or main parachute resulting in a high energy impact with the ground destroying the Launch Vehicle. 4) Ensure personnel are familiar with relevant fabrication techniques. 1) Ensure Recovery System Lead has direct and free access to experienced personnel on and off the team. 2) Provide real-time feedback of the design decisions to ensure all recovery-related requirements are meet with at least a 5% margin wherever possible. 3) Ensure proper manufacturing techniques are utilized during the fabrication of the recovery system. eorgia Institute of Technology 160 of 196

161 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW High-Risk Task Potential Impact on Verification of Field Equations & Control Logic Project L.S.I.M. 1) Unsuccessful flight demonstration 2) Flight Experiment does not function properly during flight 3) Flight Experiment encounters a flight anomaly that results in excessive draw and damage to the Flight Avionics, Power Supply, and/or Launch Vehicle Mitigation 1) Develop multiple paths to achieve the end goal of developing thee robust control logic that is required for the successful demonstration of the Flight Experiment. 2) Ensure Flight Systems personnel have direct and free access to experienced personnel on and off of the team. 4) Ensure personnel have direct and free access to the simulation and analysis tools necessary for the development (and subsequent verification) of the control logic Low-to-Moderate Risk Tasks The low-to-moderate risk tasks are considered to be those risks that pose a risk to either the project schedule and/or project budget but little to no risk of not meeting the Mission Success Criteria in Table 5. The risks and mitigations are provided in Table 50. eorgia Institute of Technology 161 of 196

162 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Table 50. Low to Moderate Risk items and mitigiations. Risk Risk Level Potential Impact on Fabrication of Launch Vehicle Sections Full-Scale Launch Vehicle Test Flight Flight Computer Fabrication round Testing & Control Logic Development Moderate Moderate Low Moderate Project A.P.E.S. 1) Schedule Impact 2) Budgetary Impact 3) Not qualifying for Competition Launch 1) Schedule Impact 2) Budgetary Impact 3) Not qualifying for Competition Launch 1) Budgetary Impact 2) Not able to collect inflight data 1) Schedule Impact 2) No Experimental Flight Data is recorded prior to the Competition Launch. Mitigation 1) Ensure Manufacturing and Fabrication Orders (MFO s) are sufficiently detailed for the task prior to starting any fabrication. 2) Ensure proper manufacturing techniques are observed during fabrication. 1) Ensure Launch Procedures are established practiced prior to any launch opportunity. 2) Have a sufficient number of launch opportunities that are in different geographical areas as to minimize the effects of weather on the number of launch opportunities. 1) Ensure proper manufacturing techniques are observed during fabrication. 2) Ensure Manufacturing and Fabrication Orders (MFO s) are sufficiently detailed for the task. 1) Ensure personnel have direct and free access to experienced personnel on and off of the team. eorgia Institute of Technology 162 of 196

163 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 10. Educational Engagement Plan and Status Overview The goal of eorgia Tech s outreach program is to promote interest in the Science, Technology, Engineering, and Mathematics (STEM) fields. The intend to conduct various outreach programs targeting middle school students and educators. The Ramblin Rocketeers will also have an outreach request form on their webpage for educators to request presentations or hands-on activities for their classroom Atlanta Makers Faire had a booth at the Atlanta Makers Fair, a fair in which various craftsman from the community and eorgia Tech assemble to show off their accomplishments. The intent of this program is to give clubs, organizations, and other hobbyists the opportunity to show others their unique creations and skills. The event is open to the entire Atlanta community and had a large attendance this year. The booth had a display of our various rockets, as well as a station for children to make their own paper rockets. Our booth had middle school aged children attend and participate in the paper-rocket activity. Figure 68. Participation at the Atlanta Makers' Faire Civil Air Patrol In the Aerospace Education program, Cadets have the opportunity to earn a Model Rocketry Badge by furthering their knowledge in the history and physics of rocketry as well as building five separate rockets ranging from non-solid fuel rockets to scale models of historic rockets. These rockets must meet specific altitude and payload requirements. The will be working again this year with a local Atlanta-based squadron to help approximately 20 children earn their own badge this spring. eorgia Institute of Technology 163 of 196

164 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW FIRST Lego League and Tech Challenge FIRST is a series of international robotics competitions for students from 3 rd -12 th grades. FIRST Lego League is an engineering competition designed for middle school children in which they build and compete with an autonomous MINDSTORMS robot. Every year there is a new competition centered on a theme exploring a real-world problem. FIRST Tech Challenge is a robotics competition designed for Figure 69: Previous FIRST Lego League outreach event. students in middle and high school where the robots can be 18x18x18 inches at the start of each match. This year the have had an educational booth at a FIRST Lego League Regional Competition at Wheeler High School which occurred on Saturday, December 8 th.at the booth students ranging from 3 rd -8 th grade were exposed to how lift is generated and participated in building a paper rocket with a straw launcher that they could take with them. The event reached 373 students, 295 of which were in the 4 th -9 th grade range, and 31 educators. Below in Figure 70 and Figure 71 are pictures from this event. eorgia Institute of Technology 164 of 196

165 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 70: FLL Regional Event at Wheel High School Figure 71: FLL Regional Straw Rocket Activity In addition to the FIRST Lego League Regional at Wheeler High School, the Ramblin Rocketeers are scheduled to have a booth at both the FIRST Tech Challenge Regional at Wheeler Middle School on Saturday, January 19 th and the FLL State Tournament at eorgia Tech on Saturday, January 26 th Atlanta Middle School Outreach eorgia Institute of Technology 165 of 196

166 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW The also plan to go to various middle schools in the Atlanta area to make presentations and demonstrations about science and physics. The intent of these programs is to teach students in the 6 th to 8 th grade range about science and rockets, as well as to spark their interest in STEM fields. Specific topics that the plan to cover in the multiple demonstrations are electricity, basic concepts of flight, and general topics relating to engineering. The will be working in conjunction with organizations like the Society of Women Engineers to reach even more middle school students, exposing them to various topics regarding space exploration and inspiring them to pursue careers in the STEM fields. Drew Charter Middle School has been contacted, and a STEM demonstration will occur there within the coming months. eorgia Institute of Technology 166 of 196

167 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW References Simon, T. M., Reitich, F., Jolly, M. R., Ito, K., & Banks, H. T. (1998). Estimation of the E ective Permeability in Magnetorheological Fluids. CRSC Technical Report CRSC-TR98-35, NC State Univ. The Dynamic Behavior of Liquids in Moving Containers: with applications to space vehicle technology. All articles. Ed. H. Norman Abramson. NASA, Washington, D.C., Niskanen, Sampo. OpenRocket vehicle Technical Documentation. 18 July Web. Apke, Ted. "Black Powder Usage." (2009). Print. < PerfecFlite. StratoLogger SL100 Users Manual. Andover, NH: Print. < Roensch, S. (2010). "Finite Element Analysis: Introduction." 2011, from Fiberglass Epoxy Laminate Sheet. MATWEB.com. search/datasheet_print.aspx?matguid=8337b2d050d44da1b8a9a5e61b0d5f85 "Shape Effects on Drag." NASA Web. 19 Nov < 12/airplane/shaped.html>. Cavcar, Mustafa. "Compressibility Effects on Airfoil Aerodynamics." (2005). Print. "Apogee Paramagnetic Oxygen as Experimental Electromagnetic Separator: Preliminary Design Review." Comp. eorgia Tech University Student Launch Initiative. Atlanta: Print. Cheng, David. Field and Wave Electromagnetics. 1st ed. Reading, MA: Addison-Wesley Publishing Company, Print. Millspaugh, Ben, Ph.D. Civil Air Patrol: Model Rocketry. Leadership Development and Membership Services Directorate, National Headquarter, Civil Air Patrol, Maxwell AFB, AL. Print. eorgia Institute of Technology 167 of 196

168 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Appendix I: antt Chart eorgia Institute of Technology 168 of 196

169 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW eorgia Institute of Technology 169 of 196

170 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Appendix II: Launch Checklist Pre-Launch Packing The night before launch go through Launch Vehicle Packing List and put all items in a designated spot. The morning of launch go through Launch Vehicle Packing List and ensure all items are still there. Load the vehicle(s) Performer Inspector Launch Avionics On Prepare Payload Bay Ensure batteries and switches are wired to the altimeters correctly. Ensure batteries, power supply, switch, data recorder and pressure sensors are wired correctly. Install fresh batteries into battery holders and secure with tape. Test the altimeters. Altimeter In Circuit Out of Circuit Altimeter 1 Altimeter 2 Insert altimeter and payload into the payload bay. Connect appropriate wires. Verify payload powers on correctly and is working properly. If it is not, check all wires and connections. Turn off payload power. Arm altimeters with output shorted to verify jumper settings. This is to check battery voltage and continuity. Disarm altimeter, un-short outputs. eorgia Institute of Technology 170 of 196

171 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Assemble Charges Test e-match resistance and make sure it is within spec. Remove protective cover from e-matches. Measure amount of black powder determined in testing. Put e-matches on tape with sticky side up. E-match Resistance E-match 1 E-match 2 E-match 3 E-match 4 Pour black powder over e-matches. Seal tape. Re-test e-matches. Check Altimeters Ensure altimeter is disarmed. Connect charges to altimeter bay. Turn on altimeter and verify continuity. Disarm altimeters. Altimeter 1 Altimeter 2 OFF ON Pack Parachutes eorgia Institute of Technology 171 of 196

172 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Connect drogue shock cord (long side) to booster section and altimeter bay (short side) Fold excess shock cord so it does not tangle. Add Nomex cloth to ensure only the Kevlar shock chord is exposed to ejection charge. Insert altimeter bay into drogue section and secure with shear pins. Pack main chute. Attach main shock cord to payload bay (long side to nose cone). Fold excess shock cord so it does not tangle. Add Nomex cloth under main chute and shock cord ensuring that only the Kevlar part of the shock cord will be exposed to the ejection charge Assemble Motor Follow manufacturer's instructions. Put on safety glasses and gloves. Do not get grease on propellant or delay. Do not install igniter until at pad. Install gasket on top of motor. Install motor in launch vehicle. Secure positive motor retention. Final Prep Turn on payload via a switch and start stopwatches. Install skin. Inspect launch vehicle. Check C to make sure it is in safe range; add nose weight if necessary. Bring launch vehicle to the range safety officer (RSO) table for inspection. Bring launch vehicle to pad, install on pad, verify that it can move freely (use a standoff if necessary). Install igniter in launch vehicle. Touch igniter clips together to make sure they will not fire igniter when connected. Make sure clips are not shorted to each other or blast deflector. Arm altimeters via switches and wait for continuity check for both. Connect shock cord to nose cone, install nose cone and secure with shear pins. eorgia Institute of Technology 172 of 196

173 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Return to front line. Launch Stop the stopwatches and record time from arming payload and launch. Watch flight so launch vehicle does not get lost. Post Launch Recovery Recover launch vehicle, document landing. Disarm altimeter(s) if there are unfired charges. Disassemble launch vehicle, clean motor case, other parts, inspect for damage. Record altimeter data. Download payload data. Trouble Shooting Test Problem Control & Mitigation Power on payload Check E-match resistance Power on altimeters Check for altimeter continuity after installing e-matches Launch Rocket Payload does not power on E-match resistance does not match required specifications Altimeters do not power on No continuity Engine does not fire Check batteries have sufficient charge, check wires are connected correctly Replace e-match before use Check batteries have sufficient charge, check wires are connected correctly Check wires are connected correctly Disconnect power, ensure igniter clips are not touching, ensure power is reaching clips,ensure motor is assembled correctly eorgia Institute of Technology 173 of 196

174 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Appendix III: Science Overview Ferromagnetism and MR fluid response Scientific Background and Mathematical Modeling To the end of accomplishing the goals of the LSIM experiment, some theoretical research and work must be accomplished in tandem with experimentation. A passive or active control system is to be developed in order to move the simulated propellant to its desired location with the magnetorheological (MR) fluid. To model the behavior of the simulant-mr fluid system, equations are being researched, modified, and developed in order to calculate magnetic fields and forces, to govern the properties of MR fluids, and to model system dynamics. In addition to equations, qualitative research has been done in the literature concerning MR fluids to suggest approaches that may be taken during experimental testing. Magnetic fields The forces on the MR fluid that will be transmitted to the simulant will depend largely on the magnetic fields that are applied to the fluid. Control of currents in a solenoid will allow for precise control of the fields. Last year, it was derived and also confirmed in the literature that the exact magnetic H field from a current loop in spherical coordinates, with the loop centered at the origin in the xy -plane and counterclockwise current, is as below (θ denotes azimuth angle): H r = CR2 cos θ α 2 E(k 2 ) β C H θ = 2α 2 β sin θ [(r2 + R 2 cos 2θ)E(k 2 ) α 2 K(k 2 )] where K and E are complete elliptic integrals of the first and second kinds, respectively, and α 2 = R 2 + r 2 2Rr sin θ, β 2 = R 2 + r 2 + 2Rr sin θ, k 2 = 1 α 2 β 2, and C = I π. I is the loop current, R is its radius, and r is the distance from the origin to the point of measurement. A solenoid simply consists of several such current loops, with the fields adding vectorally. While the above expressions are extremely nonlinear and difficult to analyze or work with, they may be simplified as needed, or modeled using a computer. eorgia Institute of Technology 174 of 196

175 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Magnetic forces After calculating the magnetic fields, in order to predict the motion of the MR fluid and simulant in the container, the forces on the MR fluid due to the field must to be calculated. In any material, the movement of atomic charges such as electrons causes the atoms to behave as microscopic magnetic dipoles, experiencing forces in magnetic fields. The magnetization vector M at a point in the material is defined as the volume density of magnetic dipole moment, i.e. m k M = lim v 0 v Each m k is the magnetic moment of the k th atom in volume v, and the sum is over all atoms. M depends on the magnetic field H at a point, and flux density B depends on the field, as follows: M = χ m H B = μ 0 (H + M) = μ 0 H(1 + χ m ) = μ 0 μ r H = μh where χ m is the material s magnetic susceptibility, μ r is its relative permeability, and μ is the absolute permeability. It is assumed that χ m, and hence μ and μ r, are approximately constant for the MR fluid. This is a very valid assumption that greatly simplifies analysis, given that the fields are not extremely large, as is evidenced in Figure 72 below taken from a paper by Simon et al. Figure 72: Plot of B field magnitude in MR fluid versus magnitude of vector μ 0 H, for iron volume concentrations of 10, 20, and 30 percent eorgia Institute of Technology 175 of 196

176 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW The force on a magnetic material can be determined by summing the forces on the dipoles in the material due to the field that it is placed in. The force on a magnetic dipole m in field B is F = (m B) Let V be the volume of a very small region of the MR fluid in which M is approximately constant. Then, letting m = MV = χ m VH = χ mv B, the force on the region is µ F = χ mv µ B B = 2χ mv B (B) µ Using equations (1), (2), and (5) for the H and B fields of a current loop, it can be seen that the force on each small region, and hence on the whole fluid, should be directly proportional to the square of the current. In addition, B (B) may be calculated using equations (1) and (2). These equations will be further developed to better understand response of the MR fluid and simulant. MR fluid rheological properties In addition to translational movement, which is governed by the preceding equations, MR fluids experience large increases in yield strength in the presence of magnetic fields. This is desirable for the LSIM system, as otherwise the sloshing propellant simulant would simply shear through the MR fluid barriers with little resistance. It is desired to characterize the rheological properties of MR fluid to understand how much resistance to movement the simulant will experience. More precisely, MR can be modeled fairly closely as a Bingham plastic, a common example of which is toothpaste. A Bingham plastic does not start flowing until a certain point of yield shear stress, after which it behaves similarly to a viscous liquid. The equation governing the shear stress of an ideal Bingham plastic, and so to model the MR fluid for future analysis, is τ = τ yield (H) + η dv dn for τ > τ yield(h) τ yield (H) is the yield shear stress of the MR fluid, and is larger for stronger H fields. η is the flow viscosity after shear, and dv is the velocity gradient in the direction normal to the plane of dn shear. This relation is shown on the next page in Figure 73, compared to a Newtonian fluid. eorgia Institute of Technology 176 of 196

177 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 73: Shear stress of ideal Bingham plastic (and MR fluid model) versus shear rate dv, compared to ideal dn Newtonian liquid Hence, if the simulant exerts such a force that MR fluid flow begins occurring, the shear stress between layers of the MR fluid should increase, keeping the simulant comparatively restrained until it settles again. If the need arises to decrease the yield shear stress for a given magnetic field, such as to make the MR fluid flow more easily, replacing a percentage of microscale ferroparticles with nanoscale particles can decrease the yield stress. Further research is still required to find the relationship between the yield strength and magnetic field, which will allow control of the yield stress acting against the simulant. However, the key observation is that there is little to no MR fluid flow below some certain shear stress, for a given magnetic field H. System Dynamics While research on the physical properties and behavior of MR fluids is ongoing, basic system dynamical modeling has already been started with variable parameters that will be determined from theory and experimentation in the future. The fluid and MR fluid mixture is assumed to operate roughly as a system with a spring, damper, and mass, where the driving force is the solenoid. The fluid is considered the mass, whose motion is restrained by a spring and damper, and driven by the MR fluid actuated by the solenoid. All system elements lie on the same x - axis, with the solenoid axis coinciding. The dynamical equation of motion in this case is eorgia Institute of Technology 177 of 196

178 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW mx = F solenoid kx bx Where m is the mass of the fluid, k and b are unknown damping placeholder constants, and x is the position of the simulant relative to some point. After some manipulation, the dynamical equation for the response of the fluid becomes: eorgia Institute of Technology 178 of 196

179 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Understanding Slosh Damping Fluid dynamics and hydrodynamic regimes of expected slosh In considering the liquid slosh, the flight regime of the vehicle is extremely important. While the experiment aims to approximate a spacecraft by manipulating MR fluid during microgravity to dampen water slosh, the realities of atmospheric flight will limit the applicability of launch vehicle test results. The extent of these flight regime limitations is revealed by three key similarity parameters: the Weber (We) number, the Froude (Fr) number, and the Bond (Bo) number. These three parameters measure the ratio of inertial to capillary forces, the effect of gravitational body forces relative to inertial forces, and the relative magnitudes of gravitational and capillary forces respectively. Finally, an understanding of the potential flow of sloshing fluid is necessary to understand the motion of fluid inside a vehicle. Flight regime However, an estimate of the flight regime of the launch vehicle near apogee must first be known. To better understand this flight regime and to confirm the microgravity requirements pulled from previous team documents, a first-order analysis of the launch vehicle s flight was computed. Neglecting drag and assuming 2-D projectile motion with instantaneous acceleration from a rocket motor, the flight profile of the launch vehicle was estimated and the characteristics of the 0.1-ee requirement from the 2009 eorgia Tech team 0.1-ee being the definition of the microgravity threshold for the purposes of the experiment were examined. The results of this simplified analysis are presented graphically in Figure 74. eorgia Institute of Technology 179 of 196

180 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 74: Microgravity time as a function of launch angle from horizon In Figure 74, the microgravity time, or t micro, was computed using equation (17). t micro = V 0sin (α) (17) g In equation (17), the 1.5 s addition represents the time from apogee to chute deployment, which by representation in Figure 74 is always less than the other half of the equation for launch angles between 60 and 90 degrees. The drogue chute deployment therefore represents the bounding time for the experiment operation in the mission profile. From the flight profile, a velocity corresponding to 0.1-ee and a maximum height can be calculated. These variables can be used for the computation of similarity parameters, as well as comparison numbers to judge the validity of the flight profile and microgravity estimates. Two comparison measures will now be observed. Among the simplest environments for creating microgravity is the free-fall drop test. This test provides microgravity times approximated by equation (18) valid to heights of 20 m with atmospheric drag. Equation (18) is nonspecific with regards to the accelerations achieved, however these are estimated by Reynolds and Satterlee (p. 435, Dynamic Behavior) to be between 10 7 and 0.2. eorgia Institute of Technology 180 of 196

181 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW t micro = h (18) The predicted times from equation (18) and from the flight profile are given in Figure 74. Table 51: Microgravity times for fall heights Height (m) Microgravity time (s) (free-fall) (free-fall, target altitude) (90 launch angle) ~7.13 Adjusting for the 1.5 second chute deployment, the 90 launch microgravity time appears to be about one half the time given for free-fall from the same maximum height given that the max ee loading specified in the reference is 0.2, or twice the requirement, this difference appears to be acceptable for a bounding and ideal case. Of course, accelerations due to aerodynamic forces will requirement additional modeling and adjustment. Similarity parameters Table 52 presents the similarity parameters relevant to the LSIM experiment calculated for the propellant simulant, water (30 C). The Weber, Bond, and Froude numbers are considered here. These numbers provide an indication of the hydrodynamic regime these regimes eorgia Institute of Technology 181 of 196

182 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW 4 Figure 75: Slosh regimes and similarity parameters3f for microgravity are illustrated in Figure 75. The Reynold s number is also included for comparison; although for this experiment the number itself is not as significant as long as the regime described by different test configurations is similar, i.e. all turbulent, all laminar, etc. A potential source of error in these computations is use of the launch vehicle velocity rather than the relative velocity of the fluid in the tank. The Weber, Froude, and Reynold s numbers are affected by this choice, which is yet to be validated. Table 52: Similarity parameters for simplified flight profile of the launch vehicle Number Equation Value Bo ρgl 2 /σ 980 We ρu 2 L/σ 1.37x10 7 Fr We/Bo 1.4x10 4 Re ρul/μ 2.023x10 7 These parameters will allow verification and comparison of ground tests with the launch vehicle test and REFP, vis-à-vis actual spacecraft and launch vehicles. 4 Reynolds, William C. and Hugh M. Satterlee. Liquid Propellant Behavior at Low and Zero. The Dynamic Behavior of Liquids p Ref Appendix XXX eorgia Institute of Technology 182 of 196

183 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Modeling Slosh In order to approximately predict the behavior of the propellant in the sloshing modes expected at apogee of our launch vehicle, a mathematical model has been developed. The aim of this model is solely to predict sloshing behavior in the absence of dampening; this is an intermediate step to modeling the effects of the MR fluid dampening. Predicting the exact distribution and dynamics of all sloshing fluid within the container, however, would require a large amount of complexity, which would only be exacerbated by attempting to add the effects of the MR fluid later on. Therefore, for analysis to be feasible, a much simpler model is proposed. Figure 76. Schematic and free-body diagram of slosh dynamic model The fluid is modeled schematically as shown above in Figure 76. It is assumed that the center of mass of the fluid in the tank behaves roughly as an object of mass m attached to a pendulumspring of spring constant k, with additional dampening effects represented by viscous dampers of constants b 1 and b 2. Also, let L be the original length of the pendulum-spring with no forces applied, and L the the amount the spring is stretched from length L (so that negative L implies compression). It is anticipated that the majority of sloshing will be longitudinal, so significant vertical motion can be expected of the fluid. Therefore, in our model, the spring may experience appreciable compression. However, a much smaller amount of lateral sloshing is predicted, so in the model, the angle θ may be assumed to be small. Therefore, throughout this analysis, it will be assumed that sin θ θ sufficiently closely for the corresponding substitution to be justified. eorgia Institute of Technology 183 of 196

184 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW The equations of motion can be written in the x and y directions, respectively, as follows: k L sin θ b 1 x k Lθ b 1 x = mx (1) k L cos θ b 2 y mg k L b 2 y mg = my (2) The equations to be developed will depend on state variables θ, θ, L, and L. Let the origin of the coordinate system be at the center of rotation of the pendulum. Then, it is known that x = (L + L) sin θ (L + L)θ (3) y = (L + L) cos θ (L + L) (4) x = L θ + (L + L)θ (5) y = L (6) x = L θ + 2 L θ + (L + L)θ (7) y = L (8) where equations 5 through 8 are found by repeatedly differentiating equations 3 and 4. First of all, substituting equations 6 and 8 into equation 2 and rearranging, it is readily found that L = g b 2 m ΔL k m L (9) Note that this equation is independent of θ, and is the same as a one-dimensional spring-massdamper system. The analysis and results from substituting equations 5 and 7 into equation 1 are significantly more complicated. First of all, carrying out this substitution and rearranging, θ = k L b 1 L m(l + L) Next, substituting equation 9 for L in equation 10, it is found that θ = (b 2 b 1 ) L m(l + L) L (L + L) θ b 1 m + 2 L (L + L) θ (10) g (L + L) θ b 1 m + 2 L (L + L) θ (11) Therefore, as k, b 1, b 2, m, g, and L are constants, it is seen that θ is a function of the state variables. Write θ = f θ, θ, ΔL, ΔL, where f: R 4 R is differentiable at any point eorgia Institute of Technology 184 of 196

185 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW θ, θ, ΔL, ΔL such that L -L, due to continuity of the partial derivatives of f at those points. The relation L = -L only occurs for the spring being compressed into a flat piece, which corresponds to all fluid molecules touching the top surface of the tank; both of these are impossible scenarios. Therefore, it is possible to linearly approximate f close to any point of interest, for easier analysis. Initially, the relation is linearized for points near the equilibrium point of the system, which is (0, 0, mg/k, 0). In this case, defining δl = ΔL mg/k, θ = f θ, θ, ΔL, ΔL θ = 0 + f θ eq f 0,0, mg mg, 0 + f 0,0, k k, 0 θ, θ, δl, ΔL (12) θ + f θ + f θ eq ΔL δl + f ΔL (13) eq ΔL eq Finally, evaluating the partial derivatives, it is found that close to (0, 0, mg/k, 0), gk θ = (kl + mg) θ b 1 m θ (14) Therefore, relations for both L and θ have been found only in terms of the four state variables θ, θ, ΔL, and ΔL, assuming that the values of θ, θ, δl, and ΔL are small. Using the above equations, whether the linear approximations (9) and (14) or the more precise but complicated form (11), further analysis by hand or by computer should yield information as to how the system should approximately behave in the absence of the MR fluid baffles. Further development should also approximate system dynamics in the presence of the baffles, allowing rough predictions to be made as to how the MR fluid baffles may impact fluid slosh. Unifying the LSIM theories Finally, it is necessary to connect the response of MR fluid to a magnetic field with the damping of slosh. According to Dynamic Behavior the damping of slosh for a ring baffle is dependent on the baffle cross-section4f5. For containers of constant geometry and liquid at constant rest height in gravity dominated slosh, empirical relationships have been illustrated between the geometry of 5 Abramson, H. Norman and Sandor Silverman. Damping of Liquid Motions and Lateral Sloshing. The Dynamic Behavior of Liquids p Ref Appendix XXX. eorgia Institute of Technology 185 of 196

186 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW a rigid baffle and the damping ratio vs. wave amplitude of slosh5f6. However, an effect not present in the dynamic analysis above is the container geometry. In the case of LSIM where the liquid height will be greater than the container radius, the damping coefficient in lateral slosh is given by: δ = 4.98ν 1/2 R 3/4 g 1/4 (1) Where ν is the kinematic viscosity, R the container radius, g the acceleration of gravity. This equation6f7, along with curve fitting with the help of tables and plots given in Dynamic Behavior, provides a means to experimentally determine the damping coefficients needed with given MR fluid response to calculate the slosh dynamics. 6 P P. 110 eorgia Institute of Technology 186 of 196

187 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Appendix IV: round Test Plan oals The LSIM ground test data will provide the basis for empirical modeling of magnetorheological fluid as a damper for liquid sloshing. All actions will be incremented to allow for a detailed model for extrapolation and interpolation of the data for future flight control systems. round Test oal round Test oal Definition 1 Create MR Fluid 2 Calibrate Sensors 3 Determine force of MR Fluid 4 Develop model for solenoid control 5 1- slosh dampening Test Sequence 1 - Creating MR Fluid MR fluid will be created using different compositions of iron powder, mineral oil, and surfactant. The iron powder will make up about 74-76% of the mixture's mass. Mineral oil will make up 20-22% of the total mass, and the surfactant will make up the remaining 1-4%. Water is then added to test the time to mixture separation and solenoids. Each mixture will be preliminarily tested by neodymium magnets. The mixture will qualify as a successful batch if MR fluid under the influence of an applied magnetic field prevents the leakage of water. Test Sequence 2 - Calibrating Sensors The team will be using two sensors for characterizing MR fluid and recording data: a load sensor and a CMOS camera that can detect light in the IR spectrum. The team will attach a known mass to the load sensor and measure the reading. For the CMOS camera, the team will emit IR light onto water and note the brightness displayed by the camera. The relative brightness will indicate a fluid's IR reflectance. eorgia Institute of Technology 187 of 196

188 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Test Sequence 3- Characterize the shear stress of MR fluid In characterizing MR fluid, the team will utilize a two-plate test for measuring the MR fluid's force and viscosity with and without a magnetic field acting upon the MR fluid. This test was chosen because of its simplicity; other tests such as a barometer test were considered for measuring the MR fluid's viscosity and force, but they turned out too complicated to realize. The two-plate test consists of two plates: a bottom plate, which is fixed to the ground and a top plate, which is free to move. A load sensor will be placed on the top plate to measure the reaction force that is generated. The plates used must not be strongly magnetic; thus, the two current choices are wood or aluminum. A control test will be performed by just having two plates together with a load sensor on the top, moving plate to calculate the force by the plates themselves. For accurate and consistent results, an automated pulling device will be used to pull the top plate. Once a control has been measured, MR fluid will be placed between the two plates and the same procedure will repeat with and without the MR fluid under a magnetic field. These tests will characterize the force that MR fluid will generate when it is under a magnetic field and when it is free of a magnetic field Test Sequence 4 - Developing solenoid control Knowing the MR fluid shear stress properties will help determine the size and strength of the solenoid used for flight testing. This will also enable the group to decide on what type of control can be used on the solenoid. At the moment, an open loop control is considered. If better coupling can be achieved between sensors and actuators, closed loop control may be considered. Test Sequence 5 1- Slosh dampening A vibration rig will be constructed such that several frequencies of vibration approximating those experienced by the launch vehicle will be exerted on the ground test rig. Using similarity parameters, the data gained from this experiment will allow predictions for dampening performance of the controlled MR fluid during the microgravity period. eorgia Institute of Technology 188 of 196

189 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Appendix V: Science MFOs and Drawings Bench Stand Take one 1/16 12 x 12 acrylic plate and laser cut out a square that is 20cm x 20cm. Figure 77. Base Plate eorgia Institute of Technology 189 of 196

190 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Take the remaining two 12 x 12 plates and cut a rectangle that is 20cm x 22cm. Laser cut a 5.3cm diameter centered in the middle of the 20cm x 22cm acrylic plates as seen in Figure 2. Figure 78. Second and Top Plate eorgia Institute of Technology 190 of 196

191 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Take four 30cm Maker Bars and place them vertically 20 cm apart in the shape of a square. Attach horizontally a 20cm Maker Bar 2 cm above each 30cm Maker Bar base and secure them with 90 Degree Maker L Brackets as shown in Figure 3 and 4. This will be done on each side of the square. Figure 3 shows a side view of two 30 cm Maker Bars attached together by a 20cm Maker Bar horizontally. eorgia Institute of Technology 191 of 196

192 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Figure 79. Side view of main structure Figure 80. Trimetric view of main structure eorgia Institute of Technology 192 of 196

193 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Attach four flat L Maker Beam brackets on the 20 cm horizontal bars as show in in Figure 5. The short end of the L bracket will be placed 1cm from the end of the 20cm Maker Bar. This will hold the 20cm x 20cm solid acrylic base plate in place. Figure 81. Top view of structure with 90 Degree L Brackets eorgia Institute of Technology 193 of 196

194 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW On each vertical 30cm Maker Bar, attach the base of a 90 degree bracket 2 cm above the 20cm horizontal Maker Bar (5cm from the bottom of the vertical 30cm Maker Bar) as shown in Figure 6. Do this for each vertical 30 cm Maker Bar facing inward into the square. These four 90 degree brackets will hold the second acrylic plate. Figure 82. Side view of 90 Degree Brackets eorgia Institute of Technology 194 of 196

195 EORIA TECH RAMBLIN ROCKETEERS CRITICAL DESIN REVIEW Four cm above the top of each 90 degree bracket (9cm from the bottom of the vertical 30cm Maker Bar), place the base of another 90 degree bracket as shown in Figure 6. These four 90 degree brackets will hold the third acrylic plate. Place the 20cm x 20cm acrylic base plate that has no hole cut in it over the four, flat L Maker Beam brackets. The place the other two 20cm x 22cm acrylic plates that have holes in them on the 90 degree brackets creating three layers of acrylic plates as shown in Figure 7. Figure 83. Test Structure with Base Plates Place the 100 ml beaker through the acrylic plates and onto the base acrylic plate. eorgia Institute of Technology 195 of 196